Effects of ice accretions on aircraft aerodynamics

Progress in Aerospace Sciences 37 (2001) 669–767

Effects of ice accretions on aircraft aerodynamics Frank T. Lyncha,*, Abdollah Khodadoustb a

Lynch Aerodyn Consulting, 5370 Via Maria, Yorba Linda, CA 92886, USA (Former MDC and Boeing Technical Fellow) b Principal Engineer/Scientist, The Boeing Company, Huntington Beach, CA, USA

Abstract This article is a systematic and comprehensive review, correlation, and assessment of test results available in the public domain which address the aerodynamic performance and control degradations caused by various types of ice accretions on the lifting surfaces of fixed wing aircraft. To help put the various test results in perspective, overviews are provided first of the important factors and limitations involved in computational and experimental icing simulation techniques, as well as key aerodynamic testing simulation variables and governing flow physics issues. Following these are the actual reviews, assessments, and correlations of a large number of experimental measurements of various forms of mostly simulated in-flight and ground ice accretions, augmented where appropriate by similar measurements for other analogous forms of surface contamination and/or disruptions. In-flight icing categories reviewed include the initial and inter-cycle ice accretions inherent in the use of de-icing systems which are of particular concern because of widespread misconceptions about the thickness of such accretions which can be allowed before any serious consequences occur, and the runback/ridge ice accretions typically associated with larger-than-normal water droplet encounters which are of major concern because of the possible potential for catastrophic reductions in aerodynamic effectiveness. The other in-flight ice accretion category considered includes the more familiar large rime and glaze ice accretions, including ice shapes with rather grotesque features, where the concern is that, in spite of all the research conducted to date, the upper limit of penalties possible has probably not been defined. Lastly, the effects of various possible ground frost/ice accretions are considered. The concern with some of these is that for some types of configurations, all of the normally available operating margins to stall at takeoff may be erased if these accretions are not adequately removed prior to takeoff. Throughout this review, important voids in the available database are highlighted, as are instances where previous lessons learned have tended to be overlooked. r 2002 Elsevier Science Ltd. All rights reserved.

Contents 1. Introduction . . . . . . . . . . . . . . . . . . 2. Icing simulation techniques . . . . . . . . . . 2.1. Simulation fundamentals . . . . . . . . . 2.2. Overall simulation effectiveness . . . . . . 2.3. Scaling techniques . . . . . . . . . . . . 3. Aerodynamic simulation considerations . . . . 3.1. Single-element lifting-surface performance 3.2. Multi-element lifting-surface performance

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*Corresponding author. Tel.: +1-714-693-8797; fax: +1-714-779-3541. E-mail address: [email protected] (F.T. Lynch). 0376-0421/01/$ - see front matter r 2002 Elsevier Science Ltd. All rights reserved. PII: S 0 3 7 6 - 0 4 2 1 ( 0 1 ) 0 0 0 1 8 - 5

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3.3. Control surface effectiveness . . . . . . . . . 3.4. Drag characteristics . . . . . . . . . . . . . . 4. Effect of initial in-flight leading-edge ice accretions 4.1. Maximum lift reductions . . . . . . . . . . . 4.2. Stall angle reductions . . . . . . . . . . . . . 4.3. Drag penalties . . . . . . . . . . . . . . . . 4.4. Trailing-edge control surface characteristics . 5. Effect of runback and ‘‘ridge’’ ice accretions . . . 5.1. Maximum lift reductions . . . . . . . . . . . 5.2. Stall angle reductions . . . . . . . . . . . . . 5.3. Drag penalties . . . . . . . . . . . . . . . . 6. Effect of large in-flight ice accretions . . . . . . . 6.1. Maximum lift reductions . . . . . . . . . . . 6.2. Stall angle reductions . . . . . . . . . . . . . 6.3. Drag penalties . . . . . . . . . . . . . . . . 6.4. Trailing-edge control surface characteristics . 7. Effect of ground frost/ice accretions . . . . . . . . 7.1. Maximum lift reductions . . . . . . . . . . . 7.2. Stall angle reductions . . . . . . . . . . . . . 7.3. Drag penalties . . . . . . . . . . . . . . . . 8. Summary and conclusions . . . . . . . . . . . . . 9. Recommendations . . . . . . . . . . . . . . . . . Acknowledgements . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . Nomenclature List of Symbols 2-D two dimensional 3-D three dimensional c chord 1C degrees centigrade cd 2-D section drag coefficient clmax 2-D section maximum lift coefficient 1F degrees fahrenheit F flap h=c non-dimensional protuberance height k=c non-dimensional roughness height L lower surface Mo freestream Mach number S slat t=c thickness ratio TT total temperature U upper surface V velocity/speed Vsig 1g stall speed x=c chord fraction a angle of attack (deg) de elevator deflection angle (deg) df flap deflection angle (deg) ds slat deflection angle (deg) y glaze ice upper horn angle (deg) Acronyms and abbreviations An Antonov

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CFD deg DHC ETW FAA GA IBL IRT LE LEWICE LTPT LWC MAC MD MDC min mod. MVD NACA NAE NASA NLF NTF RAF RANS

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computational fluid dynamics degrees DeHavilland Canada European Transonic Wind Tunnel (US) Federal Aviation Administration General Aviation interactive boundary layer (NASA Glenn) Icing Research Tunnel leading edge (NASA) Lewis Ice Accretion Code (NASA Langley) Low Turbulence Pressure Tunnel liquid water content (g/m3) mean aerodynamic chord McDonnell Douglas McDonnell Douglas Corporation minutes modified median volume diameter (mm) National Advisory Committee for Aeronautics National Aeronautical Establishment (of Canada) National Aeronautics and Space Admini stration Natural laminar flow (NASA) National Transonic Facility Royal Aircraft Factory Reynolds Averaged Navier Stokes

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Ref. RJ RN, Rey. No. S&C SLD TAI

Reference Regional jet Reynolds number Stability and control supercooled large droplets thermal anti-ice

1. Introduction Degradation of the aerodynamic effectiveness of wings, tails, rotors, inlets, etc., due to contamination of these surfaces in any form can have serious consequences, particularly if not known and accounted for. Typical contamination sources include damage incurred during ground operations from foreign objects thrown up from the ground, or from encounters with ground support equipment, plus a number of in-flight sources such as encounters with hail, birds, insects, etc. However, the source which has undoubtably received the most attention is ice accretions on these surfaces. Numerous experimental results have shown that even quite small ice accretions at critical locations can result in substantial reductions in maximum lifting capability and control surface effectiveness, control surface anomalies, quite noticeable increases in drag, and, in some cases, reduced engine performance and stability. Decades of operational experiences have revealed many situations in flight as well as on the ground when ice can accrete on ‘‘ice-protected’’ aircraft. These include: *

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Leaving selected areas or components (which are susceptible to ice accretion) unprotected by ice protection systems. This can be a result of either a lack of aerodynamic criticality, and/or as a consequence of design optimization studies which consider the cost and availability of resources needed for ice protection versus other possible design solutions such as increasing aerodynamic surface sizes, etc. The initial, residual and inter-cycle ice accretions inherent in the use of de-icing systems. Delay in activating anti-icing systems (or any other ice protection system). Runback ice formations aft of areas protected by anti-icing systems when water droplets are not completely evaporated, either by design, or as a consequence of a very intense icing encounter. Ice accretions due to water droplet impingement aft of leading-edge ice protection systems. Upper-surface frost and/or ice accumulations caused by exposure to adverse weather conditions on the ground. Frost or ice which forms on both upper and lower surfaces of wings in proximity to the fuel tanks when an aircraft is parked in high atmospheric humidity

TE ted teu TIP TU

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trailing edge (elevator) trailing-edge-down (elevator) trailing-edge-up (NASA/FAA) Tailplane Icing Program Tupolov

conditions after fuel (which remains in the tank) has been ‘‘cold soaked’’ in flight. Much has also been learned over these many decades about the many varied forms in which ice can accrete on aircraft surfaces. Lessons gleaned from natural icing encounters in flight and on the ground, from aircraft flying behind tankers dispensing water droplets to simulate icing cloud conditions, and from a great number of tests conducted in icing tunnels, have taught us that ice accretes in many locations, sizes, and shapes, and with an assortment of ice surface roughnesses. The size of in-flight leading-edge ice accretions are by-andlarge proportional to the liquid water content (LWC) of the icing cloud, the velocity of the aircraft, and the duration of the icing encounter. Conversely, the size tends to be inversely proportional to the dimensions (leading-edge radius, etc.) of the aircraft component subjected to the icing encounter. Similarly, ice shapes can vary anywhere from relatively small, nearly uniform-thickness buildups to a wide variety of large, very irregular, so-called ‘‘horned’’ and ‘‘lobster tail’’ ice shapes. These shapes are dependent again upon the icing cloud conditions, duration of icing encounter, and aircraft surface dimensions, but also very strongly influenced by the ambient temperature. Likewise, the degree and type of resulting ice surface roughness/ irregularity are strongly dependent upon the size of the water droplets being encountered, the ambient air temperature, and the effectiveness of de-icing systems in removing all the ice. Similarly, runback ice and ground ice accretions vary widely. With the very large variety of forms and sizes in which ice can accrete on aircraft surfaces in real operational conditions, the challenge facing researchers and aircraft designers has been to establish an effective process for defining the accretion process and physical characteristics of these ice accretions for any aircraft surface at any flight or meteorological conditions, and to determine which are the most harmful accretions. Typically, some flight testing in natural icing conditions is required as part of the aircraft certification process for new aircraft designs in order to demonstrate the effectiveness of ice protection systems as well as overall aircraft performance and handling characteristics. However, such testing is not a practical approach for the bulk of the effort involved in systematically assessing the physical characteristics of ice accretions, for the

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development and validation of ice protection systems, or for conducting tests to measure the aerodynamic effects of ice accretions. In addition to the major economic and technical disadvantages inherent in the use of an aerodynamic design process not centered around the use of validated ice accretion simulation and scaling techniques in conjunction with ground test facilities, a design process requiring flight testing in natural icing conditions would be plagued with seasonal limitations, and the large uncertainty and risk involved in finding an appropriate range of icing conditions. Such testing would also not be sufficiently general, nor would testing in the accompanying atmospheric conditions be suitable for quantifying aircraft performance effects. Studying ice accretion characteristics by flight testing behind tankers avoids some of these disadvantages, but also introduces some important icing simulation issues of its own due to evaporation effects in the subsaturated conditions. Plume size, droplet size and spectra limitations are also icing tanker issues. Hence, in order to realistically address the wide range of real world in-flight ice accretion issues that must be considered in the aircraft design process, much effort has been focused for the past half century now on developing and trying to validate analytical/empirical simulation and scaling techniques. These techniques have been used in conjunction with icing tunnel tests, for both the development of ice protection systems, and to define the simulated in-flight ice shapes which are used to assess a range of associated aerodynamic impacts using conventional low-speed wind tunnels and flight testing. And, interestingly enough, with the exception of one very publicized fatal accident of an ATR-72 [1,2] and a number of incidents thought to be associated with encounters with the supercooled large droplets (SLD) occuring with freezing rain and freezing drizzle (for which there is still no formal certification requirement), there does not appear to be any evidence that any other accidents or incidents attributable to in-flight icing in at least the last couple of decades or more were traceable to inadequacies in these icing simulation and scaling techniques utilized in the design and certification process in spite of the many known limitations (i.e., inaccuracies) inherent in these techniques. Applications of these methods do, however, tend to be rather time consuming, and, there are some concerns that use of these methods may at times result in ice protection system designs that are not as energy efficient as they might be. As a consequence, high priority has been given to the development of essentially stand-alone, primarily computational (CFD), advanced simulation and scaling methods which would ostensibly permit icing design tasks to be accomplished more effectively and expeditiously in the future without compromising safety in any way. However, some significant CFD challenges still remain to be overcome in this regard.

While the primary objective of this review is the assessment of ice accretion effects on fixed-wing aircraft aerodynamic characteristics based on a survey of available experimental results, some understanding of the important factors and limitations involved in computational and experimental icing simulation techniques is helpful when considering the utility of these many measurements of the aerodynamic effects of simulated ice accretions. Consequently, the next section of this review provides somewhat of an overview of ice accretion prediction fundamentals, corresponding experimental techniques, scaling rules, and the various forms of ice accretions and critical parameters which influence them. More specifies on these methodologies for those interested are provided in a 1998 review in this journal by Kind et al. [3]. Following the discussion of icing simulation issues, the remainder of this review is then focused on assessing, correlating, and summarizing a range of measured aerodynamic effects caused by various forms of simulated ice accretions obtained at a wide variety of test conditions using numerous ground-test facilities, and, where available, flight testing. To start this assessment, some of the key aerodynamic testing simulation variables and governing flow physics situations which must be considered in order to realistically evaluate the relative merits of various experimental measurements are highlighted and discussed. Some crucial controlling flow physics topics discussed include attachment line conditions and related leading-edge flow conditions, various airfoil/wing/tail stall mechanisms, and spanwise variations in stall initiation on 3-D wings and empennages both with and without ice. The latter situation clearly illustrates the importance of understanding the prevailing flow physics for each flow situation being studied in order to intelligently interpret test results for assessing icing (and other) effects. A thorough understanding of testing simulation and flow physics strengths and weaknesses is crucial for interpreting test results and establishing their limitations, and a range of these testing simulation and flow physics situations and limitations are discussed for both single- and multielement airfoil/wing/tail geometries. Several important practical considerations which must be taken into account when planning test programs to assess ice accretion effects, or when utilizing test results for specific configuration applications, are also examined. These include the typical lack of general applicability of test results obtained with many specific 3-D geometries, uncertainties involved in defining the most critical ice buildups (including any residual accretions) for specific applications, and potential unknowns in reliably defining just what flow mechanisms are controlling stall characteristics for a wide range of practical aircraft geometries. While the focus in this review is on fixed wing aircraft, and, in particular, on the lifting surfaces

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(i.e., the wing and tail) of these vehicles, it is important to note that many of the lessons learned can also provide insights for icing situations with other applications such as rotors, propellers, intakes, etc. Ensuing sections then provide the actual review, assessment, and correlation of a large number of experimental measurements addressing the aerodynamic effects of various forms of mostly simulated in-flight and ground ice accretions, augmented where appropriate by similar measurements for other analogous forms of surface contamination and/or disruptions. For in-flight icing, the various pertinent experimental results are first sorted or grouped by the type of contamination, starting with the smallest accretions which would be of the order of small roughness size (with a nominal roughnessheight-to-chord ratio o20
104), representing the initial stages of an icing encounter or the required ice buildup inherent in the use of de-icing systems. Runback ice, or other (such as SLD) ice accretions located just aft of a leading-edge ice protection system, are considered next. Large rime and glaze ice accretions, including ice shapes with rather grotesque features, are the third inflight ice accretion category addressed. Following this, the effects of surface roughness associated with hoar frost accretions on wing upper surfaces that can be accumulated on the ground are addressed, as well as underwing frost associated with cold-soaked fuel. In each of these four accretion categories, test results are also grouped or sorted according to the complexity of the test configuration whenever possible, following a descending order of generality. For example, results obtained with single-element airfoils, wings, tails, etc., are reviewed and assessed before results obtained with more complex multi-element configurations, which are more difficult yet to generalize. The effects of test Reynolds number on the results will be illustrated whenever possible by grouping and assessing relatively low Reynolds number (o5
106) data separately from corresponding higher Reynolds number wind-tunnel and flight data more appropriate to larger (transport) aircraft. This separation is necessary because even though wind tunnel and flight studies carried out at low Reynolds numbers may well be pertinent for the flight regime occupied by some components (i.e., tailplanes, etc.) of general-aviation and smaller commuter-type aircraft, the use of such results by themselves to establish performance trends for higher Reynolds number applications can be very risky because of some well-known low Reynolds number anomalies such as laminar bubbles and transition variations which can occur on the baseline un-iced geometry. Such aberrations also raise issues regarding the generality of such results, even for other low Reynolds number applications. Consequently, concerns regarding the dangers of relying on low Reynolds number data are brought up/ repeated a number of times throughout the ensuing

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review because of the importance attached to this (often overlooked) factor. Incidentally, CFD methods have not been utilized in this review to either correlate or expand the existing experimental database on the aerodynamic effects of various ice accretions. This is because even the most advanced of these methods such as Reynolds Averaged Navier Stokes (RANS) have not yet been demonstrated to be reliable for this purpose, especially relative to determining whether a flow is separated or not (even on an uncontaminated surface). Claims of good agreement between CFD and experimental results involving separation onset/progression characteristics typically involve post-test computations wherein a number of adjustments (turbulence model, grid characteristics, dissipation, constants, etc.) can be made to facilitate the agreement. For example, having one turbulence model work best for one ice shape and another one for a different ice shape is not unusual. Also, obtaining ‘‘good’’ predictions of global (integrated) forces without agreement in pressure distributions (i.e., indicating that the real flow physics are not being properly modeled) has also been seen.

2. Icing simulation techniques The range of in-flight ice accretion types for which representative predictions and simulations are desired varies from the very initial stages of the ice accretion process up to the large, very irregular, and rough ice accretions. Properly simulating the initial stages of the ice accretion process is critical in the design and validation of effective ice protection systems, as well as in assessing the aerodynamic effects of relatively small ice accretions such as those associated with delayed activation of an ice protection system or inherent in the use of de-icing systems where some ice thickness is allowed/required to build up before it is (can be) broken loose. For these conditions, parameters of most concern are icing impingement limits, accumulation rates, collection efficiencies, and the severity of the ice surface roughness. Approaching the other end of the spectrum, a large variety of ice accretion types must be dealt with including the ‘‘ridges’’ located somewhat aft of the leading edge (LE) resulting from SLD encounters. And, while rime and glaze ice are the most commonly known, there are so many variations of these that a separate vocabulary has been devised to describe the various physical features [4,5]. Rime ice forms at temperatures below about 101C, and at cloud LWCs at the lower end of spectrum, where a high rate of heat loss from the impinging droplets causes them to freeze upon impact. These ice shapes normally have a single horn, tend to conform somewhat to the surface geometry, and have a relatively smooth surface. Glaze ice, on the other hand,

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forms at temperatures between about 31C and freezing, and at higher LWCs, where there is insufficient heat transfer to remove all of the latent heat of the impinging water droplets, and so some unfrozen water remains. Glaze ice accretions are normally larger, typically develop pronounced double horns and/or ‘‘lobster tails’’, and have very rough surfaces, but there are a myriad of variations possible over a range of air temperatures, LWCs, and surface geometries. In between these rime and glaze ice accretions lies a wide variety of ‘‘mixed’’ ice accretions. Obtaining representative predictions and/or simulations of these larger ice shapes is clearly desired in order to enable realistic assessments of the aerodynamic consequences of such accretions. However, while accurate predictions and simulations may be desired, it is important to put the accuracy realistically achievable (now and in the foreseeable future) with current simulation methodology in proper perspective in light of some of the fundamental limitations inherent in the methodology employed. 2.1. Simulation fundamentals The objective of in-flight icing simulation is to accurately represent the time-dependent ice accretion process which occurs when an aircraft flies through a cloud containing super-cooled water droplets, or encounters freezing rain or drizzle. It has been well established that three fundamental factors must be taken into account concurrently for realistic simulation. They are the representation of the aerodynamic flowfield (potential and viscous) characteristics on and around the body and growing ice accretion, the establishment of the water droplet trajectories with subsequent impingement characteristics and limits, and, lastly, the thermodynamics of the freezing/ice growth process. In the analytical simulation process, successive thin ice layers are formed on the surface, and are followed by flowfield, droplet impingement, and ice growth recalculations. The frame of reference adopted to date for addressing these factors in analytical as well as ground-test (icing tunnel) studies of the ice accretion process has been the standard Galilean transformation of fixed body with intended or assumed uniform onset flow, and with the water droplets at the freestream velocity prior to encountering the aircraft flowfield [3]. 2.1.1. Flowfield determination A close representation of the flowfield around and on the aerodynamic surface and growing ice shape, either computationally or experimentally, serves two important functions. The off-body flowfield is clearly important for proper simulation of water droplet trajectories, while the (viscous) flow characteristics on the surface are critical in establishing the convective heat transfer levels and distributions which have a major

impact on ice accretion rates as well as design requirements for ice protection systems. Since the boundary-layer transition process, and the existence of any separated flow regions, have a major influence on establishing local heat transfer rates, then any flowfield (or geometry) variables which are known to influence transition and separation such as Reynolds number and freestream turbulence levels should be carefully considered and reproduced to the extent possible. Transition location is more critical in the early stages of ice buildup, and for the design of ice protection systems, while separation becomes more of an issue as ice shapes become more convoluted with additional icing exposure. Choice of a most appropriate flowfield method for this purpose is still being debated. Ideally, and realistically, the method selected should be capable of handling complex aircraft and ice geometries, and flow situations, including large flow separations caused by the existence of ice horns, etc. And, very importantly, the method should also be able to account for the effects of substantial (and often irregular) ice surface roughness. Candidate flowfield prediction methods normally suggested to provide these capabilities range from linear panel methods with either integral boundary layer [6] or interactive boundary layer (IBL) representation [7] up to RANS [8]. While it might have been presumed that (RANS) methods would be the obvious choice to capture the necessary physics, particularly the separated flow regions, this is clearly not yet the case. Reliable predictions of separation onset and progression characteristics for a range of geometries and conditions remains a most needed, but elusive, goal for RANS, and that is the situation for smooth surfaces without high local curvatures. The challenge would be even more formidable for rough, irregular ice shapes with very high local curvatures. Establishing suitable field grids for the irregular ice shapes has also been a formidable challenge for the structured-grid Navier Stokes methods [9]. Important non-physical modifications of these ice shapes such as ‘‘fairing out’’ the deep cavities and smoothing surface irregularities have been necessary in order to avoid highly skewed and/or crossing grid lines. Panel methods (with IBL), on the other hand, are theoretically well suited for analysis of these complex geometries because no field grids are required, but they are even more limited (than RANS) in their ability to adequately address flows with any significant regions of separated flow or strong viscous/inviscid interactions. Extensive, somewhat arbitrary ‘‘smoothing’’ of irregular ice shapes is also necessary in order to obtain IBL solutions for these ‘‘geometries’’ [7]. 2.1.2. Droplet trajectories Establishing representative water droplet trajectories is the second step in the process for simulating ice accretions which occur when an aircraft penetrates

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a cloud containing supercooled water droplets, freezing drizzle, or freezing rain. These trajectories are required in order to determine the amount (i.e., total collection efficiency), distribution (local collection efficiency), and extent (impingement limits) of the water which strikes the various aircraft surfaces (or developing ice shapes) for a prescribed icing cloud content. The basic mathematical formulation for droplet trajectory calculations in the presently used frame of reference was developed over 50 years ago by Langmuir and Blodgett [10], although various researchers have subsequently implemented a number of refinements and enhancements [11–15]. In this process, the trajectory of a droplet is determined by integrating a differential equation representing the force balance on the droplet involving inertial, drag, buoyancy, and gravitational forces. Trajectories for a series of droplets are determined starting from an upstream (freestream) position until the droplets either collide with or bypass the surfaces. A fundamental assumption inherent in the use of this process is that the proportion of water in the approaching flowfield is low enough that the influence of water droplets on the flowfield can be ignored [16,17]. Physical parameters of prime importance in droplet trajectory simulation are the size(s) of the droplets, the overall cloud content, and the size of the surface (leading-edge radius) subject to collision with the droplets. Smaller and/or more slender surfaces tend to collect much more ice and have more aft impingement limits because the droplets have less chance to be deflected by the (smaller) leading-edge flowfield influence. Droplet size is, however, the single most important variable in droplet trajectory simulation, largely because of the impact on the balance between droplet inertia and drag forces. Size also has an influence on whether droplet terminal fall velocities become significant enough that they are no longer negligible. Droplet sizes of interest for icing vary from 10 to 50 mm (diameter) range routinely addressed to date in the certification process at the low end, up to about 400 mm encountered in freezing drizzle, and even larger in freezing rain. Droplet sizes encountered in the latter two categories are now often referred to as SLD [18]. At the smaller sizes normally addressed in the certification process, the terminal fall velocities are quite generally accepted as being negligible [19]. At the low end, droplet drag forces dominate the inertia forces, so these droplets tend to closely follow the streamlines. However, as the droplet size increases, the inertia forces (proportional to the cube of the droplet diameter) become dominant over the drag forces (proportional to the square of the diameter). Under these circumstances, the droplets do not tend to follow the streamlines, and are therefore much more likely to impact the surface, with resultant greater collection efficiencies and often more downstream impingement limits. Further, with the much larger

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SLDs, terminal fall velocities which occur in natural icing conditions might have a small impact. For example, a 500 mm droplet would be expected to have a fall velocity of about 2 m/s. This might be expected to move upper-surface impingement limits even a little further downstream. Computational and icing tunnel simulations of the overall cloud contents found in nature can only be approximations, since the conditions encountered in the natural icing environment are believed to be quite nonuniform and irregular. For example, precipitation typically occurs when the cloud droplet population becomes unstable, and some droplets grow at the expense of others by collision and coalescence as a consequence of having different fall velocities. For the purpose of icing simulation, the cloud content is usually defined by the LWC, and either a droplet diameter distribution, or a single droplet size selected to be representative, which is normally referred to as the median volume diameter (MVD). One common distribution definition utilized to date is called the Langmuir-D distribution [10]. Resultant collection efficiencies and impingement limits may be a little different for these two approaches to modeling the cloud contents, but the ice accretions predicted usually end up being about the same. The existing conventional (for wind tunnels and computational studies) frame of reference used for droplet trajectory (and icing) simulations has been widely used and accepted for the past half century or so. This widespread acceptance can be explained by the absence of any glaring discrepancies between ice accretions established using this framework and those observed in-flight at ‘‘corresponding’’ conditions for normal droplet sizes with single element wings, etc., that were not thought to be explainable by some known shortcoming in the simulation of the thermodynamics of the freezing/ice growth process [20]. There have, however, been some concerns raised more recently regarding the adequacy of these techniques for SLD conditions, and for some aspects of ice accretions obtained with subscale model tests of multi-element geometries in icing tunnels. For the SLD conditions, the current concern is the need to account for loss of mass because of drop break-up due to shear forces before reaching the airframe, and then due to drop splashing and bouncing. Gravity effects on SLD are also a concern for icing wind tunnels. For multi-element high-lift geometries, the concern has been with regard to some of the ice accretions observed on the downstream elements in icing tunnels. To illustrate this point, typical ice accretions observed during a test of an advanced high-lift system geometry [21,22] are depicted in Fig. 1. While there have been a few (although not very many) observations and reports [23] of ice accretions occurring on flap LEs during flight in natural icing conditions,

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Fig. 1. Ice accretions on multi-element airfoil in icing tunnel [21,22].

there are no known observations or reports of ice forming around the LE of the main element (aft of the deflected leading-edge device) as occurs in icing tunnels. Initial speculation focused on the absence of any droplet terminal fall velocity component in the simulation process as the possible culprit, but subsequent studies of droplet trajectories [24] have suggested that the observed ‘‘disconnect’’ may also possibly be explained to some degree by ‘‘scale differences’’. Another possibility is that aircraft incidence angles at realistic operating conditions do not normally reach the levels where these ‘‘downstream’’ accretions seem to be more prevalent and/or pronounced in icing tunnel tests. Further

investigations to resolve this matter are certainly warranted, including some determination as to just how representative are the ice shapes formed on the flap LEs in icing tunnels. 2.1.3. Thermodynamics/freezing process For many years, it was assumed that once droplet impingement characteristics were established, the actual ice accretion process was governed almost exclusively by the convective heat transfer characteristics at the surface. The method widely used to simulate the accretion process was initially formulated by Tribus et al. [25] in the late 1940s, and later enhanced by

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Messinger [26]. The implementation of this method in the American, Canadian and European ice accretion prediction codes [3,6] involves a control volume approach. In it, each segment of the surface is assigned a control volume located on the surface and extending beyond the boundary layer, and a mass and energy balance is performed on each control volume using the first law of thermodynamics to determine the quantity of ice accreting within it. Any water which does not freeze in a control volume is assumed to run back and freeze in a subsequent control volume. This methodology produced the popular concept of a ‘‘freezing fraction’’, which is the ratio of the liquid freezing within a control volume to the total amount of liquid entering the control volume. Unfortunately, even though this methodology is typically adequate for rime ice predictions, it was shown some time ago by Olsen and Walker [20] that this freezing fraction concept/approach does not work well for glaze ice accretions for a number of reasons. For instance, the absence of any momentum balance results in important water surface tension and external flowshear effects on the runback droplets not being accounted for [27–29]. And, any imprecise knowledge of local heat transfer characteristics and effects becomes much more limiting in the simulation of these accretions. The heat transfer/freezing process for glaze and mixed ice accretions is complex, and is governed by many factors in addition to the temperature difference between the aircraft/ice surface and impinging water droplets. For example, whether the boundary layer on the surface is laminar, transitional, turbulent, or separated, has a strong influence on the heat transfer rates, as does the existence of any water layer on the surface. The state of the boundary layer is, in turn, strongly dependent upon the Reynolds number, surface sweep, surface/ice geometry, surface roughness, and freestream turbulence/ background noise levels. And, for a particular boundary layer state, heat transfer rates are a function of local skin friction level, Prandtl number, surface velocity, etc. Surface roughness also has a major impact on increasing heat transfer rates with transitional and turbulent boundary layers. While an accurate determination of local heat transfer characteristics is not that important for predicting representative rime ice accretions, it is if consequential predictions of glaze (and mixed) ice accretions are to be achieved. This requirement represents a formidable problem. Perhaps the toughest part lies in being able to effectively account for both roughness and surface tension effects, which are often very closely intertwined. Contributing to the difficulty in understanding these complex flow phenomena, especially for larger aircraft, is the fact that most of the experimental investigations carried out to explore these phenomena have been conducted with small 2-D single-element models at quite low Reynolds numbers in existing icing tunnels (which often have elevated

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turbulence levels caused by the water droplet dispensing rig, etc.). Important factors in the freezing process, especially in the early stages of glaze ice accretions, are strongly influenced by the particular surface viscous flow features peculiar to these low Reynolds numbers on 2-D single-element models, thereby causing the observed ice accretion characteristics at these conditions to likely be of limited value for assessing other geometries and conditions (i.e., higher Reynolds numbers). For example, the significant runs of laminar and transitional boundary layers at the lower Reynolds numbers will result in different accretion physics than would be experienced on 3-D swept surfaces at higher Reynolds numbers where the surface boundary layer may be all (or nearly all) turbulent. Even 2-D test measurements [30] have documented that modest increases in Reynolds number result in higher heat transfer rates over and just downstream of ice-type roughness. Unfortunately, ground test facilities do not presently exist which would permit the detailed study of ice accretion processes at the noticeably higher Reynolds numbers and lower freestream turbulence levels representative of flight for larger aircraft, although the use of so-called hybrid models (discussed in Section 3.2) has enabled some quite useful insights in this regard. The surface conditions in the initial stages of the ice accretion process are often described as closely resembling uniformly distributed roughness, with measured roughness heights typically several times larger than the boundary layer thickness close to the LE of an un-iced surface [31]. However, rather than causing a quick transition to a fully developed turbulent boundary layer as was initially thought, detailed (2-D low Reynolds number) experimental studies have shown that this relatively large distributed roughness instead started a transitional boundary layer which exists for a significant chordwise distance before attaining a fully developed turbulent state [32]. Corresponding measurements have revealed that much-higher-than-expected heat transfer rates existed in this rough transitional boundary layer region. Experimental studies of glaze and mixed ice accretion characteristics have also shown, however, that this stage with relatively uniform distributed roughness does not last for very long. A rapid ensuing roughness growth takes place which then leads to a very nonuniform surface roughness distribution which is difficult to model effectively. This is a significant concern since good estimates of roughness size and characteristics are needed as the first step in establishing reliable heat transfer coefficients. As indicated earlier, water surface tension effects have also been shown to play an important role in the accretion process for glaze and mixed ice formations, and surface roughness effects play a substantial role in this aspect as well by impeding the movement of water droplets along the surface. For rime ice accretions, it is

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believed that there is little, if any, movement of surface droplets, since freezing creates some rough ice which greatly increases surface tension adhesion forces which, in turn, prevent aerodynamic forces from moving the droplets. On the other hand, studies of the glaze ice accretion process by Olsen and Walker [20] (at low Reynolds numbers) did reveal the presence of two distinct regions in the early stages. One is a smooth region near the stagnation line characterized by a thin film of water over a uniform layer of ice, while the second is a region somewhat downstream where there is an immediate change into a rough zone, and there is no film of water present. And, ice accretion rates are much higher in this region due to the effects of roughness. It was also observed that the boundary between the smooth and rough regions propagates upstream with increasing time, at colder temperatures, and with increasing freestream velocity and LWC. Regarding the details of the low Reynolds number glaze ice accretion process, it was seen that as water droplets impinge and wet the surface, they coalesce to form larger beads, which, as a consequence of aerodynamic forces acting upon them, run back slowly along the surface until they either reach the point where the surface tension forces balance the aerodynamic forces, or the base of the water beads freezes, forming an ice substrate which helps hold the water beads in place. The freezing of these large droplets, greatly enhanced by the attendant high heat transfer rates associated with the large roughness, starts the formation of the characteristic glaze ice horns. Once the horns start to protrude, they then ‘‘catch’’ an even greater number of droplets relative to surrounding areas, and the horn growth rate increases. Unfortunately, even though the understanding gained from these detailed studies of the glaze ice accretion process on 2-D single-element models at quite low Reynolds numbers does provide some important insights into what is important for effective analytical modeling of these complex flow phenomena, this existing database does not provide much of a basis to guide the development and validation of future, primarilyFCFD modeling of glaze and mixed ice accretions for the wide range of surface geometries and Reynolds numbers needed. For example, just for starters, it would be expected that the aerodynamic forces acting on the coalescing and coalesced water droplets would be noticeably different if turbulent attachment line conditions existed, rather than the starting laminar boundary layer present in the existing 2-D low Reynolds number tests. And, as if the computational modeling challenges for glaze and mixed ice accretions were not already difficult enough, they are even further complicated (at all Reynolds numbers) if other known factors influencing these accretions such as droplet splash back effects, the shedding of surface

droplets (especially near freezing temperatures), and the effects of unsteady flow phenomena involving vortical flow structures, etc., are to be accounted for. 2.2. Overall simulation effectiveness Clearly, the two issues/questions needing to be addressed are what improvements to current icing simulation capabilities are required (i.e., just how precise or accurate do the droplet trajectory and simulated/ predicted ice shapes and roughness characteristics have to be), and what improvements are realistically attainable. To start with, it is clear that modifications to permit realistic predictions of SLD impingement characteristics, including any terminal fall velocity component present, are definitely needed, and should be possible. However, with regard to the need for other improvements, a number of factors need to be considered. For one, it is important to recognize that icing simulation methods at best can only provide approximate estimates of ice accretions, since it is really impossible to accurately represent the non-uniform and irregular conditions encountered in the natural icing environment. For instance, just how important can it be to go to great detail to represent very irregular surface roughness characteristics on approximate ice shapes when these roughness characteristics have been observed on a very limited number of geometries in an icing tunnel environment at relatively low Reynolds numbers. Also, as indicated previously, other than one exception thought to be associated with an SLD encounter, there is no evidence that any other fatal accidents in modern times attributable to in-flight icing were caused by inadequacies in these commonly used icing simulation (and scaling) techniques. This, together with the existence of (high Reynolds number) test results from years ago [33] which indicated that even very small ice accretions can have a detrimental effect on maximum lift capability nearly as large as that seen with large horned ice shapes, has led to speculation that perhaps any large glaze ice shape (and roughness) may provide a representative order-of-magnitude degradation in maximum lifting capability for lifting surfaces (except, perhaps, for SLD conditions). This possibility needs to be explored since, if true, it leads to the conclusion that additional improvements to the icing simulation methods (other than for SLD conditions) may not be needed, at least from a safety perspective. At this point, there does not appear to be a compelling justification for pursuing improvements to current ice accretion predictions techniques (for the sake of science). This issue will be revisited after the subsequent review of test results addressing the aerodynamic effects of various simulated ice accretions. If it turns out (i.e., can be shown) that there is a real need for pursuing further refinements, etc. to these icing

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simulation methods (in addition to the SLD capabilities), the technical community may be faced with the reality that just ‘‘wanting’’ or ‘‘needing’’ these improvements does not make them possible (in the foreseeable future). It is clear that the focus would have to be on glaze ice shape and roughness predictions, since that would be the area in most need of improvement. However, based on the earlier discussion of the most formidable CFD challenges involved in providing more accurate glaze ice accretion predictions, it would not be reasonable or prudent to presume that these advancements could be achieved any time soon.

2.3. Scaling techniques As indicated earlier, flight testing in either natural icing conditions or in simulated clouds produced by tankers is not a viable approach for systematically assessing the physical characteristics of a broad range of ice accretions, or for the development and validation of ice protection systems. And, since it is often not possible to test full scale models of major aircraft components such as wings or even tailplanes in icing tunnels except for the very smallest of aircraft, a high priority has been put on the development of scaling techniques in order to enable the use of subscale models in existing icing tunnels to study this wide range of icing issues. The objective for scaling procedures is to permit the definition of similar ice accretions, both in shape and surface roughness, on geometrically similar aerodynamic surfaces under different flight and meteorological conditions, thereby negating test facility operational limits. It should be understood without saying that the starting point for any viable ice accretion scaling laws or techniques must be the consideration of the same fundamental factors essential for basic icing simulation, namely the flowfield, droplet trajectories and corresponding collection efficiencies and impingement limits, plus the thermodynamics of the freezing process. It then follows that scaling techniques will tend to have similar strengths and weaknesses. Matching of many of the important nondimensional parameters which should be duplicated for meaningful scaling, especially those controlling droplet impingement characteristics, the mass of water impinging on the surface, and the thermodynamics of the freezing process, can be accomplished by maintaining the scale between the (scaled) test object and droplet sizing/spacing (matching LWC), and maintaining the same ambient temperatures. However, the companion requirement to hold Mach number, Reynolds number, and Weber number constant between simulations is, unfortunately, not satisfied by such relatively simple scaling. Even if the Mach number requirement could be waived, specifying that both Reynolds number and Weber number be held constant

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requires different scale velocities, both higher than the reference value when subscale models are used [34,35]. Most of the early scaling methods developed by a number of different organizations worldwide [36–41] have typically employed similar approaches for scaling droplet and ‘‘accumulation’’ parameters, but somewhat different ones for selection of thermodynamic scaling parameters, although usually based on Messinger’s freezing fraction approach (with all associated limitations). As would be expected, these techniques produced reasonably good scaling results for rime ice accretions, but no one has ever claimed success for scaling glaze and mixed ice accretions. Subsequent improvements to include surface tension effects yielded some improvements, but still short of that hoped for, since water droplet impact and film dynamics were not addressed [27]. Scaling methods wherein Weber number is held constant (at the expense of Reynolds number and Mach number simulation) have provided some encouraging results for glaze and mixed ice accretions [42], albeit for a very limited range of conditions at low Reynolds numbers in an icing tunnel. Since these scaling methods are sometimes employed, or at least intended, to experimentally define ice accretions for subsequent vehicle performance assessments, it is important to understand the deficiencies of even the best of these current scaling techniques [29,43,44]. Some important limitations naturally arise due to the previously discussed inadequacies of the current analytical simulation methods (for glaze ice) for modeling the flowfield and freezing/accretion process upon which the scaling methods are based. Another very important shortcoming is the lack of any validation results at the higher Reynolds numbers of interest. Obtaining ‘‘similar’’ ice accretions for two different low Reynolds number conditions in an icing tunnel does not automatically imply in any way that the resultant ice shapes are accurately duplicating shapes which would be produced in flight at a higher Reynolds number reference condition. Other concerns with the relevancy of existing scaling methods arise because these methods are basically only applicable to 2-D single-element geometries at this time, and therefore not obviously directly applicable to swept and multi-element configurations, although some surprisingly good representations of rime and mixed ice accretions on swept wings in flight have been obtained with these 2-D methods. Further, the constant Weber number scaling method for glaze ice accretions has some serious limitations because of the noticeably higher freestream velocities required when using subscale models in an icing tunnel. These higher airspeeds generate higher total temperatures which can, in turn, produce nonrepresentative accretion characteristics when temperatures approach freezing. Use of the aforementioned hybrid airfoil models with closer-to-full-scale LEs, and redesigned aft sections, has

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been successfully used in a number of cases to overcome some of the aforementioned shortcomings [45–50], especially for holding and cruise configurations during which extended exposure to icing occurs. However, the use of this concept does have some limitations. For one, the flowfield around the LE of these designs at elevated angles of attack can become non-representative because of premature separation on the upper surface. Also, this concept is not really applicable to high-lift geometries, and these hybrid models are not useable for determining the various performance effects of the ice accretions formed. Consequently, considering the many deficiencies and limitations outlined, current scaling methods just like the aforementioned icing simulation methods, do not do a very good job of defining actual ice accretion shapes for many practical 3-D geometries at flight conditions, except perhaps for less disruptive rime ice and some mixed ice accretions on single-element surfaces. Similarly, however, it remains to be seen whether these imperfect scaling techniques might actually be adequate from a safety perspective.

3. Aerodynamic simulation considerations Detrimental aerodynamic effects caused by ice accretions which are of most concern to aircraft designers and operators, as well as regulatory agencies, include: *

Reductions in lifting capability of lifting surfaces, such as wings and tail surfaces, especially loss of control due to reductions in maximum lift capability (i.e., stall margins) and associated pitching moment changes.

*

* *

Loss or rapid changes/reversals of control surface effectiveness on wings, empennages, etc. (especially for unpowered surfaces). Noticeable drag increases. Disruption of airflow to engines.

For each of these, there are a number of key flow physics characteristics and testing simulation features which must be carefully evaluated if existing experimental results for ice accretion (and other forms of contamination) effects are to be properly interpreted and/or applied. These key factors and considerations are reviewed in the following subsections for four important situations. 3.1. Single-element lifting-surface performance Establishing meaningful incremental lift effects due to ice accretions on this class of geometries is typically more dependent upon achieving/identifying representative aerodynamic lift levels and characteristics on the baseline (un-iced) configuration than on the iced one, particularly when using ground-test facilities. This is because the maximum lift levels achieved by the un-iced designs are usually quite sensitive to the leading-edge flow physics, particularly the extent of laminar flow, which in turn is a strong function of the particular airfoil or wing design, the test Reynolds number, the test facility flow quality, model installation effects, etc. This sensitivity is well illustrated by test results obtained over a range of Reynolds numbers in two pressurized wind tunnels for a representative 2-D wing airfoil section [23,51,52]. The measured maximum lift levels are compared in Fig. 2, where it can be seen that the levels obtained in the two tunnels differ noticeably at the lowest Reynolds number, with the lower level being

Fig. 2. Maximum lift differences attributable to dissimilar tunnel flow quality.

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experienced in the NASA Langley Low Turbulence Pressure Tunnel (LTPT). A comparison of the surface pressure distributions in the leading-edge region revealed that a laminar bubble was present at the higher angles of attack in the LTPT, thereby limiting the maximum lift level attainable, but did not exist in the NAE facility with its higher turbulence and noise levels. It would therefore be expected that using the lower Reynolds number LTPT measurements for this baseline (un-iced) airfoil would yield a lower maximum lift penalty for leading-edge ice accretions than would be experienced in the NAE facility at this lower Reynolds number, and would result in a significant underestimation of the maximum lift penalty which would occur in flight at higher Reynolds numbers where laminar bubbles would not exist on the baseline. Hence, one must be very careful about using low Reynolds number test results to predict high Reynolds number ‘‘consequences’’. A number of examples illustrating this concern are presented in the following sections. However, low Reynolds number data on iced configurations may be useful in indicating trends if interpreted with care with full regard to the physics of the flow. Two other important factors to be considered when assessing 2-D airfoil measurements of maximum lift penalties caused by ice accretions include being able to differentiate between the different flow physics limiting maximum lift on various airfoil types, and assuring that wind-tunnel test measurements are not overly compromised by inadequate side-wall boundary-layer control. It is well documented that some airfoil types experience a leading-edge type of stall characterized by a rather abrupt loss in lift beyond stall. Others, such as relatively thin airfoils, have a leading-edge type of stall (at a much reduced lift level) without the abrupt lift loss beyond stall, while other designs tend to have a much-lessabrupt trailing-edge type of stall. And, of course, there are variations and combinations of each. It should not be a surprise if the effects of leading-edge ice accretions, etc. on maximum lift levels were different for these different types of stall characteristics, likely being more severe for the configurations experiencing an abrupt leading-edge stall. As a case in point, it has been shown that the sustainable leading-edge peak suction pressure coefficient is reduced dramatically with even the smallest amount of leading-edge roughness [51,52]. Next, the need for effective side-wall boundary-layer control to maintain 2-D flow conditions for 2-D airfoil tests has been well established [53]. Again, this control is likely more important for attaining representative maximum lift values for the baseline un-iced airfoil. In this case, the boundary layer in the juncture flow region at the sidewall, which is more susceptible to separation than the 2-D-like boundary layer in the center of the airfoil, can well dictate the maximum lift level achievable. On the other hand, with leading-edge ice accretions, etc. on

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the airfoil itself, the flow in the juncture region could conceivably be relatively less critical, but that is pure conjecture, and not to be counted upon. The proper assessment of ice accretion effects on 3-D wings, tails, winglets, etc. becomes somewhat more complex because, in addition to all of the aforementioned factors critical in assessing 2-D airfoil effects (with the exception of the sidewall-boundary-layer control issue), it becomes necessary with 3-D configurations to determine and account for the spanwise variation in stall initiation on the un-iced baseline configuration, because that can be a critical factor in establishing subsequent (maximum) lift and pitching moment changes caused by the ice accretions. Well designed wings, as well as other lifting surfaces, do not normally have the stall occur simultaneously across the whole span. Typically, a particular spanwise location is selected to be critical for stall initiation to achieve desired stall characteristics. As a result, regions of the wing away from this area are less critical with regards to stall initiation. Consequently, whereas ice accretions at the critical spanwise location will cause an immediate earlier stall, other locations along the span may possibly accrue some small amount of ice before these sections would become critical in initiating stall. It is important to note that with relatively large wind-tunnel models, tunnel wall effects can also modify those characteristics [54]. Further, the installation of wing-mounted engines (nacelles and pylons), fences, etc. can also significantly modify stall behavior. Hence, when planning tests or assessing test results for the effects of ice accretion (and other forms of contamination) on the lifting capability of 3-D wings, tails, winglets, etc., it is essential that potential spanwise differences in relative criticality for stall initiation be taken into account. Typically, these effects have not been considered or accounted for to date in publications [55,56] reporting on the measured effects of various disruptions on the maximum lift capability of 3-D configurations. Therefore, these previous flight-test results for burred rivets, insect debris, chipped paint, etc., are not readily usable for quantitative assessments. In addition, it is important to note that ice accretion effects on the lifting characteristics of thin, highly swept wings suitable for supersonic cruise represent a quite different set of flow physics and consequences to deal with, although this situation is probably not as critical because these configurations typically do not experience the type of stall encountered by subsonic aircraft. 3.2. Multi-element lifting-surface performance Appropriate assessment of ice accretion effects on the lift characteristics of multi-element high-lift geometries is considerably more complex than for single-element designs because of the wider range of geometries, flow

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physics features, and flow conditions needing to be addressed. To start with, it is necessary to consider both takeoff and landing configurations/conditions in order to address the effects of ground frost accumulations as well as in-flight ice accretions. And, for each, it is necessary to understand the relative criticality and important interactions between the various elements for any critical flight conditions such as maximum lift, and lift at takeoff and approach attitudes. It is also important to identify where (i.e., which elements, etc.) ice can potentially accrete in each of these situations, but, as discussed previously, there are still some unresolved issues remaining with regard to identifying realistic ice accretions on elements aft of the leadingedge element. In order to put the complexities involved in assessing ice accretion effects on multi-element high-lift geometries into perspective, it is instructive to examine the aerodynamic characteristics of a representative 2-D three-element (slat, main, and flap) geometry in the landing approach configuration, and then consider how the limiting flow physics for conditions of interest can change with configuration modifications. The baseline geometry selected has been the subject of considerable research [54,57–63], both experimental and computational, and is well suited for this purpose. Both the slat and flap are deflected 301. This geometry, as well as the measured element (and total) lift characteristics for a range of angles of attack leading up to maximum lift conditions in the NASA Langley LTPT, at a freestream Mach number of 0.2, and a (stowed) chord Reynolds number of 9
106, are shown in Fig. 3. Incidentally,

Fig. 3. Element lift distribution with multi-element high lift airfoil.

these results were all obtained using a good state-of-theart sidewall boundary-layer control system. However, there is no guarantee that all sidewall viscous influences have been avoided at critical conditions. It is important to note that the requirement for effective sidewall boundary-layer control is much more crucial for multielement geometries since the sidewall boundary layer experiences the total adverse gradient from the three elements in this case. At these testing conditions, and with the basic slat and flap riggings employed, it can be seen that the maximum lift attainable is limited by an eventual reduction in lift on the main element. A close examination of the surface pressure distributions reveals that the lift reduction on the main element beyond maximum lift is caused by an unloading of the aft part of the main element, which, in turn, appears to be caused by a rather dramatic unloading of the flap itself. However, surface skin friction measurements show that neither the flap nor main element surface boundary layers are even close to separation at these conditions, nor is there any flow separation on the slat upper surface, even though the peak suction pressures near the LE are quite high, corresponding to a local Mach number of nearly 1.2. Instead, it appears, based on a review of off-body flow field measurements, that the initial flap unloading is somehow brought about by the very rapid spreading and merging of shear layers and wakes (from the main element and slat) above the flap (although there is also the possibility that some separation in the sidewall juncture region may be contributing to this limiting phenomena). In contrast, at lower-angle-of-attack approach conditions (aB81), wake spreading and merging are of much less consequence, but the flow on the flap upper surface is very close to separation near the trailing edge (TE). The flow on the upper surfaces of the slat and main element are very noncritical at this condition. With the preceding set of flow physics characteristics, certain ice accretions are likely to have more serious consequences than others. Reflecting on maximum lift conditions, any ice accretions which thicken the slat and/ or main element wakes are certainly expected to result in some reductions in maximum lift. Also, any measurable ice accretions on the slat are expected to result in an earlier, leading-edge-type stall (similar to the thin-airfoil type), since the peak velocities on the un-iced slat at maximum lift are close to experimentally observed limits. With the flap being so uncritical in terms of flow separation at maximum lift conditions, any small ice accretions on the LE of the flap, even if they existed at flight conditions, do not pose that much of a threat. However, at approach conditions where the flow on the flap is close to separation, any such accretions could lead to a loss in lift and flap buffeting. At the same time, any ice accretions on the slat or main element at these approach conditions are not nearly as critical. This

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whole situation can, nevertheless, change quite significantly with possible slat and flap rigging changes, or even with a freestream Mach number change. For instance, if the slat were deflected further (i.e., overdeflected), then the maximum lift attainable would not be nearly as sensitive to initial ice accretions on the slat. The opposite would be true if the slat were deflected less (i.e., under-deflected). On the other hand, if the flap were deflected further, or if the negative overhang were increased, then any ice accretions on the LE of the flap would have a more adverse impact. The same would apply at approach conditions with respect to the flap rigging changes. Further, an increased freestream Mach number at maximum lift conditions could result in even more serious penalties for leading-edge ice accretions. At takeoff conditions, where the slat and flap are normally not deflected as much as for approach, stall is normally initiated by flow separation at the LE. Hence, any ice accretions on the slat would be expected to have the greatest adverse impact, although the magnitude of the impact would be reduced if overspeed (i.e., using higherthan-the-minimum required speed) was being utilized. Consequently, it is absolutely essential that the flow physics controlling stall, or any other situation involving potential flow separation, be carefully considered when assessing the effects of various potential ice accretions on the lifting characteristics of multi-element configurations. Analogous with the situation for single-element lifting-surface configurations, the proper assessment of ice accretion effects on the lifting characteristics of 3-D multi-element high-lift systems becomes noticeably more complex. Again, it is absolutely essential to determine and account for where stall, or potential separation on the flap at approach conditions, is initiated spanwise on the un-iced baseline configuration. Further, because of the much higher angles of attack encountered with high-lift systems, any installation of wing-mounted engines can have an even larger, often very dominant, impact on stall initiation. In this latter case, it would therefore be quite difficult to provide any very generally applicable conclusions regarding ice accretion effects on maximum lift characteristics for such configurations.

3.3. Control surface effectiveness Ice accretions on wing and tail surfaces ahead of trailing-edge control surfaces such as ailerons, elevators, and rudders, will cause reductions in control surface effectiveness, and can also lead to some undesired anomalies (i.e., reversals, floating, etc.) with aerodynamically balanced control surfaces [64]. The severity of any such losses in basic effectiveness or characteristics is controlled by the following three quite-fundamental fluid dynamic properties:

*

*

*

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Health of the approaching boundary layer at the control surface hinge point on the un-iced baseline. Severity of the required flow turning associated with the required control surface deflection. Degree of degradation of approaching boundary layer or viscous layer health brought about by any ice accretions.

Several factors are important in establishing the health of the boundary layer approaching the control surface. These include the issue of suction versus pressure side of the surface, proximity to stall, existence of leading-edge devices, as well as Reynolds number. With regard to the ability of the flow to withstand the turning (i.e., additional adverse pressure gradient) established by the required control surface deflection, the magnitude of the required deflection is obviously of paramount importance, and Reynolds number is also important, but, in addition, there are some indications that the maximum deflection angle sustainable without flow separation may, in some cases, be reduced by the presence of any appreciable spanwise viscous flow component. And, lastly, the degree of degradation of the approaching boundary layer induced by ice accretions is a strong function of the size, location, shape, and surface roughness characteristics of the ice, as well as being somewhat dependent upon Reynolds number. With the wide range of possibilities in each of the aforementioned lists of influential variables, there is a very broad range of ice accretion effects on control surface effectiveness and characteristics possible. At one end of the spectrum there could be a small reduction in effectiveness caused by an increase in boundary-layer displacement thickness at the LE of the control surface. Such a situation might well occur if the basic lifting surface were operating at a relatively low, non-critical angle of incidence well below stall conditions. It could also occur if the ice accretion was relatively small, not too rough, and located forward where the boundary layer would have a chance to recover somewhat from the disruption. Another possibility would be if the control surface deflection required was relatively low. However, at the other end of the spectrum there could be a large loss in effectiveness or rapid change in hinge moments brought about by a massive flow separation ahead of the control surface. This situation might be expected, for example, if the basic lifting surface (suction side) were close to stall, if the ice accretions were more substantial, rough, and/or located such that the boundary layer was close to separation as it approached the control surface, and/or if a relatively large control surface deflection was commanded. And, in between these two extremes, and with different combinations of these important variables, lie a vast range of possible adverse effects. So, again it is absolutely essential that the prevailing flow physics characteristics be carefully considered when

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assessing the effects of ice accretions on control surface effectiveness. Generalizations are not appropriate, except, perhaps, at both ends of the spectrum. The same caution again applies, even more so if possible, when assessing ice accretion effects on the potential abnormal behavior of aerodynamically balanced control surfaces. This situation is made even more complex by the many variations in system design and logic employed for this type of control surface. 3.4. Drag characteristics Establishing meaningful incremental drag increments due to various forms of ice accretions is likewise dependent upon an appropriate evaluation of many of the same or similar flow physics and testing simulation considerations as required for the preceding aerodynamic characteristics. To start with, flow conditions on the un-iced baseline must be understood and accounted for regarding where boundary-layer transition occurs, as well as any propensity toward boundary layer separation, since two potentially important sources of increased drag due to ice accretion are adverse transition movements and the onset of separation. For singleelement wings and tail surfaces, the airfoil type, thickness ratio, test Reynolds number, and test facility flow quality are all important variables in this determination. Consideration of baseline characteristics is particularly crucial when assessing drag penalties due to relatively small, roughness-like ice accretions where the difference between causing separation or not can have a large impact on the magnitude of the incremental drag penalty. Similarly, test Reynolds number, as well as the size, degree of roughness, and location of the ice accretion, are all logically important variables in determining the incremental drag penalties for such ice accretions. On the other hand, such careful consideration of the baseline flow situation, ice accretion details, and test condition, may well not be as important when defining drag penalties caused by the larger, more irregular ice accretions. However, in this case, distinctions between rime and glaze ice accretions are certainly worth noting. Although the principal source of drag caused by ice accretions is increased profile drag, some increases in skin friction due to transition movement (on small aircraft) and increased roughness, and to liftdependent drag (from ice-induced changes to the spanload distribution) can arise.

4. Effect of initial in-flight leading-edge ice accretions Many observations of the ice accretion process in flight in natural icing conditions, as well as in icing tunnels, have revealed that the initial ice accumulation on the LE of aerodynamic surfaces closely resembles

uniformly distributed roughness in the form of small hemispheres, with a disturbance height nominally equal to the thickness of the ice buildup. This roughness typically extends over the first several percent of the local wing or tail chord on both upper and lower surfaces for airfoil thickness ratios of most interest. The aerodynamic consequences of this initial roughness-like buildup can be important for any aircraft design, but are particularly critical for aircraft employing de-icing systems, since an ice buildup of more than a millimeter is typically necessary before it can be consistently removed. However, it is not unusual to see recommendations that ‘‘observable’’ thicknesses (i.e., 14 –112 in) of ice be allowed to accumulate before operating pneumatic boot airframe ice protection systems. Small buildups must also be addressed for any delayed activation of an anti-icing system, and for unprotected surfaces as well. Test results available since the early 1930s [65] have revealed that disturbances caused by small (k=co20
104 ) surface protuberances in this leading-edge region of single-element lifting surfaces are more critical in terms of causing significant aerodynamic performance and control degradations than similarly sized ones located slightly aft of the LE region. Although we are not aware of any good quantitative data available from full-scale aircraft flight testing for leading-edge roughness effects which would permit an assessment of these effects on configurations that have turbulent attachment line conditions, a wide range of wind-tunnel test results from many experimental studies carried out over the past 50 years or so do exist to permit a pretty good indication of some of the potential effects of such initial ice accretions on important aerodynamic characteristics for both single- and multi-element lifting surfaces. It is of interest to note, however, that most all of these data were obtained using ‘‘sharp particles’’ for roughness rather than the (droplet) hemispheres believed to occur in natural icing conditions. The roughness density employed may also be less than that normally occurring in natural icing conditions as well. Whether or not these differences might be consequential or not relative to other differences which may exist such as the boundary layer state in the leading-edge region has not been established, but it is assumed herein that they are not critical factors considering the wide range of ‘‘roughness effects’’ observed. Reductions in maximum lifting capability and stall angle, as well as drag increases due to such roughnesses, are used primarily herein to illustrate the adverse aerodynamic effects of these initial ice accretions. Some more qualitative assessments of leading-edge roughness effects on trailing-edge control surface effectiveness and/or characteristics are also provided, although there is, unfortunately, very little data available on these subjects upon which to base any such assessments. Considering the large number of general aviation and commuter aircraft flying around

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with de-icing systems, this lack of data is certainly of concern. 4.1. Maximum lift reductions 4.1.1. Single-element lifting surfaces Even though quite a number of wind-tunnel test investigations have been carried out over the past decade or so specifically to help quantify the effects of small roughness-like leading-edge ice accretions on the aerodynamic performance and control characteristics of single-element lifting surfaces, the most extensive set of test results on this subject are those published by Abbott and von Doenhoff [66] over 50 years ago. In this experiment, leading-edge roughness effects were established for over 100 different airfoils at low speed (freestream Mach number o0.2) and at a chord Reynolds number of 6
106. A single nondimensional (k=c) roughness height of about 4.6
104 was used throughout, which would correspond to a 1 mm ice (roughness) buildup on an airfoil/wing/tail chord of just over 2 m. The roughness was applied using carborundum grains, and extended from the LE to 8% chord on both upper and lower surfaces, a fortunately quiteadequate coverage for initial ice accretion modeling. This test program was conducted in the early version of the NASA Langley LTPT. Freestream flow quality was adequate (for these purposes) at that time, although no sidewall boundary layer control was used. This should be kept in mind, but it is not as serious a limitation as it might be if the test were for multi-element geometries. The measured impact of this particular leading-edge roughness application to this large group of airfoils is quite instructive in that it clearly illustrates how the response of a particular airfoil to the addition of leading-edge roughness is strongly influenced by the flow physics controlling the stall process on each airfoil. It can be seen from Fig. 4 that a very wide range of (percentage) reductions (or increases) in maximum lift capability can occur. Roughness actually improved or had very little effect on the maximum lift level attainable by configurations which had a basic (unroughened) maximum lift (coefficient) capability of about 1.0 or less, or a thickness ratio of o8%, i.e., configurations with a very pronounced thin-airfoil-type leading-edge separation. However, for thickness ratios of 9% or higher, or for geometries with a basic maximum lift capability above about 1.1, there is a very definite reduction in the maximum lift capability of anywhere from about 15% to nearly 40%, which, not surprisingly, does not correlate well with simple parameters such as either the basic maximum lift capability or the airfoil thickness ratio (or leading-edge radii). While the foregoing set of data is quite revealing in terms of establishing the broad range of leading-edge roughness effects than can be incurred on 2-D airfoils at

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basically incompressible conditions, and at a constant Reynolds number and (nondimensional) roughness height, it is also important to be aware of the additional variations which may occur as a consequence of changing Reynolds number, varying the nondimensional roughness height, operating at conditions where compressibility effects can modify the results, and/or having 3-D geometries with finite aspect ratio such as swept wings and tails. Fortunately, there are a variety of windtunnel test results available that allow in-depth insights into some of these, such as the effects of varying test Reynolds number and (nondimensional) roughness height. For Reynolds number effects, it is instructive to first look at the (chord) Reynolds number range up to about 5–7 million. That is because most of the wind tunnels in the world are limited to this range. Very few facilities have the capability to produce data at significantly higher Reynolds numbers more representative of flight conditions with large vehicles. In this lower Reynolds number range, there are four good sources of test results [33,51,67,68] which illustrate how significantly the losses in maximum lift capability due to leading-edge roughness can vary for a given 2-D configuration over this range. Clearly, the most comprehensive set of test results addressing this issue are the data from over a half century ago provided by Loftin and Smith [67] where leading-edge roughness effects for 15 of the same airfoils used by Abbott and von Doenhoff [66] were determined at low-speed conditions (Mach number o0.15) at Reynolds numbers ranging from 0.7 to 6 million using the same roughness configuration (i.e., extent) and height (i.e., k=cE4:6
104 ) as used by Abbott and von Doenhoff. For 12 of the 15 airfoils tested, there was a very pronounced increase in the (percentage) maximum lift penalty due to the roughness as the Reynolds number was increased from 1 to 2 million and above. Results for these 12, which had thickness ratios varying from 9% to 15%, are illustrated in Fig. 5 where it can be seen that the average penalty increased from just over 10% at 1 million to 17% at 2 million, and then to over 25% at 6 million Reynolds number. This increase in indicated penalty at the higher Reynolds number is, to a large degree, attributable to an increase in the maximum lift capability of the unroughened airfoils with increasing Reynolds number in this range. However, there were, not surprisingly, also some very noticeable increases in the maximum lift capability with a number of the airfoils in the roughened state as the Reynolds number was increased, although there were also some where there was little Reynolds number effect in this roughened state. The three other sets of 2-D test results which address Reynolds number effects on the loss in maximum lift capability due to LE roughness in the Reynolds number range up to 5–7 million actually cover a wider Reynolds

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Fig. 4. Maximum lift penalty caused by leading-edge roughness [66].

number range, but attention is focused first on this lower range. All three sets of test results were obtained in the NASA Langley LTPT. One was the data provided by Ladson [68] for the 0012 airfoil, while the other two used supercritical (i.e., aft-loaded) airfoils with thickness ratios around 11%. Ref. [68] data for the 0012 airfoil were obtained with the same roughness height (i.e., k=cE4:6
104 ) used previously in Refs. [66,67], but in this case, the roughness only extended back to 5% chord on both the upper and lower surfaces. Test results for this configuration at lowspeed conditions (Mach number=0.15) are available at 2, 4, and 6 million Reynolds

number. Of the other two, one is the data presented by Morgan et al. [33] with roughness (k=cE4:5
104 ) extending to 5% chord on both upper and lower surfaces. In this case, test results are available for 3 and 7 million Reynolds number at 0.2 Mach number. The other set of applicable test results are those presented by Lynch and colleagues [51,52] with roughness extending back to 6% and 7.5% chord on the upper and lower surfaces, respectively. For this case, two different roughness sizes (k=c¼ 0:7
104 and 5.3
104) were investigated, and test results are available at 0.2 Mach number for 2.5 and 5 million

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687

Fig. 5. Reynolds number effects on maximum lift penalty [67].

Fig. 6. Reynolds number effects on maximum lift penalty [33,51,52,68].

Reynolds number. Incidentally, this is the same configuration discussed in the preceding section where a laminar bubble exists at higher angles of attack at 2.5 million Reynolds number on the unroughened airfoil, but not at 5 million. The results from all three of these sources in terms of the percentage loss in maximum lift caused by the leading-edge roughness in this lower Reynolds number range are depicted in Fig. 6. In each

case, there is a very discernible reduction in the indicated penalty at the lowest/lower test Reynolds numbers, with the reduction being most pronounced, interestingly enough, with the smallest roughness size. Clearly, therefore, the preponderance of the data available addressing Reynolds number effects in this lower Reynolds number range on the indicated incremental loss in maximum lift capability due to leading-edge

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Fig. 7. Reynolds number effects on maximum lift penalty [68,69,70].

roughness (i.e., initial ice accretions) show that test results obtained at chord Reynolds numbers much below 5–6 million are most likely not meaningful or useful for higher Reynolds number applications (i.e., X5–6 million). In fact, such test results would most likely be quite misleading in a nonconservative direction. Incidentally, Ref. [67] test results for the 0012 airfoil need to be kept firmly in mind later on in Section 6 when assessing the usefulness of the results from a significant number of low Reynolds number investigations of the incremental effects of larger ice accretions. Having established that test results addressing leading-edge roughness effects on maximum lift capabilities obtained at Reynolds numbers much less than 5–6 million are not appropriate for applications at Reynolds numbers of 5–6 million (and above), the obvious question remaining is just how high is high enough to obtain representative results for the higher Reynolds number applications. There are five sets of test results available (again 2-D) that address this issue, and the indications are definitely mixed. There are the three ‘‘older’’ sets of test results that indicate that 5–6 million is definitely not high enough. These are the Ladson results [68] for the 0012 airfoil, the Loftin and Bursnall results [69] for a 9% thick airfoil (NACA 63-009), and the Abbott and Turner results [70] for a 22% thick airfoil (NACA 63(420)-422). While the same roughness height (k=cE4:6
104 ) was used for the 9% and 12% thick airfoil tests, three different roughness heights ranging from k=cE0:6
104 to 3.1
104 were employed in Ref. [70] investigation with the 22% thick airfoil. The results from these three sources in terms of the percentage loss in maximum lift capabilities are

shown in Fig. 7. For the 9% thick airfoil, the penalty indicated at 6 million Reynolds number is substantially below that incurred at Reynolds numbers of 15–25 million, indicating that 6 million is clearly not adequate for this configuration. Incidentally, in this case, all of the variation with Reynolds number is caused by changes in the unroughened airfoil maximum lift level. There is essentially no variation in the roughened airfoil maximum lift level over the range of test Reynolds numbers. With the 12% thick airfoil, an increase in the penalty is evident as the Reynolds number is raised from 6 to 12 million (which is as far as the data go). However, in this case, ‘‘all’’ of the increase is caused by a (further) reduction in the maximum lift level of the roughened airfoil at the highest Reynolds number. Incidentally, the penalty at 9 million Reynolds number may actually be larger than shown, since it is not clear that the real maximum lift value for the unroughened airfoil was reached. For the 22% thick airfoil, the indicated penalty at 6 million Reynolds number is also substantially below those incurred at the higher Reynolds numbers, but, rather than flattening out (at 15 million) like the penalty did for the 9% thick configuration, there is, in general, a continuing increase in the penalty all the way up to 26 million, clearly indicating that 6 million Reynolds number is not adequate for this configuration either. However, on the other side, there are two sets of test results [33,51] indicating that the penalties due to leading-edge roughness obtained at 5–6 million Reynolds number can be representative for higher Reynolds number applications. Those are the results for the two ‘‘supercritical-type’’ airfoil configurations around 11% thick discussed in the preceding paragraph. The

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689

Fig. 8. Reynolds number effects on maximum lift penalty [33,51,52].

indicated penalties for these configurations at chord Reynolds numbers of 5 million and above are illustrated in Fig. 8. For these cases, it can be seen that there is essentially no change in the maximum lift penalty due to leading-edge roughness as the Reynolds number is increased above 5–6 million. There is also one other piece of evidence that supports these results for airfoils with thickness ratios around 11%. That is the data for the unroughened 12% thick airfoil contained in Ref. [68] where the variations in maximum lift level from 6 to 25 million Reynolds number for this configuration are clearly less than seen with the corresponding 6%, 9%, and 18% thick airfoils. Overall, the trend is definitely for the adverse effects of leading-edge roughness/initial ice accretions on changes in maximum lift capabilities to be reduced at ‘‘lower’’ Reynolds numbers, but defining just how low is too low for meaningful applications of test results to larger aircraft remains somewhat of an open issue. For the purpose of this review, 5 million has been established as the lower limit suitable for ‘‘higher’’ Reynolds number applications in the thickness ratio range of most interest, largely based on the Refs. [51,52] test results which were for a relatively ‘‘modern’’ wing airfoil section. However, it is clear that data obtained at this Reynolds number could still be quite misleading for some geometries in a very non-conservative way. Next, with regard to the effects of variations in nondimensional roughness heights at ‘‘high’’ Reynolds numbers, the known related quantitative data for chord Reynolds numbers of 5 million and above are displayed in Fig. 9 as a function of roughness height, with the data from Refs. [66–69] represented by the range of results observed for airfoil thickness ratios of 9% and above.

Also, shown is the often used ‘‘Brumby curve’’ [55,56] which was originally intended to represent the effects of these small leading-edge roughnesses, but, as can be seen, it is obviously inadequate. From Fig. 9, it can also be seen that the compilation of all the test results obtained subsequent to Refs. [66–69] does indicate the anticipated general reduction in the magnitude of the maximum lift loss with decreasing nondimensional roughness height, but the reduction is reasonably gradual. Therefore, even at the smallest roughness height, which corresponds to about 0.1 mm (0.004 in) with a 2 m chord, the minimum loss observed is about 10%, but the maximum loss does range up to nearly 30%, although most of the results do not exceed 20%. It should be noted, however, that this data ‘‘band’’ is established with a relatively limited set of test results for just a few geometries, and hence may well not be as broad as necessary to cover a comprehensive group of possible or likely aerodynamic designs. As a case in point, the established data ‘‘band’’ only takes in about 60% of the test results which made up the wider range depicted for the results from Refs. [66–69] at that particular roughness height. Also, it is worth noting that the Ref. [71] results, which are at the low end of the ‘‘band’’, were obtained with a much coarser roughness spacing than all the others. Consequently, with the potential for even greater (percentage) losses in maximum lift than indicated, especially at the smaller roughness heights, it would be extremely risky at this time to assume that the consequences of any very small buildups of leading-edge ice (roughness) on singleelement lifting surfaces would be anything less than the 20% to nearly 40% maximum implied by the data band of Fig. 8 unless indicated by very representative,

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Fig. 9. Roughness size effects on maximum lift penalty at high Reynolds numbers.

high quality test results for the specific geometry under consideration. In other words, it should never be assumed that there is any such thing as ‘‘just a little bit’’ of leading-edge ice buildup. A corresponding collection of known test results for leading-edge roughness effects at low Reynolds numbers on lifting surfaces having thickness ratios of 9% and above, and having an unroughened maximum lift coefficent of at least 1.0, is summarized in Fig. 10. Test (chord) Reynolds numbers for this set of data varied from 1 to 3.5 million. In comparing this set of results to the data ‘‘band’’ limits established for the high Reynolds number case, it can be seen that some of the low Reynolds number data do fall within the same ‘‘band’’ or an extension of it. However, consistent with the trends illustrated in Figs. 5 and 6 addressing Reynolds number effects, most of the test results do fall substantially below the high Reynolds number data lower limit. And, the range of results from Ref. [67] at low Reynolds numbers is at a substantially lower level that the range of results seen in Fig. 9 from Refs. [66–69] at higher Reynolds numbers. The results do tend to indicate (not surprisingly) that a greater nondimensional roughness height is required at low Reynolds numbers to offset problems with the un-iced baseline geometry in order to simulate the (incremental) penalty that would be incurred at higher Reynolds numbers. But, the problem is deciding on how much greater is enough in each case. And, anyone thinking that they can determine this a priori is, unfortunately, just kidding themselves. Hence, as indicated previously, relying on low Reynolds

number test results alone to be representative of higher Reynolds number conditions can be extremely risky. Unfortunately, there are really no applicable data available to permit a realistic assessment of either roughness size or Reynolds number influences on the effects of small leading-edge roughnesses on lifting surfaces having thickness ratios o9% and/or having unroughened maximum lift coefficients o1.0. The only data available for configurations in this category, in addition to the Ref. [66] results shown previously in Fig. 4, are the Ref. [69] 2-D results for three 6% thick airfoils at 6 million Reynolds number (with a k=cE4:6
104 ), and the Refs. [77,78] 3-D low Reynolds number (B1.8 million) test results (with a k=cE7:7
104 ) for a swept (441 tailplane geometry having a 9% thick defining airfoil section perpendicular to the quarter-chord line. While the defining airfoil section was 9% thick, the ‘‘effective’’ aerodynamic thickness was clearly less, as the unroughened maximum lift coefficient was below 0.8. In this case, as well as with the three 6% thick airfoils in Ref. [69], adding small LE roughness actually increased the maximum lift capability a little (up to as much as 10%), consistent with the Ref. [66] results for configurations in this category. One possible clue provided by these results regarding likely Reynolds number influences on the effects of small leading-edge roughnesses on some of these thinner geometries are the Ref. [69] unroughened maximum lift characteristics for the three 6% thick airfoils. With each of these, the maximum lift level was essentially unchanged as the Reynolds number was increased from

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691

Fig. 10. Roughness size effects on maximum lift penalty at low Reynolds numbers [76].

3 to 6 to 9 million, but had started to increase by 15 million. By comparison, with the two 9% thick airfoils evaluated, the unroughened maximum lift level started to increase above 6 million Reynolds number (as reflected in the results shown previously in Fig. 7). Hence, it might be anticipated that any significant variation in the maximum lift penalty (or gain) caused by small LE roughnesses with increasing Reynolds number on at least some of these thinner surfaces might well be delayed until above 9 million. However, on the other hand, some very recent test results obtained by Papadakis et al. [80] over a range of Reynolds numbers for an 8% thick full-scale tail with a ‘‘1.6 min ice shape casting’’ (i.e., larger roughness and more irregular shape than the small roughnesses just addressed) provide a quite different perspective. In this case, the ice accretion led to a small (B2.5%) reduction in the maximum lift capability at 1.36 million Reynolds number, but this penalty grew to over 20% at 5.1 million Reynolds number. And, contrary to the Ref. [69] results for the unroughened 6% thick airfoils, the maximum lift level for the clean configuration in this case was increased about 15% as the Reynolds number was raised from 2.2 to 5.1 million. Also, contributing to the large growth in the maximum lift penalty as the Reynolds number was increased was a 6% reduction in the maximum lift level

with the ice shape installed. So, once again, we are reminded of the dangers involved in generalizations, and in trying to use low Reynolds number test results to predict consequences at higher Reynolds number conditions. As we have seen, there are sufficient test data available on leading-edge roughness effects to permit a general assessment of some of the consequences of changing test Reynolds number or varying the nondimensional ice roughness height on single-element airfoils at largely incompressible conditions, particularly for the thicker (X9%) configurations. But, on the other hand, there is only one set of test results [68] that provides some insight as to how operating at conditions where compressibility effects can influence the maximum lift achievable would modify the magnitude of any roughness penalties, and there are no known (useful) data that would permit an understanding of potential differences between all the results obtained for 2-D airfoils versus what might be expected on comparable 3-D swept geometries. With regard to compressibility effects, the test results for the 0012 airfoil provided by Ladson [68] do indicate (as would be anticipated) a substantial reduction in the magnitude of the indicated penalties at a condition where the performance of the unroughened baseline airfoil (i.e., the maximum lift) is compromised/reduced

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by such compressibility effects. In this particular case, because of the significant reduction in the maximum lift capability of the baseline airfoil, the indicated penalty at 0.3 Mach number at 6–9 million Reynolds number was only about 16% compared to around 27% at 0.15 Mach number at the same Reynolds numbers. It should be kept in mind, however, that these results are for an airfoil design which has substantial upper-surface leading-edge suction peaks at stall. Reductions of this magnitude would likely not occur for other geometries with lesser suction peaks (such as NLF, thicker designs, etc.), although those designs would likely not have as large penalties at incompressible conditions. Now, with regard to 2-D versus 3-D, in addition to the absence of data, there is no unanimity in the direction of the trends suggested by various technical rationalizations. One of the complexities involved in the 2-D versus 3-D dilemma is that some run of laminar flow most likely existed on the baseline (unroughened) geometry for all of the test results available, while turbulent attachment line conditions would most likely exist on most lifting surfaces of interest on larger aircraft. With the existing uncertainty regarding direction of trends in both situations, aircraft designers should again exercise caution when using the foregoing test results to predict maximum lift penalties when either of these situations exist. 4.1.2. ‘‘Single-element’’ lifting surfaces with trailing-edge control surfaces Understanding and coping with the adverse effects of initial and inter-cycle leading-edge ice accretions on the maximum lift capability of ‘‘single element’’ wings and

tails with trailing-edge control surfaces deflected is extremely important. Not only are there the loss in lift and possible loss of control caused by any premature stalling to contend with, but, in addition, there can also be a loss of control due to a number of possible control surface anomalies which can occur at these maximum lift (stalling conditions). Loss or reversal of control surface effectiveness is an issue for any control surface concept. But, in addition, aerodynamically balanced control surfaces are also susceptible to floating, oscillations, etc. at these maximum lift conditions with ensuing loss of control a real possibility. Unfortunately, however, there are surprisingly little data available to guide the assessment of maximum lift degradations caused by initial and inter-cycle ice accretions with deflected trailing-edge control surfaces. The only known applicable data available are relatively low Reynolds number test results for four horizontal tail designs with (and without) deflected elevators [71,77,78,81]. And, of these, one of the configurations is the aforementioned 441 swept tailplane design from Refs. [77,78] which has thinairfoil-type stall characteristics, and only one [81] has data for trailing-edge-down (ted) elevator deflections (i.e., opposing the lift) which are needed to address a required pushover maneuver. The results for these four geometries in terms of the percentage loss in maximum lift caused by the initial or inter-cycle leading-edge ice accretions are illustrated in Fig. 11 for the range of elevator deflections tested. With the exception of the one configuration having thin-airfoil-type stall features, the trend is clearly in the direction of having an increased penalty (percentagewise) as the elevator deflection goes

Fig. 11. Variations in maximum lift penalty observed with trailing-edge elevator deflection.

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693

Fig. 12. Roughness size effects on maximum lift penalty for multi-element airfoil geometries at high Reynolds number.

from trailing-edge-up (i.e., aiding lift) towards tailingedge-down (i.e., opposing the lift). Incidentally, for the two sets of test results reported on in Refs. [77,78], it was acknowledged in Ref. [78] that ‘‘only upward elevator deflections (negative de ) were tested’’, and that ‘‘later it was realized that the most critical cases are for aircraft that need elevator down (fixed tailplane) for trim at flap extension’’. Interestingly, quite a range of roughness characterizations are represented in the results shown in Fig. 11. In Ref. [71], roughness extended back to 20% chord on both upper and lower surfaces, whereas in Refs. [77,78], the roughness was more localized to the ‘‘leading edge’’. On the other hand, the results from Ref. [81] are for a representative inter-cycle ice accretion remaining on the tail between pneumatic de-ice boot operation after 15 min with the boot having been cycled every 3 min up to 12 min. With regard to the test results summarized in Ref. [81], some data were obtained at a higher Reynolds number (i.e., a little above 4 million), but the results are not complete enough to draw any very quantitative conclusions regarding Reynolds number influences on the maximum lift characteristics with elevators deflected. Clearly, this is another area where more studies are warranted. 4.1.3. Multi-element lifting surfaces Again, in contrast to the relatively large database which exists for single-element geometries, much less data is available to help quantify the effects of initial and inter-cycle leading-edge ice (roughness) accretions on the aerodynamic characteristics of the much more complex (mechanical and aerodynamic) multi-element lifting surfaces routinely employed during takeoff and landing. In fact, there are only two somewhat limited sets of relevant high Reynolds number test results [51,82], and two limited sets of low Reynolds number results [74,83]. But, the little that is available, used in conjunction with lessons learned with single-element geometries (i.e., there

is no single answer), does provide some very instructional insights. Starting with the high Reynolds number results, both are from 2-D tests conducted in the NASA Langley LTPT utilizing an ‘‘effective’’ sidewall-boundary-layer control system [53]. In the first [51,52], a four-element (slat, main, and two-segment flap) geometry with leading- and trailing-edge device deflections appropriate for approach/landing conditions was employed, and the primary roughness coverage applied to the slat was based on a water droplet impingement analysis at approach conditions. The effects of extending the roughness to the entire upper surface of the slat were also investigated. The second test [82], which utilized the same basic model as the first, was conducted with a three-element geometry (slat, main, and single-segment flap) representative of a takeoff configuration. Although this particular test was primarily carried out to address the impact of ground frost accumulations, one ‘‘frost’’ configuration evaluated had just the entire upper surface of the slat covered. And, since results from the first test indicated that there was no discernible difference in the effect of roughness on maximum lift capability with this roughness coverage relative to the primary uppersurface coverage on the slat, the results from this second test are included in this particular assessment. Both sets of test results are shown in Fig. 12. By comparing the Refs. [51,52] results for the four-element geometry with the corresponding Refs. [51,52] single-segment airfoil results shown previously in Fig. 9, it can be seen that the percentage losses in maximum lift capability experienced for this approach/landing geometry (of from 7% to 11%) are about one third of those incurred with the single-element airfoil. This ratio just happens to be the inverse of the ratio of the (unroughned) maximum lift capabilities. Interestingly, this same proportional relationship holds with the Ref. [82] results for the takeoff geometry. What this relationship signifies is that the

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Fig. 13. Roughness size effects on maximum lift penalty for multi-element airfoil geometries at low Reynolds number.

absolute loss in maximum lift (coefficient) capability with these particular related geometries is about the same for all of the particular rigging variations tested at a given nondimensional roughness height. It will also be subsequently shown that the (absolute) reduction in angle-of-attack margin to stall is about the same in each case as well. Just based on this small sample of high Reynolds number test results, the following important inferences can be drawn: *

*

Small leading-edge ice accretions would, in general, be more serious at takeoff conditions (compared to approach/landing), since the percentage loss in maximum lift capability would be greater at the same time that the normal operating speed margins (percentagewise) to stall are smaller. Configurations with ‘‘hard’’ LEs (i.e., no slat or Krueger) would, in general, be more vulnerable than configurations having leading-edge devices extended at both takeoff and approach/landing conditions, since the percentage loss in maximum lift would most likely be greater.

However, having seen the large band of results possible with a variety of single-element lifting surfaces, and recognizing the likelihood of different limiting flow physics situations with other multi-element configurations, it would definitely be expected that the magnitude of percentage losses in maximum lift capability for other multi-element configurations and applications could well be noticeably different (higher or lower) than those shown in Fig. 12. And, potential differences between these 2-D test results and real world 3-D designs must

not be forgotten. So, again, aircraft designers need to be extremely cautious when estimating potential penalties for small leading-edge ice accretions on multi-element designs. Justification for this caution is underscored after examining the results obtained from the two (2-D) low Reynolds number investigations of leading-edge roughness effects on multi-element airfoil geometries that are illustrated in Fig. 13. Each of the investigations explored these roughness effects on configurations with singlesegment flaps, and both with and without leading-edge slats. One of the studies [83] addressed takeoff geometries, while the other [74] examined an approach/landing geometry. The test Reynolds number range for these tests was 2.6–2.7 million, and a sidewall boundary layer control system was used with each. From the low Reynolds number results shown in Fig. 13, together with the related high Reynolds number results, it can be seen that the leading-edge roughness penalty for the Ref. [74] approach/landing geometry with slats is not very different from the high Reynolds number approach/landing configuration results, and, not surprisingly, the maximum lift penalty due to roughness for the approach/landing geometry without a LE device was almost double the E10% penalty experienced when a slat was used. However, on the other hand, the low Reynolds number test results for the takeoff geometry with slats extended do indicate maximum lift reductions more than twice that expected based on the high Reynolds number results for the takeoff geometry. While some scatter band is certainly expected, deviations this large are alarming, especially since the trend seen with single-element airfoils was that low Reynolds

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number tests typically produced lower maximum lift penalties at a given nondimensional roughness height. With regard to maximum lift penalties with takeoff configurations not employing leading-edge devices, the Ref. [82] low Reynolds number results do show the sizable increase expected. Overall, it remains somewhat of a question as to whether these low Reynolds number test results showing quite high maximum lift penalties due to small amounts of LE roughness on multi-element takeoff-type geometries are meaningful and applicable. While the magnitude of the indicated penalties are clearly larger than what might be anticipated based on the compilation of measured penalties for single-element geometries, and they cannot be reconciled with available high Reynolds number results for takeoff geometries, it would be very risky to assume that these results were not possibly representative for some takeoff-type geometries, because the magnitude of the maximum lift reductions are equivalent to eliminating all of the typical minimum takeoff speed margins (to stall). Hence, these results should certainly provide added incentive to strictly enforce the ‘‘keep the wing (leading edge) clean’’ mandate, especially during takeoff. 4.2. Stall angle reductions Reductions in the angle-of-attack at which stall occurs caused by initial leading-edge ice accretions on wings are of major significance relative to the proper operation of stall warning systems. Such reductions can either reduce

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the margin from stall warning to actual stall to less than desired and needed, or, worse yet, stall could occur before the flight crew received any warning of an impending stall.

4.2.1. Single-element lifting surfaces As with the review of maximum lift reductions, the most comprehensive set of test results on airfoil stall angle reductions due to leading-edge roughness are derived from the results published by Abbott and von Doenhoff [66]. The compilation of the measured reductions (or increases) in stall angle from this data set are shown in Fig. 14 as a function of airfoil thickness ratio, and vary from a maximum reduction of just over 61, to actual increases in stall angle of attack of slightly over 31. The increases in stall angle at the lowest thickness ratio (6%) are consistent with the improvements in maximum lift capability observed for these designs. However, in the more commonly used thickness ratio range from 9% up to perhaps as much as 15%, aside from a very few exceptions, most of the configurations tested incurred a sizable reduction in stall angle, ranging from about 11 up to as much as 5–61. At airfoil thickness ratios above 15%, an even wider range of results are seen. The overall broad data band observed again serves as a reminder as to how the response of a particular lifting surface to the addition of leading-edge roughness is governed by the flow physics controlling the stall process, and how this flow physics is altered by the roughness. Clearly, many variations exist.

Fig. 14. Stall angle reductions caused by leading-edge roughness [66].

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Fig. 15. Roughness size effects on stall angle reductions at high Reynolds numbers.

Consistent with the changes observed in maximum lift penalties with variations in nondimensional roughness height, the stall angle reductions similarly change. But, there is also another variable that comes into play in establishing the magnitude of stall angle reductions for practical aircraft wings and tails, and that is the lift curve slope, which, in turn, is dependent upon the wing or tail aspect ratio. The influence of both of these variables can be seen in the compilation of high Reynolds number test measurements of stall angle reductions shown in Fig. 15 for geometries having thickness ratios ranging from 9% to 15%. First, there is the variation with roughness height for 2-D geometries other than Refs. [66–69]. While the variation is substantial, it is less than the large variations possible with just different airfoils. Then, with regard to planform effects, it can be seen that the stall angle reductions observed in the test of the 3-D tail configuration are noticeably higher than the upper limit of the Refs. [66–69] results. It is important to note (see Fig. 9) that this was not the case with the measured maximum lift reductions. Hence, while the 2-D airfoil test results are quite useful in showing the existence of a wide range of possible reductions in stall angle due to small leading-edge ice accretions, the magnitudes cannot be directly used for real aircraft applications without first making appropriate adjustments for planform effects. The corresponding collection of available and useable low Reynolds number test results addressing stall angle

reductions on single-element lifting surfaces caused by leading-edge roughnesses similar to initial and intercycle ice accretions are illustrated in Fig. 16. As can be seen, this particular collection of test results is somewhat of a ‘‘mixed bag’’. On one hand, many of the 2-D low Reynolds number test results do clearly fall below the 2-D high Reynolds number data ‘‘band’’ as would be anticipated as a consequence of the generally lower maximum lift penalties at these low Reynolds numbers. However, on the other hand, there are also a few of the 2-D low Reynolds number test results which fall above this ‘‘band’’. And, while the test results for 3-D finite aspect ratio configurations are by and large at the upper limits of the low Reynolds number data, the expected 3-D effect (i.e., the consequences of reduced lift curve slopes) is not as pronounced as seen at the higher Reynolds numbers (in Fig. 15). All together, this particular collection of low Reynolds number test results are further evidence yet that any conclusions based on low Reynolds number test results are most likely not really very meaningful or useful for higher Reynolds number applications. 4.2.2. Multi-element lifting surfaces Very little test data exist at either high or low Reynolds numbers to document stall angle reductions caused by leading-edge roughness on the leading-edge element of multi-element high-lift geometries, and the

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Fig. 16. Roughness size effects on stall angle reductions at low Reynolds numbers.

Fig. 17. Roughness size effects on stall angle reductions for multi-element airfoil geometries.

little that there is was all obtained from 2-D testing. These available results are presented in Fig. 17 where they can be compared with the (high Reynolds number) data bands established for single-element airfoils. It can be seen that stall angle reductions observed for the multi-element geometries, particularly at high Reynolds numbers, tend to be somewhat larger than the corresponding single-element results for the same basic

(stowed) airfoil design. This trend would seem to be logical since higher leading-edge velocities would be experienced on a leading-edge device at maximum lift conditions than exist on a single-element design [61]. It would then seem to follow that the range of possible reductions in angle-of-attack margin to stall that might be incurred with multi-element airfoil high-lift designs could be as large as that observed with single-element

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designs. And, finally, it should also be remembered that since all of the multi-element airfoil data was obtained from 2-D testing, any absolute values extracted from this limited database must be adjusted for planform (and other configuration) effects before trying to apply them to practical 3-D aircraft geometries. 4.3. Drag penalties Increases in drag caused by initial leading-edge ice accretions on wings and tails are of concern because they can lead to potentially dangerous reductions in aircraft climb and acceleration rates, aircraft range, and speed. For example, accidents and incidents have occurred when flight crews have used autopilots not augmented by an autothrottle during exposure to icing, thereby allowing the airspeed to bleed down to stall entry, often without warning. In addition, knowledge of the potential magnitude of such drag increases will be critical in determining whether aircraft performance monitoring, as proposed in some quarters [84], can be a viable and practical means of alerting flight crews to the presence of these initial ice accretions. However, as will become evident, there is a real scarcity of high-quality test data available to enable a real credible documentation of these drag increases since some of the (few) high Reynolds number test programs investigating these leading-edge roughness effects were focused strictly on lifting characteristics. 4.3.1. Single-element lifting surfaces As indicated earlier, many factors are involved in determining the magnitude of drag penalties brought

about by the addition of small amounts of leading-edge ice (roughness). In addition to the roughness drag itself, adverse (i.e., forward) movement of the transition point will contribute in cases where turbulent attachment line conditions do not exist, as will excessive thickening of the boundary layer or even separation onset over the aft part of the lifting surface. And, of course, many variations and combinations of these and other factors can exist, resulting in a sizable range of possible penalties, dependent upon the lifting surface geometry, the prevailing flow conditions, test Reynolds number, etc. Once again, the most comprehensive set of test results available to demonstrate the wide range of drag penalties possible is the early data provided by Abbott and von Doenhoff [66]. However, when assessing these results, it is important to keep in mind that the test Reynolds number is (only) 6
106, and these are tests of 2-D airfoils. At this test Reynolds number (and tunnel flow quality), transition movements due to the roughness are certainly involved. And, in considering 2-D data, only changes to the basic 2-D parasite drag of the lifting surface are involved. There are none of the normal 3-D effects associated with realistic aircraft lifting surfaces included. In analyzing drag penalties from this and subsequent data sets, two conditions are used. The first is the percentage increase in the minimum parasite drag level, while the second is the percentage increase in drag at a (higher) lift coefficient representative of flight at a speed 30% above the 1-g stall speed of the unroughened test geometry. The compilation of percentage increases in minimum parasite drag levels from the Abbott and von Doenhoff data set [66] is shown in Fig. 18, where it

Fig. 18. Minimum parasite drag increase due to leading-edge roughness [66].

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Fig. 19. Drag increase at 1:3 VS;1g due to leading-edge roughness [66].

can be seen that at this one particular test Reynolds number and roughness height, increases in minimum parasite drag level vary anywhere from about 50% up to nearly 200%. The analogous drag penalties at the higher lift-coefficient conditions, which are much more representative of realistic low speed flight conditions where drag penalties due to ice accretion could be consequential, are displayed in Fig. 19. At this condition, the range of drag penalties due to leading-edge roughness tend to fall into a somewhat narrower (although still large) band, varying from a little over 50% up to about 120% for airfoil thickness ratios of 15% and less. In addition to the relatively large and not-readily correlated variations observed in the airfoil profile drag penalties caused by small (i.e., k=cE4:6
104 ) initialice-accretion-like leading-edge roughness for quite a variety of different airfoil geometries at 6 million Reynolds number, there are, for the same small roughness, further significant, quite diverse, and also notreadily correlated variations in these penalties as the Reynolds number is both decreased and increased. The largest source of data addressing these additional variations are the Ref. [67] test results for 15 of the airfoils used in Ref. [66] where these drag penalties were determined at Reynolds numbers ranging from 0.7 to 6 million. Incidentally, of the 15 airfoils tested, other than one 9% and one 18% thick designs, the thickness ratios were either 12% or 15%. The measured increases in the minimum parasite drag and the drag at the higher lift condition are provided in Figs. 20 and 21, respectively, for chord Reynolds numbers from 1 to 6 million. There are some interesting differences between the trends seen

at the two conditions. Looking first at the minimum parasite drag penalties shown in Fig. 20, the range of measured penalties for these 15 airfoil designs clearly grows as the test Reynolds number is increased. And, it can also be seen that although there are quite a few configurations for which there is very little change in the measured penalty between 2 and 6 million Reynolds number, there are a couple where the penalty increases noticeably as the Reynolds numbers is increased in this range, and there are also some where the drag penalty noticeably decreases. On the other hand, as can be seen in Fig. 21, the range of drag penalties at the higher lift condition clearly decreases with increasing Reynolds number. And, also in contrast to the generally more subdued changes seen between 2 and 6 million Reynolds number with the minimum parasite drag penalties, there are quite substantial variations in the drag penalties at these condition for all except one of the test configurations as the Reynolds number is increased from 2 to 6 million. While the prevalent trend at this higher lift conditions is for a substantial reduction in the drag penalty as the Reynolds number is increased from 2 to 6 million, there were, however, four configurations, including the 9% thick airfoil, where there was a very noticeable increase in the drag penalty over this Reynolds number range. What these results clearly indicate is that test results obtained at chord Reynolds numbers much below 5–6 million are likely to be very misleading in terms of indicating likely results for higher Reynolds number applications (X5–6 million). The question remaining to be answered is whether or not drag penalties caused by small leading-edge

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Fig. 20. Reynolds number effect on minimum parasite drag increase [67].

Fig. 21. Reynolds number effect on drag increase at 1:3 VS;1g [67].

roughness (i.e., initial ice accretions) at a test Reynolds number of 5–6 million are appropriate for inferring the drag penalties that would exist at even higher Reynolds numbers. While there is only a very limited amount of data available which addresses this issue, there are three sets of test results which are quite revealing in this

regard. To start with, there are the results presented by Ladson [68] for the 0012 airfoil, again using the same roughness height (i.e., k=cE4:6
104 ) employed in Refs. [66] and [67], with a test Reynolds number range from 2–12 million. Next are the results presented by Loftin and Bursnall [69] for a 9% thick airfoil, again

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having the same roughness height (i.e., k=cE4:6
104 ) employed in [66] and [67], and where the test Reynolds number was varied from 6 to 25 million. The third set of useful results are those presented by Abbott and Turner [70] for a 22% thick airfoil with three different small roughness heights (k=cE0:6; 1.1, and 3.1
104) where the test Reynolds number was varied from 6 to 26 million. Although this latter set of test results for a 22% thick airfoil are not likely directly applicable to modern aircraft (wings and/or tail) designs, the results do provide an interesting link to the Ref. [67] results for the 18% thick airfoil. And, even though the test results obtained for the 9% thick airfoil [69] are also likely not directly applicable to most modern wing designs, they are most pertinent for tailplane icing concerns. These three sets of test results, both in terms of the increases in minimum parasite drag caused by the leading-edge roughness applications, as well as the corresponding drag penalties at the higher lift condition, are presented in Figs. 22 and 23, respectively. It can be seen in Fig. 22 that the minimum parasite drag penalties for both the 0012 and 22% thick airfoils appear to have essentially leveled out at 6 million Reynolds number, but that is not the case for the 9% thick airfoil where the minimum parasite drag penalty is reduced by more than half as the Reynolds number is increased from 6 to 25 million. This latter result was a bit unsuspected considering that the Ref. [67] results for the 9% thick airfoil (see Fig. 20) had indicated very little change in the penalty as the Reynolds number was increased from 2 to 6 million. Lastly, it is interesting to note that the drag penalty from Ref. [68] for the 0012 airfoil actually falls below the Ref. [66] data band at 6 million Reynolds number.

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One possible explanation for this would be the reduced extent of roughness coverage (i.e., 5% versus 8% chord). Next, in examining the results for the higher lift conditions as shown in Fig. 23, it can be seen that there are some substantial and contrasting changes in the indicated penalties due to leading-edge roughness for these three airfoils as the Reynolds number is increased above 6 million. Clearly standing out is the doubling of the drag penalty for the 9% thick design at 15 million Reynolds number. Curiously, this increased drag penalty has the appearance of being an extension of the trend seen for the 9% thick airfoil in the Ref. [67] results from 2 to 6 million Reynolds number (see Fig. 21). Similarly, the overall reduction in the drag penalty incurred with the 22% thick airfoil as the Reynolds number is increased from 6 to 26 million also has the appearance of being an extension of the trend seen from Ref. [67] for the 18% thick design. However, on the other hand, with the 12% thick airfoil, it can be seen that the trend above 6 million Reynolds number is obviously not an extension of the lower Reynolds number test results, Hence, once again, the often-nonpredictability of these Reynolds number trends is clearly illustrated. Consequently, until such time as additional high Reynolds number data addressing these leadingedge roughness effects might become available, what these three sets of test results, together with those from Ref. [67], clearly indicate is that any use of test results for assessing drag penalties caused by small initial ice accretions obtained at Reynolds numbers much less than full-scale flight values could very likely lead to erroneous conclusions regarding flight characteristics.

Fig. 22. Reynolds number effects on minimum parasite drag penalty.

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Fig. 23. Reynolds number effects on drag increase at 1:3 VS;1g :

While it would seem to be quite obvious that the drag penalties caused by leading-edge roughness would, in almost all cases of interest, increase with increasing roughness height, there is, unfortunately, clearly insufficient data available to make any kind of a meaningful quantitative assessment of the magnitude of these variations for a range of airfoil geometries. The only known source of data directly addressing this issue are the Ref. [70] test results for three 22% airfoil designs where the roughness height (k=c) was varied from about 0.6
104 to 3.1
104. Test results were obtained at 26 million Reynolds number for all three airfoils, and the Reynolds number was varied from 6 to 26 million for one of them, the one just referred to in addressing Reynolds number effects. The test results (i.e., drag penalties) for this configuration with the three roughness heights were shown in Figs. 22 and 23 where it can be seen that there are only small increases in the minimum parasite drag penalty over this limited range of (small) roughness sizes, but the roughness size effects become noticeably more pronounced at the higher lift condition, even with the quite-small roughness heights utilized in this test program. Incidentally, these same trends were also evident in the results for the other two airfoils at 26 million Reynolds number. As indicated, the roughness sizes utilized in this investigation were quite small, so there is no knowing what would have happened at larger (nondimensional) roughness sizes of more practical interest, although the trend started at the higher lift conditions is certainly ominous, at least for this airfoil thickness. There are four other sets of test results

available [72–74] for airfoils with thickness ratios ranging from 12% to 15% where drag penalties associated with larger (non-dimensional) roughness sizes, varying from about 8
104 to over 50
104, were assessed. Unfortunately, these are all low Reynolds number (2–3.3 million) results, and only a single roughness height was used for each. What these results indicate are drag penalties near the lower limit of the Ref. [67] band of results (for a k=cE4:6
104 and Reynolds number of 6 million). In fact, the result obtained with a 0012 airfoil [73] with the largest roughness size (i.e., k=c > 50
104 ) actually falls below the Ref. [67] band at the higher lift condition, once again illustrating the danger of using low Reynolds number test results for higher Reynolds number applications. 4.3.2. Multi-element lifting surfaces Fundamentally, it would be expected that the profile drag penalties (percentagewise) caused by the existence of leading-edge roughness on the leading-edge element of multi-element high-lift geometries would be much less than for single-element geometries. First, the basic drag levels of the multi-element geometries are much higher because of the coves, flowing slots, etc. And, secondly, the adverse effect of the roughness on the forward element would likely be diminished somewhat by both the presence of the downstream element. (i.e., pumping up the forward element, and by having only a small impact on the profile drag of the downstream elements) Regrettably, however, there appears to be only one

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known published source of test data which provides any insight into the absolute magnitude of these drag penalties on multi-element geometries, and, those results somewhat unfortunately, are from a low Reynolds number (2.6 million) investigation [74]. Be that as it may, however, these low Reynolds number test results do indicate that the percentage increases in drag for multi-element (two- and three-element) geometries at both minimum parasite drag conditions, and at speeds 30% above the 1-g stall speed of the unroughened airfoil, are less than the single-element penalties by factors of anywhere from 4.5 to at least 6.5, with the reductions being greatest for the three-element geometries. While these reductions are of the general magnitude anticipated, it is impossible at this time to accurately quantify what realistic increases in profile drag might be like at high Reynolds number conditions. 4.4. Trailing-edge control surface characteristics Changes in control surface effectiveness and characteristics caused by the roughnesses associated with initial and inter-cycle ice accretions which can possibly lead to aircraft control difficulties clearly need to be understood and accounted for. Regrettably, however, there is very little data available to help enable this understanding. The only known data presently available are some elevator hinge moment measurements made for three previously described (in relation to Fig. 11) horizontal tail designs at relatively low Reynolds numbers [77,78,81]. And, as previously indicated, only one of these data sets [81] has results for the critical trailing-edge-down elevator deflections used for pushover maneuvers. Be that as it may, these results are helpful in providing some good insights into possible and/or likely control surface anomalies that might be caused by small leading-edge roughnesses. However, they also reinforce the concern that there are, in reality, a wide range of possible effects depending on lifting surface and control surface geometries and type, the type of initial or inter-cycle ice accretion involved, Reynolds number, etc. Before addressing the lessons learned from these test results, it is useful to define the sign convention being used. Herein, the rather standard convention adopted is that trailing-edge-down elevator deflections (i.e., opposing the download on the tail) are defined as being positive, as are elevator hinge moments which tend to rotate the elevator trailing-edge down. Conversely, trailing-edge-up elevator deflections (augmenting a download) are defined as being negative, the same as hinge moments which tend to rotate the elevator trailing-edge up. Although this sign convention is, as indicated, rather conventional, it is not universally used as evidenced by the use of the opposite sign convention for elevator hinge moments in Refs. [77,78].

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After examining the three sets of available and applicable test results addressing leading-edge roughness effects on elevator hinge moment characteristics, there is one inescapable conclusion. In the case of these three geometries at a relatively low Reynolds number, it is very clear that any really noteworthy changes are typically tied very directly to noteworthy changes in the tail stall characteristics brought about by these small initial and/or inter-cycle ice accretions. And, the converse is true as well, i.e., if the leading-edge roughness has little effect on the tail stall characteristics (including the maximum lift level), then the elevator hinge moments are impacted very little as well. This latter situation is illustrated very well by the Refs. [77,78] test results for the 441 swept tail design which, at the test Reynolds number, had thin-airfoil-type stall characteristics in the unroughened state. In this case, the addition of small (i.e. k=cE7:7
104 ) roughness to the LE had, as seen previously in Fig. 11, almost no effect on the maximum lift level (and stall characteristics). Consistent with this, the elevator hinge moments changed very little as well. Another lesson learned from these same test results is that if the stall characteristics are quite mild, i.e., no big drop-off in lift at the stall, then the elevator hinge moments did not change dramatically either at the stall. In this case, with elevator deflections of 01 and 201, the hinge moments became more and more positive as the tail incidence/download increased prior to stall, but then just kind of flattened out at and beyond the ‘‘stall’’. A quite different set of stall characteristics and resulting elevator hinge moment characteristics occurred with the other two tailplane geometries examined, which are both nominally 12% thick. These configurations are the ‘‘thickened’’ 181 tailplane design from Refs. [77,78], which was tested with and without the same small ‘‘leading-edge’’ roughness (i.e., k=cE7:7
104 ) used with the 441 swept tailplane, and the Twin Otter tail geometry from Ref. [81], which was tested with and without the somewhat larger and rougher inter-cycle ice accretion. In both cases, and at all the elevator deflections evaluated (i.e., 01, 101, and 201 for the 181 tail, and 01, 101, 201, +101, and +14.21 for the Twin Otter tail), the elevator hinge moments without the leading-edge contamination became gradually more positive (or less negative) as the tailplane incidence/ download was increased prior to stall (but not as steep as with the 441 tail). Then, however, concurrent with the stall, the rate of increase (toward more positive) becomes noticeably greater. In fact, the increase is quite abrupt with the Refs. [77,78] ‘‘thickened’’ 181 tailplane. This abrupt change is consistent with the very abrupt/sharp lift loss experienced by this configuration at stall. On the other hand, the much less abrupt increase seen with the Twin Otter elevator hinge moments goes with the much tamer stall characteristics exhibited by this

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configuration. Now, when the leading-edge contamination is added to both of these configurations, the resulting elevator hinge moment characteristics look much like those for the respective unroughened geometries except that the significant increase in hinge moments (in the positive direction) occurs at a lower tail incidence/download as a consequence of the maximum lift penalty (i.e., lower stall angle) associated with the addition of the leading-edge roughness, etc. Incidentally, this premature stall does result in an earlier change in the sign of the elevator hinge moments from negative to positive (i.e., control surface reversal) with positive elevator deflections. Although these three sets of tailplane test results have provided some good insights into some possible types and degrees of changes to elevator hinge moment characteristics that can be caused by initial, residual and inter-cycle leading-edge ice accretions, two fundamental concerns still remain. One is that all three of the experimental studies were carried out at Reynolds numbers around 2 million, and we have previously seen many instances where results obtained at this relatively low Reynolds number are not representative for many, more practical, higher Reynolds number applications. Second, these three geometries only represent a small sample of existing or likely tailplane (or wing) designs, and hence other types and degrees of changes are likely. As a case in point, the aforementioned very recent test results obtained by Papadakis et al. [80] for an 8% thick full-scale horizontal tail do show a different combination of characteristics. In the clean (unroughened) state, some of the results look much like a thin lifting surface design (like the 441 swept tailplane), with a maximum lift coefficient below 1.0, and quite tame stall characteristics. However, the elevator hinge moment characteristics at 5.1 million Reynolds number do not look like the Refs. [77,78] 441 tailplane results at all. In fact, for elevator deflection ranging from 151 to +151, the hinge moment variations more closely resemble something between the two nominally 12% thick designs just addressed, with a very pronounced change in trend toward being more positive (or less negative). However, in this case, this change occurs a few degrees before achieving maximum lift conditions rather than concurrent with it as was the case with the two 12% designs. But, consistent with the test results for the two 12% thick designs, when the 1.6 min ice shape casting was added, this change in the slope of the elevator hinge moments occurred at a 3–41 lower incidence angle than occurred in the clean state, although it is not clear that the ‘‘stall’’ angle was actually reduced that much. So, one more time there is additional evidence available to illustrate the dangers involved in generalizations, and in trying to use low Reynolds number test results to predict consequences at higher Reynolds number conditions.

5. Effect of runback and ‘‘ridge’’ ice accretions This category of ice accretions is known to occur in either of two situations. First, typical runback ice accretions occur when all of the impinging water droplets are not evaporated by the heat produced by a leading-edge anti-icing system, and the water which is not evaporated runs back and subsequently freezes on the cold, unheated surface just aft of the anti-icing system coverage. This can occur either by design (to minimize engine extraction losses) or as a consequence of a very intense icing encounter. A wide variety of runback-type ice accretion shapes, sizes, and roughnesses have been observed. In some cases, the water which runs back freezes in the form of relatively smooth rivulets, which are probably not overly detrimental to the aerodynamic characteristics. On the other hand, runback-type ice accretions can result in a spanwise forward facing step (i.e., ridge) just aft of the ice protection system coverage. These would be expected to cause significant degradations in the aerodynamic characteristics, depending, of course, on their size, etc. Accretions similar to those can also occur during icing encounters involving larger-than-normal droplet sizes which have more aft impingement limits. The possibility of these ‘‘ridge’’ ice accretions occurring due to droplet impingement aft of a de-icing boot (or anti-icing system) has long been recognized. Johnson in 1940 [85] pointed out that ‘‘in certain cases, ice has built up on the de-icer cap strips to such an extent (that) no benefit has been derived from the operation of the de-icing boot’’. In fact, albeit at low Reynolds number (1.6 million), his wind-tunnel test data indicated greater maximum lift losses and drag increases for these ‘‘ridge’’ ice accretions than occurred with the full leading-edge ice shape (which, unfortunately, was not described). These results were a cause for real concern in those days as most de-icing boots at that time only extended back to about 5% chord. Unfortunately, however, very little attention was given to this potentially very serious problem area between then and the time of the October 1994 American Eagle ATR-72 accident [1,2] where such accretions occurring in supercooled large droplet (SLD) conditions are thought to have been the primary cause. The only known efforts (somewhat) addressing this topic in this time frame were the icing tunnel studies by Bowden in the 1950s [86], the flight testing conducted in various environmental conditions with the University of Wyoming Beachcraft Super King Air twin-turbo prop aircraft from 1978 to 1994 where some most noteworthy observations/results from a number of large droplet icing encounters were reported [87–90], and the very limited wind-tunnel test results reported by Welte et al. [91] in 1991. Sadly, in the aftermath of this crash, more than half century after Johnson [85] published his results, this lesson was

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repeated. Ashenden et al. [92], in their low Reynolds number wind tunnel tests of an airfoil with simulated ice shapes, came to the same conclusion, i.e., that a freezing drizzle ice shape consistent with de-icing boot operation (i.e., ridge ice at 6.3% chord) resulted in a more severe performance degradation than occurred without ‘‘de-icing boot operation’’. Unfortunately, in spite of the potentially serious consequences of such ‘‘ridge’’ ice accretions, there is a real dearth of good quantitative data available for these runback and ridge ice accretions. To start with, there does not appear to be any really applicable quantitative high Reynolds number data presently available for single-element lifting surfaces, but, worse yet, there does not appear to be any data at any Reynolds number for this type of ice accretion on multi-element high-lift geometries (for approach/landing). There are some low Reynolds number test results available for singleelement lifting surfaces, but there are concerns that for these runback and ridge ice accretions somewhat aft of the LE, it would not be unrealistic to expect that there could be some significant differences in the consequences between having laminar or ‘‘transitional’’ flow ahead of these ridge ice shapes at low Reynolds number test conditions versus having fully developed turbulent flow (i.e., turbulent attachment line conditions) at high Reynolds number flight conditions. Hence, the lack of any good quantitative data at high Reynolds number (flight) conditions for these runback and ridge ice accretions represents a serious deficiency in the presently available database. This concern will become quite apparent when some of the (large) magnitudes of the aerodynamic penalties incurred with simulated ridge ice shapes at low Reynolds numbers are seen. 5.1. Maximum lift reductions 5.1.1. Single-element lifting surfaces There are several sets of low Reynolds number test results available that provide some interesting insights into the possibly quite-large aerodynamic consequences of these spanwise runback and/or ridge ice accretions. One of these is the previously mentioned data published by Jacobs in 1932 [65], and the others are results obtained by Lee and Bragg [93–95], as well as Papadakis and Gile-Loflin [79], in the aftermath of the ATR-72 accident. Jacobs’ results were obtained with small protuberances mounted at various chordwise locations on a 0012 airfoil at 3.1 million Reynolds number, whereas the Refs. [93–95] results were obtained at 1.8 million Reynolds number with forward-facing quarterround shapes mounted at various chordwise locations on four different airfoils with thickness ratios ranging from 12% to 15%. One of these four airfoils was the 23012 configuration, selected ostensibly to be representative of (turboprop) commuter type aircraft (such as the

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ATR-72, etc.). The other three are modified versions of the 63A415 and 63A212 airfoils, and the NLF414 configuration. The Ref. [79] results were obtained with four different protuberances at 2 million Reynolds number using the same modified 63A212 airfoil design (i.e., Twin Otter tail) investigated in Refs. [95]. There is also a limited amount of data obtained at 1 million Reynolds number. Jacobs’ results [65] were obtained with protuberance heights (k=c) ranging from 4
104 to 125
104. The Refs. [93,94] results for the 23012 airfoil are for protuberance heights of 56
104, 83
104, and 139
104 but they also provided results with a ‘‘small’’ roughness strip (k=c¼ 14
104 ) for comparative purposes. Results for the other three airfoils used by Lee and Bragg [93,95] were obtained at a single k=c¼ 139
104 : Incidentally, for the Refs. [93,95] results with simulated ridge ice accretions aft of 2% chord, a transition strip (k=c¼ 7
104 ) was placed ahead of the simulated ice shape at 2% chord. What this strip/trip really simulates is uncertain. And, finally, with the Ref. [79] results, protuberance heights (k=c) of 41
104, 107
104, and 115
104 were utilized. The variation in indicated maximum lift penalty caused by chordwise movements of small protuberances/roughnesses (i.e., k=cp20
104 ) for the 0012 airfoil in Ref. [65] and 14
104 for the 23012 airfoil in Refs. [93,94] are illustrated in Fig. 24. At these small protuberance/roughness heights, it can be seen that the most critical location for these disturbances is indicated to be very close to the LE. This trend for small disturbance heights is consistent with the airfoil test results presented by Morgan et al. [33] over a Reynolds number range from 3 to 12 million where the maximum lift penalty caused by a roughness strip (k=cE4:5
104 ) extending form 3% to 5% chord was only about 85% of that which occurred when the roughness extended all the way to the LE. Incidentally, for the record, even though test results were obtained at 12 million Reynolds number, turbulent attachment line conditions would not have existed because of the (2-D) configuration, low tunnel turbulence level, etc. What all of these test results appear to indicate is, that for quite small runback or ridge ice accretions (i.e., k=co20
104 ), the maximum lift penalties would be less than those caused by initial leading-edge ice accretions (roughness) of the same size. Things change significantly, however, as the size of the protuberance or simulated ridge ice shape is increased, both in terms of the magnitude of the maximum lift penalty indicated, and the trend with chordwise location. Starting with the results for the 0012 and 23012 airfoils from Refs. [65] and [93,94] shown in Fig. 25 for the somewhat larger (i.e., k=cX50
104 ) protuberances/simulated ridge ice sizes (believed to be plausible for these types of accretions), some extremely large maximum lift penalties are indicated, particularly for

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Fig. 24. Variation of maximum lift penalty with chordwise location of small disturbances.

Fig. 25. Variation of maximum lift penalty with size and chordwise location of simulated runback and ridge ice shapes.

simulated accretion locations somewhat aft of 5% chord where runback and ridge ice accretions would be expected to occur. Possible maximum lift reductions of over 80% are indicated for the 23012 airfoil, more than twice the maximum seen for the initial leading-edge ice accretions addressed in the previous section. And, the results for the 0012 airfoil are actually not that much

lower than those for the 23012 configuration either where comparable results (in terms of disturbance height and location) are available. Clearly, in this case, the absence of data from Ref. [65] for disturbance locations between 5% and 15% chord is a major limitation for this application. However, be that as it may, we are still left with two sets of potentially applicable

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Fig. 26. Variation of maximum lift penalty with airfoil geometry and chordwise location of larger runback and ridge ice shapes.

low Reynolds number test results (for the 23012 and 0012 airfoils) which indicate that some potential runback and ridge ice accretions can lead to maximum lift reductions of at least 70%, and possibly more than 80%. This is downright scary. And, most interestingly, these magnitudes are not that inconsistent with the 1982 incident experienced with the King Air aircraft [87] where after encountering icing conditions which contained some large droplets, ‘‘a pronounced buffet was experienced’’ at an indicated airspeed 40% above the normal stall speed for this aircraft. Penalties caused by the larger ridge ice simulations (i.e., k=c¼ 139
104 ) with the other three airfoil geometries investigated by Lee and Bragg [93,95] at 1.8 million Reynolds number are also quite substantial, but not as severe as seen with the 23012 and 0012 airfoil geometries. The resulting penalties for these configurations are shown in Fig. 26 along with the corresponding 23012 test results. It can be seen that the reductions in maximum lift capability are in the 50–60% range in contrast with the 80% plus range seen with the 23012 configuration. In searching for explanations for the different characteristics, a number of possibilities exist. To start with, from the information presented in Fig. 26, it can be seen that the larger penalties are incurred by configurations having the higher basic airfoil maximum lift capability, and vice versa. It is also very likely that some of differences in the penalties might be related to differences in the levels of leading-edge suction pressures

on the clean airfoil, i.e., the higher the suction pressure at maximum lift conditions, the greater the adverse consequences of the simulated ridge ice accretions. The opposite might be expected as well, and the NLF airfoil (by design philosophy) and the thicker airfoil could fit in this category. Or, some of the differences might be anomalies associated with the low test Reynolds number. For example, it was seen earlier (in Fig. 5) that there was little impact of increasing the test Reynolds number from 1 to 6 million on the penalty caused by leading-edge roughness on the 23012 airfoil. However, on the other hand, for the NLF airfoil, which has the lowest indicated penalties at this relatively low Reynolds number, Addy and Chung [96] subsequently demonstrated a 25% higher maximum lift level for this same (clean) NLF airfoil at higher Reynolds numbers (up to 10 million). So, it would be somewhat natural to expect a higher penalty for this ridge ice simulation at higher Reynolds numbers on this NLF airfoil. Some additional instructive insights into factors which can influence the penalties incurred with runback/ridge ice accretions (and simulations) are provided by the Ref. [79] test results. For one, the results shown in Fig. 27 help illustrate that there is another significant variable for these ridge ice accretions in addition to size, location, lifting-surface geometry, Reynolds number, etc., and that is the details of the aft geometry of the forward-facing protuberance/simulations. This can be seen from the differences in the resulting penalties

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Fig. 27. Variation in maximum lift penalty with geometry and chordwise location of runback and ridge ice shapes.

between the first three protuberances which have very similar heights. It is also interesting to note that the corresponding Ref. [95] test results for this particular airfoil obtained with the forward-facing quarter-round protuberances do not indicate the greater penalty that would be anticipated with the larger protuberance height (i.e., k=c ¼ 139
104 versus 107–115
104). This could possibly be as a consequence of the transition-fixing technique employed in the Ref. [95] results for all protuberance locations aft of 2% chord, or it could just indicate the forward-facing quarter-round simulation is not the most adverse possible simulation for these runback/ridge-ice accretions. Unfortunately, very little seems to be known (i.e., published) about the actual ridge ice shapes encountered in flight other than the King Air flight crew observations [90] that ‘‘the drizzle drops freeze as sharp ‘‘feathers’’ or ‘‘glaze nodules’’ at or just beyond the de-icing equipment’’. Another valuable insight provided by the Ref. [79] test results has to do with the effect of varying test Reynolds number on the indicated penalties. In this case, the penalties indicated at 1 million Reynolds number for a particular protuberance were significantly lower than those indicated at 2 million Reynolds number. For example, with the larger protuberance heights, the indicated penalties at 1 million Reynolds number were as much as a third lower. But, with the smaller protuberance height (i.e., k=c¼ 41
104 ), some reductions were much more yet. However, probably the most important lesson to be learned from the Ref. [79] test

results is that most of the maximum-lift-loss penalties indicated for the larger runback/ridge ice accretion simulations are substantially larger than those indicated for/with a noticeably larger 45 min simulated glaze ice accretion on this same airfoil at the same Reynolds number. As a consequence of the shear magnitude of the potential penalties involved, as well as considering some of the uncertainties remaining, some high Reynolds number investigations of a range of potential runback/ ridge ice accretions are clearly warranted. Studies of ways to ensure that these accretions do not occur, or are avoided, would also certainly appear to be warranted. Probably the only two ground test facilities suitable for meaningful high Reynolds number experimental investigations of the effects of these runback/ridge ice accretions on aircraft which are large enough to have turbulent attachment line conditions in flight are the NASA National Transonic Facility (NTF) and the European Transonic Wind Tunnel (ETW). Little is to be gained from further low Reynolds number investigations of these potential effects such as the experiment summarized in Ref. [97] which was conducted at a (low) Reynolds number of 1.25 million with a quite large model of the 0012 airfoil in a relatively small wind tunnel, and with sparsely located small protuberances (i.e. k=c¼ 35
104 ). And, realistically, any further 2-D tests at all, even at high Reynolds numbers (especially with too-large models) are destined to leave key questions unanswered because

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of the leading-edge boundary-layer state simulation issue. 5.2. Stall angle reductions With the very large maximum lift reductions indicated for these simulated runback/ridge ice accretions from the available database, it follows naturally that there are corresponding very large reductions in stall angle indicated as well. For example, stall angle reductions indicated for the 23012 airfoil in Refs. [93,94] with the simulated ridge ice accretion placed at 10% chord range from about 91 up to 131 as the height (k=c) of the ridge is increased from 56
104 up to 139
104. As would be anticipated, these reductions are about twice the maximum seen from the 2-D results for initial leadingedge ice accretions. With the Ref. [79] test results for the Twin Otter tail airfoil section, stall angle reductions of up to 6 degrees are indicated. Although not as large as the Refs. [93,94] test results for the 23012 airfoil, they are consistent considering the reduced protuberance height and less critical airfoil design. Again, these results clearly point to the need for some new, well thought out studies to assess their applicability because of the very large magnitudes involved, and potentially very serious consequences implied. 5.3. Drag penalties Drag increases resulting from these simulated runback/ridge ice accretions are of obvious interest and concern from a basic flight safety perspective, but also as a possible means of alerting flight crews to the presence of these ice accretions. Unfortunately, the data available are not particularly well suited for assessing either of these situations. To start with, there is always the low Reynolds number concern. But, more than that, with the Refs. [93,94] results there is the problem that, with the very large maximum lift penalties incurred (with the 23012 airfoil), the approaches used previously of assessing increases in both the minimum parasite drag level and the drag levels at a higher lift coefficient are not possible. In reality, with these large maximum lift reductions, the corresponding drag increases would probably be of secondary importance. However, to give some idea of the magnitude of the profile drag increases incurred in this investigation, at a lift coefficient of 0.2, which is about the maximum for which an assessment can be made, the profile drag increases caused by the simulated ridge ice accretions at 10% chord vary from a factor of four for the smallest disturbance (i.e., k=c¼ 56
104 ) to more than a tenfold increase for the largest (i.e., k=c¼ 139
104 ). In comparison, with the Ref. [79] test results where some approximations can be made because of the lower maximum-lift-loss penalties, drag increases for the various simulations vary

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anywhere from a factor of around two to nearly ten at the selected higher lift condition. For Jacobs’ results [65], the absence of any data for disturbance locations between 5% and 15% chord, which is where such runback/ridge ice accretions would be expected to occur, remains a serious limitation for this application. Incidentally, the King Air flight test results [87–90] do provide an additional perspective in terms of possible total aircraft drag penalties which can be incurred as a consequence of icing encounters with large droplets on this size and type of aircraft. A number of cases were reported having drag increases of a factor of over two, up to a maximum of 2.3. However, there is no way of breaking these increases down in terms of the contributions from the various aircraft components including all of the ‘‘unprotected’’ areas. Consequently, probably the only thing which can be said with any certainty at this time is that a wide range of (large) drag increases is anticipated with these runback/ridge ice accretions depending upon the lifting surface geometry, flight conditions, and runback/ridge ice geometries, locations, severity, etc.

6. Effect of large in-flight ice accretions Compared to the other categories of ice accretions being addressed herein, far more effort for the past years and decades has been focused on this category, both in terms of developing methodology for predicting the physical characteristics (i.e., size, shape, surface roughness, etc.) of these larger and often-more-irregularly shaped ice accretions, and in trying to assess the aerodynamic penalties associated with them. This relatively very large emphasis on these larger ice accretions, especially glaze ice accretions with horns, etc., is quite natural and easy to understand if one makes the thought-to-be obvious assumption that aerodynamic penalties will be somehow proportional to the size and grotesqueness of the ice accretion. However, how true this is, especially relative to the runback/ridge ice accretions, remains to be seen. One of the most important objectives for a significant portion of the research which has gone on was to help define ‘‘critical’’ ice shape features, where ‘‘critical ice shapes’’ are defined in Ref. [98] as ‘‘those with ice accretion geometries and features representative of that which can be produced within the icing certification envelope that result in the largest adverse effects on performance and handling qualities over the applicable phases of flight of the aircraft’’. Unfortunately, other than identifying glaze ice accretions as the principal area to focus on, the authors of Ref. [98] concluded (in late 2000) that existing guidance material pertaining to the determination of ‘‘critical ice shapes’’ was too general and did not provide a working definition of ‘‘critical ice shape’’ or

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a description of engineering practices to be followed in the determination of such shapes. Significant additional new research was suggested as a means of eventually enabling these definitions to be made. Although the results from many of the experimental investigations conducted to date addressing the aerodynamic effects of these large in-flight ice accretions on lifting surfaces are potentially applicable to both wings and horizontal tails, a number of studies specifically focused on separately determining either wing or horizontal tail characteristics have also been needed/ pursued as a consequence of some obvious as well as some more subtle differences in geometries, operating environment, requirements, etc. Four of the factors which can lead to different ice accretions and consequences on horizontal tails compared to wings on any given aircraft are the typically smaller surface size, the corresponding reduced Reynolds number, the sometimes nondimensionally thinner defining airfoil sections used to ensure satisfactory high speed longitudinal control, and the absence of a high-lift system. The difference in physical size is important because it causes the (smaller) tail to be a more efficient ice collector. And, any reductions in nondimensional thickness ratio will further enhance this. On the other hand, the reduced Reynolds number, and possibly reduced thickness ratio, may in some cases reduce the magnitude of some of the aerodynamic penalties imposed by these large leadingedge ice accretions. However, probably the biggest concern with tailplane icing is the potential loss of longitudinal control which can result from premature tailplane stall, elevator hinge moment anomalies, etc. A number of studies have focussed strictly on this aspect starting with the studies in the 1940s [99,100] addressing uncontrollable pitching oscillations encountered on the Vickers Viking aircraft with ice on the tailplane LE. Other noteworthy studies specifically addressing tailplane icing phenomena include the work of the SwedishSoviet Working Group in the late 1970s [77,78], and the results from the NASA/FAA Tailplane Icing Program (TIP) started in the mid-1990s [101]. Similarly, there have been a number of the experimental investigations of the effects of these large ice accretions which are much more pertinent to the wing characteristics. Clearly, the studies of ice accretion effects on multi-element highlift airfoil configurations starting with the first report published by the Swedish-Soviet Working Group [74] are in this category. Also, realistically, the results from the numerous experimental studies focussed on the drag penalties caused by these large ice accretions are much more relevant for the wing (in terms of total aircraft drag) because of the relatively much larger surface area. Not surprisingly, with the much greater emphasis that has been placed on studying the aerodynamic effects of this category of larger ice accretions, there are more varied sources of potentially applicable test results. Not

only are there a wide assortment of test results available from conventional wind tunnels using simulated ice accretions, but, there are also data available from flight testing conducted in natural icing conditions, flight tests carried out with simulated large ice accretions for both research and certification programs, as well as from a lot of testing in icing tunnels such as the Icing Research Tunnel (IRT) at NASA Glenn. All of this additional testing in flight and in icing tunnels has definitely provided numerous quite unique, very insightful, and useful perspectives on a number of the critical performance and flight control issues central to properly assessing the range of aerodynamic effects which can be caused by this category of larger in-flight ice accretions. However, on the other hand, there are some inherent technical shortcomings associated with each of these additional test categories that significantly limits the amount of truly useful more-quantitative-type data that can be extracted from the test results. Flight testing in natural icing conditions provides some wonderful examples of these pros and cons. One of the benefits derived from such testing has been in providing some quite graphic examples (i.e., ‘‘wake-up calls’’) illustrating just how severely these ‘‘larger’’ ice accretions can impact aircraft performance and flight control characteristics. These include the following: *

*

*

*

Johnson in 1940 [85] reported on a flight experience with a Lockheed Electra wherein the aircraft ‘‘was iced to the extent that the indicated high speed with all the power available to the pilot with carburetor heat applied, was only 90 mph at 5000 ft’’. He also reported that ‘‘practically full power was required for landing and only the pilot’s skill prevented serious consequences in this particular case’’. The aforementioned report by Morris in 1947 [99] of uncontrollable pitching oscillations having occurred during flight with the Vickers Viking aircraft when there was ice on the tailplane LE. In what appears to be the first reported situation where this type of testing was conducted with the primary objective of determining somewhat quantitatively the effects of these larger ice accretions on aircraft performance characteristics, Preston and Blackman in 1948 [102] reported an 81% increase in the parasite drag of a ‘‘twin-engine airplane in level cruising flight’’ after 40 min in meteorological conditions which would at very most be considered ‘‘moderate icing’’ by current standards. Leckman in 1971 [103] reported a 34% increase in stall speed having occurred on a Cessna Centurion aircraft with just a 14 in rough ice accretion on the wing LE. A significant reduction in the rate of climb with this ice accretion was also reported. Then, following from the lessons learned from this testing in natural icing conditions, Leckman pointed out

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*

*

*

that ‘‘elevator effectiveness in the landing flare may be impaired by ragged ice accumulations on the LE of the horizontal stabilizer’’, and that resultant ‘‘airflow disturbances at the LE could result in a separated airflow under the elevator and a corresponding loss in elevator effectiveness’’. He went on to point out that ‘‘these effects would be aggravated by the increasing downwash angles associated with larger flap deflections’’, and concluded that ‘‘as ice accumulations become more severe, the wing flap deflections should be minimized as a precautions against partial stabilizer stalling’’. This was 30 years ago! In the first joint report published by the SwedishSoviet Working Group in 1977 [74], results from flight testing the Antonov An-12 aircraft in natural icing conditions are presented which indicated a 31% reduction in the maximum lift capability, a 31 reduction in the stall angle, and a 50% reduction in the lift-to-drag ratio, all caused by an ‘‘ice deposit shape on the LE of the wing having a maximum thickness of 20 mm’’, which is described as only moderate icing, since ‘‘operational experiences show that more severe icing conditions may take place in practice’’. They also reported that the ‘‘horizontal flight speed of the TU-124 and TU-134 were reduced by 40–50 km/h with ice roughness of 6–8 mm formed on the wing and tail units’’, and that the speed decreased more sharply in the beginning of the icing process. This led to the recommendations to ‘‘never ignore even light icing as it can be extremely dangerous’’, and to ‘‘always remember the main general recommendation of timely activation of the anti-icing system’’. They further pointed out that ‘‘for the approach icing case the test results indicated a drag increase due to icing of such magnitude that in the one-engine-out case a go-around may not be possible within the limits specified in the Airworthiness requirements’’. Other flight test results obtained with the An-12 [104], seemingly from about the same time period in the early to mid 1970s, revealed a ‘‘significant corruption of lateral controllability characteristics at low airspeeds due to the ice accumulation’’ on the wing LE. Ice accreted on the LE of the wing caused an aileron hinge-moment reversal and accompanying ‘‘snatching of the yoke’’ when the trailing-edge-down aileron deflection reached only about one-third of the full range at a condition (i.e., lift coefficient) ‘‘far from stall’’. And, this reversal occurred even earlier (i.e., lower deflection angles) at speeds closer to stall. In the third joint report issued by the Swedish-Soviet Working Group in 1985 [78], results were presented from flight testing on the An-18, IL-18, An-24, and Yak-40 aircraft at landing flap settings which showed a definite lightening of (control) stick forces when ice

*

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was allowed to build up on the LE of the horizontal tail. In some of the cases, there was a relatively sudden and sharp decrease and subsequent reversal of the stick forces which occurred with very modest ice accretions. Results obtained from flight tests of the NASA Twin Otter aircraft during 1983–1986 [105–107] showed total airplane drag penalties of up to 75%, and increases in wing section drag up to 120%, after flying through natural icing conditions.

While there are undoubtedly numerous additional sources of data from flight testing in natural icing conditions that are either not well publicized or are not available (i.e., not in the public domain), the foregoing chronology does appear to highlight most if not all of the possible more adverse repercussions which can be caused by these larger ice accretions. Another very important benefit derived from the flight testing conducted in natural icing conditions is the insight provided in terms of identifying the types of ice accretions and some of the corresponding environmental conditions in which they form that result in the most serious degradations of performance and flight control characteristics. These insights tend to fall into two groups. First, there are the mostly very qualitative and somewhat more obvious ones (today) provided by the earlier (i.e., mostly pre-1980) flight testing. These include Johnson’s [85] observation that ‘‘ice accreted at 401F was so smooth that no detrimental effect could be noticed in the airplane performance’’, and Leckman’s [103] conclusions that ‘‘glaze (clear) type of ice formation is most adverse with respect to its effect on aerodynamic drag and stall’’, and that ‘‘double horn glaze ice shapes make the drag contribution unusually high’’. Other sources/insights that fall in this group include the observations provided in the first report by the Swedish-Soviet Working Group [74] that ‘‘in spite of only half the thickness, the horn-shaped ice will be much more detrimental to the aerodynamic characteristics than the wedge-shaped ice’’, as well as those provided in the third report [78] that ‘‘the form of the ice on the tailplane LE, and apparently its surface conditions (smooth, rough, wavy) play an extremely important role’’. Some of the test results provided in [78] indicated that ‘‘if horn shaped, the ice created a severe deterioration of the aircraft stability, but, when ice conformed to the section shape, there was practically no influence’’. However, they also showed an example on another aircraft where the wedge shape had ‘‘a considerable impact’’. As indicated, however, most of these insights/ observations should be (at least) somewhat self apparent. That is also true of some of the observations provided in connection with the second group of later (i.e., starting in the early 1980s) flight test results obtained with the NASA Twin Otter aircraft. For

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example, in Ref. [105], it is stated that ‘‘glaze icing, which generally occurs at total air temperatures just below freezing, results in rough, irregular ice formations with flat or concave surfaces facing the airstream’’, and that ‘‘this type of icing causes the largest aircraft performance penalties in terms of the increase in aircraft drag’’. This is compared to ‘‘rime icing which generally occurs at lower total air temperatures below freezing, and results in smoother, more pointed ice formations facing the airstream, which cause much lower drag penalties as compared to similar amounts of glaze ice’’. But, in addition to these more qualitative findings, the Twin Otter test results in this second group do provide some very instructive specifics illustrating how some relatively small differences in environmental conditions can lead to some very significant differences in the effects on aircraft performance (i.e., total aircraft drag). The summaries of the flight test results obtained with the NASA Twin Otter [105–107] contain three very interesting sets of comparisons in terms of relating aircraft drag increases to environmental conditions, icing encounter durations, etc. (with the de-icing system turned off). In the first set [105], icing encounter specifics and the resulting drag increases are examined for two glaze ice accretions and one rime ice accretion involving three separate flights (in addition to the un-iced baseline testing). Total temperatures for these encounters were about 91C for the rime ice case, and about 31C and 21C for the glaze ice cases. This particular rime ice encounter involved a relatively high LWC and a relatively longer duration, making it a rather severe rime ice encounter. And, even though one of the glaze ice accretions was characterized on the basis of LWC and encounter duration as being ‘‘a relatively severe (glaze) icing encounter’’, the amount of ice accreted during the rime icing encounter was about twice that for either of the glaze icing encounters. In spite of this difference, the rime ice drag penalty was only 15%, compared to 47% and 62% for the two glaze ice accretions. Regarding the drag penalties for the two glaze ice accretions, the larger penalty is associated with the more severe icing encounter, but the magnitude of the increases is nearly three times greater than might be expected by just ratioing the accretion totals, indicating a noticeably non-linear progression. The second set of comparisons [106] is for what is described as a severe glaze icing encounter, and two mixed (i.e., mixed rime and glaze) icing encounters. Incidentally, a stereo photography system was used to document some of the in-flight ice shapes involved, but, unfortunately, this documentation was not provided for the glaze ice shape of (most) interest in this case. Total temperatures for the three cases were 3.51C for the glaze icing encounter, and 4.81C and 5.21C for the two mixed icing encounters. As a consequence of offsetting differences in LWC and icing encounter

durations, both of the mixed ice accretions had about the same amount of ice accreted, and this was about 50% greater than with the glaze ice accretions. However, despite this difference, the total airplane drag penalty incurred with the glaze ice shape was double that experienced with the two mixed ice accretions (i.e., 75% versus 31% and 38%). The difference in measured wing section drag penalties was even more dramatic yet. At a representative condition, the wing section drag penalty for the glaze ice of 120% is over six times greater than that seen with one of the mixed ice accretions. Other than the (less than 21C) total temperature difference, the only other envisioned contributor to these very large drag differences might conceivably be the higher maximum droplet diameters present in the glaze icing encounter (i.e., 41 mm versus 27–30 mm). This possibility remains an open issue. Looking next at the third set of comparisons involving Twin Otter flight test results [107], the comparison in this case is between a moderate glaze icing condition and what is described as a moderate-to-heavy mixed icing condition. A 4.51C difference in temperature between the two flights was what led to the formation of a glazetype ice accretion on one and a mixed-type ice accretion on the other. With this difference in shapes, even though the mixed icing encounter lasted 20% longer, and had a LWC more than double that of the glaze icing encounter, the glaze ice shape resulted in overall drag increases 15–20% higher than those incurred with the mixed-type ice accretions. So, in all three of these comparisons, a relatively small difference in ambient temperature, when it resulted in the formation of a glaze-type ice accretion was of paramount importance in establishing appropriate drag penalties. There was also a hint provided in the second set of comparisons that larger droplet sizes (i.e., greater than 40 mm) might also lead to ice shapes which cause even higher drag penalties. Although this possibility would certainly seem to be plausible, there just does not appear to be any additional flight data available to really substantiate or quantify this effect. As it is, the only other reported ‘‘results’’ available for aircraft encounters with larger droplet sizes [87–90,108,109] are for situations where the ice-protection systems were being operated during such encounters. Hence, the insights provided relate primarily to the effects of potential runback or ridge ice accretions associated with these larger droplets, and/or the likely larger inter-cycle ice accretions. Both sets of flight results obtained with turboprop aircraft employing de-icing systems, the aforementioned results obtained with the Super King Air [87–90] and the Ref. [108] results obtained with an F-27, indicated (not surprisingly) noticeably higher drag penalties associated with these conditions. On the other hand, results obtained with a Boeing 777 [109], which utilizes anti-icing on the wing, indicated ‘‘no noticeable effects on airplane performance

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and handling qualities’’ with the size of droplets encountered, i.e., 50–85 mm. Hence, in addition to addressing the potentially very large adverse effects of ridge-type ice accretions which might form in large droplet encounters, additional attention also needs to be focused on the potential consequences of the larger inter-cycle ice accretions which would likely occur as well with any de-icing concept. Notwithstanding the valuable insights which have been enabled by flight testing in natural icing conditions, there are, as indicated previously, some inherent limitations involved with this type of testing. Some practical limitations were mentioned in Section 1 such as cost, schedule, seasonal issues, and the risk involved with finding appropriate (i.e., the most critical) icing conditions. Clearly, the flight test results obtained with the NASA Twin Otter aircraft [105–107] showing that as little as a 21C temperature difference can make a big difference help put this risk in perspective. However, in addition to these, there are some other very important limitations as well which can be identified quite easily by taking note of what is typically not reported from testing conducted in natural icing conditions. Uppermost in this grouping would be a good definition of the actual accreted ice shape and surface conditions, a good definition of the ice shape and surface conditions existing subsequently when the actual performance and flight control characteristics measurements are most often made, and at least some reliable indication/ measurements of the increases in stall speed which would be incurred. Typically, available descriptions of the actual ice shapes accreted in natural icing conditions and their surface condition/roughness characteristics are based on a combination of visual observations and/or photographs taken by the flight test crew, both of which fall well short of providing a useful quantitative description of the shape, let alone the surface condition/roughness characteristics. The only known exceptions to this (definitely inadequate) method of ‘‘defining’’ the actual ice shapes were the utilization of stereo photography during some of the Twin Otter flight testing [106], and some reported cases such as those in Ref. [78] when temperatures were cold enough that the aircraft could be landed with the accreted ice shape largely intact (i.e., not melted or shed) so as to permit appropriate measurements, castings, etc. to be made. However, even though the use of stereo photography did enable a fairly good definition of some of the more 2-D ice shape characteristics, there was still no adequate description of surface conditions and small scale 3-D features. Also, in this particular demonstration/application of the technology, definitions of the most critical (i.e. glaze) ice shapes encountered were not obtained/ reported. And, with regard to making the necessary ice shape, etc., measurements after landing, unfortunately this only works with the much less bothersome rime ice

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accretions which typically form at temperatures below 101C. For the more dangerous glaze ice accretions, which form at temperatures just below freezing, these accretions would likely be melted away in descending to land (at typically warmer temperatures). The analogous second typically missing ingredient from reports of flight testing in natural icing conditions is a good definition of the ice shape and surface conditions existing when ‘‘corresponding’’ performance and/or stability and control measurements/evaluations are actually made. This requirement arises because very often it is necessary to exit the natural icing conditions before undertaking these tests, either for safety reasons, and/or in order to find suitable atmosphere conditions for making meaningful (incremental) performance and flight-control-related measurements/evaluations. However, in the time it takes to do this and prepare for the desired testing, erosion/sublimation and even possibly some shedding of the accreted ice can take place. There have been a number of reports acknowledging such changes having taken place. For example, in Ref. [78], it was reported that ‘‘some ice came off before starting measurements’’. Occasional shedding of the ice accretion in this time interval has also been acknowledged in other reports as well, included cases involving loss of ice caused by airloads and wing flexure. Erosion or sublimation of accreted ice shapes is probably much more common, however, and although the geometric changes may be more subtle, such changes can nevertheless potentially still lead to significant differences in their effects on aerodynamic characteristics. One common way in which this erosion or sublimation occurs is associated with melting of the ice, and this can occur either as a consequence of climbing above the icing cloud to where sunlight can be an issue, or by descending below the cloud to where warmer air exists. Erosion caused by ice crystals has also been reported/ suggested [90]. One of the forms in which this erosion has been observed to occur is a reduction in the surface roughness. For example, it is reported in Ref. [106] that ‘‘conventional photographs taken during this time indicate that some ice sublimation or erosion occurred’’, and ‘‘although the general ice shape remained, the characteristic roughness was reduced’’. This is for a situation when a little over 20 min was taken to exit the icing cloud and take the performance measurements. Although testing procedures have been adopted in some cases [107] in an attempt to minimize this erosion, it is hard to totally avoid. Hence, the typical result from testing in natural icing conditions is a lack of a good definition for both the actual accreted and the subsequent ice shapes and surface conditions, plus an absence of any good data addressing what the consequences of the differences in ice shapes and surface conditions might be. This is not an enviable state of affairs, nor is it a good basis for determining the specifics

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for appropriate simulated ice shapes (and surface conditions). While all of the foregoing limitations discussed are very important considerations when assessing the true value of flight testing in natural icing conditions, perhaps the most important shortcoming of this type of testing has been, with only two known exceptions [74,103], the inability to determine the stall speed increases/penalties associated with the various ice accretions formed in these conditions. Reasons for this, primarily flight safety considerations, are most understandable. For example, oftentimes these icing encounters are experienced at relatively low altitudes where adequate ground clearance for sensible stall testing is clearly not available. Also, concerns regarding instrumentation status, visibility, etc., have been a factor, as have been concerns with not really knowing what to expect. One likely typical reaction to all these concerns has been that ‘‘the aircraft was not permitted to stall’’ as stated in Ref. [87]. Interestingly, in the two cases where stall speed increases/penalties were reported, the indicated ice accretions were certainly not of the form that would have been expected to be the most detrimental. For the 34% increase in stall speed (i.e., 45% reduction in the maximum lift achievable) reported in Ref. [103] for the Cessna Centurion aircraft, the ice shape which caused this was described as a 14 in rough ice accretions on the wing LE, certainly not a large double-horned glaze type ice accretion. Hence, the magnitude of this particular penalty for such an ice accretion naturally leads to some questions when subsequently examining wind-tunnel measurements of maximum lift penalties obtained with simulated ice shapes as to whether or not the tunnel results are at all indicative of what can really occur in flight with the thought-to-be much more critical large glaze ice accretions. And, lastly, the extreme reluctance of experienced flight test crews to stall aircraft in natural icing conditions should certainly be an even further incentive for others to avoid these potentially very dangerous situations. Turning next to flight testing conducted using simulated large ice accretions, there are some different pros and cons, but also some that are closely related as well. Assuming that the simulations are appropriate (and this is a major assumption), this type of testing avoids a number of the limitations inherent with flight testing in natural icing conditions, but it also introduces some new concerns. On the positive side, flight testing with simulated ice accretions avoids scheduling and seasonal issues, concerns about having to find the more critical icing conditions, and not knowing what the ice shape/surface conditions are for the measurements and evaluations being made. Curiously, however, although there have undoubtedly been many instances where flight testing with simulated ice accretions has been a part of the development and certification process for

specific aircraft addressing a variety of performance as well as stability and control issues, the only published results available are from investigations focused on potential longitudinal stability and control problems associated with tailplane icing. Such applications started with studies conducted with the F-27 and F-28 in the 1970s [110], and include subsequent studies carried out on a variety of Russian aircraft [78], as well as a number of studies conducted utilizing the NASA Twin Otter aircraft [107,111–118]. Some of the lessons that have been learned from these studies are that with these configurations, almost any ice accretion on the LE of the horizontal tail leads to some lightening of the stick forces in a landing condition, and that the degradation is related to the severity of the contamination. As an example, results reported in Refs. [107,114,116] illustrated that the degradation got progressively worse in going from an inter-cycle type ice accretion, to a failedboot type of accretion, and finally to a double-horned glaze ice type accretion. A noteworthy aspect of the results reported by the Swedish-Soviet Working Group in Ref. [78] was that in two of three cases where the flight results obtained with simulated ice shapes (referred to as ‘‘imitators’’ in this case) could be compared to the ‘‘corresponding’’ results obtained from testing in natural icing conditions, the correlation was deemed unacceptable. This report should be discounted, however, because in the one case when the correlation was judged unacceptable and a comparison of the so-called ‘‘imitator’’ geometry with the ice shape reported from the natural icing encounter was provided, the ‘‘imitator’’ was obviously not an adequate simulation. In fact, whereas the natural icing shape was of the doublehorned type, the ‘‘imitator’’ did not have horns at all. The consequence of this is that there is really no flight data available to realistically assess the ability of simulated ice shapes (no matter how defined) to reliably produce the aerodynamic effects incurred with ‘‘corresponding’’ natural ice accretions, and there is no obvious way to accomplish this validation either. Similar to the situation existing with flight testing in natural icing conditions, some of the important limitations associated with flight testing with simulated ice accretions become apparent by taking note of what is not reported. In this case, the most serious omissions are the lack of any really quantitative-type data in terms of maximum lift degradations, stall margin reductions, etc. from any of the tailplane icing studies, and the absence of any available data addressing aircraft/wing stall speed or drag increases associated with large ice accretions. Again, flight safety concerns would certainly seem to be an important element in the absence of stall speed and drag data, most notably serious concerns about taking off with large horned ice shapes on the wing. Hence, this category of testing has provided little additional information to assist with the quantitative assessment

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of the various aerodynamic effects caused by these larger ice accretions other than to establish that the doublehorned glaze ice accretions were more detrimental from a tailplane stall perspective than smaller less obtrusive accretions with the tailplane designs on the aircraft tested. In a very similar manner, numerous tests conducted in icing tunnels have also enabled/provided a number of most helpful insights into a number of factors critical in establishing the magnitude of possible adverse aerodynamic effects/penalties caused by larger ice accretions. However, once again, there are likewise some unique and inherent limitations associated with this kind of testing as well. Applicable results from a number of these investigations are reported in Refs. [119–132]. The earliest of these appears to be the research conducted by Gray and von Glahn in the 1950s and early 1960s [119– 122] utilizing a number of different airfoils. Following this, there appears to have been somewhat of a hiatus in this type of research until the early 1980s when it started again in earnest [123–132]. Some of most important insights gained from the results produced by all this testing, as well as some of the important utilizations of these results, include the following: *

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The results from such testing in the form of the actual ice accretions observed have been the principal source of guidance in defining the simulated ice shapes used in subsequent flight tests and (conventional) wind tunnel testing to evaluate the various aerodynamic effects of these large ice accretions on both single-element and multi-element-high-lift airfoils. This testing has enabled assessments to be made of the effectiveness of these intended-to-be simulations of the actual icing tunnel accretions in duplicating the drag penalties caused by the actual shapes. Unfortunately, in the only known (published) study addressing this issue [123] conducted nearly 20 years ago, the results were not exactly encouraging. In this particular study for a single-element airfoil, two rime ice and two glaze ice accretions were addressed, and both smooth and subsequently roughened (with an abrasive) versions of each were evaluated. For one of the rime ice accretions, the roughened simulation produced results very similar to those seen with the actual ice shape. In this case, results obtained with the smooth version were not at all representative. However, with the second rime ice simulation and the two more critical glaze ice simulations, none of the results produced by the simulations (either smooth or roughened) were satisfactory at the lower angles of attack where drag increases would be of consequence. Clearly, this issue is most definitely worthy of more attention. Test results obtained in icing tunnels, most specifically the IRT at NASA Glenn, are the largest source

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of (somewhat quantitative) data available addressing the range of potential drag penalties that can be incurred with these larger ice accretions. For example, consistent with the previously discussed results from the flight testing conducted with the NASA Twin Otter aircraft [105–107], results from a number of investigations conducted in the IRT have demonstrated that the ice shapes which cause the largest drag penalties on single-element lifting surfaces are the glaze ice accretions formed at (total) temperatures just below freezing conditions. Certainly amongst the first to come to this determination were Gray and von Glahn in 1953 [119] who reported that ‘‘glaze ice formations on the upper surface near the LE of the airfoil caused large and rapid increases in drag, especially at datum (total) air temperatures approaching 321F and in the presence of high rates of water catch’’. They compared these results with ‘‘ice formations at lower temperatures (rime ice) did not appreciably increase the drag coefficient over the initial (standard roughness) drag coefficient’’. These results were obtained with a NACA 651-212 airfoil. Similar results have been observed in a number of subsequent investigations using a variety of different airfoils [122,124,125,129–132]. One of the more graphic examples of this (localized) occurrence of high drag penalties at temperatures just below freezing was reported by Olsen et al. [125]. Their results obtained with a NACA 0012 airfoil are depicted in Fig. 28, where it can be seen that at the higher speed condition, the rapid increase in drag penalty as temperatures approach freezing conditions coincides with the upper ice horn becoming noticeably more vertical. The same correlation also held true at the lower speed condition. In addition, it is significant to note (from Fig. 28) that even though the product of the airspeed, LWC, and the time (duration) of the icing encounter was held constant between these two cases, the ice accretion formed at the higher speed condition resulted in noticeably higher drag penalties. Although not a factor in the results shown in Fig. 28 where the angle of attack was held constant (at 41) throughout, higher speeds also cause even more detrimental ice accretions because ice is accreted at lower angles of attack where more of the ice accretion occurs on the upper (suction) surface. Other noteworthy results/insights obtained from tests conducted in the IRT regarding possible drag penalties associated with these larger ice shapes are that the penalties typically increase with increasing LWC (as might be expected), and, adding credence to the trends indicated by the Twin Otter flight test results pertaining to droplet size effects, the available data from the IRT [125,132] also indicate that these penalties do increase as the droplet size increases,

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Fig. 28. Ice accretion temperature effect on parasite drag penalty for NACA 0012 Airfoil in IRT [125].

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although the largest droplet sizes investigated (i.e., 40 mm) were small compared to what we now consider to be ‘‘large’’ droplets. Interestingly, Olsen et al. [125] observed that these droplet size effects appear to be greatest whenever the temperature effect is greatest, i.e., with the most detrimental glaze ice shapes. This observation should not be overlooked. Although a large majority of the testing conducted to date in the IRT has focused on determining the drag penalties associated with a wide variety of large ice accretions, the much more limited amount of data pertaining to the corresponding effects on maximum lift capabilities are nonetheless quite informative. For example, results obtained for single-element airfoils by Berkowitz et al for a Boeing 737-200 wing section [126–128], and by Addy [132] for a 14% thick GA NLF airfoil [132], both demonstrate that the largest reductions in maximum lift capability with these types of airfoils are also caused by the glaze ice accretions formed at temperatures just below freezing conditions. One set of representative results from Ref. [132] for the GA NLF airfoil are shown in Fig. 29. However, in contrast with these results, other test results reported in Ref. [132] for a 8.7% thick ‘‘business jet’’ airfoil with a relatively low basic maximum lift capability (B1.1) do not show as much of a temperature effect (see Fig. 30), once again

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illustrating the inadvisability of generalizing. Incidentally, while the trends indicated by the results shown in Figs. 29 and 30 are thought to be appropriate and meaningful, the absolute magnitude of the penalties indicated may not be for a number of reasons, not the least of which are doubts about the appropriateness of the maximum lift levels indicated for the clean (un-iced) configurations. An additional noteworthy aspect of the Ref. [132] results presented in Figs. 29 and 30 for both the GA NLF and ‘‘business jet’’ airfoils is that the penalty caused by the ice accretion in the first 1–2 min represents a significant part of the penalty eventually incurred after 22 min. This should not be surprising, however, in light of the magnitude of the penalties (seen previously) that are caused by initial leadingedge (roughness) ice accretions. A further benefit derived from testing in icing tunnels is the number of very helpful and thought-provoking insights provided relative to the effects of ice accretions on the multi-element high-lift geometries used on many aircraft types during takeoff and landing. A good example of this is the data reported by Potapczuk et al. [126–128], illustrating types of ice accretions, areas where they can form when testing in an icing tunnel, and the corresponding effects of these accretions on the aerodynamic characteristics

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Fig. 29. Ice accretion temperature and duration effects on maximum lift penalty for GA NLF airfoil in IRT [132].

Fig. 30. Ice accretion temperature and duration effects on maximum lift penalty for ‘‘business jet’’ airfoil in IRT [132].

(lift, drag, and pitching moments) of these designs. What these results revealed was that once again the largest reduction in maximum lift level was caused by

ice accretions formed at temperatures just below freezing conditions. However, in this case, the temperature effect on drag penalties incurred was,

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in general, not nearly as pronounced as that seen with single-element airfoils. One interesting phenomena encountered during this particular investigation was a flow separation which occurred on the upper surface of the flap over a wide range of angles of attack with the flap deflected 151 that was caused by ice which formed around the LE of the flap. The result was a significant loss in lift and increase in drag. It remains to be seen, however, as to whether these results are truly meaningful or not. That is because the real detrimental ice accretions on the flap LEs were formed at incidence angles that are perhaps too high to be really representative of likely flight conditions (in icing situations). In addition to the previously mentioned (in Section 2.1) concern about the accuracy of simulation/scaling techniques used when testing multi-element lifting surface geometries, there are some additional concerns associated with testing in icing tunnels that also need to be taken into account in order to put test results obtained (or not obtained) from these facilities into proper perspective. One large concern has been regarding tunnel wall interference effects contaminating test results obtained with the relatively large models often used (mostly in the past) in these facilities in order to limit some of the uncertainties associated with scaling techniques, and to achieve more meaningful Reynolds numbers. Examples of this situation are the test results from Refs. [119–124] where model airfoil chords ranging from 1.36 m up to over 2.4 m were used in the IRT, which has a cross section of 1.82
2.74 m2. Clearly, excessive wall interference would limit the usefullness of these test results, especially at higher angles of attack. In the next group of studies conducted in the IRT addressing aerodynamic effects of these larger ice accretions [125–130], much smaller models with chord lengths of about 0.46 m [126–128] and 0.53 m [125,129,130] were used. However, while the use of these smaller models greatly alleviated concerns about wall interference, the Reynolds number resulting from the use of these smaller models in an atmospheric tunnel such as the IRT is certainly lower than desired/needed. And, if a higher tunnel velocity is used to raise the Reynolds number, then compressibility effects contaminate the test results, especially for assessing maximum lift penalties. The chord length of about 0.9 m (36 in) used in the most recent test results reported by Addy [131–133] could be the best compromise, but even in this case, concerns still remain with regard to wall effects and compressibility effects at maximum lift conditions, and Reynolds number effects at all conditions. Hence, it appears that when testing for a range of aerodynamic effects in an icing tunnel such as the IRT, there is a ‘‘choice of poisons’’ involved if trying to extract quantitative results, i.e., having the data contaminated

by either uncorrectable tunnel wall interference effects, having an inadequate test Reynolds number, and/or by having inappropriate compressibility effects. And, there are always the quite-high freestream disturbance levels which exist in these facilities due to the presence of the spray rig, etc., in the circuit, as well as questions regarding the accuracy of the scaling methods used. Another limitation often mentioned relative to trying to use test results from icing tunnels for reliable representation of maximum lift characteristics is the nonuniformity of the icing cloud over the span of the model, and the corresponding limitations of force balance measurements of lift, etc. This would be much less of an issue for the measurement of section drag characteristics. At this point, the numerous lessons learned from flight tests as well as tests in icing tunnels such as the IRT addressing the aerodynamic effects caused by these larger ice shapes, plus the insights gained from the preceding reviews of the various effects caused by initial in-flight leading-edge ice accretions as well as runback and ridge ice accretions, have provided a solid foundation for effectively scrutinizing the total database currently available for quantifying some of the more critical consequences that can be caused by this category of larger ice shapes, and for identifying important deficiencies in the current database needing attention. Some of the more important lessons learned that should be most applicable and helpful in this particular assessment include the following: *

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Leading-edge ice accretions can have very different effects on ‘‘thinner’’ versus ‘‘thicker’’ lifting surfaces, particularly with regard to maximum lift characteristics/penalties. Low Reynolds number test results can often be quite misleading, mostly in a non-conservative manner, usually due to Reynolds number effects on the uniced baseline geometry. Uncorrectable tunnel wall effects and/or unrealistic compressibility effects can also lead to erroneous indications. Roughness effects can be very significant with many geometries, usually in a harmful way, but sometimes also in a helpful manner (for thinner geometries). Larger protuberances (perhaps analogous to the horns which occur with glaze ice accretions) tend to be more detrimental to maximum lift characteristics when they are located somewhat aft of the LE. For many applications, the most detrimental large ice accretions appear to be the glaze ice shapes which form at temperatures just below freezing, and high LWC, larger droplet sizes, and higher speeds tend to accentuate this effect even further. In this situation, small differences in temperature can lead to significant differences in the consequences.

F.T. Lynch, A. Khodadoust / Progress in Aerospace Sciences 37 (2001) 669–767 *

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There does not yet appear to be a good understanding of just what constitutes an adequate or effective ‘‘simulated ice shape’’. Ice accretions which form on some of the downstream elements of multi-element high-lift geometries in an icing tunnel, especially on the LE of the main element behind an extended slat or Krueger flap, may well not be representative of what occurs during flight in natural icing conditions.

6.1. Maximum lift reductions 6.1.1. Single-element lifting surfaces Including three very recent additions, there are now six known sources of potentially applicable ‘‘high’’ Reynolds number data available to provide some guidance for quantifying the reductions in maximum lift capability that can be incurred by this broad class of larger ice accretions. In chronological order, they are as follows: *

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The previously mentioned flight test results for a Cessna Centurion aircraft [103] wherein a 14 in rough ice accretion on the wing LE caused a 34% increase in stall speed (i.e., a 45% reduction in the maximum lift level achievable). The also-previously mentioned flight test results for an An-12 aircraft [74] where an ice accretion consistent with a ‘‘moderate’’ icing encounter led to a 31% reduction in the achievable maximum lift level. Test results from the LTPT obtained over a range of Reynolds numbers for a 11% thick airfoil with a somewhat ‘‘generic’’ glaze-type ice accretion [33]. These results are particularly useful because there is a direct comparison made with corresponding measurements obtained with simulated leading-edge frost (i.e., roughness). Test results obtained from the LTPT as part of the NASA Modern Airfoils (ice accretion) program for the previously discussed GA NLF 14% thick airfoil over a range of Reynolds numbers and Mach numbers with various simulated ice shapes [96,132]. These results also enable a comparison to be made with results obtained from testing in the IRT with ostensibly the same ice accretions. Further test results obtained from the LTPT as part of the NASA Modern Airfoils program, this time for the previously discussed 8.7% thick ‘‘business jet’’ airfoil with various simulated ice shapes [133]. Similarly, a comparison with supposedly corresponding results from the IRT can be made as well. Results obtained from a test of a 8% thick full-scale business jet T-tail model with a range of simulated ice shapes conducted in the NASA 40
80 ft2 Wind Tunnel [80].

Fig. 31. An-12 flight test ice accretion [74].

So, the available database consists of two flight test results for not-the-most-critical ice accretions, two sets of conventional wind tunnel test results for relatively thin airfoils (i.e., 8% and 8.7% thick), and two sets for thicker airfoils (11% and 14% thick). Looking first at the two available flight test measurements, these results are helpful in terms of trying to establish a lower limit for realistic maximum lift reductions that can be incurred in flight with critical glaze ice accretions (on thicker lifting surfaces). Certainly, the 45% reduction experienced on the Cessna Centurion for just a 14 in rough ice accretion on the wing LE [103] is most useful in this regard. Similarly, the 31% reduction reported for the An-12 [74] with the ice shape described (see Fig. 31) is also pertinent, especially when the type and size of ice accretion involved are considered. By comparing this relatively small (less than an inch) ice accretion with the ice shapes illustrated in Figs. 28 and 29, it can be seen that the An-12 ice shape and size are indicative of an accretion developed in a relatively short time duration at quite low temperatures (i.e., well below freezing conditions). That being the case, the corresponding penalty associated with a worst case (most critical) ice accretion would be expected to be much larger than 31%. Hence, considering these two available examples, it should not be unrealistic to expect that the lower limit for maximum lift reductions caused by the most critical ice accretions on ‘‘thicker’’ wings at flight conditions would be (as a minimum) close to 50%. Considering next the two sets of high Reynolds number test results from the LTPT obtained for thicker airfoils using simulated large ice shapes, there are again some very useful results/insights provided, but there are also some uncertainties remaining relative to just how well these particular results might represent the penalties that would/could be experienced in flight with the most critical of ice shapes. Starting with the investigation involving the testing of a ‘‘generic’’ glaze ice simulation on the 11% thick airfoil [33], the ice accretion ‘‘simulated’’ was derived from IRT measurements for a 0012 airfoil [125], but the orientation of this accretion

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Fig. 32. Maximum lift penalty for simulated glaze ice accretion [33].

on the 11% thick airfoil LE was ‘‘modified’’ somewhat. The Ref. [125] ice accretion (and associated accretion conditions) along with the simulation selected for the LTPT test are shown in Fig. 32, together with the associated test results. Also shown for comparative purposes are the previously discussed penalties caused by just LE roughness (k=c ¼ 4:5
104 ) on this same airfoil. Looking first at the Ref. [125] ice accretion, although the LWC is significant, the remaining parameters, notably the accretion temperature, duration/ time, speed, angle of attack, and droplet size, are not quite what would be anticipated for a most critical ice accretion condition. However, how all this is impacted by the subsequent reorientation/rotation of this (simulated) ice shape on the 11% thick airfoil that moves the upper horn more toward the upper surface of the airfoil is a big unknown. Whether this reorientation/rotation would simulate a reduced accretion angle of attack and/ or a (more critical) higher temperature (or something else) is certainly debatable. There is also an issue as to how well the simulated (i.e., smoothed) shape with ‘‘an extremely coarse grit’’ added that was selected for this study actually represents the actual ice accretion with its 3-D roughness. And, lastly, there is the ever-present concern that inadequacies in the tunnel sidewall boundary layer control capability could be masking some characteristics (i.e., such as Reynolds number effects). Notwithstanding all of these open issues, the test results obtained are still instructive in providing some insights into the magnitude of penalties that can be experienced with these larger ice shapes. As can be seen in Fig. 32, the maximum lift penalties incurred in this

case with this particular ice accretion are only about 40%, and do not vary much with Reynolds number over the range tested. Relative to the corresponding penalties incurred with the small leading-edge roughness on this same airfoil, it can be seen that this 40% reduction is about half again as much as that experienced with the roughness. The second set of high Reynolds number test results from the LTPT addressing the effects of simulated ice accretions on thicker airfoils, this time for the 14% thick GA NLF airfoil [96,132], are presented in Fig. 33. Results are shown for both 6 and 22.5 min ice accretions, and most importantly, results are shown for both smoothed (2-D) ice accretion simulations and for simulations having the actual (IRT measured) 3-D roughness characteristics represented. Also included are some comparable results from the IRT for the two (3-D) ice accretions. Important lessons to be retained from the test results depicted in Fig. 33 include the following: *

*

Maximum lift reductions of around 50% are experienced for the 22.5 min ice accretion having the 3-D roughness characteristics represented. It is important to mimic the roughness characteristics of these larger ice accretions when defining the simulated ice shapes to be used for either flight testing or testing in conventional wind tunnels. Penalties indicated from the use of smoothed simulations are noticeably lower than those obtained with the 3-D roughness simulated. Unfortunately, it is not known how well the typically used distributed roughness would represent the actual 3-D roughness effects.

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Fig. 33. Maximum lift penalty for actual and simulated glaze ice accretions on GA NLF airfoil [96,132].

*

*

In this case, indicated Reynolds number effects on the maximum lift penalties are small over the range of test Reynolds numbers covered. However, as indicated previously in Section 5, the maximum lift level attained in this case for the baseline un-iced airfoil is 25% higher than that achieved at 1.8 million Reynolds number for the same airfoil as reported in Ref. [93]. Interestingly, however, the iced maximum lift levels for similar shapes are remarkably similar. The (percentage) maximum lift penalties indicated from the IRT tests for the two (rough) ice accretions are similar to those measured in the LTPT with the 3-D roughness simulated, even though the absolute values of maximum lift in all cases are somewhat different.

However, once again, it is necessary to consider some of the limitations inherent with this set of LTPT test results for the GA NLF airfoil in order to place the foregoing observations in proper perspective relative to just how representative the indicated penalties might be of those that would be experienced in flight with the most critical ice accretions. In this case, the concerns are primarily with regard to the size of the model used in the LTPT, and the ice accretion conditions utilized. Rather than the more commonly accepted size having a 22–24 in chord (p0.6 m), such as used in the Ref. [33] investigation, a 36 in chord model was used in this instance, causing the test results to be contaminated with uncorrectable wall interference and nonrepresentative

compressibility effects. Perhaps the most obvious consequence of the excessive wall interference existing with this size model is the very early (nonrepresentative) separation present on the upper surface of the un-iced (baseline) airfoil approaching the TE. Even prior to reaching 41 angle of attack, this separation appears to exist over the last 20% of the chord. This is certainly indicative of a serious flow problem in the sidewall juncture region. Hence, it is not likely that either the indicated baseline or the iced airfoil maximum lift levels and variations with Reynolds number are really representative. How this would impact the magnitude and variation of the indicated penalties is, unfortunately, not known. With regard to the ice accretion conditions selected for this study, the main areas of concern relative to not having the most critical type of ice accretions are the relatively low LWC, droplet size, and flight velocity. The ambient temperature is also a little low as well. Consequently, considering all of the concerns, limitations, and uncertainties associated with the only two available sets of high Reynolds number wind tunnel results for thicker airfoils, it is not really possible to reliably predict the maximum lift penalties likely to be encountered in flight under the most critical icing conditions, but they are probably in excess of 50%. Clearly, some new well focused research is needed to provide more insights into just how large might these penalties be that could be encountered in flight.

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A somewhat analogous situation also seems to exist relative to trying to use the two sets of high Reynolds number wind tunnel test results available for ‘‘large’’ ice accretions on ‘‘thinner’’ lifting surfaces to estimate penalties likely to be encountered with these surfaces under the most critical icing conditions. That is because once again, in spite of the fact that there are some very useful insights enabled by both sets of test results, there are, unfortunately, aspects of each that tend to limit their general applicability. Starting with some of the results obtained and provided by Addy [133] from the test of the 8.7% thick ‘‘business jet’’ airfoil in the LTPT, representative maximum lift penalties for three different ice accretions simulations are shown in Fig. 34. The three ice shapes simulated were castings of 2 and 22.5 min accretions formed in the IRT at a total temperature just below freezing (i.e., glaze icing conditions), and a casting of a 16.7 min rime/intermediatetype accretion formed at 111C (11.71F) in the IRT. Corresponding results from the tests of this airfoil in the IRT with the 2 and 16.7 min accretions are also provided in Fig. 34 for comparative purposes. Important points to be noted from these results are as follows: *

There is a very significant Reynolds number effect on the indicated penalties, with the lower Reynolds number results obtained at 3.5 million being well below those obtained at the higher Reynolds numbers. This effect, which is most noticeable with the two shorter duration accretions, is due entirely to changes on the baseline un-iced airfoil.

*

*

*

The largest penalty measured with this relatively thin airfoil was just over 50% with the 22.5 min accretion at a total temperature just below freezing. Ambient temperature is a much more important factor in establishing the magnitude of these penalties than the duration of the icing encounter, as evidenced by the similarity of the results for the 2 and 16.7 min accretions, and the big difference between the 16.7 and 22.5 min accretions. Maximum lift penalties (percentagewise) indicated from the IRT tests for the two (smaller) ice accretions tested there are again similar to those measured in the LTPT, but, also once again, the absolute values of maximum lift differ significantly.

When trying to put these results into perspective, however, it is necessary to keep in mind that these results were also obtained with a 36 in chord model in the LTPT, and, as a consequence, are most likely somewhat contaminated with uncorrectable wall interference effects like the previously discussed results for the GA NLF airfoil. The same is likely true of the IRT test results as well. Hence, while perhaps the trends indicated are helpful, the absolute values/levels are almost certainly in error. Results obtained from the test of the 8% thick fullscale business jet T-tail model [80] fit in very well with the business jet airfoil test results in a number of regards. Three basic ice shapes were assessed in this program. One was an ice-shape casting of a 1.6 min accretion from the IRT, while the other two were LEWICE predictions

Fig. 34. Maximum lift penalty for actual and simulated ice accretions on ‘‘business jet’’ airfoil [133].

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Fig. 35. Maximum lift penalty for simulated glaze ice accretions on ‘‘business jet’’ T-tail model [80].

for 9 and 22.5 min accretions at a total temperature of 2.51C. Both smooth and rough (with a k=cE24
104 ) versions of these two larger accretions were tested during this program. Also included were spoiler-type protuberances intended to approximate the effect of the two larger ice accretions. Pertinent results from this program are displayed in Fig. 35, from which the following observations can be made: *

*

*

*

Similar to the business jet airfoil test results, there is a very significant Reynolds number effect on the indicated penalties. In this case, there is a very noticeable drop off in the indicated penalties below about 4–5 million. However, in this case, there are variations with both the un-iced baseline and the ice shapes with Reynolds number that contribute. Similar to the results obtained with the GA NLF airfoil [96,132], penalties indicated from the use of smoothed simulations are definitely lower than those obtained with roughness ‘‘simulated’’. The largest penalty observed with the (largest of the) simulated ice shapes was a little less than 40%, somewhat lower than that seen with the business jet airfoil test results. The attempt to approximate the effect of the 22.5 min ice shape simulation with a spoiler-type protuberance was partially successful in that it resulted in about the right order-of-magnitude penalty, but not close enough to be really useful.

Once again, however, there are aspects of this test program which limit the usefulness and/or general applicability of these particular results. For one, the use of predicted ice shapes rather than actual measured ice shapes for these glaze ice accretions raises doubts because of the significant differences observed in a number of cases between predicted and actual ice shapes such as illustrated in Ref. [131]. Also in this case, since increased Reynolds numbers are accompanied by an increased freestream Mach number, there is evidence that compressibility effects, primarily on the baseline uniced tail, limit the maximum penalty observed at the higher Reynolds numbers. Thus, for one reason or another, whether it be not having the thought-to-be most critical ice accretions, or being concerned with the simulations utilized, and/or having reservations about the test arrangement, conditions, etc., the foregoing six sources of high Reynolds number test results available for these ‘‘larger’’ ice accretions do not enable a meaningful determination to be made of the likely upper limit of maximum lift penalties that might be experienced at flight conditions for any of the test geometries, let alone a wider range of ‘‘thick’’ and ‘‘thin’’ lifting surfaces, other than the previously stated estimate that they are probably in excess of 50% for the thicker surfaces. Unfortunately, the results from numerous (additional) low Reynolds number test programs conducted over many years which focused on this broad class of larger ice accretions

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provide little in the way of additional insights into the potential magnitude of the upper limit of these penalties other than to perhaps reinforce the >50% estimate for thicker surfaces. This can be seen by examining the synopses of many of these test programs that are provided in (generally) chronological order in Tables 1 and 2. In each case, the subject lifting surface, the targeted ice accretion conditions, the simulated ice shapes utilized, and the aerodynamic test conditions and results (maximum lift penalties and stall angle reductions) are documented if they were available/ published. Incidentally, all of the illustrations of the ice shape simulations have the suction side of the lifting surface shown as the upper surface whether the results are for a wing or tail section. And, further, any place there is an asterisk signifies that there are some extenuating circumstances involved that need to be noted in order to place the indicated results in proper perspective. When examining the synopses/results provided in Tables 1 and 2, there are a number of critical observations to be made. To start with, in addition to the ever present concern over the likelihood that such low Reynolds number test results will not properly simulate higher Reynolds number conditions, probably the biggest concern with all the results shown is not having the most critical ice accretion conditions possible. Starting with the premise that the most critical ice accretion conditions involve temperatures just below freezing, high LWCs, maximum icing exposure durations/times, higher speeds (with corresponding lower accretion angles of attack), and larger droplet sizes, there just are not any results provided for these combinations of conditions. Other significant observations include the following: *

*

Whenever a direct comparison can be made between results obtained with both LEWICE-based glaze ice simulations and IRT-based simulations, the LEWICE simulations were not as detrimental. There are two examples of this shown in Table 2. It is of interest to note that in one of these cases [147], the penalty incurred with the 45 min LEWICE-based glaze ice simulation was only slightly more than that observed with a much shorter duration IRT intercycle ice accretion. Having essentially the same penalties for noticeably different glaze ice simulations such as occurred with the LEWICE-based and ‘‘S&C’’ glaze ice simulations evaluated in Refs. [81,101] as well as the 22.5 and 45 min glaze ice simulations studied in Ref. [146], it might be tempting to conclude that the details of the glaze ice simulations are not that important. However, the Ref. [147] results tend to dispel that notion indicating that there are some aspects of these glaze ice simulations that are definitely more critical than

*

*

others. Incidentally, the asterisk shown by the results for the first ice accretions shown from Refs. [126–128] is because some sublimation of the ice accretion had occurred before the aerodynamic test measurements could be completed. In contrast to some of the previously discussed high Reynolds number test results [80,96,132] which indicated that simulation of the surface roughness characteristics of these larger ice accretions is important (because they are detrimental), the results from these low Reynolds number tests are certainly mixed in that regard. There are cases where there is no penalty indicated for adding roughness [134,135,136,139–141,146], cases where there is a relatively small adverse effect indicated [143,146], and even cases where adding roughness had a favorable effect [136]. However, based on the conclusions reached in Section 4 that low Reynolds number indications of roughness effects are often not realistic, the recommendation is to largely discount these low Reynolds number results, and rely more on the high Reynolds number indications. And, speaking of roughness, the asterisks shown with the Ref. [146] results having a roughness with k=c ¼ 4
104 are a reminder that these results were obtained with so-called ‘‘loose grit’’ rather than with sandpaper as is more common. There is a big difference indicated between the rime ice results obtained in Refs. [134,135] and those in Ref. [136], with a noticeable penalty for the former and an improvement (i.e., increased maximum lift) in the latter case. These differences occur with rather subtle differences in the ice shapes, leading to the suspicion that there may be some low Reynolds number ‘‘anomalies’’ involved here.

In addition to this rather hodgepodge of low Reynolds number test results obtained with a variety of simulated ice shapes, there have been two somewhat systematic low Reynolds number investigations conducted fairly recently using protuberances to ‘‘simulate’’ glaze ice horns in an attempt to identify the most detrimental features of these glaze ice accretions. One was a study conducted by Papadakis et al. [149,150] using spoiler-type devices on a NACA 0011 airfoil to address horn angle, location and horn height effects. Pertinent findings from this study are summarized in Fig. 36. In this case, two parameters are used to describe the ‘‘maximum lift’’ penalties caused by these devices since their use resulted in what is sometimes referred to as a ‘‘long bubble stall’’, where the lift curve slope decreases very noticeably at some angle of attack, but the lift still continues to increase somewhat beyond that. Therefore, penalties based on both the maximum lift level actually achieved as well as the lift level where the lift curve slope decreases are used to describe the effects

F.T. Lynch, A. Khodadoust / Progress in Aerospace Sciences 37 (2001) 669–767 Table 1 Additional low Reynolds number test results addressing maximum lift penalties for ‘‘larger’’ ice shapes

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Table 2 Further low Reynolds number test results addressing maximum lift penalties for ‘‘larger’’ ice shapes [75,138,142,144,145]

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Fig. 36. Maximum lift penalties for spoiler glaze ice simulations on NACA 0011 airfoil [149].

of these spoiler-type devices. It can be seen in Fig. 36 that based on the penalty utilizing the maximum lift demonstrated, there is not a strong effect of the upper horn angle except for the extreme case with the longer horn in the vertical (901) position. However, if the penalty based on the lift level where the lift curve slope decreases very noticeably is used as being more descriptive of the true adverse effect, then there is a significant influence of upper horn angle, with larger angles (i.e., more vertical displacement) being significantly more harmful (as would be expected). In this case, penalties well in excess of 50% are indicated, even for quite moderate horn angles. This strong influence of upper horn angle also seems to be apparent in the second set of results, those obtained by Kim and Bragg [151,152], where thicker protuberances were used to study the effects of upper horn variations on the aerodynamic characteristics of an NLF 0414 airfoil. Interestingly, in this case, with the thicker protuberances and thicker airfoil, there were no complications involved due to any long bubble stall in defining the maximum lift penalties. Results from this investigations are illustrated in Fig. 37, where it can be seen that the strong influence of upper horn angle on the maximum lift penalties incurred seems apparent. However, in this case, this observation must be qualified knowing that horn angle increases were also accompanied by some aft movement of the horn (as shown in Fig. 27). Consequently, the trend shown in terms of increased penalties with increasing horn angle is likely a combination of effects. Nevertheless, with the very

modest aft movement involved, at least up to horn angles of 401, it is likely that the predominant factor in the increased penalties observed is the increased horn angle. Now, with these results in hand, especially regarding the adverse effect of increasing the upper horn angle, and keeping in mind that having higher accretion speeds, LWC, and droplet sizes, as well as temperatures closer to freezing, will result in increased upper horn angles (and sizes), it is instructional to go back and reexamine some of the low Reynolds number results summarized in Tables 1 and 2, as well as the previously discussed high Reynolds number test results. For example, in Tables 1 and 2, it can be seen that in most of the cases where the largest maximum lift penalties were experienced, such as the LEWICE-based glaze ice shape used in Refs. [81,101], there is a very prominent horn angle involved. Also, a reduced horn angle would seem to explain why a lower penalty was seen with the LEWICE-based simulation used in Ref. [147] compared to the IRT-based glaze ice shape. Consequently, based on all of the foregoing high and low Reynolds number test results available at this time, it is quite apparent that the maximum lift penalties likely to be encountered in flight by thicker single-element lifting surfaces under the most critical icing conditions are well in excess of 50%. However, just how high they are, as well as whether they are as large (or larger) than those encountered with the potential runback/ridge ice accretions addressed in Section 5, will remain unknown until such time as some well conceived high Reynolds number test programs are carried out for both types of ice accretions.

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Fig. 37. Maximum lift penalties for protuberance simulations of glaze ice accretions on GA NLF airfoil [151].

6.1.2. Single-element’’ lifting surfaces with trailing-edge control surfaces It should be understood (without saying) that trying to get a handle on the upper limit of maximum lift penalties that could be incurred by ‘‘single-element’’ wings and tails with trailing-edge control surfaces deflected at flight conditions is fraught with all the same difficulties preventing such an assessment being made for configurations not having control surfaces. And, to compound the situation even further, there are not that many sources of applicable test results available either. However, as was the case when examining these characteristics for the effects of initial leading-edge ice accretions (in Section 4), the trends indicated from the results available do provide some helpful insights. Looking first at the test results available for thinner lifting surfaces (i.e., those having un-iced maximum lift coefficients o1.0 without control surfaces deflected), there are test results available for four configurations (lifting surface/simulated ice accretion combinations), including one set of high Reynolds number data for the previously discussed 8% thick business jet T-tail configuration [80]. These results, along with low Reynolds number results for the same configuration [153], and three other tail configurations [77,78,154,155], are shown in Fig. 38. Incidentally, ice shape simulations utilized in these investigations are illustrated in Table 1 for Refs. [77,78], and in Fig. 35 for Refs. [80,153].

Important items to be noted from the results summarized in Fig. 38 are as follows: *

*

*

As has been noted before, there can be big differences in the penalties incurred between configurations with these thinner lifting surfaces (at low Reynolds numbers). While both the high and low Reynolds number test results for the business jet tail models indicate penalties which become quite substantial, especially at the critical elevator trailing-edge-down conditions, the low Reynolds number results for the other three indicate quite small penalties at the conditions investigated. Analogous to some of the results shown earlier in Fig. 11 for initial leading-edge ice accretions, both sets of test results for the business jet tail with the 22.5 min LEWICE glaze ice shape show that the maximum lift penalty (percentagewise) increases significantly as the elevator deflection moves toward becoming more trailing-edge down. This occurs as a consequence of a relatively constant incremental maximum lift penalty becoming a larger percentage of an un-iced maximum lift level that is decreasing as the elevator is deflected in this direction (opposing the download). An example of some of the difficulties involved in trying to interpret low Reynolds number test results is provided by the two sets of low Reynolds number

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Fig. 38. Variations in maximum lift penalty with trailing-edge elevator deflection for thinner surfaces with large ice shapes.

results obtained for the business jet tail model without the elevators deflected. In this case, the penalties indicated for the (rough) 22.5 min LEWICE ice shape at a Reynolds number of 1.36
106 are conspicuously different for the results from the test of the full scale model [80] versus those from the test of the 25% scale model, with the penalty indicated from the test of the 25% scale model being about a third larger. And, even more important, both sets of low Reynolds number test results indicate penalties less than those indicated at higher Reynolds numbers. Turning next to the corresponding penalties incurred by thicker lifting surfaces, the only known applicable test results available appear to be low Reynolds number results for an NLF-0414 airfoil [148] and three tail geometries, namely the thickened LE design from Refs. [77,78], and the Twin Otter [81] and F-28 [110] horizontal tails. Results from these four investigations are summarized in Fig. 39, while the simulated ice accretions utilized in these investigations are all depicted in Tables 1 and 2. Incidentally, the results shown for the NLF-0414 airfoil in Fig. 39 are presented as though it were a tail configuration in terms of the control surface deflection terminology utilized. It is also important to note that many of the test results illustrated in Fig. 39 are ‘‘flagged’’. This signifies that the actual penalties could be higher than indicated because the angle of

attack range employed in the tests of the un-iced baseline geometry was not large enough to ensure that the maximum lift level had been attained. Finally, in examining the test results summarized in Fig. 39, the following points should be noted:

*

*

*

Some very large penalties, exceeding 80%, were encountered with the LEWICE glaze ice shape employed in the tests of the Twin Otter tail [81] at the critical trailing-edge-down elevator deflections. And, as indicated, the actual penalties could be even a littler larger. When there were variations in test Reynolds number, such as with the Twin Otter test results, the penalties indicated at the higher Reynolds numbers were generally larger. However, once most of the tail lifting capability has been lost because of these ice accretions, changes to the un-iced baseline maximum lift level have little impact on the indicated (percentage) penalties. At these low Reynolds numbers, the penalties indicated at the trailing-edge-down elevator deflections for the inter-cycle ice accretions are only about half of those incurred with the larger glaze ice accretions. Whether this relationship would hold at higher flight Reynolds numbers is another question however.

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Fig. 39. Variations in maximum lift penalty with trailing-edge elevator deflection for thicker surfaces with large ice shapes.

Consequently, based on the limited applicable test results available, indications are that maximum lift penalties caused by these larger ice accretions on tails at trailing-edge-down elevator deflections (or on wings with trailing-edge-up aileron deflections) can become quite large. Therefore, similar to the prevailing situation with single-element lifting surfaces, some well conceived high Reynolds number test programs utilizing the most adverse ice accretions likely are needed to better document just how large these penalties can be in flight. And, perhaps even more important, existing efforts aimed at educating pilots and flight crews of the dangers involved in extending wing flaps for landing if there is any chance that ice has accreted on the tail must be continued. Although this danger is and has been known by many, as evidenced by the rather ‘‘subtle’’ quote of ‘‘the famous old Swedish pilot, Count von Rosen’’ referred to in Ref. [156] that ‘‘there can not be such an idiot as to use full flap if there is the slightest risk for ice on the tailplane’’, it must be known and remembered by all. 6.1.3. Multi-element lifting surfaces Once again, in what has become an all too familiar theme, there is very little really applicable data available

addressing the possible range of maximum lift penalties which could be caused by ‘‘large’’ in-flight ice accretions on multi-element high-lift wings. However, be that as it may, the available results, if viewed in the framework of lessons already learned, do enable a good understanding of important trends and relative criticality for different geometries. Addressing first high-lift systems without any leading-edge devices, i.e., the so-called ‘‘hard’’ LE designs, what is available is four sources of low Reynolds number test results. The earliest of these were the results reported by Johnson [85] in 1940 wherein the incremental loss in maximum lift capability caused by the ice accretions evaluated (see Table 1) was essentially the same with the flaps either extended or stowed. This resulted in the maximum lift penalty being reduced from about 37% with the flaps stowed to 22% with them extended, which is the inverse of the ratio of the un-iced maximum lift coefficients. This is the same phenomenon seen earlier in Section 4 with leading-edge roughness effects on multielement geometries. Not surprisingly, however, this exact relationship does not always prevail as evidenced by the results from the other three available (low Reynolds number) data sources [74,135,148], but the general trend of reduced (percentagewise) penalties with

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Fig. 40. Variations in maximum lift penalty for high-lift configurations with ‘‘Hard’’ LEs.

the flaps deflected does prevail, especially at higher flap settings appropriate for approach/landing conditions. This can be seen by examining the results from these other three sources illustrated in Fig. 40. Incidentally, the ice accretion simulations employed in these investigations were portrayed in Tables 1 and 2. And, in all of these cases, the use of simulated ice shapes was limited to the main airfoil LE. Any ice accretions on the flap LE were not considered. It is also interesting to note that the results displayed in Fig. 40 from Ref. [148] are the same as those used in Fig. 39. This acts as a reminder that the test results for ‘‘single-element’’ lifting surfaces with trailing-edge control surfaces shown in Figs. 38 and 39 indicating reduced penalties for negative elevator deflections (i.e., aiding the tail download) provide further evidence of the trend of having reduced (percentagewise) penalties with flap extension. However, once again, analogous to the situation existing with single-element lifting surfaces (including those with trailing-edge control surfaces), additional appropriate experimental studies are needed to determine just how large these ‘‘reduced’’ penalties might be for the most critical ice accretions possible at flight conditions. This means that the testing must be done at higher, more representative (of flight) Reynolds numbers as evidenced by the fact that the penalties derived from the Ref. [148] test results are most likely unrealistically low because flow separation was present on the flap of the un-iced

baseline even at 51 flap deflection at the low test Reynolds number. Looking next at the limited amount of applicable data available addressing the effects of ‘‘larger’’ ice accretions on high-lift systems having leading-edge devices, the trend of reduced maximum lift penalties continues, although there are again the seemingly ever present concerns about probably not having the most critical ice accretions and/or test conditions represented. Most of the data available in this area comes from two low Reynolds number investigations. These are the results contained in the first report of the Swedish-Soviet Working Group [74] for a geometry with a vane flap at a landing position, and the previously mentioned results obtained by Potapczuk et al. [126–128] in the IRT with a Boeing 737-200 wing section/high-lift system with flap settings appropriate for holding or initial approach conditions. Illustrations of the high-lift geometries and the simulated ice shapes employed in these two investigations are shown in Figs. 41 and 42, respectively, while the ice accretions conditions and the test results are summarized in Table 3. In examining the earlier set of results [74], it can be seen that the maximum lift penalties indicated for this landing configuration are all less than 25%. Also evident in these results are indications that the more critical ice accretions are those occurring closest to freezing

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Fig. 41. Four-element high-lift geometry and ice ‘‘imitators’’ [74].

conditions, and those formed at higher angles of attack. The indicated temperature effects are important as they are clearly consistent with a number of other findings, but the indicated accretion angle-of-attack effects are likely of little practical significance. To start with, it can be seen from Fig. 41 that these higher accretion attitudes are the ones which result in the more detrimental ice accretions that form on the LEs of the main element and flap vane. However, the angles of attack involved correspond to operating speeds very (too) close to stall. In this case, more realistic accretion conditions corresponding to typical speed margins (to stall) would be more like a couple of degrees or less. And, at these attitudes, such detrimental accretions do not appear to occur. That being the case, it would be concluded from these results that ice accretions on the slat are realistically the only ones to be concerned with. Some very similar conclusions, trends, etc. are also evident from the second set of low Reynolds number test results [126–128]. For one, even though the un-iced maximum lift levels existing with the non-landing type flap settings addressed in this test are well below that seen in Ref. [74] (with a landing flap setting), still, the highest penalties observed with the ice accretions investigated are around 25% as well. Also, it can be seen in Table 3 that once again the most detrimental accretions (with everything else being equal) are those occurring closest to freezing temperatures, and at the higher angles of attack (which are probably not that representative for holding conditions, etc.). Hence, once again, and especially for the lower flap settings where the slat LE is sealed against the main element upper surface, and there are very small flap gaps, the quite obvious

conclusion is that ice accretions on the slat are of prime importance. Other than these two 2-D low Reynolds number sets of test results, the only other known/available sources of test results addressing the effects of these ‘‘larger’’ ice accretions on multi-element high-lift systems with leading-edge devices are from a very low Reynolds number test of a complete 737 model at a landing flap setting [157], and from a 2-D high Reynolds number test of some relatively small simulated ice accretions on a previously introduced (in Sections 2 and 3) three element high-lift geometry, also in a landing configuration [158]. Results from the 737 model test indicate a maximum lift loss of approximately 15% caused by a simulated ice accretion on the ‘‘wing LE’’. Unfortunately, however, because of a lack of specifics provided regarding the icing simulation used, together with the quite low Reynolds number for a really meaningful 3-D high-lift test addressing maximum lift characteristics, it is difficult to make use of this result. The 2-D high Reynolds number test results do, however, provide an interesting insight into Reynolds number effects on the relative contributions of ice accretions on the slat versus those on the downstream elements, even though the icing simulations used are certainly not the most detrimental possible (especially on the slat). Results from this particular investigation are provided in Fig. 43, where it can be seen that while there is very little Reynolds number effect over the range investigated (from 5 to 16 million) when the subject (rough) ice accretions are simulated on all three surfaces, there is a very definite Reynolds number influence on the results when only the main element and flap ice accretions are

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Fig. 42. Five-element high-lift geometry and IRT ice accretions [126–128].

modeled (representing the use of anti-icing on the slat). The results shown also reinforce previous findings regarding the need to simulate the roughness characteristics of the ice accretions being modeled. This would be especially critical on smaller, smoother shapes. And, most importantly, the results again very clearly illustrate the dominant criticality of any ice accretions allowed to form on the slat versus those formed on downstream elements. Further, the results would appear to indicate that ice accretions on high-lift configurations with leading-edge devices are not very important if the LE device is effectively anti-iced. However, to put these results into perspective for failure cases, or if the leading-edge device is not anti-iced, the following need to be considered: *

The maximum lift penalty indicated by these tests of around 10% with ice on the slat is not much more

*

than seen earlier (in Fig. 12) for an approach/landing geometry at high Reynolds numbers with just roughness on the slat. By comparing the simulated ice accretion on the slat used in this investigation with others observed in an icing tunnel test with this same high-lift geometry (see Fig. 1 in Section 2), it is clear that what was tested here is certainly not the worst case possible.

Thus, based on the data available addressing maximum lift penalties caused by ‘‘larger’’ ice accretions on high-lift systems having LE devices, the largest penalties encountered to date are around 25% if ice is allowed to form on the extended leading-edge device. This is noticeably less than seen for high-lift geometries not having leading-edge devices, and clearly much less than incurred by single-element geometries. If ice is not allowed to form on the leading-edge device, then any

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Table 3 Low Reynolds number test results addressing maximum lift penalties caused by ‘‘Larger’’ ice shapes on multi-element high lift configurations with leading-edge devices

Fig. 43. High Reynolds number test results for ice accretions on multi-element airfoil [158].

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penalties caused by (realistic) ice accretions on downstream elements would appear to be relatively minor. However, because of well-founded concerns that the most detrimental possible ice accretions were likely not addressed in these investigations, some well conceived new studies are needed to establish the upper limit of these penalties, both with and without the LE device ice protected. As part of these studies, attention needs to be focused on establishing just what are the most critical (and realistic) ice accretions which can form on the downstream elements. Items needing to be addressed include: *

*

*

*

Establishing incidence angles consistent with realistic flight operating conditions. Considering higher LWCs, temperatures close to freezing, and possibly longer exposure times. Addressing a range of droplet sizes, especially larger ones. Assess the impact of variations in aircraft flight path (i.e., descending for approach/landing, and ascending for landing climb or takeoff climb).

It would be expected that current state-of-the-art CFD capabilities for multi-element airfoil geometries [159] would certainly be most helpful for a number of these studies, especially for addressing droplet trajectory issues. 6.2. Stall angle reductions In reviewing the database available pertaining to the stall angle reductions caused by a wide range of ‘‘larger’’ ice accretions on a variety of aerodynamic lifting surfaces, it is very important to keep in mind all the

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concerns expressed regarding the limitations of the available database for maximum lift penalties, because all the same concerns obviously apply here as well. These include likely not having the most critical ice shape/accretion possible, not having very many good quality high Reynolds number test results, having inappropriate model/tunnel interactions, etc. However, once again, while the database available may not be appropriate for determining with any degree at certainty the largest stall angle reductions possible, it is useful for examining potential trends, especially for the ‘‘iced’’ geometries. 6.2.1. Single-element lifting surfaces Consistent with the very large range of maximum lift penalties caused by these larger ice accretions, there is likewise a broad range of stall angle changes indicated in this category from the data available, including some cases with simulated rime ice accretions where the ice shape enables an increase in the stall angle of attack. Starting with the existing database for 2-D airfoil geometries, but not including test results where there was an increase in the stall angle of attack because they are obviously not relevant when trying to identify the worst/most critical consequences, the available data are presented in Fig. 44 as a function of the corresponding incremental loss in maximum lift capability. It can be seen that stall angle reductions up to (at least) 131 have been observed, which is similar to the level encountered to date with the simulated runback/ridge ice accretions addressed in Section 5, and about twice that seen on 2-D airfoils with the initial leading-edge-roughness-type ice accretions reviewed in Section 4. While there is the expected general correlation of the measured stall angle

Fig. 44. Stall angle reductions caused by larger ice accretions on 2-D single-element airfoils.

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reductions with the incremental (not percentage) loss in maximum lift (coefficient) capability, there are also some significant variations within this general trend as well. For reference, also shown in Fig. 44 is a lift curve slope of 0.1, which would correspond to the 2-D case in the linear range (i.e., before the roundover approaching stall). Having a preponderance of test results falling above this reference line would, for one, seem to indicate that there is typically more roundover in the lift curve prior to stall with the un-iced baseline cases than there is with the airfoils having simulated ice shapes. This is certainly a characteristic which might well be more prevalent in low Reynolds number test results, and/or in 2-D results obtained without any sidewall boundary layer control. Another potential contributor to some of the variations seen could be having uncorrectable wall interference effects. In assessing the database summarized in Fig. 44, it is also important to note that there is only one set of high Reynolds number test results available/included [96,132], and these are believed to be contaminated with uncorrectable wall interference effects caused by having a model too large for the test facility. This lack of reliable high Reynolds number test results is an important limitation in that low Reynolds number test results are often not representative, both in terms of maximum lift levels achieved, as well as with regard to stall characteristics. With regard to maximum lift levels achieved, it should also be noted in Fig. 44 that the largest stall angle reductions displayed come from

test results [81,101] where there is no assurance that the maximum lift level of the un-iced baseline airfoil was reached. Hence, the actual stall angle reductions (at these low Reynolds numbers) could well be larger than indicated. However, the real question to be answered is how large might these stall angle reductions be with the most critical ice accretions at higher Reynolds numbers more representative of flight, and without being encumbered by uncorrectable wall interference (floor and ceiling as well as sidewall viscous effects). There is a similar scarcity of reliable results available for assessing stall angle reductions incurred with larger ice accretions on 3-D wings and tail. What is available is shown in Fig. 45. By comparing these results with the 2-D results in Fig. 44, it can be seen that with this collection of data, the stall angle reductions incurred with the 3-D geometries follow the same general trend as the 2-D results, but are, in general, not any larger than those seen with the 2-D configurations. This is contrary to what would normally be expected with the lower lift curve slope typically associated with these finite aspect ratio geometries. It is also contrary to what was seen earlier (in Fig. 15) with the data available for initial leading-edge ice accretions. The cause of this apparent disparity for most of the results probably lies with the low test Reynolds numbers involved since low Reynolds number anomalies are often even more prevalent with 3D maximum lift characteristics. And, while there are two sets of ‘‘high’’ Reynolds number test results included in

Fig. 45. Stall angle reductions caused by larger ice accretions on 3-D single-element wings and tails.

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Fig. 46. Stall angle reductions caused by larger ice accretions on tails with elevators deflected.

Fig. 45, the one flight test data point for the An-12 [74], and the test results for the 8% thick full-scale business jet T-tail model [80], neither configuration is really that appropriate for establishing more general trends. The An-12 has a fairly thick wing, and the ice accretion involved was certainly not at all representative of a critical accretion, while the business jet tail model had an un-iced maximum lift coefficient of only 0.94 at the test Reynolds number. So, the ‘‘well conceived’’ high Reynolds number testing recommended to better establish the upper limit of maximum lift penalties that can be caused by the most detrimental large ice accretions is likewise needed to establish the upper limit of corresponding stall angle reductions feasible at flight conditions, since that cannot be done now with data presently available. 6.2.2. ‘‘Single-element’’ lifting surfaces with trailing-edge control surfaces The available and (somewhat) applicable test results for these configurations with control surfaces deflected for both 2-D and 3-D geometries are illustrated in Fig. 46. Again, unfortunately, there is only one set of high Reynolds number test results available, and these are for the full-scale business jet tail model with the low un-iced maximum lift level. Overall, the results appear to be very similar to those displayed in Figs. 44 and 45 for single-element geometries without such control surfaces deflected. And, there does not appear to be anything out of the ordinary happening with the critical trailing-edgedown elevator deflections. The only results indicating that stall angle reductions might be noticeably greater for these deflections are the Ref. [153] very low Reynolds number data. But, the Ref. [80] high Reynolds number data for the same business jet tail geometry tend to dispel that notion, and, in the process, provide yet

another example of the inadvisability of relying on low Reynolds number test results. So, to overcome the existing state of affairs where the available database for these characteristics is predominantly from low Reynolds number testing, appropriate new high Reynolds number testing of representative geometries with the most critical ice accretions feasible is again needed. 6.2.3. Multi-element lifting surfaces A very similar situation exists with the database available for these geometries in that there are very few reliable high Reynolds number test results, and, typically, the ice accretions evaluated are not for the most detrimental icing conditions possible. For example, the only high Reynolds number test results available [158] fall in this category of simulating less-than-themost-severe-possible ice accretions. However, again, these results are useful in identifying likely trends. In examining the existing database which is summarized in Fig. 47, it can be seen that there tends to be more results below the 0.1 lift curve slope than seen with the previous two types of lifting surfaces. What is noteworthy is that these data points below the reference slope are mostly for configurations having LE devices (extended), and those above the reference slope are mostly for ‘‘hard’’ LE designs (similar to the previous lifting surfaces). This indication that configurations having LE devices will have proportionately significantly smaller stall angle reductions needs to be further explored (at high Reynolds numbers with more critical ice accretions). Another important observation to be made from the data shown in Fig. 47 is with regard to the Ref. [158] results representing a case with slat anti-icing activated. In this case, even though there is a reduction in the maximum lift level caused by having simulated ice accretions on the LEs of the main element (behind the

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Fig. 47. Stall angle reductions caused by larger ice accretions on multi-element high-lift configurations.

slat) and flap, there is not a reduction in stall angle, once again indicating that the most consequential ice accretions would be those which could be allowed to form on the slat. 6.3. Drag penalties It is helpful when sorting through the large database of test results available addressing the drag penalties which can be caused by these larger ice accretions to keep in mind some of the lessons already learned regarding leading-edge geometry/size effects, what constitutes the most critical icing conditions in terms of causing the largest penalties, concerns over the effectiveness of simulated-versus-actual ice accretions, the importance of simulating roughness characteristics, as well as issues with the performance testing conditions in terms of Reynolds number, wall effects, etc. For example, as pointed out previously, the most critical ice accretions have been found to be those which are formed at total temperatures just below freezing, and additionally, have a high LWC, larger droplet sizes, longer icing durations, and higher aircraft speeds and lower angles of attack (while the ice is being accreted). Hence, test results obtained at less critical conditions are of no real use in establishing the upper limit of drag penalties which might be encountered, nor are test results presented without the associated accretion conditions identified. Also, because of differences

reported in drag penalties measured with simulated ice shapes compared to the penalties seen with actual icing tunnel ice accretions, it is appropriate to be somewhat skeptical of test results obtained using simulated ice shapes other than those which are castings of actual ice accretions. This would be particularly true for ice shape simulations not having the roughness simulated (or defined). And, further, errors or misinterpretations likely to be introduced by the use of low Reynolds number test results need to be factored into the assessment, especially so when considering drag penalties expressed in terms of percentage increases above the measured baseline un-iced drag levels. Caution also needs to be exercised when assessing/interpreting icing tunnel results obtained with small and/or thin lifting surface models where the ice accretions are most likely unrealistically large for representing more typical larger and/or thicker (flight) geometries. Another very important factor needing to be considered when assessing drag penalties caused by these larger ice accretions is to recognize that drag penalties for a given ice accretion will be different at different operating conditions/speeds (i.e., lift coefficients). Unfortunately, all too often, test results are presented at a fixed or different angles of attack, or as a function of angle of attack, when, in reality, what is most needed is to be able to assess drag penalties for various operating conditions (i.e., reference speeds). Regrettably, this is not possible in many cases, either because of insufficient drag

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Fig. 48. Variation in drag penalty with operating conditions for NACA 0012 airfoil with simulated IRT glaze ice shape [139–141].

measurements, and/or not having a reliable basis for establishing appropriate reference speeds (i.e., based on the un-iced stall speeds). This is especially true of many of the earlier studies in the IRT [119–124] where, because of the very large size of the models, stall speeds were not established (and would not have been representative even if measured). An example of the variations in the incremental drag penalties which can occur with a given large ice shape at different operating conditions is provided in Fig. 48 using the 2-D test results published by Bragg et al. [139– 141] for a 0012 airfoil with a glaze ice shape simulating an IRT accretion which occurred at 41 angle of attack. The ice shape and accretion conditions simulated were previously documented in Table 2. Although these particular test results were obtained at a relatively low Reynolds number (1.5 million), they are nevertheless quite useful in that the resulting maximum lift penalty was substantial (54%), and the un-iced baseline profile drag levels/variations are sensible (which is definitely not always the case with the total available database). It can be seen from Fig. 48 that there is a large variation in the incremental drag penalty which occurs over the range of possible operating conditions (i.e., speed margins to stall with the un-iced baseline). The penalty increases quite rapidly as operating speeds are reduced approaching the (increased) stall speed associated with the ice accretion. This is very important, especially if the flight speed has been reduced for climbing, approach, etc. It can also be seen from Fig. 48 that assessing drag penalties at a constant angle of attack (accretion condition) such as

was done with the Ref. [125] test results shown previously in Fig. 28 results in an under-prediction of the incremental penalty which would occur at a constant aircraft speed. These results also lead to the conclusion that the most detrimental (largest) drag penalties are those which will occur after ice has accreted at high(er) speed/lower angle of attack conditions, and the aircraft is then slowed for climb, approach, etc. In addition to having larger ice accretions at higher speeds, etc., more of the accreted ice forms on the upper surface of the wing at lower angles of attack. Finally, the results shown in Fig. 48 provide some basis for adopting the approach that there are two components of the drag penalty, one being a parasite-drag-type increment which stands out at higher speeds/lower angle of attack conditions (and is probably a strong function of the glaze ice horn height and angle), and a second component which varies with lift and becomes predominant at flight speeds approaching the iced-configuration stall speed. It is also important to note that while there are a number of (previously mentioned) flight test results reported documenting drag increases experienced from flight in natural icing conditions, these results are not very helpful for assessing the upper limits of drag penalties possible. To start with, results reported include the contributions of (undocumented) ice accretions on aircraft components in addition to the wing and/or tail (i.e., landing gear, fuselage, engine cowling, protuberances, etc.) But, much more important, it would appear that, for obvious crew safety considerations, less-than-the-most-critical icing and operating

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conditions were involved in all the reported cases. For example, with the previously quoted case where the Twin Otter experienced a 75% total aircraft drag increase (with a glaze ice shape on the wing), the corresponding wing section drag increase reported was only 120%. This is almost an order of magnitude less than the section drag increases seen in Fig. 28 with the Ref. [125] test results, and those just seen at some conditions in Fig. 48 with the Refs. [139–141] results. 6.3.1. Single-element lifting surfaces Development of a methodology for predicting the drag penalties caused by large ice accretions on singleelement lifting surfaces has been a long standing objective and focus of the icing research community. One prime avenue pursued has been the development of mostly empirical correlations relating measured drag penalties to either accretion conditions/characteristics and airfoil geometry, or the physical characteristics of the ice accretions, most notably the (glaze ice) upper horn height and inclination. Early on, Gray [122], using data from tests of five different airfoils in the IRT, proposed a correlation based on using icing time, air speed, temperature, LWC, impingement efficiencies, as well as airfoil chord, angle of attack, and leading-edge radius of curvature. Incidentally, almost all of the test results used in the development of this correlation had drag penalties measured only at the ice accretion angle(s) of attack. Two decades later, however, Olsen et al. [125] demonstrated that Gray’s correlation did not adequately predict the drag penalties measured for various ice accretions on a NACA 0012 airfoil in the

IRT. In fact, the agreement was quite poor. Other correlations, also based on ice accretion terms, such as those developed in the early 1980s by Bragg [160], Miller et al. [161], and Flemming and Lednicer [162], appeared to overcome some of the shortcomings of Gray’s correlation, but these still all had limited applicability. More recently, efforts aimed at developing drag penalty correlations have concentrated more on associating such drag increases with the physical characteristics of the ice accretion in the hope of obtaining a correlation that is ‘‘universally applicable’’. Toward this end, Cook [163] demonstrated that ‘‘measured ice shape maximum height and angle may be correlated to measured drag due to ice accretions’’. This finding is reflected in the recent correlation put forth by the 12A Working Group [98] as illustrated in Fig. 49. In this case, the correlation parameter selected represents the nondimensional height of the ice horn (or protuberance) normal to an undisturbed streamline. It can be seen from Fig. 49 that the drag penalties for the ice accretion conditions selected (i.e., ap3:2) generally increase with the correlation parameter utilized, but, there is a large amount of scatter in the results, clearly indicating that some fundamentals are definitely missing. Unfortunately, however, even if the scatter in correlations such as depicted in Fig. 49 could be greatly reduced, such correlations are of very limited practical value to aircraft developers and operators because, in reality, as was illustrated previously in Fig. 48, there is no single drag penalty for any particular ice accretion even for any one airfoil geometry, yet alone for different geometries, etc. This is an extremely fundamental point,

Fig. 49. 12-A working group correlation of airfoil drag increases due to ice accretions [98].

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well known to aerodynamic designers of aircraft. Interestingly, the authors of the first joint report published by the Swedish-Soviet Working Group back in 1977 [74] were well aware of this fact. Much of the relatively large database of IRT ‘‘single point’’ (i.e., fixed a) incremental drag measurements is also, unfortunately, of very limited value for establishing realistic drag penalties for a range of operating conditions. These results have, however, been very useful in establishing trends such as the pronounced ‘‘peaking’’ of drag penalties at total temperatures just below freezing conditions, illustrated by the Ref. [125] test results, as well as identifying the detrimental effects of increased air speed, droplet size, and LWC. And, the indications from the limited number of (fixed a) test results obtained at angles of attack above the accretion condition that drag penalties at these conditions are increased significantly above those at the accretion condition are also most instructive. But, for some other trends indicated by these ‘‘single point’’ drag increments, caution is suggested before presuming they apply to a range of operating conditions. An example of this would be the indication from a number of test results [119,123,124,135] etc. that the increases in drag penalties with accretion time tend to be more linear in contrast to maximum lift penalties where a significant part of the eventual penalty occurs very early in the accretion process. However, as will be seen subsequently, this observation is misleading. So, in order to place the drag penalties caused by large glaze (and inter-cycle) accretions into an aircraft operational framework rather than a research perspec-

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tive, test results available where drag penalties for mostly simulated ice accretions have been measured over a range of lifting conditions, and, corresponding maximum lift characteristics (and penalties) have also been established, are used herein to assess drag penalties at specific operating conditions (i.e., speeds relative to the reference un-iced stall speeds). The (potentially) applicable available database includes two primary sets of high Reynolds number test results, and ten sources of lower Reynolds number results. Except for one set of low Reynolds number test results obtained with actual ice accretions in the IRT [126–128], all the others are from tests in conventional wind tunnels using simulated ice shapes. High Reynolds number test results available are from the LTPT test of the 14% thick GA NLF airfoil [96,132], and the test of the 8% thick full-scale business jet T-tail model in the NASA 40
80 ft2 Wind Tunnel [80]. While tail drag increases due to icing per se are not really a high priority issue (compared to the wing), the latter set of test results, when interpreted as though they were for a wing, are helpful in identifying drag penalties which can be experienced on ‘‘thinner’’ (wing) configurations, and for demonstrating Reynolds number and other effects. Turning first to the Refs. [96,132] tests results obtained for the 14% thick GA NLF airfoil in the LTPT, the indicated incremental drag penalties for the four simulated glaze ice accretions evaluated are shown in Fig. 50. These four are 6 and 22.5 min accretions having the actual (IRT measured) 3-D roughness

Fig. 50. Drag penalties for simulated glaze ice accretions on GA NLF airfoil [96,132].

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characteristics represented, and smoothed (2-D) simulations of the 2 and 6 min accretions. Accretion conditions and illustrations for these simulated ice shapes were provided previously in Fig. 33 with the corresponding maximum lift penalties. From the test results illustrated in Fig. 50, the following observations can be made: *

*

*

*

At the higher speed conditions, the indicated drag penalties are roughly proportional to the accretion times. There is a very pronounced growth of the drag penalties at lower speed conditions approaching the (increased) iced configuration stall speeds. At these conditions, it appears as though the maximum penalties likely (without stalling) will not be proportional to the accretion times, but they will occur at different speeds with different ice shapes. The Reynolds number effect on the indicated drag penalties is rather inconsequential over the range tested (i.e., from 4.6 to 10 million). There is a very noticeable increase in the drag penalty incurred with the larger 22.5 min accretion as the Mach number is increased from 0.21 to 0.29. Interestingly, though, there is actually a small reduction in the indicated penalty for the 6 min simulations at this same Mach number.

An instructive perspective is gained by examining crossplots of the results shown in Fig. 50 at fixed speed

conditions as a function of the corresponding maximum lift penalties incurred with these ice simulations. Such a presentation for the lower Mach number results (i.e., M0 p0:21) is provided in Fig. 51 where it can be seen that the drag penalties at a given speed are clearly related to the maximum lift penalties which, in turn, are an indication of the proximity to stall. These results would seemingly indicate that the incremental drag penalty incurred by the 6 min simulated rough ice accretion at a speed 30% higher than the un-iced configuration stall speed would be about the same as the penalty incurred by the 22.5 min accretion at a speed 50% above the un-iced stall speed. Hence, it is clearly necessary to specify the operating conditions (i.e., proximity to stall) when defining drag penalties for any ice accretion. A similar presentation comparing the drag penalties observed at 0.29 Mach number to those seen at M0 p0:21 is shown in Fig. 52. Here, the significant potential adverse impact of the drag increase suffered by the larger ice accretion at 0.29 Mach number on the magnitude of drag penalties possible (prior to stall) is most obvious. Unfortunately, these data are the only known set of airfoil test results where such potential Mach number effects for a given ice shape have been investigated. This is important to note because caution must (again) be urged in making quantitative evaluations using the results from this test program (at any Mach number) because of the previously discussed premature separation problem present at very modest angles of attack with the (36 in

Fig. 51. Relationship of drag penalties and maximum lift losses with simulated glaze ice accretions on GA NLF airfoil [96,132].

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Fig. 52. Compressibility effects on relationship of drag penalties and maximum lift losses for GA NLF airfoil [96,132].

Fig. 53. Drag penalties for simulated glaze ice accretions on ‘‘business jet’’ T-tail model [80].

chord) un-iced baseline configuration. This premature separation caused by tunnel wall effects has a very obvious (adverse) impact on the un-iced configuration baseline drag polars. Because of this, it is feasible that the indicated drag penalties (at most all conditions) might well be lower than would be the case without these wall effects.

Looking next at the other primary source of available high Reynolds number test results addressing drag penalties for larger (simulated) glaze ice accretions, the indicated drag penalties from the Ref. [80] test of the full-scale business jet tail model at a Reynolds number of 5.1 million are shown in Fig. 53 as a function of speed ratios relative to the un-iced baseline stall speed for the

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three ice shapes addressed. The subject ice accretion conditions being simulated, and illustrations of the two larger accretions, were shown previously in Fig. 35 with the corresponding maximum lift penalties. In Fig. 53, results are presented for both smooth and roughened versions of the two larger LEWICE-defined accretions (9 and 22.5 min duration), as well as the roughened version of the shorter duration IRT ice accretion. From these results, it can be seen that the major impact (increased drag penalty) of simulating roughness characteristics occurred at the lower speeds, primarily reflecting the increased maximum lift penalty. There is little effect of roughness indicated at the higher speed conditions. It can also be seen from Fig. 53 that, in general, the drag penalties increase in a rather linear manner with accretion duration after the initial 2 min accretion. It should be noted also that the indicated drag penalties incurred with this 3-D, relatively low aspect ratio configuration, are in general about double those seen with the Refs. [96,132] airfoil test results even though the corresponding maximum lift penalties are lower than the airfoil results. Clearly, some of this increase may well be associated with the much thinner design (i.e., 8% versus 14%), but it is also likely that some part of the growth is due to non-viscous liftdependent drag increases. The significant effect (i.e., reduction) of reducing the Reynolds number from 5.1 to 1.36 million on the indicated drag with this configuration for the two larger (roughened) ice accretions is illustrated in Fig. 54. However, here again, the reduction occurs mostly at lower speeds, and is, to a large degree, associated with the reduced maximum lift penalties incurred at this lower Reynolds number.

The crossplots of the drag penalties at fixed speed conditions as a function of the corresponding maximum lift penalties measured for this business jet tail configuration provide an additional useful perspective regarding the potential inappropriateness of test results obtained with non-roughened ice accretions and/or those obtained at low Reynolds numbers, albeit these results are for a quite thin lifting surface. In Fig. 55, these crossplots are illustrated for the high Reynolds number test results obtained with the three rough ice accretions at speed ratios of 1.3, 1.5, and 1.8. In addition, results are also provided in Fig. 55 at 1.3 for the two smooth ice shapes tested at the same Reynolds number, and for the two larger rough ice accretions at a much lower Reynolds number. Similar comparisons at speed ratios of 1.5 and 1.8 are shown in Fig. 56. At all three conditions, it can be seen that even though the test results at low Reynolds numbers as well as those for the smooth ice simulations have reduced drag penalties, the corresponding reductions in maximum lift penalties are disproportionately larger, resulting in indicated drag penalties which are too high relative to the maximum lift penalties. It is worth noting, however, that the trend regarding smooth simulated ice shapes was not apparent in the Refs. [96,132] high Reynolds number test results for the thicker GA NLF airfoil. The remaining database of almost exclusively low Reynolds number test results for 2-D airfoils is presented in Figs. 57–59 together with the Refs. [96,132] high Reynolds number results for speed ratios of 1.3, 1.5, and 1.8, respectively. Incidentally, all of these low Reynolds number results are (believed to be) for relatively low Mach numbers (i.e., o0.20), so, these

Fig. 54. Reynolds number effects on drag penalties for simulated glaze ice accretions on‘‘business jet’’ T-tail model [80].

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Fig. 55. Relationship of drag penalties and maximum lift losses with simulated glaze ice accretions on ‘‘business jet’’ T-tail model [80].

Fig. 56. Reynolds number and roughness effects on relationship of drag penalties and maximum lift losses on ‘‘business jet’’ T-tail model [80].

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• •

Fig. 57. Relationship of drag penalties and maximum lift losses with simulated glaze ice accretions on various 2-D airfoils at speed ratio=1.3.

Fig. 58. Relationship of drag penalties and maximum lift losses with simulator glaze ice accretions on various 2-D airfoils at speed ratio=1.5.

additional results do not provide any new information regarding the drag increases seen at higher Mach numbers (i.e., 0.29) with the Refs. [96,132] test results.

In examining the results displayed in Figs. 57–59, there are probably three primary observations to be made. For one, there is no indication from these

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Fig. 59. Relationship of drag penalties and maximum lift losses with simulator glaze ice accretions on various 2-D airfoils at speed ratio=1.8.

low Reynolds number data for thicker airfoils that test results obtained with smooth versions of glaze ice accretions yield disproportionately high drag penalties such as seen with the thinner business jet tail model. The second, and more important observation to be made is that in almost all cases, the lower Reynolds number results do yield higher drag penalties for a given maximum lift penalty than seen with the Refs. [96,132] higher Reynolds number results. In conjunction with this, it is important to note that in all three of the low Reynolds number data sets where results are available at more than one Reynolds number for the same ice shapes, the higher (but still relatively low) Reynolds number results (shown as solid symbols) do generally indicate reduced drag penalties for a given level of maximum lift penalty, consistent with the business jet tail results. The third observation of significance is noting the similarity of the indicated parasite drag penalties (i.e., approaching 0.10) just prior to stall at both the 1.3 and 1.5 speed ratios, even though the ice shapes and maximum lift penalties are quite different. This again demonstrates the importance of defining the operating conditions being considered when attempting to identify drag penalties caused by large glaze ice (or any other) accretions. Incidentally, keeping in mind that the Refs. [96,132] indicated drag penalties are likely somewhat understated, a best-guess representation of the parasite

drag penalties likely caused by larger glaze ice accretions on typical wings at Reynolds number representative of flight conditions (but at incompressible conditions) based on the current available database would be somewhere between the two fairings indicated in Figs. 57–59. The foregoing review of available data addressing the drag penalties caused by larger glaze ice accretions on single-element lifting surfaces (primarily airfoils/wings) has enabled a number of important observations, but, at the same time, has made it very evident that there are some major holes in the database needed to realistically predict these drag penalties for the range of geometries and potential operating conditions needed. Regarding priorities, it is clear from a safety perspective that avoiding stall in the iced configuration is top priority. After that, it is important to avoid operating even close to stall if ‘‘large’’ drag penalties are to be avoided. However, for both of these, the previously suggested ‘‘well conceived’’ high Reynolds number test programs are needed to first establish the upper limit of maximum lift penalties possible with the so-called (most) critical ice shapes. At the same time, believable (i.e., no more 36 in chord models in the LTPT) drag penalties at high Reynolds number conditions need to be determined for the range of possible operating conditions for these ‘‘critical’’ ice accretions. This would have to include elevated Mach numbers, and determining penalties

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likely on realistic 3-D finite aspect ratio designs where drag sources in addition to just the parasite drag contribution are involved. 6.3.2. ‘‘Single-element’’ lifting surfaces with trailing-edge control surfaces Drag penalties caused by large glaze ice accretions on these geometries are of relatively minor significance considering the very short duration of such occurrences. While there are (at least) a couple of sources of test results available addressing this situation, namely, the Ref. [80] results from the test of the full-scale business jet T-tail, and the Ref. [81] results obtained with a section from the Twin Otter tail, no analysis of these test results is included herein because of the very low impact of any such drag penalties. Again, avoiding (tail) stall is the top priority. 6.3.3. Multi-element lifting surfaces Even though the available database of test results addressing drag penalties caused by larger ice accretions on multi-element high-lift geometries is quite limited, and not well suited for enabling good quantitative estimates of the largest drag penalties possible, it is, nonetheless, very useful for identifying some important trends/tendencies regarding these penalties. There are three sources of applicable test results, from Refs. [74,126–128,158], and most of the results are for configurations having leading-edge slats. All were 2-D

tests, so, as usual, the penalties indicated are only the parasite drag contribution. The indicated drag penalties at 1.3 times the un-iced configuration stall speeds are first summarized in Fig. 60 for Reynolds numbers of 5 million and below. For clarity, test results for either the IRT or simulated ice shapes involving noticeable or meaningful accretions on the LE of the flap are represented by solid symbols, and results for configurations not having leading-edge devices are denoted by flagged symbols. For the record, the ice shapes evaluated in these investigations were shown previously in Table 1 and Fig. 41 for the Ref. [74] geometries, Fig. 42 for Refs. [126–128], and Fig. 43 for Ref. [158]. Upon reviewing the indicated parasite drag penalties shown in Fig. 60, there is one principal lesson to be learned, and that is, without doubt, flow deterioration and/or separation on the upper surface of the flap is the most important factor involved in establishing the magnitude of drag penalties incurred. First, for the cases where there was some meaningful ice accretion on the flap LE, it can be seen that the penalties indicated are about double most of those indicated without such accretions. Second, for the one data point from Ref. [74] which has a much higher drag penalty indicated than the other results obtained without ice on the flap LE, it must be noted that this particular data point is for a configuration having a single-segment flap deflected 401. That being the case, it is very likely, particularly at the low test Reynolds number, that flow separation on the flap was imminent

Fig. 60. Drag penalties caused by larger ice accretions on multi-element high-lift airfoils at speed ratio=1.3.

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in the un-iced condition, and, hence, the ice accretion on the LE of the main element would most certainly have led to/triggered a large flow separation on the flap. Consistent with this presumption, it can be seen from Fig. 60 that the indicated drag penalty for this geometry with the flap only deflected 201 is in line with the rest of the test results (not having ice on the flap LE). Consequently, it seems quite apparent that drag penalties incurred with larger ice accretions on multielement high-lift geometries will, in general, also be strongly influenced by what flow separation margins exist on the flap of the un-iced configuration at the flight conditions under consideration. Clearly, many possibilities exist. Other issues needing to be considered relative to the results shown in Fig. 60 are concerns that the ice accretions involved are likely not the most critical possible, plus the ever-present concern regarding the adequacy of the test Reynolds numbers. However, it is conceivable that these two factors may be somewhat offsetting. Clearly, by definition, more critical ice accretions would have greater drag penalties, but the little data available illustrating Reynolds number effects on drag penalties clearly indicates that these penalties might well be reduced at higher Reynolds number more representative of flight on larger vehicles. The data indicating the possibility of reduced drag penalties at higher Reynolds numbers are the Ref. [158] data shown in Fig. 61 (for the not-too-large accretions considered). It can be seen that for the three icing simulations

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considered, there is an order-of-magnitude reduction in the drag penalty as the Reynolds number is increased from 5 up to 16 million. It remains to be seen, however, what these Reynolds number effects would be with larger, more critical ice accretions. 6.4. Trailing-edge control surface characteristics With leading-edge roughness representative of initial ice accretions, it was found that any really noteworthy changes in elevator hinge moment characteristics were tied very directly to noteworthy changes in the tail maximum lift level and stall characteristics brought about by these small ice accretions. And, conversely, it was found that if these initial leading-edge ice accretions had little effect on the tail maximum lift level and stall characteristics, the elevator hinge moments were likewise impacted very little. Not surprisingly, the same thing happens with larger ice accretions, the only difference being that when changes do occur, they happen at a lower incidence angle consistent with the (generally) somewhat larger losses in maximum lift and stall margin that occur with these larger ice accretions. The available database of test results where the effects of larger ice accretions on elevator hinge moment characteristics has been determined is predominantly low Reynolds number, and comes from the same sources which provided the material reviewed previously regarding leading-edge roughness effects on such characteristics. Results are available for three relatively thin tail

Fig. 61. Reynolds Number effect on drag penalty [158].

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geometries and two thicker ones. Both sets of test results for thicker geometries, the 181 tailplane with thick LE from Refs. [77,78] and the 2-D section from the Twin Otter tail from Ref. [81], are at low Reynolds numbers. The same is true for the two thinner geometries addressed in Refs. [77,78], the 441 tailplane and the 181 tailplane with normal LE. Unfortunately, the only high Reynolds number data available are the Ref. [80] test results for the 8% thick full-scale business jet tail. With the thicker tail geometries, as was pointed out previously, the elevator hinge moments without any leading-edge ice become gradually more positive (or less negative) as the tailplane incidence/download is increased prior to stall, but then, right around stall, the rate of increase (toward more positive) becomes noticeably greater. Then, when the larger ice shapes are added, the S3 shape on the thickened 181 tail and the 45 min LEWICE and ‘‘S&C’’ shapes on the Twin Otter tail section (see Tables 1 and 2 for details), the resulting elevator hinge moment characteristics closely resemble those for the respective baseline geometries except that the noticeable increase in hinge moments in the positive direction occurs at a much lower incidence angle consistent with the stall angle reductions illustrated earlier in Fig. 46. Interestingly, in the one set of test results where trailing-edge-down elevator characteristics were evaluated [81], the increase in hinge moments occurring around stall at these deflections is greater than for trailing-edge-up conditions. However, in no case is the change any more severe than seen with the un-iced baseline, it just occurs earlier. Also, once again, the results from these two investigations illustrate that the magnitude of the change in elevator hinge moments around stall is closely related to the baseline stall characteristics, with the Refs. [77,78] results providing a more abrupt change, consistent with a larger more abrupt lift loss at stall with this geometry. The elevator hinge moment characteristics measured for the three thinner tail geometries with larger ice accretions produce results somewhat different from those seen with the thicker designs. For example, with the 441 tailplane addressed in Refs. [77,78], the hinge moments on the un-iced baseline prior to stall increase (i.e., became more positive) at a greater rate than seen with the thicker designs, but there is very little change in slope seen around the maximum lift condition, consistent with a very gentle (i.e., unexciting) roundover of the lift curve at these conditions. Going along with this, when the R3 ice ‘‘imitator’’ (see Table 1) is added, there is very little change in elevator hinge moment characteristics seen over the wide range of incidence angles investigated. Pretty much the same thing holds true with the results seen when the S2 ice ‘‘imitator’’ is added to the 181 tailplane with normal LE. For both the un-iced baseline configuration and the geometry with the ice ‘‘imitator’’ added, there is no significant change in slope/

trend seen around maximum lift conditions. In this case, the steeping of the hinge moment rate of increase which starts long before stall for both configurations seems to happen a couple of degrees earlier with the ice ‘‘imitator’’ installed. With the third set of test results available for thinner tail geometries, the data provided in Ref. [80] for the 8% thick business-jet tail geometry, particularly at the high Reynolds number test condition with the largest (rough) ice accretion evaluated (illustrated in Fig. 35), the hinge moments do have some characteristics which start to resemble those seen with the thicker designs (at lower Reynolds numbers), but there are still some significant differences remaining. Probably the biggest difference remaining is that, in this case, a rather abrupt increase (more positive) in hinge moments occurs about 51 prior to stall with the un-iced baseline rather than more coincident with it as occurs with the thicker designs. The primary similarity with the test results for thicker tail geometries is that the increase which occurs with the simulated ice accretion, although not quite as abrupt as seen with the un-iced baseline, occurs about 5–61 earlier than on the un-iced baseline. This is very different than what occurred with the ice accretions evaluated in Refs. [77,78] on the two other thinner tail geometries. There is one other bit of related information available on this subject which should be noted, and that is the presentation of flight-measured stick forces obtained with larger ice accretions on the tails of several Russian transports (i.e., An-10, An-24, IL-18, IL-62, and Yak40) that is contained in Ref. [78]. In general, these results illustrate the large and sudden changes that can occur, particularly with flap extension, as well as a very noticeable lightening of stick forces, and, for some, an earlier sign reversal. In concluding this section, what has again been demonstrated herein is that there are a wide range of possible effects depending on lifting surface geometries and types, the severity of the ice accretions simulated, test Reynolds numbers, etc. And, our biggest concerns with the current database have to be that none of the test results addressing elevator hinge moment characteristics were obtained with the most critical ice accretions possible, and there is no high Reynolds number data whatsoever for thicker tail designs. These deficiencies need to be addressed if a proper understanding is to be obtained of the most adverse possible types of control surface problems which could be encountered in flight.

7. Effect of ground frost/ice accretions Ground frost, which accumulates in a relatively rough thin layer on the upper surface of wings (and horizontal stabilizers) while the aircraft is parked, or during subsequent taxi operations, can have a serious adverse

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impact on aircraft aerodynamic characteristics and operating margins if not removed before takeoff. These adverse effects have been described by a number of investigators such as Zierten and Hill [164], von Hengst and Boer [165], Brumby [56], etc. In fact, as documented in Refs. [56] and [165], a number of aircraft incidents/ accidents have been attributed to the failure to adequately remove such so-called hoar frost accretions from the wing upper surface. Similar accretions on the upper surface of the tail are of much less consequence since that is not the critical surface during takeoff. Although the roughness characteristics of ground ice accretions can vary significantly depending on atmospheric conditions, etc., typical ground frost accretions are often similar in roughness characteristics and thickness (order of one millimeter) to the in-flight leading-edge buildups necessary for removal by de-icing systems. However, instead of just being confined to the first several percent of the wing chord on both upper and lower surfaces as was the case with the in-flight accretions, these ground frost accretions can cover the entire exposed upper surface of the wing (slat, main element, and flap). Some variations in this upper surface coverage exist depending on whether or not the leadingedge devices (typically slats) and trailing-edge flaps are retracted during ground operation prior to takeoff in these adverse weather conditions. If left extended, a more extensive roughness coverage would exist. And, in addition to aerodynamic performance effects, these wing-upper-surface ice accretions are also of concern for aircraft with aft-fuselage-mounted engines due to the possibility of engine ingestion of this ice as it subsequently breaks loose. Another form of ice accretion which can occur when an aircraft is sitting on the ground is what is referred to as ‘‘in-spar’’ frost/ice, which most often forms on the lower surface of wings in proximity to fuel tanks containing ‘‘cold soaked’’ fuel, but it can also form on the wing upper surface as well with some configurations [166], and would have to be removed in these cases. However, analogous to the situation with upper surface frost on the horizontal tail, lower surface ice accretions are of relatively minor consequence to the aerodynamic characteristics of the wing at takeoff conditions as shown by Bragg et al. [167]. While this type of ice accretion could conceivably be an issue for any aircraft that have ‘‘wet’’ tails, i.e., fuel in the tail, it is felt that such an occurrence would be extremely rare. Whereas there was a real scarcity of good quantitative data available from full-scale aircraft flight testing to help validate the aerodynamic effects of the other forms of ice accretions considered up to here, there are, fortunately, several sets of flight test results existing for contemporary jet transports addressing the aerodynamic consequences of upper-surface ground frost/ice accretions. The mere existence off these data is certainly a clear indication of the widespread concern regarding the

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consequences of failing to thoroughly remove these ice accretions prior to initiating takeoff. Incidentally, eliminating these ice accretions by applying highly viscous ground anti-icing fluids does not necessarily remove all performance degradations. That is because the ‘‘rough’’ residual from these fluids that remains on the wing during takeoff can itself, lead to performance degradations, which are especially worrisome for smaller aircraft having ‘‘hard’’ LEs. 7.1. Maximum lift reductions Reductions in maximum lift capability are most critical at takeoff conditions since speed margins to stall are typically smaller at these conditions than at any other time in the flight envelope. 7.1.1. Single-element lifting surfaces Not surprisingly, considering that small leading-edge disturbances are far more critical in terms of causing reductions in maximum lift than those located aft of the LE, maximum lift penalties for having the entire upper surfaces of single-element lifting surfaces covered with roughness simulating ground frost accretions are very similar to those seen with just leading-edge roughness. To illustrate this similarity, results from two high Reynolds number wind tunnel tests [71,168] and an MD-87 flight test program [169], all with the entire airfoil or wing upper surface covered with roughness, are presented in Fig. 62, together with the previously established band of corresponding high Reynolds number penalties observed for having just leading-edge roughness (from Fig. 10). As can be seen, the penalties with the entire upper surface covered with roughness are well within the range established for just leading-edge roughness effects. In fact, in one of these investigations [71], penalties were measured for both types of roughness coverage, and the additional penalty due to having the entire upper surface roughened was an order of magnitude less than the basic penalty for just the leading-edge roughness. Also confirming the non-criticality of aft-located roughness were further results from the other wind tunnel investigation [168] which illustrated that having only the aft 50% of the airfoil upper surface covered with roughness (k=cB104 ) resulted in maximum lift penalties of only 1% or less. A very key lesson provided by the rather unique set of flight test results shown in Fig. 62 is that sizable maximum lift penalties due to upper surface (leading-edge, etc.) roughness do persist down to very low roughness heights (i.e., k=c ¼ 2
105 ). Following the persistent trend now anticipated, available test results where the effect of upper surface roughness on maximum lift penalties has been measured over a range of Reynolds numbers again indicate that low Reynolds number test results are not representative

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Fig. 62. Upper surface roughness (ground ice/frost) size effects on maximum lift penalty at high Reynolds numbers for single-element lifting surfaces.

of higher Reynolds number conditions. It can be seen from the available data summarized in Fig. 63 that the indicated penalties at low Reynolds numbers are clearly lower than those experienced at higher Reynolds number conditions. That being the case, other available low Reynolds number measurements of these upper surface roughness effects such as those presented in Ref. [165] at 0.9 million Reynolds number, Ref. [137] at 1.3 million, and those presented in Refs. [74,171] at 2.6 million, are of little value when needing to establish the upper limit of penalties likely/possible at flight conditions for a variety of aircraft types. For the record, the Ref. [165] test results for the F27 do fall below the lower limit of the band established at high Reynolds numbers for just leading-edge roughness, and the penalties indicated by the Refs. [74,172] test results for the NACA 65A215 airfoil tend to be toward the lower end of the band, as do the results from Ref. [137]. 7.1.2. Multi-element lifting surfaces The magnitude of reductions in maximum lift capability caused by ground-frost-type accretions on the upper surface of multi-element airfoils/wings at takeoff-type conditions is well documented at high Reynolds numbers for configurations employing leading-edge devices (primarily slats). Extensive flight test results from Refs. [164,169] are available for four contemporary jet transports (737, 757, 767 and MD87). In addition, there was also one 2-D high Reynolds number wind tunnel investigation [82] which addressed these reductions as well as just leading-edge roughness

effects, thereby permitting a direct comparison, as well as a semispan model test of the Fokker F29 configuration [173,174] where upper surface frost/roughness effects both with and without a slat were measured. Incidentally, the upper surface frost (0.5 mm roughness) coverage simulated during the 737, 757, and 767 flight test programs assumed that the leading-edge device as well as the flap were retracted when the frost was being accreted, while the MD-87 flight test and the wind tunnel test used a frost/roughness coverage that assumed that these elements were extended during this time, thereby resulting in a somewhat greater chordwise extent of roughness near the LEs of the main element and flap. This difference should be rather inconsequential, however, since these areas are not at all critical at takeoff conditions as shown in Ref. [61]. These high Reynolds number test results for configurations with leading-edge devices extended are shown in Fig. 64, along with the bands representing the range of penalties observed at high Reynolds numbers for single-element lifting surfaces for both leading-edge roughness and having the entire upper surface covered with roughness. In addition, the band obtained from low Reynolds number wind tunnel tests of the 737, 757, and 767 with upper surface frost accretions is shown. Several important conclusions can be reached based on the results shown. First, consistent with the trends observed with leading-edge roughness effects, percentage losses in maximum lift incurred with upper-surface frost accretions on multi-element lifting surfaces are noticeable smaller than with single-element designs.

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Fig. 63. Reynolds number effects on maximum lift penalty caused by upper surface roughness (ground ice/frost) on single-element airfoils [170].

Fig. 64. Maximum lift penalties caused by upper surface frost/roughness on multi-element high-lift configurations with leading-edge devices at high Reynolds numbers.

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Next, percentage losses in maximum lift capability for multi-element designs caused by upper surface frost decrease with increasing flap deflection (i.e., increasing basic maximum lift). It is also very significant to note that the low Reynolds number wind tunnel test results for the 737, 757, and 767 indicate that comparable penalties for upper surface frost accretions do not occur until an order of magnitude higher nondimensional roughness height than happens at high Reynolds number flight conditions, another indication that extreme caution should be used when attempting to directly apply low Reynolds number test results to higher Reynolds number flight conditions. Another important conclusion to note is that the reductions in maximum lift do persist down to quite-small roughness sizes, indicating that even very thin coatings of frost are of concern. With regard to the test results for the F29, it can be seen that they tend to fall in line with the wind tunnel test results for the 737, 757, and 767, perhaps indicating that an MAC Reynolds number of 5 million may not have yielded results representative of flight in this case. It was indicated earlier that maximum lift penalties at takeoff conditions are most critical because speed margins to stall are typically the lowest at these conditions. To illustrate the seriousness of any reductions in maximum lift capability at takeoff conditions, the maximum lift penalty which would result in a stall at minimum takeoff speeds (i.e., 13% higher than 1-g stall speeds) assuming no overspeed is employed is highlighted in Fig. 64. It can be seen that if overspeed is not employed on takeoffs where upper surface frost is present (at the accumulations addressed), then half or more of the stall margin can be eroded. Even more of the margin would be eroded with larger accumulations. Lastly, although the data is not shown, the high Reynolds number wind tunnel test results [82] indicate that the additional penalty for having the entire upper surface covered with frost/roughness is less than half the penalty for just the LE roughness. Furthermore, additional results from this same test program indicate that if frost was just removed from the slat upper surface, then the reduction in maximum lift caused by the remainder of the frost was less than a third of the penalty incurred when the whole upper surface was rough, again indicating that the focus in ground de-icing must start with the LE of the wing. Unfortunately, there is much less high Reynolds number data available addressing upper surface frost/ roughness effects on the expected-to-be-more-critical wing designs not having leading-edge devices (i.e., ‘‘hard’’ LEs). Unlike the situation for configurations with leading-edge devices, there are no flight test results available. What is available are results at (MAC)

Reynolds numbers around 5 million from semispan model tests of the Fokker 100 and (aforementioned) F29 [173,174]. This is certainly not the most confidenceinspiring high Reynolds number database, especially considering the indications provided by the F29 test results at this Reynolds number with a LE slat. However, be that as it may, the results from these tests plus the available and applicable database from tests at lower Reynolds numbers are provided in Fig. 65. Incidentally, when tests were conducted both with and without leading-edge devices, both sets of results are shown so as to illustrate the trend of higher (percentage) penalties for ‘‘hard’’ leading-edge designs. In these cases, test results with leading-edge devices are shown as solid symbols. From the results presented in Fig. 65, there are some important conclusions to be drawn. For one, in both cases where configurations both with and without leading-edge devices (slats) were evaluated, the (percentage) maximum lift losses caused by having the entire upper surface covered with frost/roughness are clearly higher for the ‘‘hard’’ LE designs. It is also quite apparent from all the data presented for ‘‘hard’’ leading-edge designs that the level of maximum lift losses incurred with these designs is higher than seen in Fig. 64 for the designs with leading-edge devices. Consequently, based on these test results, much more, if not all, of the stall margin is eroded for ‘‘hard’’ leading-edge designs by modest accumulations of upper surface frost/roughness unless overspeed is employed. Whether this situation might be even more critical or require even more overspeed (to avoid) at flight Reynolds numbers remains to be seen, and should be investigated. Somewhat similar to the results for configurations having leading-edge slats, the limited (low Reynolds number) database available for configurations having ‘‘hard’’ LEs also indicate that the maximum lift penalties are much reduced if just the forward part of the airfoil is kept clean. In fact, the reductions indicated are greater than the corresponding reductions seen (at low Reynolds numbers) in the same investigations with configurations having leading-edge slats. For example, the results presented for the NACA 652A215 airfoil (with flap) in Refs. [74,172] at a Reynolds number of 2.6 million indicate that if the forward 18% of the wing chord is kept clean, the maximum lift penalties are only about 13214 of those incurred with the whole upper surface roughened. When only the forward 5% of the wing chord was kept clean, the indicated reductions are somewhat less, but still significant. So, at least based on low Reynolds number test results, it would seem clear that the absolute highest priority in ground de-icing should be in ensuring that the LEs of ‘‘hard’’ leading-edge designs are kept clean. Accident statistics would also lead to the same conclusion.

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Fig. 65. Comparison of maximum lift penalties caused by upper surface frost/roughness on multi-element high-lift geometries with and without leading-edge devices.

7.2. Stall angle reductions 7.2.1. Single-element lifting surfaces Since the maximum lift reductions caused by uppersurface ground frost/ice accretions fall in the same range or band as those caused by the initial leading-edge ice (roughness) accretions, it would be expected that the associated stall angle reductions would be similar as well. And, that is indeed the case, as can be seen from the available high Reynolds number measurements of stall angle reductions for upper-surface frost accretions shown in Fig. 66. In examining the data, it is of interest to note that the one set of test results for a 3-D configuration, namely the MD-87 flight results, do not stand out above the 2-D data as was the case for the results with the 3-D tail with leading-edge roughness (see Fig. 15). This difference can be explained, at least in large part, by the significantly higher aspect ratio of the MD-87 wing. Hence, in terms of accounting for the effects of lifting surface aspect ratio on stall angle reductions, it is clearly more important for relatively low aspect ratio geometries. There are also a number of low Reynolds number test results available addressing stall angle reductions caused by upper surface ice/frost on single-element geometries from most of the References mentioned in conjunction with the low Reynolds number maximum lift losses

illustrated in Fig. 63. However, since the maximum lift penalties measured at these low Reynolds numbers are not believed to be representative of higher Reynolds number flight conditions, neither would the indicated stall angle reductions. Consequently, these results are omitted herein. Although not shown, there is one most noteworthy aspect of all the low Reynolds number indications of stall angle reductions caused by upper surface roughness, and that is the results for the one 3-D geometry available [137], which had an aspect ratio of just under five, were at the upper end of the band of indicated stall angle reductions. This result, although conceivable a low Reynolds number anomaly in this case, is consistent with the trend highlighted earlier at high Reynolds numbers for LE roughness effects, and would seem to support the notion that aspect ratio effects are most important for relatively low aspect ratio designs. 7.2.2. Multi-element lifting surfaces As was the case with the maximum lift reductions caused by ground-frost-type accretions on the upper surface of multi-element airfoils/wings with leading-edge devices at high Reynolds number takeoff-type conditions, a representative range of corresponding stall angle reductions for these configurations is also well documented based on the four sets of flight test results on

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Fig. 66. Upper surface roughness (ground ice/frost) size effects on stall angle reductions at high Reynolds numbers for single-element lifting surfaces.

contemporary jet transports, and the one high Reynolds number wind tunnel investigation of these effects. The relatively narrow band of stall angle reductions determined from these five sets of test results are presented in Fig. 67 along with the band of corresponding high Reynolds number test results (from Fig. 66) for singleelement airfoils. Similar to the results presented earlier for stall angle reductions with initial in-flight leadingedge ice accretions, the reductions experienced with multi-element designs having leading-edge devices tend to be slightly larger than the corresponding singleelement designs. When examining the results shown in Fig. 67, it is worth remembering that stall angle reductions of three degrees were experienced in-flight with the extremely small roughnesses evaluated on the MD-87, and a reduction of over four degrees was experienced on the 737 with just 0.5 mm roughness. This certainly raises a concern over the possibility of stall angle reductions noticeably higher than these on smaller aircraft with similar relatively small frost (roughness) accretions. For configurations not having leading-edge devices, the (all low Reynolds number) database available shedding any light on stall angle reductions caused by ground-frost-type accretions is very limited. In addition, to make matters even worse, some of the (limited) results available indicate reductions several times more than those seen for configurations with leading-edge devices at high Reynolds numbers. The results available for these

‘‘hard’’ leading-edge designs are illustrated in Fig. 68, and, as was done in Fig. 65 for the related maximum lift penalties, where corresponding results are available with leading-edge devices, these results are shown as solid symbols. From the results shown, it would be difficult to find any prevalent trend in the results for ‘‘hard’’ LEs versus those for configurations having leading-edge devices, again illustrating that the higher (percentage) maximum lift penalties for ‘‘hard’’ leading-edge designs are primarily a consequence of the lower basic (i.e., clean) maximum lift levels of these designs. It can also be seen that in contrast to the relatively narrow band of reductions indicated at high Reynolds numbers for configurations with leading-edge devices, the low Reynolds number database for geometries both with and without leading-edge devices portray a much broader range of potential stall angle reductions caused by small upper-surface ground frost/ice roughness accretions. In fact, stall angle reductions observed in these low Reynolds number tests range from o11 to over 101, an alarmingly wide band, with relatively inconsequential effects at the lower end, to potentially disastrous effects at the upper end. Certainly, some well conceived high Reynolds number testing is warranted, particularly for the worst ‘‘hard’’ leading-edge configuration(s), to document the reductions likely at flight conditions if such frost/ice accretions are not removed prior to takeoff.

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Fig. 67. Upper surface roughness size effects on stall angle reductions for multi-element high-lift configurations with leading-edge devices at high Reynolds numbers.

Note: • All configurations have upper surface roughness 0-100%.

Fig. 68. Comparison of stall angle reductions caused by upper surface frost/roughness on multi-element high-lift geometries with and without leading-edge devices at low Reynolds numbers.

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7.3. Drag penalties Because of the much longer chordwise extent of the roughness brought about by upper-surface ground frost/ ice accretions, it would seem reasonable to expect that associated drag penalties on a given configuration would be somewhat higher than those incurred with initial inflight leading-edge type ice accretions having the same roughness height. However, unfortunately, there does not appear to be any data set available which would permit direct substantiation of this expectation. 7.3.1. Single-element lifting surfaces There appear to be only two (known) sources of available test results which have any data, high or low Reynolds number, potentially useable for assessing the drag penalties caused by upper-surface ground frost/ice accretions on single-element geometries. They are some very low Reynolds number (i.e., 0.9 million) results for the Fokker F27 [165], and some very limited higher Reynolds number results for the NACA 0012 and RAF 34 airfoils from Ref. [168]. Unfortunately, neither one of these sets of test results are very helpful. The results for the F27 are definitely suspect because of the low test Reynolds number, since, as was shown earlier in Figs. 20 and 21 when addressing Reynolds number effects on the drag increases caused by initial (and residual) leadingedge ice/roughness accretions, penalties measured at such a low Reynolds number are typically not representative of higher Reynolds number flight conditions, nor can the direction of the Reynolds number trend be predicted. As it is, the indicated drag penalties are right at the bottom of the band of results for leading-edge roughness effects shown in Figs. 18 and 19 (even though the nondimensional roughness height employed is larger in this case by a factor of three. The test results for the two airfoils reported in Ref. [168], albeit at higher Reynolds numbers, are also of limited value because only minimum parasite drag penalties are provided. There are no results given for the more critical lower speed conditions. The minimum parasite drag penalties in this case also fall toward the bottom of the band shown in Fig. 18, although part of the reason for this is likely the lower nondimensional roughness height utilized in this case (i.e., 1.25
104 versus 4.6
104). Interestingly, in this case, the variations in drag penalties with increasing Reynolds numbers for the two airfoils have opposite trends (i.e., decreasing with the RAF 34, but increasing with the 0012). Consequently, with this very limited amount of test data, it is not possible to make any really meaningful quantitative assessments of the range of drag penalties possible with these upper surface ground frost/ice accretions, but it would be surprising if this range was not at least as large as the very wide band documented for the leading-edge type roughness accretions.

7.3.2. Multi-element lifting surfaces Flight test results available for the 737, 757 and 767 from Ref. [164] together with test results from one 2-D high Reynolds number wind tunnel test [82] provide a very good indication of the range of absolute drag penalties incurred with upper surface ground frost on configurations having leading-edge devices. These flight test results for a roughness height of 0.5 mm, and wind tunnel results for a nondimensional roughness height (k=c) of about 0.7
104, are shown in Fig. 69 at both minimum drag conditions and at representative takeoff V2 speeds. It can be seen that in all cases, as would be expected because of the more adverse flow conditions on the upper surface of the wing/airfoil at this condition, the drag penalties at takeoff V2 speeds are always significantly higher than those incurred at minimum drag conditions. At takeoff V2 speeds, the flightmeasured drag penalties vary from about 0.01 up to over 0.04, compared to a range of from about 0.005 up to about 0.012 at minimum drag conditions. It is comforting to see that the 2-D high Reynolds number wind-tunnel test results do appear to be consistent with the flight measurements, especially when differences in reference area are taken into account. Finally, it is of interest to note that the drag penalties caused by these relatively small upper surface frost/ice accretions are, again as would be expected, much less than the penalties caused by large in-flight glaze ice accretions. With it already rather conclusively established that the frost and ice accretions close to the LE of lifting surfaces are clearly the most critical in terms of causing reductions in the maximum lift capability and associated stall angles, drag measurements made with an alternate roughness coverage on the slat during the 737 flight test program [164] give us a good indication that the same is also true with regard to drag penalties at takeoff climb conditions. For this alternate roughness coverage, ‘‘the simulated frost was partially removed from the slats to evaluate the benefits of the 737 ground mode slat thermal anti-ice (TAI)’’. The subsequent drag measurements obtained with this alternate coverage at the flap 1 and 5 riggings, which had suffered the largest absolute drag penalties with the basic roughness coverage at takeoff climb speeds, indicated that these drag penalties were, on average, just about cut in half. This is yet another set of (high quality) test results that provide even more substantiation (if it were needed) that the focus in ground de-icing must start with the LE of the wing. In contrast to the good, high-quality database available for high-lift configurations with slats, there is very little data of any kind available for assessing these drag penalties for high-lift systems having ‘‘hard’’ LEs. And, what little is available, such as the very low Reynolds number F27 test results reported in Ref. [165], are, by now, more than just suspect. So, as was the

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Fig. 69. Drag penalties caused by upper surface ground frost/ice (roughness) accretions on multi-element takeoff geometries.

situation with single-element lifting surfaces, there is really no good basis at this time for estimating the drag penalties likely to be caused by ground-frost type accretions on these configurations at critical takeoff conditions. This is obviously another big hole in the database that needs to be plugged. Lastly, with regard to any drag penalties that would be caused by wing-lower-surface ‘‘in-spar’’ frost accretions at takeoff conditions, available high Reynolds number test results [82] indicate that these accretions are of little practical consequence. With the same (nondimensional) roughness height, the drag penalty indicated for lower surface frost was about a factor of 50 less than that observed with upper surface frost/roughness. Even with a roughness height on the lower surface about six times greater than that used on the upper surface, the indicated drag penalty for the lower surface roughness was about a factor of 15 less.

8. Summary and conclusions The foregoing is a systematic and comprehensive review, correlation, and assessment of wind-tunnel and flight-test measurements available in the public domain

which address aerodynamic performance and control degradations caused by various types of ice accretions on the lifting surfaces of fixed wing aircraft. The intent has been to define the range of possible consequences which can occur at flight conditions, especially the worst which could be encountered. A part of this assessment has also been to identify critical voids in the available database that need to be rectified. CFD techniques have not been used herein to either augment or help correlate existing test results because of known inherent limitations of current state-of-the-art techniques such as RANS methods, etc., most specifically with regard to the inability to effectively predict separation onset characteristics, which are at the core of establishing the incremental effects of ice accretions. Four types of ice accretions have been considered. The first category addressed are the initial small accumulations that occur on the LE of aerodynamic surfaces either before the activation of the ice protection system and/or between de-icing system cycles. Next are the runback and/or ridge ice accretions which can form just aft of the ice protection system, especially during large droplet icing encounters. Following these, the large and often-more-irregularly shaped ice accretions which form during longer encounters either as a consequence

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of not having an ice protection system, and/or having a failure of the ice protection system, are studied. The last category considered are the ground frost/ice accretions which can accumulate on the upper surface of wings while the aircraft is parked, or during subsequent taxi operations. In addressing the various consequences of these different types of ice accretions, efforts have focussed on assessing resulting maximum lift reductions, the corresponding stall angle reductions, resulting drag penalties, and trailing-edge control surface anomalies. Based on this review, the four types of ice accretions have been classified as follows: *

*

*

Dangerous because of possibility of being underestimated and/or misunderstoodFClearly, heading this category are the initial small leading-edge ice (roughness) accretions and inter-cycle ice buildups occurring on both wings and tails. Misconceptions abound in terms of the thickness of this initial (or inter-cycle) ice accretion which can be allowed before any serious degradation of aerodynamic effectiveness occurs. However, what this review and assessment has again highlighted is that there are high Reynolds number test results available which demonstrate that even very small (i.e., 2–3 mils) buildups can result in very noticeable reductions in maximum lift capability and corresponding stall angles. Further, for the ice thicknesses required for operation of state-of-the-art de-icing systems, maximum lift reductions of up to 40% can occur for single-element surfaces. Dangerous because of potential for catastrophic reductions in aerodynamic effectivenessFLeading the way in this class are the runback/ridge ice accretions which can form just aft of a leading-edge ice protection system due either to the impingement of larger-than-normal water droplets in this area, or by the runback and subsequent freezing of water which is not evaporated by a leading-edge anti-icing system. Initial (low Reynolds number) test results indicate the possibility of maximum lift losses in excess of 80% for some single-element geometries. Clearly, more needs to be known about the consequences of these types of ice accretions at full scale flight conditions. Dangerous because upper limits of potential aerodynamic consequences are not really definedFThis category belongs to the larger glaze ice accretions which, ironically, have been the focus of a majority of studies conducted to date addressing the aerodynamic consequence of ice accretions. To date, maximum lift losses approaching 60% for singleelement lifting surfaces have been indicated for these accretions. However, it is clear that the maximum possible reductions in aerodynamic effectiveness have yet to be adequately defined. The penalties associated with such accretions at flight-like Reynolds numbers,

*

with appropriate wind tunnel model installations, and with simulations representing the most severe icing environment in terms of ambient temperature, droplet sizes, LWC, duration, and accretion altitude and speed, have yet to be established. Dangerous because of portion of flight operation envelope involvedFThis is clearly the province for ground frost/ice accretions that are not removed prior to takeoff, since margins to stall are typically at a minimum during takeoff compared to the rest of the flight envelope. The data available indicate that the penalties for these accretions are very similar to those incurred with the initial in-flight leading-edge and inter-cycle buildups. However, because of the reduced margins at takeoff, these ground frost/ice accretions can result in much if not all of the margin being eroded/erased.

In reviewing the wide variety of test results available for the foregoing types of ice accretions, it has been shown that the aerodynamic consequences of these ice accretions, in addition to being generally dependent upon the ice accretion features, are also strongly dependent on the geometry of the various baseline lifting surfaces. It has been shown that a given type/size of ice accretion can cause quite disparate effects on different lifting surfaces. Examples of this include the following: *

*

*

Adverse ice accretion effects, especially those associated with stall characteristics, are much larger on non-dimensionally ‘‘thicker’’ (single element) lifting surfaces which have baseline maximum lift coefficient capabilities in excess of 1.0. Corresponding ice accretion effects on thinner surfaces with lower maximum lift capabilities are often minimal. ‘‘Unfortunately’’, the superior performing thicker surfaces are much more prevalent on fixed wing civil aircraft types of most interest. Smaller, dimensionally thinner surfaces (with smaller leading-edge radiuses) such as tails are more efficient ice collectors than (dimensionally) thicker surfaces such as wings. Consequently, (percentage) degradations on these smaller surfaces are often larger for the same icing encounter. This can be an important issue when addressing tail stall concerns. Percentage penalties in terms of maximum lift reductions and drag increases are typically roughly inversely proportional to the number of surface elements, (i.e., the baseline performance capabilities). For example, configurations with leading-edge devices such as employed on most larger transport aircraft are the least impacted. Conversely, configurations having ‘‘hard’’ LEs typical of most smaller transports/aircraft typically incur noticeably higher penalties. And, it follows that single-element

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geometries are the worst. Accident statistics clearly substantiate this order of criticality. Testing conditions and ice accretions simulation issues also emerged as critical factors in establishing the appropriate magnitude of aerodynamic penalties caused by the various ice accretions addressed. For example, whenever test results were available for both simulated shapes (computational or otherwise) and actual ice accretion shapes, roughness, etc., more often than not, the penalties were greater with the actual ice accretion. With regard to testing issues, much evidence exists to firmly establish that testing at a sufficiently high Reynolds number to adequately represent flight conditions is a critical requirement if the measured incremental aerodynamic penalties caused by the various forms of ice accretions are to be representative/meaningful. Many low Reynolds number test results (i.e., usually below about 5 million) have produced penalties noticeably lower than would occur at flight conditions, certainly not a desired result. While Reynolds number effects on the performance (and control) of the baseline un-iced geometry predominate in many cases, especially when addressing maximum lift characteristics, there are also a number of examples of Reynolds number effects on the iced geometry as well, particularly for smaller ice accretions. Schemes aimed at extensively utilizing the absolute values of low Reynolds number test results of the maximum lift characteristics for larger ice shapes as a consequence of the apparent general insensitivity of these results to variations in Reynolds number could be very risky. This is because it is not known as to whether or not this insensitivity seen in 2-D test results could have came about as a consequence of strong sidewall boundary layer interaction effects controlling the stall rather than the ‘‘pure’’ aerodynamics of the larger ice shapes. Also, trying to use this approach with 3-D geometries where different spanwise stations may be critical at different Reynolds numbers, or where very different flow physics may be controlling at high and low Reynolds numbers, would present further major ‘‘challenges’’. Other findings of significance from this review involve ice accretion effects on control surface characteristics and drag penalties. With regard to control surface anomalies which can be critical in tail stall situations, these anomalies (including reversals) are in most all cases strictly a consequence of an earlier stall occurring with the contaminated surface. There are no real mysteries and/or subtleties involved. With regard to drag penalties associated with the larger glaze ice accretions, it is concluded based on the results presented herein that long-standing efforts aimed at identifying a single drag penalty based on parameters defining the geometry of the ice shape are misguided. Not only is there not a single penalty for a given ice shape on

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a particular lifting surface, even if one did exist it would be different for each surface geometry. Lastly, after reviewing many test results and associated commentaries, a number of which have been available for many decades, it is concluded that important lessons often learned by one generation regarding the various aerodynamic consequences of various forms of ice accretions have not been adequately passed on to, and/or accepted by, later generations, leading to accidents which seemingly could have been avoided if ‘‘past lessons learned’’ had been disseminated, taken advantage of, and acted upon. Prime examples of these failures include knowledge of tail icing consequences and procedures for minimizing the likelihood of tail stall (and associated control surface anomalies), the large effect of possible runback/ridge ice accretions, the significant adverse consequences of small initial and inter-cycle ice accretions associated with the use of deicing systems, and the importance of having an adequate test Reynolds number. In addition, lessons learned in terms of what constitutes acceptable wind tunnel/model installations for the determination of the consequences of various ice accretions have often been ignored for any number of reasons. Clearly (continuing) education programs designed to ensure that important lessons learned in the past are adequately disseminated, utilized, and acted upon, are a must.

9. Recommendations Presuming that the top level objective is to eliminate aircraft accidents caused by icing which could have been avoided if flight crews, operators, etc., had been alert to the potentially serious consequences of various forms of ice accretions, and hence taken appropriate precautionary actions, there are some important steps remaining to be taken. Foremost amongst these should be the following: *

*

*

Establishment of continuing/enduring training programs by government research and regulatory organizations to ensure that all designers, operators, flight crews, etc., are actually aware of all lessons which have been learned regarding the potentially serious consequences of almost any possible ice accretions, and know what actions are needed in case such ice buildups do occur. To be sure that the correct possible aerodynamic consequences are adequately defined, there are additional well conceived (primarily high Reynolds number) experimental programs still needing to be carried out. Actions need to be taken to ensure accountability for the sizeable adverse aerodynamic consequences of the

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*

*

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initial and inter-cycle ice accretions associated with the use of de-icing systems. Decisions need to be made regarding how to deal with the consequences of possible large droplet icing encounters. A strong focus should be placed on developing new certifiable ice detection systems/concepts that can reliably alert flight crews as to the onset of any ice accretions.

As indicated, there are a number of high priority, well conceived wind-tunnel test programs still needed to ensure that the upper limit of realistic aerodynamic penalties which could be incurred by representative lifting surfaces at flight conditions are properly quantified. These are as follows: *

*

*

*

For the small initial ice accretions which will always exist when de-icing systems are employed, or when there is a delay in activating anti-icing systems, it should be determined if penalties which would be incurred with the densely distributed (droplet) hemispheres believed to occur in natural icing conditions are much different from those indicated by the existing (large) database based almost exclusively on the use of less dense distributions of ‘‘sharp particles’’ to represent these initial accretions (roughness). A representative range of droplet sizes should be considered. Meaningful high Reynolds number test results (i.e., without uncorrectable wall effects, etc.) are needed for representative inter-cycle ice accretions in order to properly establish the magnitude of aerodynamic penalties inherent in the use of de-icing systems (which are all too often ignored). Since existing low Reynolds number test results indicate potentially catastrophic aerodynamic penalties can result from the runback/ridge ice accretions believed to be formed during large droplet icing encounters, it is clearly imperative that both the shapes, etc. of such ice accretions possible as well as the magnitude of the resulting aerodynamic penalties at flight-Reynolds-number-type conditions be determined. However, since properly representing the state of the leading-edge boundary layer is most certainly a critical requirement for establishing believable aerodynamic consequences for these accretions, the aerodynamic testing required for many vehicle types can realistically only be carried out in large cryogenic test facilities such as the NTF and/or ETW where leading-edge attachment line conditions existing in flight on swept wings can be duplicated. Although the existing available database of aerodynamic penalties indicated for large glaze ice accretions has enabled many very useful insights into the magnitude and form of such penalties, it is,

*

however, lacking in terms of identifying the likely upper limit of penalties which could be encountered in flight with the most critical ice accretions possible. That is both because of usually not really representing the most critical ice accretion conditions possible, and often not having unchallengable testing conditions, model installations, etc. Consequently, some new meaningful high Reynolds number test results are needed in order to better document the upper limit of penalties possible with large glaze ice accretions having the most critical accretion conditions in terms of droplet sizes, LWC, ambient temperature, aircraft speed and altitude, duration of encounter, etc. This is clearly an area where large droplet effects need to be established. Meaningful high Reynolds number test results are also needed for takeoff geometries having ‘‘hard’’ LEs in order to better establish (and illustrate) the potentially catastrophic effects of not adequately removing any ground frost/ice accretions with these configurations.

For each of the foregoing categories of additional tests needed, a sufficient number of lifting surface geometries should be addressed so as to represent the range of penalties possible considering both dimensional as well as nondimensional differences with existing as well as possible new designs.

Acknowledgements The authors wish to acknowledge and express our utmost appreciation for the very valuable contributions of Gene Hill and Jim Riley from the FAA, Mike Bragg from the University of Illinois, and our editor, Barry Haines, in terms of the constant encouragement to continue with this (long) review, supplying additional new pertinent data and other information, and in providing most helpful and insightful critiques of draft versions of the paper throughout its preparation. We would also like to recognize the significant value added to this review by Gene Addy from NASA Glenn in providing yet-to-be published high Reynolds number test results for a variety of ice accretion shapes. Note to reader While the authors have made every effort to review and include herein all known available and pertinent test results applicable to defining the important consequences of various types of ice accretions on aircraft aerodynamics, it is very possible because of our language and other limitations that we may have missed some. If there are additional pertinent data that are available which would be useful in either modifying or

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confirming positions established herein, researchers or others knowledgeable of such information are encouraged to contact the authors. Then, if appropriate, an addendum to this paper would be prepared reflecting the additional lessons learned from these ‘‘new’’ results.

References [1] Anon, Aircraft Accident Report, ATR Model 72–212, Roselawn, Indiana, October 31, 1994, NTSB/AAR-96/01, vol. 1, 1996. [2] Marwitz J, Politovich M, Berstein B, Ralph F, Neiman P, Ashenden R, Bresch J. Meterological conditions associated with the ATR72 aircraft accident near Roselawn, Indiana on october 31, 1994. Bull Am Meteorol Soc 1997;78(1):41–52. [3] Kind RJ, Potapczuk MG, Feo A, Golia C, Shah AD. Experimental and computational simulation of in-flight icing phenomena. Prog Aerosp Sci 1998;34:257–345. [4] Hansman RJ, Turnock SR. Investigation of surface water behavior during glaze ice accretion. AIAA J Aircr 1989;26(2):140–7. [5] Hansman RJ. Micro physical factors which influence ice accretion. Proceedings of the First Bombardier International Workshop on Aircraft Icing/Boundary-Layer Stability and Transition, Montreal, Canada, 1993. p. 86–103. [6] Wright WB. Users manual for the improved NASA lewis ice accretion code LEWICE 1.6. NASA CR-198355, 1995. [7] Cebeci T. Calculation of flow over iced airfoils. AIAA J 1989;27(7):853–61. [8] Chung J, Choo Y, Reehorst A, Potapczuck M, Slater J. Navier-stokes analysis of the flowfield characteristics of an ice contaminated aircraft wing. AIAA Paper 99-0375, 1999. [9] Joongkee C, Reehorst AL, Choo YK, Potapczuk MG. Effect of airfoil ice shape smoothing on the aerodynamic performance. AIAA Paper 98-3242, 1998. [10] Langmuir I, Blodgett KB. A mathematical investigation of water droplet trajectories. Army Airforce TR No. 5418, 1946. [11] Bourgault Y, Habashi WG, Dompierre J, Boutanios Z, DiBratolomeo W, An Eulerian approach to supercooled droplets impingement calculations. AIAA Paper 97-0176, 1997. [12] Lozowski EP, Oleskiw MM. Computer modeling of timedependent rime icing in the atmosphere. CRREL Report 83-2, USAF, 1983. [13] Ruff GA, Berkowitz BM. Users manual for the NASA lewis ice accretion code (LEWICE). NASA CR-185129, 1990. [14] Cransdale JT, Gent RW. Ice accretion on aerofoils in two-dimensional compressible flowFa theoretical model. RAE TR 82128, 1983. [15] Brunet L. Conception et discussion d’un modele de formation du givre sur des obstacles varies. ONERA Note Technique 1986-6, 1986.

763

[16] Brun EA. Icing problems, recommended solutions. AGARDograph, 16, 1957. p. 11. [17] Potapczuk M, Gent R, Guffond D. Review, validation and extension of ice accretion prediction codes. AGARD, AR-344, 1997. p. 5.1–5.13. [18] Brahimi MT, Tran P, Chocron D, Tezok F, Paraschiviou I. Effect of supercooled large droplets on ice accretion characteristics. AIAA Paper 97-0306, 1997. [19] Trunov OK. Icing of aircraft, the means of preventing it, Translated from Russian, foreign technology division. WPAFB, Ohio, 1967. p. 14. [20] Olsen W, Walker E. Experimental evidence for modifying the current physical model for ice accretion on aircraft surfaces. NASA TM 87184, 1986. [21] Shin J, Wilcox P, Chin V, Sheldon D. Icing tests on an advanced two-dimensional high-lift multi-element airfoil. AIAA Paper 94-1869, 1994. [22] Miller D, Shin J, Sheldon D, Khodadoust A, Wilcox P, Langahls T. Further investigation of icing effects on an advanced high-lift multi-element airfoil. AIAA Paper 951880, 1995. [23] Hill EG. private communication, February 2001. [24] Addy HE. Private communication, July 2000. [25] Tribus MV, Young GBW, Boelter LMK. Analysis of heat transfer over a small cylinder in icing conditions on Mount Washington. ASME Trans 1949;70:871–6. [26] Messinger BL. Equilibrium temperature of an unheated icing surface as a function of airspeed. J Aero Sci 1953;20(1):29–42. [27] Bilanin AJ, Anderson DN. Ice accretion with varying surface tension. AIAA Paper 95-0538, 1995. [28] Anderson DN. Rime-, mixed and glaze ice evaluations of three scaling laws. AIAA Paper 94-0718, 1994. [29] Anderson DN. Methods for scaling icing test conditions. AIAA Paper 95-0540, 1995. [30] Bragg MB, Lee S, Henze CM. Heat transfer and freestream turbulence measurements for improvement of the ice accretion physical model. AIAA Paper 97-0053, 1997. [31] Shin J. Characteristics of surface roughness associated with leading-edge ice accretion. AIAA J Aircr 1996;33(2):316–21. [32] Kerho MF, Bragg MB. Airfoil boundary-layer development and transition with large leading-edge roughness. AIAA J 1997;35(1):75–84. [33] Morgan HL, Ferris JC, McGhee RJ. A study of high-lift airfoils at high Reynolds numbers in the Langley lowturbulence pressure tunnel. NASA TM-89125, 1987. [34] Bilanin AJ. Proposed modifications to ice accretion/icing scaling theory. AIAA J Aircr 1991;28(6):353–9. [35] Anderson DN. Evaluation of constant-weber-number scaling for icing tests. AIAA Paper 96-0636, 1996. [36] Hauger HH, Englar KG. Analysis of model testing in an icing wind tunnel. Report SM14933, Douglas Aircraft Company, 1954. [37] Sibley PJ, Smith Jr., RE. Model testing in an icing wind tunnel. Report LR10981, Lockheed Aircraft Corporation, 1955. [38] Dodson EO. Scale model analogy for icing tunnel testing. Document D66-7976, Boeing Airplane Company, Transport Division, 1962.

764

F.T. Lynch, A. Khodadoust / Progress in Aerospace Sciences 37 (2001) 669–767

[39] Jackson ET. Development study: the use of scale models in an icing tunnel to determine the ice catch on a prototype aircraft with particular reference to concorde, SST/B75T/RMMcK/242, British Aircraft Corporation Ltd, Filton Division, 1967. [40] Armand C, Charpin F, Fasso G, Leclere G. Techniques and facilities used at the Onera Modane centre for icing tests. AGARDograph AR-127, 1978. [41] Ruff GA. Analysis, verification of icing scaling equations. AEDC TR-85-30, vol 1 (Rev), 1986. [42] Bilanin AJ, Anderson DN. Ice accretion with varying surface tension. AIAA Paper 95-0538, 1995. [43] Anderson DN. Rime-, mixed and glaze ice evaluations of three scaling laws. AIAA Paper 94-0718, 1994. [44] Anderson DN, Ruff GA. Evaluation of methods to select scale velocities in icing scaling testing. AIAA Paper 990244, 1999. [45] von Glahn UH. Use of truncated flapped airfoils for impingement, icing tests of full-scale leading-edge sections. NACA RM E56E11, 1956. [46] Saeed F, Selig S, Bragg MB. A hybrid airfoil design method to simulate full-scale ice accretion throughout a given Cc-range. AIAA Paper 97-0054, 1997. [47] Saeed F, Selig MS, Bragg MB. Design of subscale airfoils with full-scale leading edges for ice accretion testing. AIAA J Aircr 1997;34(1):94–100. [48] Saeed F, Selig S, Bragg MB. Hybrid airfoil design procedure validation for full-scale ice accretion simulation. AIAA J Aircr 1999;36(5):769–76. [49] Saeed F, Selig MS, Bragg MB. Hybrid airfoil design method to simulate full-scale ice accretion throughout a given a range. AIAA J Aircr 1998;35(2):233–9. [50] Tezok F, Brahimi MT, Paraschiviou I. Investigation of the physical processes underlying the ice accretion phenomena. AIAA Paper 98-0484, 1998. [51] Lynch FT, Valarezo WO, McGhee RJ. The adverse aerodynamic impact of very small leading-edge ice (roughness) buildups on wings and tails. AGARD-CP496, Paper 12, 1991. [52] Valarezo WO, Lynch FT, McGhee RJ. Aerodynamic performance effects due to small leading-edge ice (roughness) on wings and tails. AIAA J Aircr 1993;30(6):807–12. [53] Paschal K, Goodman W, McGhee R, Walker B, Wilcox PA. Evaluation of tunnel sidewall boundary-layer-control systems for high-lift airfoil testing. AIAA Paper 91-3243, 1991. [54] Lynch FT, Crites RC, Spaid FW. The crucial role of wall interference, support interference, and flow field measurements in the development of advanced aircraft configurations. AGARD CP-535, 1993. [55] Brumby RE. The effect of wing ice contamination on essential flight characteristics. SAE Aircraft Ground DeIcing Conference, 1988. [56] Brumby RE. The effect of wing ice contamination on essential flight characteristics. AGARD CP-496, Paper 2, 1991. [57] Valarezo WO, Dominik CJ, McGhee RJ, Goodman WL. High Reynolds number configuration development of a high-lift airfoil. AGARD CP-515, 1992. [58] Valarezo WO. Topicsin high-lift aerodynamics. AIAA Paper 93-3136, 1993.

[59] Chin VD, Peters DW, Spaid FW, McGhee RJ. Flowfield measurements about a multi-element airfoil at high Reynolds numbers. AIAA Paper 93-3137, 1993. [60] Klausmeyer SM, Lin JC. An experimental investigation of skin friction on a multi-element airfoil. AIAA Paper 94-1870, 1994. [61] Lynch FT. subsonic transport high-lift technologyFreview of experimental studies. AIAA Overview of High Lift Aerodynamics, 1995. [62] Lynch FT, Potter RC, Spaid FW. Requirements for effective high lift CFD. 20th ICAS, 1996. [63] Ying SX, Spaid FW, McGinley CB, Rumsey CL. Investigation of confluent boundary layers in high lift flows. AIAA J Aircr 1999;36(3):550–62. [64] Dow Sr., JP. Roll upset in severe icing. Federal Aviation AdministationFAircraft Certification Service, 1995. [65] Jacobs EN. Airfoil section characteristics as affected by protuberances. NACA Report 446, 1932. [66] Abbott IH, von Doenhoff, Albert E, Stivers LS. Summary of airfoil data. NACA Report 824, 1945. [67] Loftin LK, Smith HA. Aerodynamic characteristics of 15 NACA airfoil sections at seven Reynolds numbers from 0.7 million to 9.0 million. NACA TN 1945, 1949. [68] Ladson CL. Effects of independent variation of Mach, Reynolds numbers on the low-speed aerodynamic characteristics of the NACA 0012 airfoil section. NASA TM 4074, 1988. [69] Loftin Jr., LK, Bursnall WJ. The effects of variations in Reynolds number between 3.0 million and 25.0 million upon the aerodynamic characteristics of a number of NACA 6-series airfoil sections. NACA Report 964, 1950. [70] Abbott Jr., FT, Turner Jr., HR. The effects of roughness at high Reynolds numbers on the lift and drag characteristics of three thick airfoils. NACA ACR No. L4H21, 1944 (Wartime Report No. L-46). [71] Mavriplis F. Icing, Contamination of aircraft surfacesFindustry’s concerns. Proceedings of the First Bombardier International Workshop on Aircraft Icing/ Boundary-Layer Stability and Transition, Montreal, Canada, 1993. p. 51–5. [72] Bragg MB, Gregorek GM. Environmentally induced surface roughness on laminar flow airfoils: implications for flight safety. AIAA Paper 89-2049, 1989. [73] Gulick BG. Effects of simulated ice formation on the aerodynamic characteristics of an airfoil. NACA WR L-292, 1938. [74] Ingleman-Sundberg M, Trunov OK, Ivaniko A. Methods for prediction of the influence of ice on aircraft flying characteristics. Report JR-1, a joint report from the Swedish-Soviet Working Group on Flight Safety, 1977. [75] Bragg MB, Khodadoust A, Kerho M. Aerodynamics of a finite wing with simulated ice. Computational and Physical Aspects of Aerodynamic Flows, Long Beach, CA, 1992. [76] Ferguson S, Mullins Jr., B, Smith D, Korkan K. Fullscale empennage wind tunnel test to evaluate effects of simulated ice on aerodynamic characteristics. AIAA Paper 95-0451, 1995. [77] Trunov OK, Ingleman-Sundberg M. Wind tunnel investigation of hazardous tail stall due to icing. Report

F.T. Lynch, A. Khodadoust / Progress in Aerospace Sciences 37 (2001) 669–767

[78]

[79]

[80]

[81]

[82]

[83]

[84]

[85]

[86]

[87]

[88] [89]

[90]

[91]

[92]

[93]

[94]

[95]

JR-2, a joint report from the Swedish-Soviet Working Group on Flight Safety, 1979. Trunov OK, Ingleman-Sundberg M. On the problem of horizontal tail stall due to ice. Report JR-3, a joint report from the Swedish-Soviet Working Group on Flight Safety, 6th Meeting, 1985. Papadakis M, Gile-Laflin BE. Aerodynamic performance of a tail section with simulated ice shapes and roughness. AIAA Paper 2001-0539, 2001. Papadakis M, Yeong HW, Chandrasekharan R, Hinson M, Ratvasky TP, Giriunas J. Experimental investigation of simulated ice accretions on a full-scale T-tail. AIAA Paper 2001-0090, 2001. Hiltner D, McKee M, LaNoe R, Gregorek G. DHC-6 twin otter tailplane airfoil section testing in the Ohio state university 7
10 wind tunnel. NASA CR-2000-20992 (2 vols), 2000. Valarezo WO. Maximum lift degradation dueto wing upper surface contamination. Proceedings of the First Bombardier International Workshop on Aircraft Icing/ Boundary-Layer Stability and Transition, Montreal, Canada, 1993. p. 104–12. Boer JN, Van Hengst J. Aerodynamic degradation due to distributed roughness on high lift configuration. AIAA Paper 93-0028, 1993. Myers TT, Klyde DH, Magdelano RE. The dynamic icing detection system (DIDS). AIAA Paper 2000-0364, 2000. Johnson CL. Wing loading, icing, and associated aspects of modern transport design. J Aero Sci 1940;8(2): 43–55. Bowden DT. Effect of pneumatic de-icers, ice formations on aerodynamic characteristics of an airfoil. NACA TN 3564, 1956. Cooper WA, Sand WR, Politovich MR, Veal DL. Effects of icing on performance of a research airplane. AIAA J Aircr 1984;21(9):708–15. Politovich MR. Aircraft icing caused by large supercooled droplets. J Appl Meteorol 1989;28(9):856–68. Politovich MR. Response of a research aircraft to icing, evaluation of severity indices. AIAA J Aircraft 1996;33(2):291–7. Ashenden R, Marwitz J. Turboprop aircraft performance response to various environmental conditions. AIAA J Aircr 1997;34(3):278–87. Welte D, Wohlrath W, Seubert R, DeBartolomeo W, Toogood RD. Preparation of the ice certification of the Dornier 328 regional airliner by numerical simulation and by ground test. AGARD CP-496, Paper 14, 1991. Ashenden R, Lindberg W, Marwitz J, Hoxie B. Airfoil performance degradation by supercooled cloud, drizzle, and rain drop icing. AIAA J Aircr 1996;33(6):1040–6. Lee S, Bragg MB. Effects of simulated-spanwise ice shapes on airfoils: experimental investigation. AIAA Paper 99-0092, 1999. Lee S, Bragg MB. Experimental investigation of simulated large-droplet ice shapes on airfoil aerodynamics. AIAA J Aircr 1999;36(5):844–50. Lee S, Bragg MB. Private communication, December 2000 (same data subsequently published in AIAA Paper 2001-2421, 2001).

765

[96] Addy H, Chung J. A wind tunnel study of icing on a natural laminar flow airfoil. AIAA Paper 2000-0095, 2000. [97] Calay RK, Hold AE, Mayman P, Lunn I. Experimental simulation of runback ice. AIAA J Aircr 1997;34(2): 206–12. [98] Anon, Report of the 12A Working Group on Determination of Critical Ice Shapes for the Certification of Aircraft. DOT/FAA/AR-00/37, 2000. [99] Morris DE. Designing to avoid dangerous behavior of an aircraft due to the effects on control hinge moments of ice on the leading edge of the fixed surface. RAE TN-Aero 1878, 1947. [100] Worrall, Cole. Wind tunnel measurements of the effect of ice formations on the hinge moment characteristics of the Viking elevator. ARC 10691, 1947. [101] Ratvasky TP, van Zante JF, Riley JT. NASA/FAA tailplane icing program overview. AIAA Paper 99-0370, 1999. [102] Preston GM, Blackman CC. Effects of ice formations on airplane performance in level cruising flight. NACA TN 1598, 1948. [103] Leckman PR. Qualification of light aircraft for flight in icing conditions. SAE Paper 710394, 1971. [104] Teymourazov R, Kofman V. The effect of ice accretion on the wing and stabilizer on aircraft performance, unidentified USSR Paper. [105] Ranaudo RJ. Mikkelsen KL, McKnight RC, Perkins PJ. Performance degradation of a typical twin engine commuter type aircraft in measured natural icing conditions. AIAA Paper 84-0179 (also NASA TM 83564), 1984. [106] Mikkelsen KL, McKnight RC, Ranaudo RJ, Perkins PJ. Icing flight research: aerodynamic effects of ice and ice shape documentation with stereo photography. AIAA Paper 85-0468 (also NASA TM 86906), 1985. [107] Ranaudo RJ, Mikkelsen KL, McKnight RC, Ide RF, Reehorst AL, Jordan WC. Schinstock WC, Putz SJ. The measurement of aircraft performance and stability and control after flight through natural icing conditions. AIAA Paper 86-9758, 1986. [108] Brown AP. Analysis of the aerodynamic effects of freezing drizzle inflight icing on a turboprop aircraft. AIAA Paper 99-3151, 1999. [109] Cook DE. Unusual natural icing encounters during boeing 777 flight tests. AIAA Paper 97-0304, 1997. [110] Obert E. Low-speed stability, control characteristics of transport aircraft with particular reference to tailplane design. AGARD CP-160, Paper 10, 1975. [111] Ranaudo RJ, Batterson JG, Reehorst AL, Bond TH, O’Mara TM. Determination of longitudinal aerodynamic derivatives using flight data from an icing research aircraft. AIAA Paper 89-0754 (also NASA TM 101427), 1989. [112] Ranaudo RJ, Batterson JG, Reehorst AL, Bond TH, O’Mara TM. Effects of horizontal tail ice on longitudinal aerodynamic derivatives. AIAA J Aircr 1991;28(3):193–9. [113] Ratvasky TP, Ranaudo RJ. Icing effects on aircraft stability and control determined from flight data. AIAA Paper 93-0398 (also NASA TM 105977), 1993.

766

F.T. Lynch, A. Khodadoust / Progress in Aerospace Sciences 37 (2001) 669–767

[114] Van Zante JF, Ratvasky TP. Investigation of dynamic flight maneuvers with an iced tailplane. AIAA Paper 99-0371 (also NASA/TM-1999-208849), 1999. [115] Ratvasky TP, Van Zante JF. In-flight aerodynamic measurements of an iced horizontal tailplane. AIAA Paper 99-0638 (also NASA/TM-1999-208902), 1999. [116] Ratvasky TP, Van Zante JF, Sim A. NASA/FAA tailplane icing program: flight test report. NASA/TP2000-209908 (also DOT/FAA/AR-99/85), March 2000. [117] Ranaudo RJ, Ratvasky TP, Van Zante JF. Flying qualities evaluation of a commuter aircraft with an ice contaminated tailplane. SAE Paper 2000-01-1676 (also NASA/TM-2000-210356), 2000. [118] Bragg MB, Hutchinson T, Merret J, Oltman R, Pokhariyal D. Effect of ice accretion on aircraft flight dynamics. AIAA Paper 2000-0360, 2000. [119] Gray VH, von Glahn UH. Effect of ice and frost formations on drag of NACA 651-212 airfoil for various modes of thermal ice protection. NACA TN 2962, 1953. [120] Gray VH. Correlations among ice measurements, impingement rates, icing conditions and drag of a 65A004 airfoil. NACA TN 4151, 1958. [121] Gray VH, von Glahn UH. Aerodynamic effects caused by icing of an unswept NACA 65A-004 airfoil. NACA TN 4155, 1958. [122] Gray VH. Prediction of aerodynamic penalties caused by ice formation on various airfoils. NASA TN D-2166, 1964. [123] Bragg MB, Gregorek GM, Shaw RJ. Wind tunnel investigation of airfoil performance degradation due to icing. AIAA Paper 82-0582, 1982. [124] Shaw RJ, Sotos RG, Solano FR. An experimental study of airfoil icing characteristics. NASA TM 82790, 1982. [125] Olsen W, Shaw RJ, Newton J. Ice shapes and the resulting drag increase for a NACA 0012 airfoil. NASA TM 83556, 1984. [126] Potapczuk MG, Berkowitz BM. An experimental investigation of multielement airfoil ice accretion and resulting performance degradation. AIAA Paper 89-0752 (also NASA TM 101441), 1989. [127] Potapczuk MG, Berkowitz BM. Experimental investigation of multielement airfoil ice accretion and resulting performance degradation. AIAA J Aircr 1990;27(8):679– 91. [128] Berkowitz BM, Potapczuk MG, Namdar BS, Langhals TJ. Experimental ice shape and performance characteristics for a multielement airfoil in the NASA Lewis icing research tunnel. NASA TM 105380, 1991. [129] Shin J, Bond TH. Results of an icing test on a NACA 0012 airfoil in the NACA Lewis icing research tunnel. AIAA Paper 92-0647 (also NASA TM 105374), 1992. [130] Shin J, Bond TH. Experimental and computational ice shapes and resutling drag increase for a NACA 0012 airfoil. NASA TM 105743, 1992. [131] Addy HE, Potapczuk MG, Sheldon DW. Modern airfoil ice accretions. AIAA Paper 97-0174 (also NASA TM 107423), 1997. [132] Addy HE. Ice accretions, icing effects for modern airfoils. NASA/TP-2000-210031 (also DOT/FAA/AR-99/89), 2000. [133] Addy HE. Private communication, 2000.

[134] Bragg MB, Gregorek GM. Aerodynamic characteristics of airfoils with ice accretions. AIAA Paper 820282, 1982. [135] Bragg MB, Gregorek GM, Lee JD. Airfoil aerodynamics in icing conditions. AIAA J Aircr 1986;23(1):76–81. [136] Bragg MB, Zaguli RJ, Gregorek GM. Wind tunnel evaluation of airfoil performance using simulated ice shapes. NASA CR 167960, 1982. [137] Wickens RH, Nguyen VD. Wind tunnel investigation of wing-propeller model performance degradation due to distributed upper surface roughness and leading edge shape modification. AGARD CP-496, Paper 11, 1991. [138] Bragg MB, Coirier WJ. Detailed measurements of the flow field in the vicinity of an airfoil with glaze ice. AIAA Paper 85-0409, 1985. [139] Bragg MB, Coirier WJ. Aerodynamic measurements of an airfoil with simulated glaze ice. AIAA Paper 86-0484, 1986. [140] Bragg MB. Experimental aerodynamic characteristics of an NACA 0012 airfoil with simulated glaze ice. AIAA J Aircr 1988;25(9):849–54. [141] Bragg MB, Khodadoust A. Experimental measurements in a large separation bubble due to a simulated glaze ice accretion. AIAA Paper 88-0116, 1988. [142] Bragg MB, Khodadoust A. Effect of simulated glaze ice on a rectangular wing. AIAA Paper 89-0750, 1989. [143] Khodadoust A, Bragg MB, Kerho M, Wells S, Soltani MR. Finite wing aerodynamics with simulated glaze ice. AIAA Paper 92-0414, 1992. [144] Khodadoust A, Bragg MB. Measured aerodynamic performance of a swept wing with a simulated ice accretion. AIAA Paper 90-0490, 1990. [145] Bragg MB, Khodadoust A, Soltani MR, Wells S, Kerho M. Effect of simulated ice accretion on the aerodynamics of a swept wing. AIAA Paper 91-0442, 1991. [146] Papadakis M, Gile Laflin BE, Youssef GM, Ratvasky TP. Aerodynamic scaling experiments with simulated glaze ice accretions. AIAA Paper 2001-0833, 2001. [147] Jackson DG, Bragg MB. Aerodynamic performance of an NLF airfoil with simulated ice. AIAA Paper 99-0373, 1999. [148] Gile Laflin BE, Papadakis M. Experimental investigation of simulated ice accretions on a natural laminar flow airfoil. AIAA Paper 2001-0088, 2001. [149] Papadakis M, Alansatan S, Seltman M. Experimental study of simulated ice shapes on a NACA 0011 Airfoil. AIAA Paper 99-0096, 1999. [150] Papadakis M, Alansatan S, Wong S-C. Aerodynamic characteristics of a symmetric NACA section with simulated ice shapes. AIAA Paper 2000-0098, 2000. [151] Kim HS, Bragg MB. Effect of leading-edge ice accretion geometry on airfoil performance. AIAA Paper 99-3150, 1999. [152] Lee S, Kim HS, Bragg MB. Investigation of factors that influence iced-airfoil aerodynamics. AIAA Paper 20000099, 2000. [153] Papadakis M, Alansatan S, Yeong HW. Aerodynamic performance of a T-tail with simulated ice accretions. AIAA Paper 2000-0363, 2000. [154] Laschka B, Jesse RE. Determination of ice shapes and their effect on the aerodynamic characteristics

F.T. Lynch, A. Khodadoust / Progress in Aerospace Sciences 37 (2001) 669–767

[155]

[156]

[157]

[158]

[159] [160] [161]

[162] [163]

[164]

for the unprotected tail of the A300. ICAS Paper 74-42, 1974. Laschka B, Jesse RE. Ice accretion and its effects on aerodynamics of unprotected aircraft components. AGARD AR-127, Paper 4, 1978. Ingelman-Sundberg M. Tail plane stallFwhat Isthat. FFAP-A-925 (aeronuatical research institute of Sweden), 1991. Reehorst A, Potapczuk M, Ratvasky T, Gile Laflin B. Wind tunnel measured effects on a twin-engine short-haul transport caused by simulated ice accretions. AIAA Paper 96-0871, 1986. Khodadoust A, Dominik C, Shin J, Miller D. Effect of inflight ice accretion on the performance of a multi-element airfoil. Proceedings of AHS/SAE International Icing Symposium, 1995. Thomas JL. (NASA Langley). Private Communication, April 2001. Bragg MB. Rime ice accretion, its effect on airfoil performance. NASA CR-165599, 1982. Miller TL, Korkan KD, Shaw RJ. Statistical study of an airfoil glaze ice drag coefficient correlation. SAE Paper 830753, 1983. Flemming RJ, Lednicer DA. High speed ice accretion on rotorcraft airfoils. NASA CR-3910, 1985. Cook DE. Relationships of ice shapes and drag to icing condition dimensionless parameters. AIAA Paper 20000486, 2000. Zierten TA, Hill EG. Effects of wing simulated ground frost on aircraft performance. VKI Lecture Series, 1987.

767

[165] van Hengst J, Boer JN. The effect of hoar-frosted wings on the Fokker 50 take-off characteristics. AGARD-CP496, Paper 13, 1991. [166] Hill EG. Private communication, August 2001. [167] Bragg MB, Heinrich DC, Valarezo WO, McGhee RJ. Effect of underwing frost on a transport aircraft airfoil at flight Reynolds number. AIAA J Aircr 1994;31(6): 1372–9. [168] Jones R, Williams DH. The effect of surface roughness on the characteristics of the aerofoils. NACA 0012 and RAF 34, ARC, R&M 1708, 1936. [169] Morelli JP. Flight test wing surface roughness effects for development of an airfoil surface contamination detection system. MDC Report 96K9377, 1996. [170] Hooker RW. The aerodynamic characteristics of airfoils as affected by surface roughness. NACA TN 457, 1933. [171] Hoerner SF, Borst HV. Fluid dynamic lift. 1985, pp. 4–16 to 4–22. [172] Ljungstrom BLG. Wind tunnel investigation of simulated hoar frost on a 2-dimensional wing section with and without high lift devices. FFA-AU-902, 1972. [173] van Hengst J. Aircraft de-icing, future technical developmentsFan aircraft manufacturer’s point of view. Wingtips (Fokker Aircraft) No. 26, 1993. [174] van Hengst J, Gent R, Hammond D, Seubert R, Wagner B. Ice accretion and its effects on aircraft. AGARD, AR-344, 1997. p. 3.1–3.36.