Problems of cabin noise estimation for supersonic transports

J. Sound F’&. (1964) I (z), 113-126

PROBLEMS

OF CABIN NOISE ESTIMATION SUPERSONIC

FOR

TRANSPORTS-f

J. A. HAY Aerodynamics

Department, Vickers-Armstrong (Aircraft) Weybridge, Surrey, England (Received

I 8 June,

Limited,

I 963)

A discussion is given of the problems associated with the estimation of the internal cabin noise levels of supersonic transports. In particular, it is shown that existing design data on boundary layer pressure fluctuations and their transmission through aircraft structures are inadequate. Use of existing data can lead to estimated internal noise levels 15 to 20 dB higher than experimental values.

INTRODUCTION

The object of this note is to discuss the problems associated with the estimation of the internal cabin noise levels of supersonic transports. In fact, what it actually presents is the extreme difficulty of making any reasonably accurate estimate of these noise levels, and it is presented mainly in the hope that it might inspire more work aimed at solving these problems. The difficulties of obtaining reasonable estimates of noise levels in aircraft cabins are not in any way peculiar to supersonic aircraft. They exist at subsonic speeds, and are merely complicated further by flight at supersonic speeds. Since we are concerned with supersonic transports, a supersonic aircraft, not in any way related to any specific aircraft project, has been taken to illustrate the problems. THE

AIRCRAFT

AND

THE

NOISE

SOURCES

The aircraft selected (see Figure I) is the sort of machine that might be used for transatlantic operation, and would carry about 28,000 lb of payload, or 140passengers, from London to New York. A slender delta configuration has been selected, which would fly at a Mach number of 2-2 at a mean cruise altitude of about 55,000 ft. The noise levels will be calculated for this cruise condition. If we consider the possible sources of cabin noise on this aircraft, we find there are a considerable number of them. They are as follows. I.

ENGINE NOISE SOURCES

This can be subdivided (a) jet noise; (b) engine noise-mainly (c) compressor noise.

into : from the turbine

Here it should be noted that, although travelling at supersonic speeds cannot

noise from rear mounted engines on an aircraft be directly radiated to the cabin through the

t This paper was read at the Symposium on Noise Problems of Supersonic Transport Aircraft, conducted by the Department of Aeronautics and Astronautics of the University of Southampton, on rgth and 20th April, 1963. 9

rr3

114

Figure

I.

Slender

delta supersonic

airliner.

Figure

2. Noise sources on a supersonic - - - : Shock waves. aircraft.

external supersonic airstream, nevertheless in cases such as that illustrated transmission paths exist from the engine to the cabin through the interior of the wing and rear fuselage, and along the structure itself. 2. SYSTEMS NOISE SOURCES (a) Pressurizing and air conditioning systems not only produce noise, but (most unfairly) use the distribution ducts to conduct it directly to the passengers ! Noise may also be produced in the ducts themselves, and may be.radiated from the duct surfaces. (b) Hydraulic systems, electrical equipment, and cooling fans all produce their quota of noise. 3. AERODYNAMIC NOISE SOURCES These include : (a) the boundary layer, which envelopes the whole aircraft ; (b) vortex noise-wings with highly swept leading edges generate vortices when producing appreciable lift and high intensity pressure fluctuations exist beneath these vortices; these may or may not be present on this type of configuration; (c) shock wave noise-caused perhaps by shock waves oscillating on the aircraft surfaces, or possibly by an increase in the turbulence of the boundary layer as it passes through the shock waves. CONSIDERATION

OF

THE

KOISE

SOURCES

Considering these noise sources, we find we can eliminate some from our calculation of cabin noise. Others we have to ignore for the time being, knowing they may well have an appreciable effect on the result, but lacking sufficient knowledge to make an adequate estimate of their effects. In this way we can assume that the engines can be ignored because the jet noise is generated behind the aircraft and cannot catch it up. The transmission path of engine and compressor noise along and within the structure is tortuous, and we can hope there will be sufficient attenuation to render these sources not too serious. Systems noise is not so easily dismissed. Figure 3 shows the difference in noise level in the forward cabin of a transatlantic jet aircraft caused by operating the aircraft air conditioning and pressurization systems. Other aircraft systems, such as those employed to

SUPERSONIC

TRANSPORT

CABIN

NOISE

115

cool specific equipment, and components of the hydraulic and electrical systems can also contribute to the internal noise levels. Unfortunately we have no satisfactory method for tackling this problem of systems noise estimation and have to rely on common sense applied in the aircraft design stage, coupled with remedial action at a later date, if required. For our supersonic aircraft exercise we will have to assume that this problem will be satisfactorily dealt with in this manner. We hope that on this type of aircraft, the shock waves on the fuselage will be reduced to one at the nose of the fuselage, and some on the rear fuselage aft of the passenger cabin, produced at the wing trailing edge and by the engine nacelles. The passenger cabin should, therefore, be free from shock wave noise. Similarly, it may be anticipated that at the low operating lift coefficients used under cruising conditions the wing will be free from detached vortex flow. With this type of configuration, with the wing between the passengers and the wing upper surface, it is unlikely to be a troublesome source in any case, but with some other configurations it could be important, therefore it is worthwhile making a brief inspection of the available

Figure 3. Typical systems noise. M = 0.85 ; H = 35,000 ft. data relating to this source. This is very scanty but some work has been done at the Royal Aircraft Establishment and by Dr. J. P. Jones at Southampton University. Some of these results are presented in Figure 4 (a) and (b). It will be seen that at a typical cruise incidence of three or four degrees the intensity of the pressure fluctuations below the vortex core are about three times as large as those on the surrounding surface. If these were to occur close to the wing root, very high noise intensities could be radiated into the wing interior, and possibly transmitted to the fuselage. The peak intensity of this noise would be around ISO dB at M = 2.2, H = 55,000 ft. However, as stated, we will assume that separated vortex flow is avoided by suitable wing design. Thus, we appear to be doing very well in having reduced our original list of six or so noise sources to only one, the boundary layer. Any satisfaction at this result is, however, ill advised. The real problems are only just starting. BOUNDARY

LAYER

PRESSURE

FLUCTUATIONS

Let us examine the available data on boundary layer noise. To simplify the problem further we will assume that the boundary layer which envelopes the fuselage is fully turbulent, and contains no separated regions. A certain amount of data, largely measured

116

J.

.I.

H.lI

Incidence Vortex

(a)

breakdown

(b)

Figure 4. (a) Typical surface distribution of space correlation coefficient under the separation vortex of a sharp edged flat plate delta wing (0.8 root chord aft of apex) ; (b) typical variation with incidence of r.m.s. surface pressure fluctuations on a sharp edged flat plate delta wing.

in wind tunnels at low Mach number but containing a few flight and high Mach number points, exists. The results of these researches can be split into two parts, the first part relating to the magnitude of the overall pressure fluctuation intensity in the boundary layer, and the second to its spectral distribution. With regard to the intensity, the various experiments have produced a fairly wide variety of results (Figure 5). It can be seen that the maximum and minimum values are

Work

performed

by

Type

of Work

.\/(pz, x

IO3

4

Willmarth Harrison Lilley and Hodgson Williams Mull and Algranti McLeod and Jordan von Gierke Webb

Low speed wind tunnel Low speed wind tunnel Low speed wind tunnel Supersonic (M = 1.4) wind tunnel In flight measurement (subsonic) In flight measurement (subsonic) In flight measurement (up to M = I. I 6) In flight measurement (up to M = 1.5)

6-0

9’5 8.0

4.75 to 6.9 1’3 3’0 to 5’0 6.0 4.2 to 5.5

p = freestream density; V = freestream p = Mean square fluctuating pressure ; q= $pVz; velocity. Figure 5. Overall r.m.s. pressure fluctuations on boundary surface-scatter of intensity data.

in the ratio of about 6: I. This is not a very good start for our calculation. However, we must select some value to start with, so we will take the most generally accepted value of dp/q = 0.006. How does the effect of flight at M = z-z at 55,000 ft compare with that of the current transatlantic jets on this basis ? Figure 6 shows that the effect of altitude compensates largely for the increased speed, and the overall boundary layer noise intensity on the supersonic transport need only be some 8 dB greater than that on current aircraft. It is

SUPERSONIC

TRANSPORT

CABIN

NOISE

I’7

worth noting that this effect also suggests that the end of the cruise may be quieter than the beginning, due to the increase of altitude that occurs during the cruise. Next, let us see how these pressure fluctuations may be distributed through the frequency spectrum (Figure 7). Here again there is wide scatter, particularly at the higher frequencies. This, however, is not the full story on boundary layer noise. 70,000

I

60,000 -

High sub-sonic jet transport cruise.,, range

‘s

%

2 Q

50’ooo~ 40,000

30,000

20,000_

-

0

0.5

IO

l-5

2.0

_I

2-5

Mach no

Figure

6. Noise level with variation

of altitude and Mach

number.

J

Figure 7. Frequency distribution of r.m.s. pressure on boundary surface-scatter of data. P(f) = Spectral density of mean square fluctuating pressure ; u = free stream or aircraft velocity; 6% = boundary layer displacement thickness ; f = frequency (cps) ; q = 4pz42;p = free stream density. Wind Tunnel : I Willmarth, 2 Harrison, 3 Bull ; Flight : 4 McLeod and Jordan, 3 Mull and Algranti, 6 von Gierke; 7 Bull (IO January 1963); 8 Mean.

There is some flight data now predominantly wind tunnel data conjunction with Professor Lilley other sources, indicate a fall off this suggest that the tunnels are

available which points to a trend not present in the so far available. Results obtained by Dr. Hodgson in at the College of Aeronautics, as well as results from in intensity at the lower frequencies (Figure 8). Does possibly not giving the correct picture at the lower

J.

118 frequencies

A. H.11

? The latest tunnel results of Dr. 31. I<. Bull at Southampton

show a tendency

in this direction.

This

effect

could

he important

University

also

in our cabin noise

predictions and needs clarifying. _4gain we must take some mean curve for our calculation and will take that used at Vickers-armstrongs, as shown in Figure 7. Having now selected a mean non-dimensional spectrum, we are faced with the problem of applying it to our supersonic aircraft. The greater part of this data has been obtained

Figure 8. Frequency distribution of r.m.s. pressure on boundary surface-Hodgson’s results. P(j) = Spectral density of mean square fluctuating pressure ; f = frequency (cps) ; p = free stream density. u = Free stream or aircraft velocity; 6” = boundary layer displacement thickness ; q = $pu*.

from low Mach number experiments, and it has been correlated on the basis of boundary layer displacement thickness as the typical length. It could equally well have been boundary layer thickness. At a given Reynolds number the total boundary layer thickness is almost the same for low Mach numbers and M = 2.2 (Figure 9), but the displacement

Distance from

Figure

9. Boundary

layer thickness

leodmg

and displacement

edge

:ftl

thickness.

M = 2.2;

H = 55,000

ft.

thickness is twice as great at M = 2.2 as at M = o. Depending on which is the more correct method to allow for Mach number effects, there is a possible 2 : I factor on frequency! For the purpose of this exercise, the standard form of the non-dimensional spectrum will be adhered to, and the compressible value of 6” taken. Taking, then, the mean value of 47/q = 0.006, the mean boundary layer nondimensional spectrum, and using the compressible value of boundary layer displacement thickness, the actual intensity and spectral distribution of the boundary layer pressure fluctuations can be calculated. The results of such a calculation at four stations on the supersonic transport fuselage are shown in Figure IO. The tendency towards more high frequency content towards the front of the aircraft, and greater low frequency excitation

SUPERSONIC

TRANSPORT

CABIN

lI:I:oo 357

ii:,

150 300

300 600

1%

Octave

119

NOISE

1200 2400

4800

9600

bands kps)

Figure IO. Boundary layer noise spectra on supersonic transport fuselage. Mid-cruise condition: M = 2.2; H = 55,000 ft.

towards the rear is typical of boundary layer noise at both sub- and supersonic speeds, and might influence the distribution of sound-proofing treatment on an actual aircraft. Having established the external distribution of pressure fluctuations, the next step is to determine how much of this is transmitted through the structure and sound-proofing treatment. STRUCTURAL

TRANSMISSION

Here we immediately encounter the next problem. The classical treatment for the transmission loss through structural panels is based on consideration of their response to sound waves. Sound waves are of large extent, compared with practical panel sizes, and move progressively across the panel, but the boundary layer pressure fluctuations are caused by eddies in the boundary layer beating on the panel as they move across it. The dimensions of these eddies are less than the boundary layer thickness, and therefore

Mdin frame

Minor

Skin

frame

Figure

0 056”

I I.

Typical supersonic aircraft structure.

Is the panel going to respond to these pressure fluctuations in the same way as it does to sound waves ? We can only assume for the time being that it will, but this should be borne in mind as a possible reason for discrepancies in the results of these calculations. Figure I I shows a typical panel structure of the sort which might be used in a supersonic aircraft. In Figure 12 is shown the soundproofing treatment assumed in this calculation.

rather smaller than the panel dimensions.

J. A. HAT

120

In lyigure 13 the generalized form for the response of panels subjected to soundvvaves is given. At the lower frequencies the response of the panel is dictated by its stiffness. As the frequency increases the fundamental resonant frequency is reached. At this frequency the panel responds to the exciting force to an extent dependent only on the damping present, and the transmission of sound waves through the panel reaches a maximum. This frequency is dependent on the edge supports of the panel, and is twice as high for a

1.9”

f r-c #
J

,

_

I -------

\ P~47~urethonefoam

o Yi’ I ‘9” 0,25”

Figure

Y h 12.

41:

-. _. ~~-

_____---_

Supersonic

,vdltIoning

duct

,'olyi-ethone

focm Final preformed plastic

transport

typical

acoustic

and thermal

trim

insulation.

panel whose edges are restrained in bending as for one supported on knife edges. Where does the typical aircraft structural panel fit between these? For our calculation we will take the degree of fixation implied in reference (I). Above this frequency there are a number of other resonances, until a band of frequencies is reached where the response of the panel is dictated essentially by its mass. At higher frequencies still, coincidence effects appear

Figure

13. Classical

panel transmission

loss characteristic.

where sound waves striking the surface at various angles of incidence traverse the surface at a speed equal to that of waves in the skin. Figure 14 shows the transmission characteristics of our sample structure. We can use this in two different ways. Firstly, account can be taken of the transmission of the structure and treatment fully allowing for the resonances. Secondly, the resonances can be ignored and we can assume we will have the full benefit of the mass effect. There is some justification for this. To begin with, a certain but ill-determined amount of natural damping exists in the structure due to friction in the structural joints. The

SUPERSONIC TRANSPORT

effect of damping is to reduce the response work at Southampton University shown (Figure I 5).

Stiffness

+--

121

CABIN NOISE

at the resonant frequency, as illustrated by some recently in a paper by Richards and Doak (2)

-----+

controlled

_____

___ iG0

0

1000 Frequency

Figure

14.

Supersonic

transport

10,000

kps)

cabin wall transmission

loss.

The next point is that our supersonic aircraft is pressurized, and some other work at Southampton University has shown the effect of pressurization may be to practically eliminate the resonant response (Figure 16). Finally, we shall see that our calculation

Frequency

Figure

15. Effect

kps)

of sound damping

tape on z-frame

panel.

over-estimates the internal noise levels, and ignoring these resonances gives an answer a little nearer the values we actually measure ! Applying this calculated transmission loss data to the calculated external boundary

J. A. I-I.11

122

layer spectra, and allowing for representative absorption within the cabin, we arrive (Figure 17) at the spectra at the four stations in the aircraft cabin. The result is very disappointing and the levels appear very high, indeed being of the order of I IO dB if account is taken of resonant panel response, and around IOO dB if it is

Figure 16. Response of panel to random excitation when under a static -0.03 Panel thickness = 0.005 in. Xfo, Y = o. Pressure differences:

pressure differential. psi; ---2.4 psi.

ignored. How do they compare with similarly calculated values on existing aircraft, and how do these compare with measured noise levels ? In Figure 18 are shown results for a prop-turbine aircraft in gliding flight, where engine noise is not a major noise source. It will be seen that the measured values are some

Figure

17. Supersonic

transport

cabin noise due to boundary

layer (at 4 stations).

15 to 20 dB below the theoretical ones. In Figure 19 are the corresponding values for a contemporary jet aircraft in the mid-cabin. Here we again find a 15 to 20 dB difference, and we also see that ignoring the resonances of the panels reduces the discrepancy, but is far from eliminating it. However, we see too that the measured transmission loss through

SUPERSONIC

TRANSPORT

CABIN

NOISE

123

the structural panel is far greater than the estimate, which could explain part of the discrepancy. This still leaves quite large differences between the estimated and measured results. What other factors go to create such a discrepancy? Firstly, the boundary layer data

60-

1200 2400 Octave

bands kps)

2400 4800

4800 9600

mid-cabin

Figure 18. Typical turbo-prop aircraft internal noise due to boundary layer. allows quite a margin for error. In Figure 20 is given the possible variation in overall intensity that could result by taking the extremes of the experimental data shown earlier. This could account for up to 5 dB of the discrepancy. We see in Figure 21 the possible effect of the scatter in experimental data relating to I

I

I

I I Predicted external /

I spectrum

--. a

9

5 1

I””

/

___----

80

2 a

7

23

60

1200 2400 2400 4800 Octave

4800 9600

bands kps) mid-cabin

Figure 19. Typical subsonic jet transport internal noise due to boundary layer. M = 0.81; H = 35,000 ft (mid-cruise).

the spectral distribution. This could account for a further 5 dB, depending on the panel principal resonant frequency and the spectrum itself. So these two alone can practically account for the remaining difference, and we still have all the structural uncertainties about the effect of edge conditions in determining the panel principal resonant frequency, the degree of damping present, the effect of pressurization and so forth. The best we can do, therefore, is to compare the theoretical results for the projected aircraft and some existing ones, as in Figure 22, and hope that the factors which reduce

J. A. HAI

40! 37 is

I

I 150 300

1:;

Octave

Figure

20.

Variation

meter k =1/(3)/q fluctuations.

of estimated

is a measure

bands

cabin

kps)

noise

of the scatter

140

, 1200 2400

I 600 I200

+iO 600

mid-cabin

Y

I 2400 4800

4800 9600

position

due to scatter

of data on intensity.

of the data on overall

boundary

The

layer

__...~

/

mean >Externai

Octave

Figure

21.

Variation

of estimated

bands

Octave 22.

mid-cabin

position

cabin noise due to scatter

150 300

Figure

kps)

Comparison

300 600

600 1200

of data on spectral

1200 2400

2400 4800

4800 9600

bands kps) mid-cabin position

of predicted

internal

noise levels.

para-

pressure

distribution.

SUPERSONIC TRANSPORT CABIN NOISE

125

the actual values some 15 to 20 dB below those calculated for the existing machines will be equally present on the supersonic transport. On this basis we see that the supersonic aircraft will probably have noise levels comparable with those of existing machines, the actual levels being 15 to 20 dB below the estimated ones. The anticipated cabin noise levels in the supersonic transport are shown in Figure 23 for four different stations in the cabin. These levels in our hypothetical aircraft are comparable with those of current

Forward cabin Mid-cabin m--mmpm-

40&L 1

4 Octave

Figure

23.

5

6

7

8

band no.

Anticipated cabin noise levels in supersonic transport at 4 stations.

aircraft, with possibly a little less low frequency, and a little more high frequency content which can be fairly readily dealt with by the application of a little more soundproofing. It is to be hoped this note demonstrates the extreme difficulty in carrying out sensible cabin noise calculations, and indicates the major sources of this difficulty, lying principally in our lack of knowledge of boundary pressure fluctuations and structural response, coupled with an almost complete lack of means for tackling systems noise.

CONCLUSIONS

It is apparent noise estimation.

that much work needs to be done to obtain The principal requirements for this are :

reliable

methods

of cabin

(i) a more precise determination of the intensity and spectral distribution of boundary layer pressure fluctuations for aircraft ; (ii) a much fuller understanding of structural response, with particular regard to (a) the effect of damping, (b) the effect of pressurization, (c) its response to boundary layer pressure fluctuations, as opposed to sound waves. This will enable a basic estimate of internal cabin noise to be made satisfactorily. When this has been achieved, we must then more thoroughly understand the sources and methods of reduction of: (iii) systems noise ; (iv) vortex noise ; (v) shock wave noise ; (vi) engine, compressor and jet noise, where applicable.

126

J. A. HA\

P. -4. FRANKEK and E. i\,I. KIX\VIK (Bolt, Beranek and Nekvman, Inc.) 19.58 ,Mrthods of flight rrhcle noise predirtion. \VXL>C l’echnical Report 5X-343, ASTl.1 Document Number ADzo5776, Novcniher 195X. 2. E:. J. RICHARDS and P. E:. Do.41~ 1963 Sonze practical applications of bomdary l(:yer pressure $uctzration work. University of Southampton A.A.S.U. Report Number 235. March, 1963. I.