Efficiency analysis of reusable aerospace launchers

Aerospace Science and Technology, 1991, n3 1, 37-46

Efficiency

analysis of reusable aerospace launchers Uu. I. Lobanovsky

Central Aerohydrodynamic

Institute, Manuscript

Lobanovsky

Abstract

Yu. I., Aerospace

TsAGI, 140160, received

Science and Technology,

Zhukovsky-3, Moscow region, Russia.

January 10; 1995; accepted May 17, 1995.

1997, no 1, 37-46.

An efficiency analysis of basic types of reusable aerospace launchers is made. The criteria used are the ratio of payload mass to launch mass of the vehicle as well as the ratio of mechanical energy acquired by the payload to fuel chemical energy. The second criterion correlates well with payload launch cost as it is followed from estimates performed. It is shown that two-stage aerospace planes with a high stage-separation speed (-4.4 = 12-12.5), which incorporate novel technologies based on using aerothermodynamic properties of hypersonic flow, are several times more effective than other known reusable launchers. In addition, such vehicles may ensure qualitatively higher flexibility of the space transportation system permitting offset launches to be performed with flight ranges of 3000-6500 km. Single-stage aerospace planes of the comparable technological level cannot be created at all.

Keywords: Aerospace plane - Aerospace engine - Rocket engine - Scramjet. RCsumk

transportation

system - Launcher

- Efficiency

- Airbreathing

fitude du rendement des lanceurs akrospatiaux rhtilisables. Dans cet article, nous presentons une etude du rendement des principaux types de lanceurs aerospatiaux reutilisables. Les critbres appliques sont le rapport de la masse de la charge utile a la masse au decollage du lanceur ainsi que le rapport de l’energie mecanique acquise par la charge utile a l’energie chimique du combustible. Ce dernier critere est en bonne correlation avec le coat par kg mis en orbite, deduit des estimations effectuees. 11 a CtC demontre que le rendement des lanceurs a deux Ctages dont la vitesse de separation des &ages est &levee (M = 12, 12,5), et qui incorporent de nouvelles technologies mettant a profit les proprietes aerothermodynamiques de l’ecoulement hypersonique, est plusieurs fois supe’rieur B celui des autres lanceurs reutilisables connus. En outre, ces vehicules sont en mesure de procurer une souplesse sensiblement plus &levee du systbme de transport spatial, permettant ainsi des lancements cc decalis )> a des distances de vol de 3000 a 6500 km. Par contre, la realisation d’un lanceur monoetage de niveau technologique comparable est absolument exclue. Mats-cl& : VChicule drospatial - Systeme de transport drospatial - Lanceur - Rendement - Propulsion aerobic - Propulsion fusee - Statoreacteur combustion supersonique (superstato).

I

NOTATIONS

0s

M MC

RO Aerospace

V h

Mach number Mach number of propulsion system changing over from the airbreathing mode to the rocket one

9

AE

(AE = 0.5 V2 + gh)

flight range capability Science and Technology,

0034.1223,

specific impulse characteristic velocity flight velocity height of trajectory gravitational acceleration specific total energy of vehicle

97/01/$

7.00/O

Gauthier-Villa

38

Y. I. Lobanovsky

Pi yLax

fuel heating value payload fraction (the ratio of payload mass to the launcher start mass) structural mass fraction (the ratio of launcher dry mass to its start mass) structural mass fraction under specified conditions propellant (fuel and oxidizer) fraction fuel fraction energy efficiency coefficient morphological parameters maximum lift-to-drag ratio Ulrust

Si n-c

area of compressor inlet compressor pressure ratio

Q mP m, m, mf f 77P

I - INTRODUCTION A substantial decrease, by more than one order of magnitude, in the cost to inject a unit payload mass to a low-Earth orbit, as compared to today’s level, is possible in the case of the creation of fully reusable aerospace launchers using liquid hydrogen fuel. However, economical criteria of efficiency depend on many conditions, which either cannot be taken into account at the first phase of concept formulation, or are determined to a large degree by non-technological factors. For a given class of launchers the structural mass fraction in total start mass is a rather constant quantity. In this case, mp, the ratio of payload mass to the launcher start mass, can be taken as an efficiency criterion of a reusable aerospace transportation system. It should be noted that this parameter of the reusable systems for practically all known realistic projects is in the neighborhood of l-2% [ 1, 21 and less as compared to 2.5-4% of expendable launchers [3]. However, when this criterion is used in comparing reusable vehicles of various classes, such as purely rocket-propelled launchers and airbreathers, the latter gain an unjustified advantage. If the payload mass is related not to the launch but to the structural mass of launcher, the estimated efficiency of rocket vehicles with respect to aerospace planes increases by 5-6 times whereas preliminary estimates show that the ratio of payload launch costs for these classes of launchers, with the annual launch rate being equal, is only 22.5 times greater than in using the mass criterion mp [4-71.

A more adequate criterion may be used, namely the energy efficiency coefficient rip, representing the ratio of the mechanical energy acquired by the payload (which corresponds to the useful work of the launcher) to the fuel chemical energy:

vp = mp AE/(fQL

where AE = 0.5 V2 + gh, V and h are the velocity and height of the payload on a low orbit, g is the gravitational acceleration, Q is the fuel heating value, f is the fuel fraction in the vehicle start mass. Because the rocket vehicles need to carry both fuel and oxidizer, and spaceplanes spend from the tanks only the first component of propellant for a part of their trajectory, for the latter vehicles the f value will be substantially greater. This results in an increase in the efficiency index of rockets in the case of the energy criterion 71pbeing used compared to the mass criterion mp and corresponds to a potentially much greater cost to develop spaceplanes. At the same time, for a specified class of vehicles the dimensionless parameter rip is generally determined by the payload mass fraction, mp. It should be noted that if the energy efficiency coefficient rip for expandable rocket launchers of Soyuz or Energia types amounts to 7-ll%, for advanced reusable systems of Sanger or Delta Clipper types, whose payload mass fractions are about the same and approximately equal to l-2% [ 1, 21, it varies from 0.7-2% in the first case and to 2-4.5% in the second one. In modern times it is difficult to find a less energetically inefficient device than these very high-technology products not having been realized yet. Thus, an analysis is needed of possible ways of increasing the efficiency (in above mentioned meaning) of future reusable aerospace transportation systems. Such an analysis should be based on a single mathematical model, adequately describing basic characteristics of the reusable launcher classes being considered, as well as on estimates of the technology level of modern projects’ basic parameters, primarily of the S&ringerand Delta Clipper launchers, which are most extensively studied. Such efforts may result in the creation of vehicles with a substantially lower payload launch cost. When comparing the launchers of various classes one should take into consideration the greater operational flexibility of the space transportation system based on an aerospace plane due to its maximum flight range capability which determines a possible decrease in payload orbit-inclination angle relative to the launch site latitude. II - ANALYSIS OF CHARACTERISTICS OF BASIC TYPES OF ROCKET LAUNCHERS Principal types of reusable chemically-fueled aerospace launchers may be classified with the help of a simple scheme. The features are the number of stages and the operation of their propulsion system in airbreathing or only rocket modes during acceleration. As is well known the range of minimum characteristic velocities to inject a payload to a lowEarth orbit with the aid of a rocket-powered vehicle amounts to AU = 9.2-9.7 km/s depending on launch Aerospace

Science

and Technology

gfficiency analysis of reusable aerospace launchers/ Etude du rendement des lanceurs ae’rospatiaux re’utilisables

site latitude and direction of flight trajectory [8]. The specific impulse of hydrogen/oxygen rocket engines increases from a value of about Isp M 3.6 km/s at sea level up to IsP = 4.5-4.7 km/s in vacuum [3, 91. In doing so, the trajectory-averaged specific impulse values ((Is,) = 4.1-4.25 km/s) approach to half the characteristic velocity for injection to a lowEarth orbit [lo, 111, hence there is a little point in considering rocket-powered vehicles with the number of stages more than two, much less in considering such aerospace planes whose propulsion system have substantially higher specific impulse. Thus, to the first approximation the set of aerospace launchers considered is exhausted by Zwicky’s twodimensional morphological matrix [ 121 depicted in Figure 1. Let us consider that the value of the parameter pl equal to 1 corresponds to a purely rocket powerplant, pl = 2 corresponds to the use of airbreathing engines as well, the parameter p2 expresses the number of vehicle stages.

I

Fig.

1. - Zwicky’s

I 1 two-dimensional

1

I 2 morphological

PI matrix.

The cell 11 of this morphological matrix corresponds to single-stage rocket vehicles usually referred to as SSTO. They can be ballistic or winged, and in the latter case - with either vertical or horizontal In any case the propellant fraction rnf takeoff. with consideration for expenses for return amounts to at least 0.90-0.91 of the launcher takeoff mass, and a payload fraction of mp = 0.01-0.02 is possible at such a vehicle structural fraction of m, M 0.08, which corresponds to the design data of the Delta Clipper launcher [2]. The development of such a vehicle may be essentially based on already proven technologies, and the sole really fundamental problem is to create a fully reusable rocket structure with the structural mass fraction equal to that of an expendable launcher [3]. However, an increase in the structural 1997, Ilo 1

39

mass of such a vehicle by N 20% (less than 2% of its start mass) results in a decrease in the payload mass to zero. The recent information about the Delta Clipper program [13] testifies that, as it develops, the possibility of solving this problem in the immediate future is believed to be not so optimistic. The structural mass fraction of the winged rocket vehicles would be noticeably greater 141, especially in the case of horizontal takeoff [14] (the m, values may exceed the data corresponding to the Delta Clipper by a factor of about 2). The results of competition of all the SST0 versions described above, which was conducted in 1991 to reveal the Delta Clipper as a winner [15], confirm the validity of the estimates briefly presented above. Thus, the realization of concept 11 from Zwicky’s morphological matrix is primarily determined by the possibility of creating a reusable hydrogen/oxygen rocket vehicle with a structural mass fraction of = 0.08. Even if it is achievable in the immediate K&re, then only be using ballistic [2, 5, 151 or aeroballistic (with horizontal landing) vehicles [6]. Cell 12 of the morphological matrix corresponds to two-stage rocket vehicles (TSTO). Contrary to expendable launchers, for which division of the vehicle into two stages results in a substantial increase in payload mass, such an approach to reusable rocket vehicles, which have to return to the launch site, is ineffective. In this case, the first stage must obligatory be winged and have a sufficiently high lift-to-drag ratio to perform a turn maneuver and return flight. This leads, as pointed out above, to a substantial growth of the structural mass fraction, which increases also in consequence of scaling down the stages. Therefore, estimates of mp values for fully reusable two-stage rocket launchers give practically the same results as for single-stage ones, whereas the complexity of the system and, therefore, its operational cost are significantly greater. It should be noted that at the aforementioned competition of the projects of reusable rocket launchers [15] TSTO-class vehicles were not presented at all. Therefore, concept 12 (TSTO) is potentially inferior to concept 11 (SSTO) and can be excluded from further consideration [ 16, 171. The situation can somewhat be changed if the landing of the first stage is permitted not only near the launch site. But in this case the complexity and operational cost of such a system increase still further. III - ANALYSIS OF CHARACTERISTICS OF BASIC TYPES OF AIRBREATHING LAUNCHERS Reusable airbreathing launchers (cells 21 and 22 of the morphological matrix) are substantially more diverse from the standpoint of their realization. The estimation of their main parameters is much more complex due to a greater uncertainty of the data on

40

powerplants and structural mass fractions as well as greater influence of aerodynamics on the launchers characteristics. In addition, the use of offset launches greatly expands the possibilities of transportation systems but diminish payload mass fraction and makes the comparison of their transport efficiency more difficult. Having eliminated two-stage rocket launchers from further considerations, the discrete classification shown in Figure 1 can be extended using a quasi continuous representation with the Mach number n/r, corresponding to the propulsion system changing over from the airbreathing mode to the rocket one as an argument [ 181, and the payload fraction mp as a function (Fig. 2). The single-stage rocket vehicles correspond to region I on the axis of ordinates, point 1 refers to a winged launcher, points 2 and 3 refer to a ballistic Delta Clipper-type vehicle [2] and an aeroballistic one [6] respectively. It should be noted that the maximum payload mass fraction declared for Lockheed aeroballistic vehicle (mP M 0.03) [6] can be achieved only with structural mass fraction being less than that of the Delta Clipper having simpler and more capacious structure shape (mp M 0.02) [2], or with the average specific impulse of the rocket engine being extremely high, (Is,) N” 4.3 km/s. Thus envisaged, in the aeroballistic launcher design there is more high technology level than that in the Delta Clipper design, which casts quite justified doubts upon the possibility of implementing the declared characteristics in the near future. Changing over to the rocket mode at a moderate Mach number (MC 5 3-4) leads to a rather small decrease of the characteristic velocity for this acceleration phase and is accompanied by the need for using winged vehicles with a corresponding increase in their structural mass fractions. This requires the use of two-stage launchers, and the payload fraction in the entire system start mass at the stage separation before hypersonic speeds slightly decreases as compared to the maximum possibilities of the rocket SST0 (see region II in Fig. 2). Point 4 corresponds to the characteristics of the Interim Hotel/An-225 system (MC = 0.8, mp M 0.01) [ 191, point 5 corresponds to the data of Boeing two-stage launcher with the rocket second stage (AJc = 3.3, mp M 0.015) [7, 201, and point 6 to an original version of the Hotol vehicle with taking into account the mass of the sled (AJ& = 5, mP M 0.02). The design structural mass fractions of the second stage are equal to 0.16 of the value for Boeing orbiter [7] and to a difficult-to-reach value of 0.13 for the winged Hotol vehicle. There is a natural trend for a gradual increase in the payload fraction with the speed of changing over to a rocket mode. Among the reusable launchers being considered, the SBnger aerospace plane [l, 4, 211 occupies a special place being a vehicle designed from the outset to perform a mission with flight range capability of about 3000 km. Such a mission radius is of interest

Y. I. Lobanovsky

either in the case of starting from Cape Canaveral (for equatorial orbit insertion) or from Europe. In addition, this concept differs from others by more conservative levels of the structural mass fraction, which make their reaching more possible with further studies. Because of this, at AJc = 6.5-7 the mp values for the Sgnger amount to 0.008 for the manned version of the second stage (point 7) or 0.022 for the unmanned one (point 8) [ 11. The Sgnger variants under study provide in specified conditions the orbit insertion of the minimum of the payload to be delivered. It is evident that as the atmospheric portion of the acceleration trajectory increases and, accordingly, the launcher complexity and the technical risk of its design, the payload fraction can be increased. The studies made in the framework of the NASP program as well as the data regarding assumed characteristics of the advanced scramjets [9, 22-241 allow one to conclude that sufficiently efficient atmospheric flight is possible up to Mach numbers of 12-15. These limits are shown in Figure 2 with the vertical dashed stripes. In so doing, the first stripe (A,!& = 12) corresponds to a moderate level of technical risk and the second one (n/r, = 15) corresponds to a significantly greater level.

Fig. 2. - Payload

fraction

versus

changing

mode

Mach

number.

Of some interest is an analysis of the potentialities of single- and two-stage spaceplanes which, by properly combining aeronautical and rocket elements, can implement the same task as the SBnger launcher at higher speeds of changing over the propulsion system to the rocket mode. From estimates it follows that a decrease in structural mass fraction of the SWger vehicle compared to the technology level of the 1970s (Space Shuttle) is of quite realistic value, amounting to 25% whereas the goal of the excessively ambitious NASP program was the reduction of this parameter by 60% [ 181. An increase in the Mach number AJ& from 7 to 12 accompanied by using advanced technologies gives the growth of structural mass fraction by about Aerospace

Science

and Technology

Efficiency

analysis of reusable aerospace launchers/

Etucle du rendement des lanceurs ae’rospatiaux re’utilisables

lo%, with maximum deviation of limit values from the average ones being 15% [ 181. Thus, among all kinds of launchers described with the cells 21 and 22 of the morphological matrix (see Fig. 1) it is worth-wile to consider advanced vehicles with the same or somewhat increased levels of the structural mass fraction relative to that of the Stinger aerospace plane, with estimates being performed under comparable conditions using a unifie.d mathematical model. Such a model was described in [25-271. It should be noted that the results of computations with the aid of this model are in good agreement with the total body of information on the Sanger vehicle. It was also shown that from the standpoint of limiting the technical risk growth as well as of drastically increasing the payload mass fraction, a two-stage aerospace plane with stage separation Mach number at M = 12-12.5 named Star Liner [25, 271 may be of much interest. Several variants of such a launcher were studied. With the aim of minimizing structural mass, along with the winged orbiter, the ballistic variant of the Delta Clipper-type vehicle [2] was considered, which also lands vertically with the final speed being canceled with the help of the engines. Similar values of w+, can be provided

by the aeroballistic

second

horizontal landing of Lockheed vehicle-type One

of

the

most

important

factors

stage with [6].

defining

the

41

The turbo-compressor engine proposed, called the synergetic jet, is able to operate in a continuous regime at Mach numbers from 0 to 5.5-6. The nearest analog among the most advanced airbreathing engines which have already passed through preliminary design phase is the air-turbo-ram jet engine using hydrogen expander cycle, one of the leading candidates for being used on aerospace planes. According to preliminary estimates, in the vicinity of the lower bound of the operating envelope the values of the specific impulse and thrust per unit inlet area (thrust T related to compressor inlet area Si) for the engine being proposed amount respectively to $p = 52-55 km/s and T/Si = 0.25-0.26 MPa. This is greater by a factor 1.25-1.3 than the corresponding parameters of turbo-expander powerplant with a compressor pressure ratio of 7r, = 3.5 at M = 0, which to the maximum degree satisfies conflicting requirements for sufficiently high propulsive-economic characteristics of such accelerating engines at hypersonic and suband transonic flight speeds [31, 321. The specific impulse of the synergetic jet at M = 5.5 (lsP = 3637 km/s) is greater by a factor 1.1-I. 15 and its thrust per unit inlet area T/Si is 1.3-1.5 times greater than the corresponding values of a such turbo-expander engine. In the average, along the trajectory the specific impulse can be 1.15-1.2 times greater, and the thrust 1.25-1.35 times greater for engines of equal size.

main properties of a space transportation system with a given flight range capability is the maximum lift-tot-drag ratio (L/D)rrlax at hypersonic speeds. Wind-tunnel tests of a model created with the aid of the computationally substantiated interference concept of aerodynamic design of hypersonic vehicles proposed by the author have shown that at M 2 7 (L/D)ll,aX values can be 15-30% greater than those of known configurations. Moreover, the layouts developed differ little in structure from classical low aspect ratio wing-body combinations [25, 271.

The superiority of the engine being proposed is no less so as compared to the turbo-ram jet engine [3l] powering the first stage of the Sangcr aerospace plane [I, 41. An incrcasc in the thrust-to-weight ratio allowing a more vigorous vchiclc accclcration and an improvement in propulsion system fuel efficiency would lead, according to cstimatcs, to a l‘uel saving during the first llight scgmcnt up IO changing over to the scramjet no less than I .4 times as compared to the use of any one of the known airbrcathing engines or their combinations.

A propulsion system of the winged first stage of the proposed aerospace plane is made up of a scramjet, operating only in supersonic diffusive combustion mode, and an airbreathing engine of the novel layout [25, 261 representing the development of the engine combining inverse and turbo-expander cycles which is presently under study at ONERA [28, 291. This makes it possible to simplify the structure significantly and improve the scramjet properties - the key element for the efficient hypersonic flight. The major scramjet parameters of the Star Liner concept correspond to moderately optimistic estimates for such engines (Is,, = 35 km/s at M = 5.5-6, and I,, = 15 km/s at M = 12.5) [22-241. In this case, its average propulsion efficiency (7) = 0.5, whereas the maximum values of this parameter are reached at M = 8-9 [25]. It should be noted, that US subscale scramjet module has exceeded in a supersonic combustion mode at M M 6.5 performance expectations by 10% during the ground tests [30].

of aerospace plane having above indicated novel technologies in the fields of hypersonic aerodynamics and airbreathing propulsion at a specified flight range capability of 3000 km, as for the Sanger vehicle (n/r, M 7) can use the fuel, intended for its cruise flight in both flight directions, as well as the fuel saved during the first acceleration flight segment, for accelerating from M = 7 to M = 12.5 and returning to the launch site, with the mission range and, to a first approximation, the payload mass fraction of the first stage being retained. Thus, the second stage size and mass are unchanged and a decrease in its characteristic velocity by 1.5-1.6 times as well as a decrease in its structural mass by 1.5-2 times due to changing over to a ballistic flight pattern result in increasing by 5-10 times the second stage payload mass fraction, and consequently, the whole Star Liner payload as compared to the Sanger-type vehicle. Under these conditions the payload fraction in the launcher start

1997, Ilo 1

It

has

been

shown

[ 25, 271

that

a

version

Y. I. Lobanovsky

42

mass accounts for about 0.10 (point 11 in Figure 2, MC = 12.5, mp M 0.10).

A decrease in the speed of the powerplant changing over to rocket mode down to iWc = 10 diminishes, according to estimates, the mp value to 0.06 (point 12), and its increase up to A& = 16 does not vary it (point 13, mp M 0.10). In the latter case the design separation Mach number exceeds the boundary set up earlier to reveal more clearly the trends being considered. A possible evolution of the payload mass fraction with variation of the Mach number of changing over to the rocket mode in indicated conditions is described by region IV. When using traditional technologies in the propulsion system layout and in the first stage aerodynamic configuration, the payload mass fraction decreases down to about mp z 0.045 (point 14). In the case of using a winged second stage, mp becomes about 0.03 (point 15), and the launcher proposed becomes conceptually very close to the Radiance [33], a French hypersonic aerospace plane project (point 9, nir, = 12, mp = 0.01-0.015 using as a fuel not only liquid hydrogen, but also kerosene, with a rocket second stage of comparatively low efficiency). Based on standards for unmanned orbiter design corresponding to the Sgnger concept, a payload mass fraction of the Radiance would be not smaller than 0.035 (point 10 in Figure 2). The trend of mp variation under these conditions is described by region III. It should also be noted that the structural mass fraction of the Radiance first stage (relative to its start mass) in the case of using only hydrogen as fuel and for the same payload fractions and equal sizes, would be, according to estimates, 5-10% less than that of the first stage of the Sgnger spaceplane [33, 341 (in this case the question of sufficiency of this amount of hydrogen for the flight mission of Radiance is not considered). Thus, at a 10% increase in structural mass fraction with increasing the Mach number A!& from 7 to 12, the distinction between data for the two sufficiently closely related European concepts are in rather good agreement with the estimates of possible scattering of their characteristics according to [18], and the Star Liner concept holds a middle position between them in this index. It is also advantageous to obtain similar estimates for a single-stage aerospace plane, hereinafter referred to as the NASP-type vehicle. However, as was indicated [9, 22-251, they are seriously complicated by high uncertainty of scramjet characteristics at M > 12 and difficult problems regarding a possible increase in the structural mass of such a vehicle due to highly severe atmospheric flight conditions at A,! > 12. Because of this, let us consider the single-stage counterpart of the Star Liner aerospace plane as a first version of such a vehicle differing from the two-stage one only by the fact that instead of separating stages and returning the first stage to the launch site the ignition of the hydrogen/oxygen rocket engine of the counterpart is

accomplished at M = 12.5 and it comes to orbit in rocket mode. It is evident that a similar flight scenario is not optimal for the single-stage aerospace plane and places it in somewhat unequal position relative to twostage system, but the provision of optimum conditions requires the technical perfection of a single-stage plane to be substantially higher [9, 251. In addition, an increase in the Mach number of the propulsion system changing over up to a practically optimum value of A& = 16 allowing more adequate comparison of these two systems will be considered too. A flight range of R, = 3000 km to offset launch is envisaged in all the cases. If the start mass of the Stinger aerospace plane is taken to be unity, the structural mass fraction of the first stage will account for m, = 0.42, its fuel (liquid hydrogen) mass fraction will be mf = 0.27 and payload mass fraction (the second stage) will be mp = 0.31 [l]. Mass distribution of the second stage is respectively as follows: m, = 0.07-0.08, rnf = 0.22 (liquid hydrogen amounting to l/7-1/8 of propellant mass, the rest being oxygen whose density is 15 times greater than that of hydrogen), mp = O.Ol0.02 depending on whether manned Horus-M or unmanned Horus-C second stage is used [ 11. According to first-approximation assumptions, the mass distribution of main components of the Star Liner first stage is the same as that of the Sanger, and averaged characteristics of the second stage is as follows: m, = 0.04, rnf = 0.17; mp = 0.10. These parameters of reusable single-stage ballistic rocketpowered vehicle Delta Clipper according to design data [2] are: m, = 0.08, mf = 0.90-0.91, mp = O.Ol0.02 depending on launch azimuth. Fractions of main components such as structure (dark grey area), fuel (white zone), oxidizer (light grey), and payload (darkened area) in the start mass of the aerospace launchers being considered are shown in Figure 3: 1 - Delta Clipper, 2 - NASP-type vehicle, 3 - SWger, 4 - Star Liner. The mass of the single-stage counterpart at instant of the ignition of the rocket engine at M = 12.5 is equal to that of the Star Liner before stage separation, ml M 0.74. Propellant consumption for accelerating the vehicle in the rocket mode and its deorbiting (5% of orbital mass) with a specific impulse of Issp = 4.7 km/s will in this case be Am, = 0.41. Thus, the total propellant consumption of the singlestage counterpart of the Star Liner is mf = 0.67 as compared to mf = 0.44 for the two-stage version. Moreover, due to increasing the rocket engine thrust and volume of propellant tanks intended for rocket mode by approximately 2.5 times (rocket propellant mass will exceed the sum of propellant and payload masses of the Star Liner second stage by more than 1.6 times), the mass fractions of the corresponding structural elements will increase from 0.04 to at least 0.06-0.08. In so doing, the structural mass fraction will increase from m, = 0.46 for both stages of the twostage version of the Star Liner up to m, w 0.49 for Aerospace

Science

and Technology

Efficiency

analysis of reusable aerospace launchers/ 43

Etude du rendement des lanceurs ae’rospatiaux re’utilisables

m 1.0

n

0.5

0.0

Fig. 3. ~ Fractions 4 aerospace launchers.

of

main

components

in the

start

mass

for

its single-stage counterpart. Consequently, the singlestage vehicle whose technological level is equal to that of the Star Liner aerospace plane can have only negative payload mass on orbit mp z - 0.16 (point 16 in Figure 2). An increase in the Mach number of changing over to rocket mode up to Mc = 16 results in decreasing propellant consumption down to mf = 0.64. As this takes place, hydrogen consumption increases by about 15%. Under the assumption that the structural mass of the single-stage vehicle is invariant, which is somewhat unrealistic, we obtain an increase in the negative payload mass up to mp = - 0.13 (point 17 in Figure 2). The mass of advanced hydrogen tanks is about 0.1-0.15 of the fuel contained [34, 351. Taking into account to a first approximation the structural mass growth due to increasing the volume of fuel tanks by 15% we obtain point 18 in Figure 2 (mP z - 0.14). Consequently, at conditions considered the effect of technological improvements allowing a 30% increase in maximum atmospheric flight speed on payload mass fraction is one order of magnitude less than that of the concept option. The data for single-stage aerospace planes form region V in Figure 2. Point 19 (A& = 25, mf = 0.56, mp z - 0.05) corresponds to a hypothetical vehicle version which up to M = 25 uses only airbreathing engines with average propulsion efficiency of about 0.4 whose structural mass fraction not increases relative to that of the single-stage aerospace plane represented with point 16 (MC = 12.5) 125, 35, 361. Region 20 for MC = 13 represents payload mass fractions of the European single-stage aerospace plane [37]. With practically the same hydrogen-tooxygen ratio as for the aerospace plane represented 1997, no 1

with point 16, it features less flight range capability (- 1650 km) and uses a somewhat simpler and less efficient propulsion system. Structural mass fraction makes up as low as 0.20-0.25 of the vehicle start mass depending on the assumed level of technical perfection and on whether or not the 10% reserves are included. Such relative structural mass corresponds at its upper boundary to the goal of unrealized NASP program. As this parameter increases up to the values comparable with those considered in the present study, even with taking into account the lesser power plant mass fraction and an approximately two-fold increase in the vehicle size, the payload mass fraction in this project of the single-stage aerospace plane becomes still less than that of the series of vehicles considered above and represented with region V. Thus, the rejection of stage separation, resulting in the necessity of launching into orbit in rocket mode a rather large passive mass of the aerospace vehicle, causes a decrease in payload mass by Am, M 0.25 at the level of structural perfection inherent in the Star Liner launcher and of the approximate level of the Sanger and Delta Clipper transportation systems. It is evident that this is too high an expense for assumed operational advantages of a single-stage aerospace plane. Let us take the above-indicated (see Figure 3) level of structural perfection as the basic one and consider the effect of structural mass variation on the value of payload mass fraction. The dependence of mp on the ratio of structural mass m,5 to its value under conditions indicated m, are shown in Figure 4 for four basic types of the aerospace launchers being studied (1 - Delta Clipper, 2 - NASP-type vehicle, 3 - Sanger, 4 - Star Liner). In all the cases excluding the purely rocket-powered Delta Clipper vehicle the flight range

0.2

0.0

-0.2

Fig. 4. - Structural mass fraction versus to the structural mass under conditions.

the ratio

of structural

mass

44

Y. I. Lobanovsky

capability is assumed to be equal R, = 3000 km. For characteristic trends to be revealed, the dependencies are plotted up to a physically unrealistic complete disappearance of the structural mass of the vehicles. Only in this case the payload mass of the single-stage counterpart of the Star Liner system becomes equal to that of its two-stage version; and the technically more perfect vehicle which changes over to the rocket mode at A& = 16 can have the greater payload mass at an obviously unrealistic structure perfection level (m,/m,

< 0.2).

At the level of characteristics considered the NASPtype vehicle is capable of launching to a low-Earth orbit a payload of mr, = 0.05 at m,/m, = 0.55-0.6, which agrees with estimates made in [ 181. It is a good job for structural specialists to find when and in what manner one can be able to reach such indexes. It should be only noted that in that case the relative payload mass of the two-stage aerospace transportation system amounts to mp M 0.2. From estimates performed it follows that the Earth is a too large planet which has too great low-Earth orbit speed for the chemically-fueled single-stage aerospace plane to be operated effectively. Therefore, the set of concepts 21 of the morphological matrix (see Fig. I) can be excluded from further consideration for the foreseeable future as well. An increase in the structural mass of the Star Liner first stage by 10% compared to that of the Sanger would lead to decreasing its payload mass by 13-14%. As this takes place, the payload mass fraction makes down 0.085-0.09 and structural mass fraction m, in start mass is no less than 50%. Thus, as the study performed shows, satisfying the condition m, 5 0.5 is necessary for efficient two-stage aerospace planes to be created. For single-stage planes and rocket-powered launchers these boundary conditions are respectively m, < 0.3 and m, 5 0.08. IV - ANALYSIS OF ENERGY EFFICIENCY OF AEROSPACE LAUNCHERS Let us go to the assessment of the energy efficiency of fully reusable aerospace launchers of different classes which is more adequate to economical criteria. Plotted in Figure 5 are dependencies of energy efficiency coefficient rip introduced above against Mach number A& of changing over propulsion system from airbreathing to rocket mode. Designations are the same as in Figure 2 but the points represent only the vehicles with positive payload being the most important for subsequent analysis. When launching the payload into a low-Earth orbit (V = 7.8 km/s, h = 200 km) using liquid hydrogen as fuel (Q = 1.2.10’ J/kg) the ratio of specific mechanical energy to fuel chemical energy is LYE/Q = 0.27 (see formula (1)) which is rather close to the fuel fraction values of all hypersonic aerospace planes

Fig. 5. - Energy efficiency versus air breathing to rocket mode.

Mach

number

of changing

from

considered at 6 5 iWc 5 16 (f = 0.29-0.36). Because of this, regions from III to V constructed in the new coordinate system are little changed. The f value for single-stage rocket vehicles is substantially lesser (f = 0.12-0.15), therefore the size of region I has increased approximately two-fold and the energy efficiency of a Delta Clipper-type launcher (points 2, np = 2-4.5%) are also about two times greater than their mass efficiency (mp = l-2%). The assessments of payload launch cost for the Sanger aerospace plane having the unmanned second stage at 20-25 launches per year give a value of $2500 per kg [4] and for the Delta Clipper vehicle at the same conditions and optimum launch azimuth it is about $1100 per kg [2, 51. The ratio of these values is about 2.25, which is in line with the ratio of the energy efficiency factors for these launchers, respectively 77PM 2% and nr, w 4.5% (points 8 and 2 in Figure 5). According to the energy criterion, the efficiency of the launchers with relatively low Mach numbers it&c increases significantly (see region II in Figure 5, np = 3-3.5%). The assessment of payload launch cost for the Boeing two-stage launcher amounts to about $2000 per kg at the same conditions [7], which is 1.25 times less than that for the Sanger aerospace plane, with the ratio of rip values for these vehicles being about 1.5. Consequently, there exists a satisfactory correlation of this objective, physically-meaningful criterion of the efficiency of space transportation systems of various classes with economic assessments of their efficiency. According to the energy efficiency criterion the Sanger-type aerospace planes (&& = 6-7, 71p= 0.72%, points 7 and 8) fall within the region of a local minimum, which is partly associated with obtaining sufficiently high flight range capability for Aerospace

Science

and Technology

Efficiency

analysis of reusable aerospace launchers/

Etude du rendement des lanceurs ae’rospatiaux re’utilisables

offset launch not ensured by the launchers considered with lower Mach numbers of the propulsion system changing over to rocket mode. In addition, if one compares not concrete launcher designs but conceptual approaches, it is more correct to consider vehicles with the same launch masses. The 1.25 times scaling of the rocket-propelled Delta Clipper and the 1.X-times scaling of Boeing launcher down to the size of the Sanger aerospace plane lead, according to estimates, to a decrease in their efficiency indexes by 1.1 times and no less than 1.7 times respectively. Thus, in the latter case the value of energy efficiency factor would amount approximately to qP M 2%, which is practically coincident with the indexes of the Sanger aerospace plane version with the unmanned second stage. The vehicle of the type proposed, the Star Liner aerospace plane, has the highest energy efficiency (VP E 9%, point ll), which is at least 2 times greater than that of any alternative reusable space transportation system and is close to that of expendable launchers. This also makes it possible to perform flight with a distance of 3000 km for offset launching. The incorporation of a subsonic refueling plane into the aerospace transportation system ensures the flight radius of the first stage to be increased up to 6500 km and removes in fact any operational limitations on the orbit inclination angles of the payloads [25, 271. It should be noted that the stage separation Mach number corresponding to the maximum of the energy efficiency of the two-stage aerospace planes being proposed only slightly exceeds the value selected previously for the Star Liner design concept, A!& = 12.5. A decrease in scramjet specific impulse at high flight speeds compared to accepted values, some increase in structural mass fraction with flight speed, as well as deterioration of aerodynamic characteristics of the vehicle because its geometry must be defined by thermal fluxes to a greater extent, can somewhat shift this maximum to lower Mach numbers. V - CONCLUSIONS Thus, the analyses made shower that, with specified design parameters being held, the most efficient reusable aerospace launchers from an energetic point of view which correlates well with economical indexes consist of a single-stage wingless rocket-powered vehicle with vertical start combined with a twostage aerospace plane with high stage-separation Mach number (MC = 12-12.5), using novel technologies based on the application of aerothermodynamic properties of hypersonic flows and a wingless orbiter. Moreover, the latter type of the vehicles, according to estimates, can be at least 2 times more efficient than any known type of reusable launchers. Besides, it is capable of ensuring a substantially higher operational flexibility by making it possible to accomplish flights over distances of 3000-6500 km. 1997, no 1

45

At comparable technology level single-stage aerospace planes cannot be created at all. REFERENCES [I] Kuczera H., Sacher P., Krammer P. - The German Hypersonic Programme, Status Report, 1991, AIIA-91-

5001. [2] Colucci F. - Launching the Delta Clipper, Space, 1991, 7, no 6. [3] Cosmonautics (Encyclopedia) - ed. by V. P. Glushko, Moscow, Soviet Encyclopedia, 1985 (in Russian). [4] Koelle D. - Singer Advanced Space Transportation System, Progress Report, 1990, AMA-90-5200. [5] Caubatz W. - Developing the Delta Clipper for Low Cost Future Space Transportation Systems, International Aerospace Congress, 1994, Moscow. [6] Fulghum D. A. - Shunk Works Design May Cut Launch Costs, Aviation Week & Space Technology, 1993, 139, no 7. [7] Hardy R., Eldrenkamp L., Ruzicka D. - The Boeing TSTO, 1993, ALU-93-5167. [8] Levantovsky V. I. - Mechanics of Space Flight in Elementary Presentation, Moscow, Nauka, 1980 (in Russian). [9] Kors D. L. - Design Considerations for Combined Air Breathing-Rocket Propulsion Systems, 1990, ALU-905216. [lo] Grey .I. - Engineering Reality Must Guide SSTO, Aviation Week & Space Technology, 1994, 140, no 2. [ 1 I] Dornheim M. A. - DC-X Proving Initial Operational Concepts, Aviation Week & Space Technology, 1993, 139, no 15. [12] Zwicky F. - Discovery, Invention, Research through the Morphological Approach, N. Y., MacMillan, 1969. [ 131 NASA Launch Preparation Funds Give DC-X Reprive. Aviation Week & Space Technology, 1994, 140, no 6. [14] Shkadov L. M., Denisov V. Eu., Lazarev V. V., Plokhikh V. P., Buzuluk V. I., Volodin S. V., Chervonenko K. A., Skipenko V. V. - Comparative Analysis of Various Concepts for Reusable Aerospace Systems, 1992, IAF-92-0865. [15] Dalby A. R. - The Delta Clipper, Ad Astra, 1991, 3, no 8. [16] Asker J. R. - Clinton Launch Policy: Upgrade ELVs, Push SSTO, Aviation Week & Space Technology, 1994, 140, no 19. [17] Anfimov N. A., Vakhnichenko V. V., Kostromin S. F. Validation of Vertical Take-Off Reusable Rocket Space Transportation System Basic Concepts, International Aerospace Congress, Moscow, 1994. [18] Froning D., Gaubatz W., Mathews G. - NASP: Enabling New Space Launch Options, 1990, AIAA90-5263. [19] Parkinson R., Webb E. - An225/Hotol, 1993, AZAA93-5169. [20] Proctor P. - Boeing Proposes Two-Stage-To-Orbit Plan, Aviation Week & Space Technology, 1993, 138, no 20. [21] Kuczera H., Hauck H., Krammer P.: Sacher P. - The German Hypersonics Technology Programme, Status 1993 and Perspectives, 44th IAF Congress, Gratz, Austria, 1993, IAF-93-V.4.269.

Y. I. Lobanovsky

46

[22] Menees G. P., Adelman H. G., Camber J.-L. Wave Combustors for Transatmospheric Vehicles, 1989, ISABE

89.

[23] Kuznetsov M. M., Neiland V. Ya., Sayapin G. N. - Study of the Efficiency of the Scramjet Engine with Detonation and Shockless Supersonic Combustion, Scientific Memoirs of TsAGI, 1992, XXIII, no 2 (in Russian). [24] Andreyev V. V., Penzin V. I. - Duct Geometry Selection for a Scramjet Integrated into a Vehicle Layout, Theory and Structure of Flying-Vehicle Engines, Proceedings of the XVIII Scientific Lectures on Cosmonautics, Moscow, Fazis Publication, 1994 (in

Russian). [25] Lobanovsky Yu. I. - Concept of an Advanced Aerospace Transportation System, Preprint of TsAGI, 1994, no 95 (in Russian). 1261 Lobanovsky Yu. I. - Advanced Aerospace Transportation System. Theory and Structure of Flying-Vehicle Engines, Proceedings of the XVIII Scientijc Lectures on Cosmonautics, Moscow, Fazis Publication, 1994

[27] [28] [29] [30]

(in Russian). Lobanovsky Yu. I. - Concept of an Advanced Reusable Aerospace Transportation System, La Recherche Ae’rospatiale, no 1996-2. Ribaud Y. - Inverse Cycle Engine for Hypersonic Air-Breathing Propulsion, 1989, ISABE 89-7111. Casamayou J.-P. - Le moteur PRIAM candidat pour le vol hypersonique, Air & Cosmos, 1993, no 1434-35. Kandebo S. - Scramjet Tests Hit Mark, Aviation Week & Space Technology, 1994, 141, no 12.

[31] Minoda M., Sakata K., Tamaki T., Saitoh T., Yasuda A. - Feasibility Study of Air-Breathing Turbo-Engines for Horizontal Take-Off and Landing Space Plane, 1989, AMA-89-2296.

[32] Buzuluk V., Volodin S., Semienov A., Dulepov N., Lanshin A. - Optimization of Main Parameters of Turborocket Engines in the Propulsion System of Single-Stage-to-Orbit Vehicle, International Aerospace Congress, Moscow, 1994. [33] Billant C., Boyeldieu E., Wagner A. - Radiance: a High Staging Speed Airbreathing First Stage TSTO Launcher, 44th IAF Congress, Gratz, Austria, 1993. [34] Waingartner S. - Stinger: the Reference Concept of the German Hypersonic Technology Program, 1993, AMA-93-51

61.

[35] Blinchevski M. Y., Kazakov M. I., Klimov V. T., Poukhov A. L. - The Concept of the Single-Staged Aerospace Plane with Horizontal Takeoff and Landing, International Aerospace Congress, Moscow, 1994. [36] Jones R., Donaldson C. - From Earth to Orbit in a Single Stage, Aerospace America, 1987, 25, no 8. [37] Weinreich H.-L., Grallert H., Parkinson R., Berry W. Studies on a Scramjet-Propelled Horizontal Launch and Landing Single-State-to-Orbit Launcher, 1993, AIAA 93-50.53.

Acknowledgements. - The author wishes to thank P. Sacher (DASA), V. Denisov, V. Orekhov, G. Pavlovets, and L. Shkadov (all of TsAGI) for helpful discussions of the issues considered in the paper.

Aerospace

Science

and Technology