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THIS DOCUMENT HAS BEEN REPRODUCED FROM MICROFICHE. ALTHOUGH IT IS RECOGNIZED THAT CERTAIN PORTIONS ARE ILLEGIBLE, IT IS BEING RELEASED IN THE INTEREST OF MAKING AVAILABLE AS MUCH INFORMATION AS POSSIBLE f i ^

NASA Technical Memorandum 74087

(RASA -mM-74U@7 HORIZONTAL TA KEOFFA FUMY REUSABLE, WITH TWO SPACE TBAvS SMALL TURBOJET BOOSTERS CO 50 P HC A031MF A01 CONCEPTNCEPT N78-11172 (NASA) CSCL 22A Unclas G3/1S 52815

A FULLY REUSABLE, HORIZONTAL TAKEOFF

SPACE TRANSPORT CONCEPT WITH TWO SMALL TURBOJET BOOSTERS

+ L, ROBERT JACKSON, JAMES A. MARTIN, AND WILLIAM J. SMALL s i

OCTOBER 1977

AMA r+s3 a, C'0 National Aeronautics and N 4 Space Administration^j- Langley Research Center Hampton, Virginia 23665

j y

2 k T E

A FULLY REUSABLE, HORIZONTAL TAKEOFF SPACE TRANSPORT CUNt;EPT WITH TWO SMALL TURBOJET BOOSTERS

L. Robert Jackson, James A. Martin, and William J. Small Langley Research Center

SUMMARY

Results of a preliminary study of a novel space transport concept are reported. The concept consists of a winged orbiter containing ascent propellants and small, twin turbojet-powered boosters used for acceleration to

Mach 3.5. Each booster contains sufficient JP fuel for ascent and fly-back functions. The concept offers a fully reusable system capable of hori7ontal takeoff from conventional runways. Other potential features -include lateral offset orbit insertion, ferry-capability, abort and recall landing, and considerable versatility in takeoff site, which may enable round trips to space from many existing airfields. Preliminary performance analyses show this space transport concept, using current structures and rocket technology, has a lower gross takeoff weight for a selected payload than either a single-stage-to-orbit vertical takeoff advanced technology vehicle, or a sled-launched vehicle which applies current rocket propulsion but requires structural advancements.

Alternatives to the baseline concept of a current technology orbiter with turbojet boosters were also studied. A refinement of advancing the orbiter structure offers a 50 percent increase in payload. A further 50 percent increase in payload is indicated by adding a lightweight scramjet under the orbiter, thus doubling the payload over thebaseline concept. An alternative to a new orbiter configuration is the concept of stretching the space shuttle orbiter. Preliminary analyses indicate that the stretched shuttle orbiter with turbojet boosters could carry the same payload as the NASA space shuttle ' 2 but with about half the gross weight.

Some principal probl-em areas warranting study to verify the space transport concept using turhojet boosters are configuration refinement, transonic aerodynamic tests, structural design, and engine perfon,.iance analyses.

INTRODUCTION

Projections for orbit;nq solar power collectinq stations and industrial processing plants (ref. 1) indie:ato the need for frequent spice Flights. The mission scenario studied in reference 1 shows that current space Shuttle launch costs must be reduced by an order or magnitude before lame scale space utilization can become a practical goal. Other desirable space transportation system features, which offer increased operational versatility include horizontal takeoff from conventional runways, lateral offset orbit insertion, and ferry capability.

The dominant factors in the space shuttle launch costs are the replace- ment of expendable propellant tanks and the recovery, refurbishment, and refueling of ,.he solid rocket mo'rors. A variety of space transportation systems have already been studied which avoid these recurring costs; however, several factors such as high technical risk, high development costs, and lack of versatility,have precluded commitment to any particular system to date.

None of the previous systems simultaneously satisfy the low cost and versatility goals.

To provide a comparative base to assess the system proposed in this paper, a brief review of the various space transportation systems studied to date is warranted. Several concepts of fully reusable vehicles with hori-

-A

_717-

i 3 I zontal takeoff (HTO) from runways have been studied. One concept, called the t aerospaceplane (ref. 2), required a very low structural mass fraction,,

scramjet propulsion, and complex air collection equipment. A second reusable

HTO vehicle concept, the airbreathing launch vehicle (ABkV, ref. 3), used

turbojets, ramjets, and scramjets with hydrogen fuel in a first stage, which

separated at a Mach number of about 10. Consequently, the first stage was

larger than the orbiter. Furthermore, the high speed and advanced propulsion

systems of the launch vehicle required advanced technology with its associated

high costs. A third reusable HTO vehicle concept, studied in several variations

from the late fifties thrcugh the mid-sixties, employed a supersonic aircraft

(designed for other missions) as a first stage (refs. 4 and 5). This large

and complex booster stage would have been very expensive. Therefore, the

three reusable HTO vehicle concepts satisfy the versatility goals but not

the low cost goals.

More recent space transportation concepts studied, that avoid the

recurring costs of the space shuttle, include single-stage-to-orbit (SSTO) concepts (refs. 1, 6, 7, 8, and 9). Two of these concepts, which represent

the two takeoff options, are shown in fig. 1. Fig. 1(a) shows a concept designed

for vertical takeoff (VTO), and fig. 1(b) show=s a concept designed for hori-

zontal takeoff from a rocket-powered sled. For a space shuttle payload of

29 500 kg (65k lb m ), the VTO vehicle would have a gross weight of 1.26 M kg

(2.77M lbm ), and the HTO vehicle a gross weight of 1.22M kg (2.68M lb ). These W m weight estimates taken from ref. 6 incorporate some technology advances.

Analyses of similar vehicles by NASA contractors show that with more accele-

rated technology the gross weight of the vehicles can be reduced to 1.14M kg

(2.51M lbm ) for the VTO concept and LOOM kg (2.2M lb m ) for the HTO sled-

launched concept. The VTO concepts employ advanced rocket engines using both

F E 3

Y S '

hydrocarbon and hydrogen fuel. An all-hydro°jen-fueled VTO vehicle would have

a relatively high gross weight of 1,63M kg (3.6M lb^ T) for a 29 500 kg (65k lb ni) payload. The sled-launched HTO concept has been proposed to enable use of

hydrogen °fuel only (by use of a two position nozzle added to the space shuttle

main rocket 'engines (SSME)) at considerably less gross weight than the all-

hydrogen VTO concept. But a sled offers limited takeoff azimuths and few,

launch sites, thus limiting operational versatility. Neither the VTO nor the

sled-launched HTO operational modes offer the convenience of a conventional

runway takeoff mode. Thus, the single-stage-to-orbit concepts may satisfy

cost objectives but lack versatility.

The object of this paper is to present a concept, with supportive analy-

tical results, for a fully reusable horizontal takeoff space transport that

has potentially low initial and operating costs. This potential stems from

use of reusable components, no solid rocket boosters, a lighter orbiter than

the advanced technology SSTO concepts (even with current structures and rocket

technology used for the proposed concept) and sm,ller boosters (with existing

structures technology and SST-type turbojet engines) than the ABLV or supersonic

aircraft boosters discussed above. However, some turbojet engine development may

be required. In addition, this space transport concept offers the potential,of

lateral offset orbit insertion, ferry capability, intact abort and recall landing,

and v ersatility in takeoff location, which in combination may enable round trips

to space from many existing airfields,

SYMBOLS

Measurements and calculations were made partly in the Internationa l System of Units and partly in U.S. Customary Units and data are presented in both 5

systems of units. Mass in the S.I. Units is given in kg, or as required,

in millions of kg (M kg).

Ac inlet area per, engine

CD drag coefficient J lift coefficient CL C lift curve slope L a Cis pitching moment coefficient

thrust coefficient CT D zero-lift drag per booster

DI engine drag (spillage, bleed, and boattail drags)

M free-stream Mach number

% mass flow rate of air based on unit inlet area

mf mass flow rate of fuel

Nb number of boosters

Ne number of engines per booster

q free-stream dynamic pressure

R stoichiometric fuel/air ratio (turbojet =0.066 and

scramjet -0.029)

booster reference wing area S orbiter wing area, reference for aerodynamic coefficients Sref wetted surface area of booster Swetted T net thrust per engine

weight change

angle of attack

expansion ratio

fuel equivalence ratio (actual fuel/air ratio divided by

stoichiometric fuel/air ratio) I E

6 f PROPOSED SPACE TRANSPORT CONCEPT

The proposed space transport concept shown in figure 2 consists of a

winged orbiter with rocket propulsion and on-board propellants and two small

turbojet-powered boosters ;!counted under the orbiter wings. The orbiter

configuration is similar to thos, shown in figure 1 for SSTO vehicles. However,

since acceleration is achieved by turbojets at relatively hidh dynamic pressure,

aerodynamic drag is of greater concern than for the rocket-powered SSTO

vehicles. Moreover, stability is of greater concerti during the boost phase

since the rockets are not in use to provide controlled flight through

gimbaling as for the SSTO vehicles. Consequently, the proposed orbiter is

more streamlined and more stable than the SSTO configurations. At staging,

the rocket engines are ignited and the orbiter has sufficient on-board

propellants to continue ascent to orbit. The rocket engines could be the

SSME's with the expansion ratio increased to 155:1 since the engines would

operate only at altitudes above 15 km (50 000 ft). Orbiter structure is based

on current space shuttle orbiter technology;'however, integral wing tanks

for lox containment could significantly reduce wing weight.

Figure 3 shows the booster concept in more detail than figure 2. The

current booster configuration is a winged body containing a cluster of four

turbojets with afterburners and is pylon-mounted to the lower surface of the

orbiter wing. An aluminum alloy heat sink structure was selected for the

boosters. Each turbojet engine has a two-dimensional, variable geometry

inlet. For shuttle size payloadsa loads of 29 500 kg (65k lbm ), supersonic

transport (SST) class engines with thrust in the 380 kN (85k lb f) range are

required for the selected number of engines. For smaller payloads in the

4500 kg (10k lbm ) range, existing J-58 crass turbojets may suffice. ^P {

a 3

Details of the SST turbojet concept (giving a thrust of 331 kN_(74K lb f ) without afterburner) and its performance are available in ref. 10. The engine

concept selected for this study is based on increased turbine inlet tempera-

tures over current technology and has a variable stator at the turbine inlet.

This later feature may not be required for the space transport application,

thereby simplifying the engine.

The twin boosters are essentially flying engine pods that also contain,

the main landing gear, sized for takeoff at the gross weight of the transport

system. JP fuel for ascent, offset range, and flyback is stored in the

boosters. Following horizontal takeoff, the boosters accelerate the transport

With full afterburner thrust to a staging Mach number of about 3.5. The stage

separation procedure may make use of booster thrust, since the thrust-to-

Weight ratio of the booster exceeds that of the orbiter at staging. After

staging, the boosters fly back to the takeoff site. The boosters are

considered to be remotely-piloted vehicles.

This proposed space transport concept combines the efficient performance

of turbojet engines with efficient: staging of the boosters, which contain the

turbojets and main landing gear. And, of equal importance, the twin booster

concept offers low inert weight of a single booster; thus a potential for low

development costs since these are proportional to the inert weight of one

booster. The staging and propulsive improvements over the SSTO vehicles enable

use of current structures and rocket technology for the orbiter at the same

payload and gross weight as the SSTO vehicles which require advanced tech-

nologies.

1 T ^ ! ! I I ! t A

ANALYSES OF BASELINE CONCEPT

The analyses methods used to determine the payload mass for a given gross

mass are described in detail in reference 8. The input assumptions in the

areas of propulsion, aerodynamics, and structures used for analyses of the

baseline transport concept are discussed in the following sections. The

consumption of the propellants was calculated by integratintj fuel flow rate

for an optimum trajectory using the Program to Optimize Simulated Trajectories

(POST), the use of which is described in reference 8. Figure 4 shows the

ascent trajectory followed for the baseline transport as well as the tra-

jectory on which turbojet performance characteristics are based. Liftoff occurs

a^ a Mach number of 0.38 at an angle of attack of 17° and at a wing loading

1120 kg/m2 (220 lb f/ft2 ). The takeoff velocity may exceed current tire

technology requiring further study to lower takeoff velocity or to increase

tire speed limits. The vehir_le accelerates until the dynamic pressure limit

of', 90 kPa (1880 psf) is intercepted at a Mach number of 1.3. The dynamic

pressure limit is followed until a pull-up maneuver is initiated prior to

staging, which occurs at a Mach number of 3.5. The booster is dropped at

staging, and the orbiter continues to orbit powered by the space shuttle

main engines on an optimized path at'lower dynamic pressure.

Propulsion Data

Table I lists turbojet thrust coefficient, sea level mass flow rate of

air, fuel equivalence ratio, andd thrust specific fuel consumption. The

thrust_ coefficient was assumed to be ,'i nt lependent of altitude and for a

°'°^ +°a ""°^" N.,m"^°was obtained from the relationship

Ne (T^-De ) - D ^T q N (1) Ac 9

In which the net thrust of the engines is reduced by the engine drag and the booster drag. The thrust, drags, and dynamic pressure were evaluated for the turbojet performance trajectory shown in figure 4. To enable calculation of turbojet fuel consumption for other trajectories, the air flow rate was first transformed to sea level conditions (increased by the ratio of air densities at'sea level to those at the turbojet performance trajectory altitude). Then, in the POST analysis, at a selected Mach number, the air flow rate was reduced by^the ratio of air density at the particular altitude to that at sea level.

In effect this procedure approximates the effect of the altitude change between the turbojet performance trajectory and the trajectory selected for space transport study. Using the air flow rate for the baseline concept the fuel flow rate is a function of Mach number and was calculated in the POST analysis from

iilf = ma RgAc Nb N e (2)

in which the stoichiometric fuel-to-air ratio, R,is multiplied by the fuel

equivalence ratio, ^. The fuel consumed is obtained by an integration of the

fuel flow rate, m f , performed by the POST analysis. For reference, table I

lists the thrust specific fuel consumption of the turbojet selected for this -

study. Thrust specific fuel consumption (as tabulated) is based on engine net

thrust minus the engine drag, which consists of spillage, bleed, and boattail drags.

The SSME with the increased expansion ratio produces 2.13 MN (478k lb ) f vacuum thrust at,a-specific impulse of 463 seconds for .in exit area of

8.34 m 2 (90 ft2). Aerodynamics

Estimated lift and drag characteristics of the orbiter configuration are given in figure 5. brag coefficient is shown as a function of Mach number for

selected angle-z of attack of 00 and 5 For a given angle of attack, 10

drag was found by linear interpolation of curves for a more complete set of

angles of attack. The lift curve slope used is shown in the figure.

The estimated zero-lift drag coefficient for the booster (based on

the reference wing area of the orbiter) is shown in figure 6. This estimate

is based on a similar wing-body configuration, studied as a hypersonic

research airplane in reference 11. The POST analysis ignored booster induced

drag.

The drag of the mated orbiter and boosters was assumed equal to the sum

of the drags of the isolated orbiter and booster: no allowance was made for

interference drags between the components. Th.e lift of the combination was

assumed to include a contribution from the booster which increased the orbiter-,

alone lift by about 36 percent. The analysis employed did not provide for

evaluation or a drag penalty associated with the additional lift.

To determine the effect of an increase in transonic an',, supr:rsonic a.•5c,

analyses were also performed with a 5U percent increase in orbiter drag from

M=0.8 to 3.5 for all angles of attack. This increase is believed to be conserva-

tive because induced drag was increased as well as drag at zero lift; whereas,

interference drag at transonic speeds is primarily a Zero-lift wave drag phenomena,

Weights and Structures

The propellant consumption was taken from the trajectory and was input

to the ODIN system of computer programs for orbiter sizing, geometry calculations,

and mass property calculations. The ODIN system, as used in this analysis, is

described and referenced in reference 8. Current space shuttle structures

technology was used for ODIN weights. However, the space shuttle wing was

designed for a qa product of about 144 kN degrees (3000 psf degrees); whereas,

the baseline space transport has a maximum qa of at least 19? kN degrees

MWr,;- :t °" .: - = 1 alp.- i 1 . lama ...... -.^:..,_.,---r^...,z+^n-.-.-,:-.-r-.^- "°'..m^m^ F_...... _. -.-.,--e...,-. ,R..-...,q i ., a ra. ....,.y r _. ,.. y ¢... +•P°"'+^ .^.,.. ^rseeaa-.---^+*m.-.-^.--

(4000 psf degrees). Moreover, the baseline space transport flies at a higher

dynamic pressure than the space shuttle. Consequently,, the estimated wing

weight of the baseline transport is probably less than necessary. However,

the baseline transport has a dry wing (no propellants in the wing), and studies

(refs. 7 and 9) have shown that the structure weight of the wing can be reduced 3.

by storing lox in the wings. The heavy lox produces a load alleviation, thus

reducing wing bending moment and the respective wing weight.

The booster weight at staging was the sum of the turbojets.: inlets, structure,

landing gear (assumed to be 3.0 percent of the gross takeoff weight of the space

transport), miscellaneous items, and fly-back fuel weight. Fly-:back fuel is based

on a range of about 460 Km (250 n.m.). The booster structural concept consists

of a center frame (see fig. 3) that serves as the pylon and support for engines,

inlet ramps, and landing gear. Conventional structure forms the fuselage shell i and wings, except the skins are thickened to provide heat sink thermal protection.

Aluminum alloy has been selected for the booster structure. Unit heat load was

calculated in the POST analysis for the ascent trajectory. Mass analyses show that

19.5 kg/m2 (4.0 lbm/ft2 ) of aluminum at a temperature rise of 167K (300°F) can

absorb the entire heat load from Mach 0 to 3.5 without benefit of radiation cool-

ing; however, with radiation, a skin weight of about 14.7 kg/m 2 (3.0 lbm /ft 2)

should suffice. A weight estimating procedure based on wing loading at gross

weight, ultimate load factor, and wing sweep was used initially to estimate

the weight of the booster structure. This weight estimating pro-

cedure indicated a weight per unit wetted area or 12.7 kg/n, 2 (2.6 lbm /ft 2)

for the conventional fuselage structure and a weight of 20.5 kg/m 2 (4.2 lbm /ft 2)

for the wing structure. Since the wetted wing and fuselage areas are nearly

equal, this results in an average structural weight of 16.6 kg/m ` (3.4 lbm /ft 2}. 12

About half this weight is for internal structure suchas ribs, spars, and fuselage ring frames. This leaves a skin weight of 8.3 kg/m 2 (1.7 lbm/ft2), which is not sufficient to absorb the heat load within the allowable temperature rise. Therefore, the heavier 14,7 kg/m 2 (3.0 lbM/ft 2 ) heat sink skin weight is,us;ed, which results in an average unit weight of 23.0 Wn", (4.7 lb^ /ft 2).

Thus, conservatively, an average structural weight per unit of wetted area of

24.4 kg/m2 (5.0 lbm /ft 2 ) was estimated for the booster structure.

RESULTS AND DISCUSSION FOR BASELINE CONCEPT I

A discussion of the results for the baseline transport using current tech- . nology is given. The optimum trajectory, weight statement, payload performance, and results of aerodynamic calculations are discussed in the following sections.

Optimum Trajectory

The optimum trajectory obtained by the POST analysis for the baseline concept is shown in figure 7 as a function of flight time. Altitude to

Mach 6 is shown in figure 4 as a function of Mach number to enable comparison between the trajectories used forthe turbojet performance and the baselinE transport. Figure 8 shows the available thrust and drag as calculated in the POST analysis.. At all Mach numbers - including transonic the thrust exceeds the drag; however, a 50 percent increase in transonic and supersonic drag would require more thrust. Either larger turbojet engines or a cluster of five or more engines =in each booster fuselage may be required instead of the engine size and cluster of four shown for the baseline concept in figures 2 and 3. As seen in figure 7, the winged hooster and orbiter space transport flies at low flight path angles unlike vertical takeoff transports.

Flight path angles are less than 6° during boost and 3 0 during rocket ascent, T ^_____I_ I I. I 1 _

similar to airline-type ascent. Staging occurred at about 12 minutes after

takeoff; and after about 9 minutes of rocket ascent, orbital velocity and

altitude are achieved. The maximum acceleration was held at 1.7g to avoid

high loads which results in little payload penalty.

Weight Statement

Table II is a weight statement for the baseline space transport concept. Or-

biter weights are from the ODIN analysis, and all orbiter weights are based on

current technology. Unit weight of tie orbiter structure plus thermal protection

system is 27.5 kg/m2 (5.65 lbIn /ft 2 ), comparable to that of the space shuttle orbiter.

Booster Weight estimates are based on the heat sink requirement and

structural weight estimating procedure discussed earlier, and on unpublished

engine and inlet weight data, calculated fly-back fuel weight, and assumed

weights for other items. The main landing gear was assumed to weigh 3.0 percent

of gross takeoff weight. An estimate of about 3.5 percent of gross takeoff

weight might be more representative of current technology.

Performance

The baseline transport concept, as described previously, has space

shuttle structures and rocket engine technology. Use of the turbojet

boosters provides shuttle payload capability as well as a fully reusable

system with conventional horizontal takeoff. The payload to orbit for the

baseline transport, as determined by the POST and ODIN analyses, is 29 800 kg

(65.5k lbm ). This performance is achieved by use of the increased expansion G ratio (€ = 155:1) for the space shuttle main engines. At the current expansion

ratio (e = 77.5:1) the payload would be reduced to 25 000 kg (55k lb m ). The

current SSME has the lower expansion ratio to satisfactorily operate at sea level;

-however, the baseline transport is launched from the boosters at 17.35 km (57 000 ft),

u. ^ W r

14 thus the larger expansion ratio may be used for improved SSME performance. The increased SSME rocket nozzle area will increase orbiter base drag requiring further study to assess the real effectiveness of the greater expansion ratio.

Table III shows comparisons of the baseline transport and various

si''ngle-stage-to-orbit concepts (see ref. 6) for a 29 500 kg (65k lb m ) payload

to; orbit. Both rocket and structures technologies are compared as well as

gross and dry weights. Gross weights reduce from a maximum value of

1.63M kg (3.60M lb m ) for the all-hydrogen VTO concept to 1.26M kg (2.77M lb m)

for dual-fuel VTO with advanced rockets to 1.22M kg (2,68M lb m ) for an all-

hydrogen orbiter with a rocket powered sled, which is essentially a M = 0.6

stage, to a minimum value of 1.18M kg (2.6M lb m ) for the HTO Mac h 3.5 turbojet

booster space transport. Only the baseline transport has a current technology

structure.

Total dry weight reduces in the same order as gross wei g ht for the

Various SSTO vehicles; however, the baseline space transport has a higher

total dry weight than the dual-fuel VTO and sled-launched HTO. Should the

sled weight be included, only the dual-fuel VTO concept would weigh less

than the baseline concept. The baseline concept has the lower weight orbiter,

so the boosters are responsible for the higher dry weight than the dual-fuel

INTO concept. Since only one booster must be developed, only half of the

booster dry weight effects development costs. Also, the booster weight

is largely turbojet weight, and since only one turbojet must be developed

mach of the total turbojet (8 engines) weight is not indicative of developY

ment cost. Moreover, the booster structure is a more economical concept

than the orbiter structure further indicating potential for low development

cost. i l

Aerodynamic Calculations

Although performance analyses used estimated aerodynamics for the

bajseline concept, aerodynamic calculations were simultaneously performed

to guide configuration definition. Of primary concern was the transonic

and supersonic drag of the mated configuration. Initial calculations were

for the orbiter and booster separately, since total drag was hasad on the

SUM of the drag of each stage. Later analysis of the mated configuration

confirmed that total calculated wave drag is nearly equal to the sum of the

wave drag of the stages. These analyses used the GEMPAK code (ref. 12) which

gives the geometry definition. GEMPAK is coupled with aerodynamic codes

used in the analysis of the baseline concept. These coupled codes include

a vortex lattice program (ref. 13) for subsonic aerodynamics, the Harris

Wave Drag Program (ref. 14) for supersonic aerodynamics, and the Gentry

Program (ref. 15) for hypersonic aerodynamics.

Results of the calculated drag for the booster are shown in figure 9 and

are compared with that estimated for the booster in the POST analysis. As

indicated;, the calculated drag is greater than the estimated values; however,

the greatest difference amounts to only 15 percent error in the total drag

of the mated configuration.

Calculated drag for the orbiter is shown in figure 10 and compared with

that estimated for the orbiter in the POST analysis. Also shown is the

-estimated drag increased by 50 percent in the M = 0.8 to 3.5 range. As

indicated, the calculated drag for the baseline transport orbiter is near

the 50 percent margin in the critical transonic range, which requires

maximum turbojet thrust.

^, T 16

Prior to the calculation of drag, an increase over the estimated value

oforbiter drag (used for the POST analysis) of 50 percent in the Mach numb'er'

range of 0.8 to 3.5 was assumed. Performance analysis with 'a resized vehicle

which increased the turbojet thrust by 25 percent with this increased drag

shows a reduced payload by 1800 kg (4k lb ), or the payload is about 28 100 kg m (62k lbm ) for the 1.18M ka (2.6M lb m ) gross takeoff weight. This payload loss

1 indicates that the performance is not extremely sensitive to transonic drag,..

only a 6.1 percent decrease in payload mass fraction results. However, i transonic and supersonic drag will be a major factor affectinq concept

feasibility since turbojet power is used for acceleration. The required

25 percent increase in thrust may be achieved either by larger turbojets or

by a cluster of five engines in each booster instead of the cluster of four

engines shown for the baseline concept in figure 3. i

Analytical methods do not indicate whether high drag is present as a

result of flow separation due to shock impingements on opposing surfaces of

the boosters and orbiter. Wind tunnel tests are required to determine if

and where flow separation exists and should it exist the necessary configuration

changes to reduce the drag. Results to date are based on estimates and

preliminary analyses. Radically different configurations may be required to

achieve satisfactory results should the wind tunnel derived drag greatly

exceed the calculated drag.

Aerodynamic calculations were also made to determine booster landing

characteristics and booster and orbiter stability data. These analyses show

the booster lands at about 124 m/s (240 knots), is stable, and has sufficient

control power to flair for landing. Subsonic stability is shown in

figure 11 for various booster configurations studied. The upper

f V"' ^F ^^ . 7-5 -.^.,.-_"'s.^--r .__,...r... _. . ^:.. ---•-..,--... --aa ,. ...,.a .. , :: ,..,, a .... ., .....Fr•-•._....„'^°1 f" r ^

......

configuration is seen to be unstable over part of the lift coefficient, CL,

range as indicated by the positive slope of the curve at low C , values.

The next lower configuration has a more rearward wing and is shown to be

stable; however, the elevons did not produce adequate control power for

landing. The third configuration has the pylon extended rearward with a

- horizontal control surface which does provide sufficient control power.

However, the vertical tail on this boom interferes with the orbiter wings.

The last configuration, shown at the bottom of figure ll,has swept trailing

edge wings with tip fins to provide a more rearward elevon location for

adequate control power.

Hypersonic, Mach 6, stability for ascent conditions is shown in figure 12

for various orbiter configurations studied. The first configuration had the

most forward wing location and was unstable. The second configuration has

trailing edge sweep and a more rearward wing than the first and is stable.

However, without application of area rule, the calculated drag increase was more

than the 50 percent increase. estimated for the POST analysis. So a third-more

streamlined-configuration was studied; it too is stable. This third configura-

tion is longer and has more wetted area than the first two orbiter configurations.

Consequently, recent analysis includes the application of area rule to shorter

configurations to achieve a drag reduction with minimal increase in wetted

area.

CONCEPT ALTERNATIVES AND REFINEMENTS

To assess the growth potential of the baseline space transport concept,

several technology improvements were studied. In addition, a concept of

stretching the current space shuttle orbiter as an alternative to an all-new

orbiter for two classes of payload is discussed.

V, Advanced Structure

Applying advanced structure (described in reference 1) to the orbiter

results in a structure that is 25 percent liqhter than the baseline structure

and a payload of about 45 400 kg (100k lb ). m As seen in figure 13, this is 50 percent more than the payload of the baseline concept.

As for the case with the baseline concept using existing structure

technology, a 50 percent increase in orbiter transonic drag (now found to be

a more realistic estimate) reduces the payload to 43 500 kq (96k lbm ) for

the 1.18M kg (2.6M lbm ) gross weight. This loss in payload due to a 50 per-

cent increase in transonic drag is only a 4.0 percent loss in payload mass

fraction.

Advanced Rockets

Replacing the improved shuttle rockets with advanced rockets offers a

payload of 49 900 kg (110k lb m ), as shown in figure 13. These rockets are

more advanced than those listed in Table III for the dual-fuel VTO, SSTO

concept. The rockets listed in Table III are separate engines, i.e,, hydrogen-

oxygen engines and hydrocarbon-oxygen engines. Whereas, the advanced rockets

used in obtaining the data in figure 13 are dual-expander engines, ref. 1.

These engines burn both hydrocarbon and hydrogen fuel through different

tt internal ports within the engine; they are considered to require more advanced

technology than the separate dual-fuel rockets. A benefit yet to be assessed

is the reduced drag due to the reduced orbiter volume made possible by use of

the denser hydrocarbon fuel. Consequently, the advanced rockets may offer a

greater payload than shown in figure 13. Moreover, the reduced volume also

means reduced dry weight and development costs.

19 A Advanced Scramjet

Airbreathing propulsion may be used for the orbiter as well as for the

boosters. However, at Mach numbers starting at about 3.5, a dual mode scramjet

is required since supersonic transport turbojets are currently limited to about

Mach 3.5 (ref. 16). Figure 14 shows the orbiter with a scramjet and SSME

rockets and the turbojet boosters. A baseline fixed geometry scramjet has

been described in reference 17. Scramjet thrust coefficient, air flow rate,

and fuel equivalence ratio are given in Tattle IV as a function of Mach number and

velocity. This data was used in the POST analysis in a like manner to that for the

turbojet. Unpublished studies of the baseline scramjet indicate that it

will weigh about 1260 kg/m 2 (258 lbm /ft `) based on engine inlet area. At

this weight, the POST and ODIN analyses indicate the payload with the scramjet

is about the same as without the scramjet.

The weight estimate of the scramjet was based on stress

analyses of rectangular and circular combustor cross sections, and the

combustor length was considered to be a variable. The structural concept

analyzed was full-depth honeycomb-core sandwich, which also included a plate-

fin sandwich for cooling all interior surfaces. At selected stations an

optimum weight and thickness of honeycomb-core sandwich were calculated based

on a minimum sum of core and face-sheet weights. The design load was the

internal pressure, and material properties were based on local temperatures.

Shown in figure 15 is the weight of a scramjet module with the baseline

rectangular combustor and with a circular combustor. The baseline scramjet

has rectangular sections throughout. At the baseline combustor length, a

40 percent reduction in weight is indicated for a circular combustor. This

is due to the fact that a circular cross section is more efficient than a

l _17

^ 9 ^ } g ! `3

rectangular section for support of the high pressure in the combustor.

For a 30 percent reduction in combustor length an added 10 percent reduction

in engine weight results, giving a total weight reduction of 50 percent.

With this advanced lightweight scramjet and an advanced structure, the

POST and ODIN analyses indicate a payload of 61 400 kg (135K lb m ) for a

gross takeoff weight of 1.18M kg (2.6M lb ). As seen in figure 13, this m ..^ payload is double that of the baseline space transport concept. However,

further study is needed on thrust and weight of the circular combustor scram-

jet and on ascent and entry cooling of the scramjet when not in use.

Stretched Shuttle Orbiter

Results presented thus far are based on new orbiter configurations;

however, a concept of stretching the current shuttle orbiter was also studied

and is shown in figures 16 and 17. Figure 16 shows a comparison of the proposed

space transport concept (labeled Spacejet) with the current space shuttle. Nu-

merous items from the present shuttle might be used in a new vehicle. For

example, the external liquid hydrogen tank . may form the forward fuselage of the

stretched orbiter as indicated in figure 17. Other possible common components

are the crew accommodations, avionics, payload bay with its composite material

doors, main rocket engines, and on-orbit propulsion system. The wing, tail, and

some of the fuselage of the stretched shuttle cannot be directly taken from the

current shuttle, because they must be larger, and operate in a different environ-

ment; however, the type of structure is the same as the shuttle orbiter.

Although this composite vehicle has not been analyzed to the depth of

the all-new orbiters, tentative results indicate a payload equal to the current

shuttle (29 500 kg (65k lbm )), at a gross weight of only 1.27M kg (2.8M lb M).

This gross weight may be reduced by use of the higher expansion ratio (155:1)

SSME nozzles. In addition, the enlarged nozzles require less orbiter stretching ^ f

f f f 21

than the current nozzle expansion ratio. Gross weight for the stretched

shuttle orbiter and turbojet boosters is estimated to be 0.73M kg (1.614 lb in )

less than the current space shuttle principally due to elimination of the

solid rocket motors. More study of the stretched shuttle orbiter concept

is,warranted, because it potentially offers a fully reusable, horizontal

• takeoff system with significantly less operating cost than the current shuttle

and lower development costs than an all-new orbiter.

Small Payload Vehicle Concepts

The proposed turbojet booster space transport concept has been sized for

both space shuttle and utility vehicle (ref. 1) class payloads. -Fi g ure 18

compares these two payload class vehicles for new orbiter configurations

and for stretched shuttle orbiters where each vehicle has current structure

and propulsion technology. Results are for dry wing orbiters; should integral

wing tanks be used for lox, orbiter size and takeoff weight may be reduced.

As seen for the 29 500 kg (65k lbm ) shuttle payload class transport, in the new orbiter is about 65.5 ft) long, and the stretched shuttle

orbiter is about 77.7 in (255 ft) long. For the 4500 kg (10 k lb m ) class

payloads, the new orbiter is only 50.3 in (165 ft) long, and the stretched in shuttle orbiter is about 61 ft) long. As seen in figure 18, the

gross weights are greatly reduced for the utility vehicle payload. In

addition, the booster turbojets may be existing J-58 class engines. There-

fore, the proposed turbojet booster transport concept is even more attractive

for the smaller payload due to the reduced size and cost.

STUDY STATUS

As indicated, the above results are based on preliminary estimates of weight,

drag, and engine performance. Trajectory analysis was current art, and preliminary 22 vehicle sizing techniques were used. Considerable effort is needed to verify the concepts presented and to determine the economic and operational merits.

Problem areas requiring early study are: booster and orbiter configuration. refinement, wind-tunnel tests particularly at transonic speeds for the individual stages and the mated configuration to improve the aerodynal^ic estimates, introductory study of staging and sonic boom considerations, structural design and analyse.;, engine performance analyses, and effect of payload class on design of the turbojet booster transport.

CONCLUSIONS

A preliminary study of a novel space transportation concept employing a winged orbiter with on-board ascent propellants and two s.-;ill turbojet boosters has been performed. This study based on estimates of aerodynamics, weights, and engine performance has led to the following conclusions:

(1) A horizontal takeoff, fully reusable space transport concept using current structures and rocket technology andl SST class turbojets in twin,

Mach 3.5 boosters appears to be,feasible. This baseline concept is pro- jected to have a payload mass fraction equal to the single-stage-to-orbit vehicles, which require advanced structures technology. Moreover, the proposed space transport concept has potentially much lower operating cost than the NASA space shuttle, similar to the SSTO,vehicle concepts and considerably greater operational versatility than the SSTO concepts.

(2) Alternative refinements to the baseline space transport concept were studied and indicate the following:

(a) Use of an advanced structure with 25 percent lowerunit-area weight than the present space shuttle orbiter is expected to provide a

50 percent increase in payload. 23

(b) Use of advanced dual-fuel rockets increase the payload by an

additional 10 percent.

(c) Use of an advanced scramiet on the orbiter with advanced orbiter structure increases the payload to twice that of the baseline space

transport.

(d) The turbojet boosters may be applied to a stretched version

of the present space shuttle orbiter rather than the completely new orbiter

of the baseline concept.

(e) The proposed space transport 'concept appears attractive for

smaller payload class utility vehicles with either an all-new orbiter or a

stretched shuttle orbiter.

(3) Further effort is warranted in configuration refinement, transonic

aerodynamics, staging techniques, sonic boom constraints, structural design

and analysis, and eng i ne performance as well as economic and operational

aspects to verify the concept of a space transport with two small turbojet-

-powered boosters. 1. Henry, Beverly Z. and Eldred, Charles H.: Advanced Technology and Future

Earth-Orbit Transportation Systems. AIAA Paper No. 77-530, May 1977.

2. BOoda, tarry: USAF Plans Radical New Space Plane. Aviation Week and

Space Technology, Vol. 73, No. 18, Oct. 31, 1960, pp. 26.

3. Morris, R. E. and Villiams, N. B.: Study of Advanced Air Breathing Launch

Vehicles with Cruise Capability — Volume 3, Aerodynamics, Propulsion,

and Subsystems (U). Final Report, 2 Jan. - 3 Oct. 1967. Lockheed

California Company, NASA CR-73196, Feb. 1968.

4. Keller, R. B. and Wheaton, R. J.: Preliminary Study of Air Breathing

First Stage Boost Vehicle. Allison Div., General Motors Corporation

Report No. EDR1169, March 17, 1959. (C)

5. Pierpont, P. Kenneth: An Aircraft Launch Vehicle for Military and

Scientific Space Missions. Presented at the AIAA Military Aircraft

Systems Meeting, Dallas, Texas, October 18-19, 1966!

6. Henry, B. Z. and [)ecker J. P.: Future Earth Orbit Transportation Systems/

Technology Imp ications. Astronautics and Aeronautics, V. 14, No. 9,

Sept. 1976, pp. 18-28.

7. Bangsund, Edward L'., et al: Technology Requirements for Advanced Earth-

Orbital Transportation Systems., NASA CR-2879, 1977.

8. Martin, James A.: An Evaluation of Composite Propulsion for Single-Stage-

to-Orbit Vehicles Designed for .Horizontal Takeoff. NASA TMX-3554, 1977.

9. Haefeli, Rudolph C., et al: Technology Requirements for Advanced Earth-

Orbital Transportation Systems. NASA CR-2866, 1977. ^ I

d n

10. Barber, Hal T., Jr. and Swanson, E. E.: Advanced Supersonic Technology

Concept AST-100 Characteristics Developed in a Baseline-Update Study.

NASA TM X-72815, Jan. 1976.

11. Hearth, D. P. and Preyss, A. E.: Hypersonic Technology -_App'roach to an

Expanded Program. Astronautics and Aeronautics, Vol. 14, No,, 12,

Dec. 1976, pp. 20-37.

12. Stack, Sharon H., Edwards, C. L. W., and Small, W. J.*, GEMPAY: An

Arbitrary Aircraft Geometry Generator. NASA TP-1022, 1977.

13. Lamar, J. E. and Gloss, B. 6.: Subsonic Aerodynamic Characteristics of

Interacting Lifting Surfaces with Separated Flow Around Sharp Edges

Predicted by a. Vortex-=Lattice Method. NASA TM D-7921, Sept. 1975.

14. Harris, Roy V., Jr.: An Analysis and Correlation of Aircraft Wave Drag..

NASA TM X-947, 1964.

15. Gentry, Arvel E. and Smyth, Douglas N.: Hypersonic Arbitrary-Body

Aerodynamic Computer Program (Mark ILI Version). Rep. DAC 61552 (Air

Force Contract Nos. F33615 67 C 1008 and F33615 67 C 1602), McDonnell

Douglas Corp., April 1968. Vol. I User's Manual (available from DDC

as AD 851 811, Vol. II - Program Formulation and Listings (available

from DDC as AD 851 812).

16. Denning, R. M. and Hooper, J. A.: Prospects for Improvement in Efficiency

Of Flight Propulsions Systems. Journal of Aircraft, Vol. 9, No. 1,

Jan. 1972.

17. Henry, J. R. and Anderson, G. Y.: Design Considerations for the Airframe-

Integrated Scramjet. NASA TM X-2895, Dec. 1973,

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TABLE 11 - TURBOJET BOOSTER SPACE TRANSPORT WEIGHT STATEMENT

Weight 7 Orbiter component kg (Tbm)

W4--g group 13 031 (28 728); Tail group 3 ,;697 (8 151) ; Body group 381!,44 (84 533) Thermal protection system 21 858 (48 188); Landing gear and docking 5 920 (13 052) Propulsion - ascent 9 417 (20 760); Propulsion - auxiliary 2123 (4 681) Prime power 1 '774 (3 912) Electrical conversion and distribution 1 ' 634 (3 602) Hydraulic conversion and distribution 1760 (3 881); Avionics 2 021 (4 445) Environmental control 1 857 (4 093) Personnel provisions 790 (1 742) Grow th - contingency _._._...__. _.._.^.. _._ _.,_ _. _ _.._^.._..._...... _ 8 164 17 998) --- Dry weight ... _ 99 309 ._ (218 935)_ Personnel 705 (1 555) Cargo - payload 29 769 (65 629) Attitude control propulsion system reserves 68 (150) Residuals -4 679 (10 316) eigh Landing w t r _ __._ _..__' _1149 734 (330 _101 . Attitude control propellant 4 536 (10 000} En try weig ht 154 270 (340 101) 1 Reserve fluids 334 (2 942) Inflight losses 1 334 (2 942) Pro pell ant - ascent 667 161 (1 470 813) Gross weight - orbiter 824 100 (1 816 797) li TABLE II - Continued ._.^. .,. ._._.._w._._.^. _... _._..Weight .^

Booster component kg (lb ) m Structure group 19 000 (42 000) Propulsion with inlets 31 800 (70 000) Landing gear 17 700 (39 000) Equipment 2 300 --(5-000). Dry weight + ^ 70 800 (156 000) Fly-back fuel __._._._.._.__._.._..._._^. 3 600 (8 000) - -°---_.. • • ^ Staged weight 14 400 (164 000) Boost fuel 102 000 (225 000) '_... _ .. _.. _...... _.. 1 Gross weight - booster 176 400 (389 000) Gross weight - two boosters 352 800 (778 000)

Gross weight space transport j: 177 000 (2 595 000) r ^

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1. Report No. 2. Government Acceesion No. 3. Recipient's catalog No TM-74087 j 4. Title and Subtitle 5 Report Date A FULLY REUSABLE, HORIZONTAL TAKEOFF SPACE TRANSPORT O ctober 1977 CONCEPT WITH TWO SMALL TURBOJET BOOSTERS 6. Performing Organization lode

7 Author(s) 8. Performing Orgamzauon Report No L. Robert Jackson, James A. Martin, and William J. Small 10, Work Unit No 9. Performing Organization Name and Address 505-11-31-02 I NASA-Langley Research Center 11. Contract or Grant No Hampton, VA 23665

13 Type of Report and Period roc eiei dr., i 12 Sponsoring Agency Name and Address Trschnir,al hlelnora ►ldum National Aeronautics and Space Administration 14. Sponsoring Agency Code Washington, D.C. 20546

15 Supplementary Notes r i

16. Abstract

Results of a preliminary study of a novel space transport concept are presented. The concept consists of a winged orbiter containinct ascent propellants and two small turbojet-powered winded boosters, used For accelerdtion to supersonic speeds. The concept offers full reuse and horizontal takeoff from numerous existing airports. With current structure and rocket technology this transport concept has lower gross weight for a selected payload than single-stage-to-orbit concepts, which require structural advancements. Dis- cussion includes alternatives to the baseline space transport concept which improve performance; such as advanced structures, dual-fuel rockets, and lightweight scramjets for the orbiter. Moreover, a concept of usinq a stretched shuttle orbiter instead of an all-new configuration is discussed for two payload-class vehicles.

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Y 17 Key Words (Suggested by Authorls)) 18. Distribution Statement Boosters 1 Turbojet Unclassified-Unlimited Orbiter I Horizontal takeoff i Space transport Subject Cateao v: 15 19, Security Classif. (of this report) 20, Security Classic. (of this page) 21. No. of Pages 22. Price' Unclassified Unclassified 48 $4.50

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