I. 

NASA Technical Memorandum 101513 

TWOlSTAGE REUSABLE 

LAUNCH SYSTEM UTILIZING 

A WINGED CORE VEHICLE 

AND GLIDEBACK BOOSTERS 

Ian 0. MacConochie 

James A. Martin 

James S. Wood 

Miles 0. Duquette 

July 1989 

- 

(flASA-TM-lo%!513) TUO-STaGE REUSABLE UUUCH B89-268 78 

6UDBBACX BOOSTERS (HAS., Langley Besearch 

Center) 21 p CSCL 22B Unclas 

t SYSTEil UTILIBIFSG A UIBCBD COBE VlzEICLE BllD 

63/16 0324583 # 

National Aeronautics and 

Space Administration 

Langley Research Center 

Hampton, Virginia 23665 
TWO-STAGE REUSABLE LAUNCH SYSTEM UTILIZING A WINGED CORE VEHICLE 

AND GLIDEBACK BOOSTERS 

BY 

Ian 0. MacConochie 

James A. Martin 

James S. Wood* 

Miles 0. Duquette* 

ABSTRACT 

A near-term technology launch system is descrlbed in which Space 

Shuttle main engines are used on a manned orbiter and also on twin strap-on 

unmanned boosters. The orbiter is configured with a circular body and 

clipped delta wings. The twin strap-on boosters have a circular body and 

deployable oblique wings for the glideback recovery. The dry and gross 

weights of the system, capable of delivering 70 klb of cargo to orbit, are 

compared with the values for the current Shuttle and a core vehicle with 

hydrocarbon-fuel ed boosters. 

INTRODUCTION 

In recent conceptual design studies of launch vehicles (Ref. l), 

emphasis has been placed on reducing operational complexity by employing 

comnonality in systems and propellants. 

has been configured in which liquid oxygen, liquid hydrogen, and current 

Space Shuttle main engines are used in both a manned core vehicle (orbiting 

stage) and its strap-on unmanned boosters. The principal objective of this 

study was to investigate the size and performance of an all-oxygen/hydrogen 

system using fixed numbers of Shuttle main engines. A parametric analysis 

was conducted to determine the optimum number of engines in the boosters 

and core vehicle in the presence of varying payload weights and volumes. 

The performance of the resulting vehicles was then compared with a recent 

study in which an oxygenlhydrogen propulsion subsystem was used in the core 

vehicle, but a lower performing, higher density hydrocarbon was used in the 

boosters. 

In this regard, a launch vehicle 

CBV 

GLOW 

9 

LOX 

LH2 

OMS 

PIL 

POST 

Ixx, 

Xcg Y 

ABBREVIATIONS AND SYMBOLS 

Circular body vehicle RCS Reaction control system 

Gross 1 iftoff weight SRB Solid Rocket booster 

Gravitational constant, 32.2 ft/sec2 SSME Space Shuttle Main Engine 

Liquid oxygen TSLS Thrust (Sealevel static) 

Liquid hydrogen T/W Thrust-to-weight ratio 

Orbital maneuver system 

Pay 1 oad 

Program to Optimize Simulated Trajectories 

Iyy, Izz moments of inertia about x, y, and z axes, respectively 

Ycg, Zcg, center of gravity locations in x, y, and z directions 

*George Washington University, Joint Institute for the Advancement of 

Flight Sciences 
2 

DESCRIPTION OF LAUNCH SYSTEM 

In the current study, the launch system elements are similar in 

geometry to the systems described in reference 2; the most visible change 

i s i n the relative sizes of the boosters and core vehicles. The same core 

and booster geometries were used for both the all-LOX/LHz and the 

LOX/LHz-RP propulsion system to enhance the accuracy of the comparisons. 

In all the designs, both the core and booster vehicles are simple 

circular shapes having ogive forebodies. The 

replicas of the core vehicle body; however, an 

of the clipped delta wing (fig. 1). 

Core Vehicle 

The manned orbiter is a circular body veh 

WinclS, similar to the core vehicle detailed in 

ooster bodies are geometric 

oblique wing is used in lieu 

cle with clipped delta 

reference 1. Core vehicle 

geoietry is shown in figures 2 and 3. The core vehicle is designed for 

vertical takeoff with the two strap-on boosters but lands horizontally. 

Crew accomnodations are located in the mid-body section, between the LH2 

propellant tank and the payload bay. The payload bay is located 

imnediately aft of the crew station, and its shape is dissimilar to that of 

the current Shuttle payload. The payload bay on the study configuration is 

circular in cross section and spans the inner diameter of the mid-body 

sect i on. 

No provision is made for a canopy or windshield in the cockpit. This 

minimizes structural cut-outs and results in a significant savings in 

structural weight and a reduction in aerodynamic drag. Flush mounted 

viewports provide exterior visual access, while additional ports allow 

observation from the cockpit into the payload bay. A remote camera system 

deployed from the nose is used for forward pilot vision during final 

approach and landing. 

The core vehicle employs tip fins for energy management and as a 

redundant method for directional control. The characteristics of the tip 

fins are detailed in reference 3. The core vehicle employs a forwardmounted dorsal fin for primary directional control. Subsonic control 

characteristics for the dorsal fin are reported in reference 4. The 

forward placement of the dorsal fin requires an active control system, but 

it eliminates the need for a large vertical control surface at the rear of 

the vehicle, thus reducing ascent drag and structural weight. 

Boosters 

The unmanned boosters used in this study are also very similar to 

their counterparts in reference 2. The propellant tanks are integral with 

the body. Relative body dimensions (ogive section dimensions and body 

radii as fractions of reference length) equal those of the core vehicle. 

For the purposes of this study, the structure of the boosters is entirely 

Shuttle-technology skin stringer aluminum. Booster propellant tank 

weights, however, represent a 10-perCent increase over current Shuttle 

external tank weights to allow for the weight of additional tank structure 

necessary to make the tanks reusable. Booster vehicle geometry is shown in 
3 

figures 4 and 5. The boosters employ oblique wings to facilitate an 

unpowered glide back to the launch site after staging. To allow for the 

unpowered return, the boosters are staged from the core vehicles at Mach 

3. As in the designs shown in reference 2, the pivot assemblies for the 

oblique wings on the boosters are mounted on linear bearings with a worm 

lead screw and motor drive so that the wing can be driven rearward during 

deployment to the approximate body station for trimned flight. Wings 

without the translational feature are currently under study for use in 

subsonic and transonic aircraft (ref. 5). The boosters shown here in the 

launch vehicle booster application are conceptual. The booster propellants 

are crossfed to the core vehicle until staging. Advanced avionics and 

propellant management subsystems are assumed, as well as electric actuators 

for control surfaces. 

ascent or return, heating is moderate, and the heat sink capability of the 

all-aluminum structure is relied upon to accomnodate the thermal load. 

Because the boosters do not exceed Mach 3 during 

APPROACH 

The engines on both the orbiter and boosters were assumed to be 

current SSMEs. Mission-related assumptions include ascent of the core 

vehicle plus payload to a 50- by 100-nautical-mile orbit from a due East 

launch. Accomnodation for two crew members and a mission duration of not 

more than 72 hours were assumed. An ascent acceleration limit of 3 g was 

also assumed. 

A weights and sizing program, described in reference 6, was used for 

estimating the weights of the various subsystems. Structural weights are 

based on size or loading or combinations of both. Other subsystem weights 

and sizes were based on such factors as mission length, crew size, power 

requirements, and control surface sizes. The types of subsystems, 

structural materials, and the general configurations were selected from the 

p r og ram. 

A propellant weight fraction is used herein. It is defined as the 

ratio of propellant weight to gross system weight. The required values 

were obtained from the Program to Optimize Simulated Trajectories (POST) 

(Ref. 7). The propellant weight fractlons from POST were inputs for the 

weights and sizing program. When supplied with these inputs, the weights 

and sizing program iterates until the specified propellant weight fraction 

is obtained. This involved the trending of propellant loads (and therefore 

tank sizes), body volumes, and a1 1 related subsystem weights. 

From the weights and sizing program outputs, a plot was made of 

propulsion system weights versus payload weights (Fig. 6). From this plot, 

two core vehicles with propulsion system weights corresponding to two and 

three SSMEs, were selected for further analysis. The two-engine core 

vehicle requires a weight allocation of 20 klb for the propulsion system, 

whereas the three-engine configuration requires 30 klb. The a1 locations 

for the propulsion subsystem weights include 7500 lb for each SSME and 2500 

lb per engine for the preSSUriZation and feed system. In the study, the 

core vehicle mass properties were then held constant, and the program was 

used again to size the boosters. Considerable iteration was necessary to 

size the boosters such that an Integral number of engines were selected for 

each booster and that the system thrust-to-weight ratio at lift-off was 1.3 

or higher. 
4 

Boosters with main propulsion subsystems having three and four SSMEs 

satisfied the minimum thrust-to-weight ratio requirement of 1.3. This 

yielded two launch system configurations--a two-engine core matched to two 

three-engine boosters (2-3-3 configuration) and a three-engine core matched 

to two four-engine boosters (3-4-4 configuration). Lift-off thrust-toweight ratios were approximately 1.5 for each configuration. 

configurations for the 2-3-3 and 3-4-4 systems are shown in figure 1. 

established, further study was conducted on the weight of the additional 

structure necessary to accomnodate the booster oblique wing. Equations for 

the weight of an oblique wing for a supersonic transport aircraft are found 

in reference 8. 

modified and applied to the glideback boosters. A provision was also made 

for the weight of the electric motors used to pivot the wing assembly. 

Launch 

After the basic core and booster vehicle subsystems and weights were 

The basic weight equations developed in reference 8 were 

RESULTS 

The two-engine core, three-engine boosters (2-3-3) combination 

resulted in a payload capability (ascent and return) of 37 klb. System 

GLOW for the 2-3-3 configuration, with a 37-klb ascentlreturn payload, was 

1,940 klb. The three-engine core, four-engine boosters (3-4-4) combination 

yielded a 70-klb payload. System GLOW for the 3-4-4 configuration, with a 

70-klb ascent and return payload, was 2,901 klb. Launch vehicle 

comparisons are made in Table 1 for the two all-LOX/LHz vehicles, the 

current Shuttle system, and three other systems from an earlier report 

(Ref. 2). The launch vehicles are ranked according to ratios of 

payload to dry weight and payload to gross weight. 

Based on the highest ratio of payload to dry weight, the five-engine 

core with RP strap-on boosters ranks number one. In this launch system, 

LOX/LH2 propellants are crossfed from tanks within the booster to the 

orbiter. Tanks dedicated to the LH2 crossfeed propellant are required 

within the RP boosters. Based on the payload-to-gross-weight ratio, the 

all-LOX/LHz 3-4-4 launch system ranks number one. 

proportional to payload-to-dry-weight ratio, then the LOX/LH2 core with RP 

boosters would be the most economical System. In reality, the all-LOX/LHz 

launch vehicles would probably yield the lowest cost system, since no new 

main engine development is required. A new RP engine would have to be 

developed for a LOX/LHz/RP propellant launch system. The five-engine, 

fully reusable core vehicle with Shuttle SRBs having filament wound cases 

ranks fourth for both payload-to-dry-weight and payload-to-gross-weight 

ratios. 

If cost is assumed to be 

Factors which could affect the rankings (Table 1) if the vehicles were 

to be normalized for mission and payload capabilities are as follows: 

(1) The current Shuttle is capable of remaining in orbit for 

approximately 8 days and can accomnodate a crew of at least 

seven. It has a 15-ft-diameter by 60-ft-long cargo bay. The 

launch vehicles with which the Shuttle is compared can 

accomnodate a crew of two for 3 days. The cargo bays are 30 ft 

in diameter by 15 ft long. 
5 

. 

(2) The vehicles compared In Table 1 have somewhat different payloads 

as an outgrowth of the many sizing constraints, principally that 

of englne size. 

equal payloads the ratios would change slightly. 

If the launch vehicles were to be resized for 

The 2-3-3 launch configuration reaches 3 g at staging without 

throttling either core or booster engines (Fig. 7). The 2-3-3 boosters 

stage 86 seconds after launch at an altitude of 86,000 ft, and the orbiter 

reaches orbit 516 seconds after launch. The 3-4-4 system only reaches 2.8 

g at staging (Fig. 8). The 3-4-4 boosters stage 93 seconds after launch, 

also at an altitude of 86,000 ft, and the orbiter reaches orbit in 528 

seconds. For both configurations, throttling of the core vehicle engines 

only becomes necessary for approximately the last 75 seconds of flight. 

Weights of the core and booster vehicles taken from the weights and 

sizing program are shown in Table 2. Moments of inertia and center-ofgravity locations are shown Jn Table 3 for various stages of the mission. 

The moments of inertia are outputs from the program and are provided for 

possible future use in dynamic analysis of separation and flight of the 

boosters and the orbiter. The Z-axis center-of-gravity locations for the 

boosters are positive values (above vehicle centerline) due to the high 

location of the oblique wing and pivot mechanism. The X-axis center-ofgravity locations in the core vehicle range between 69.3 and 71.6 percent 

of body reference length during entry and are within trimnable limits for 

the original CBV design (Ref. 4). 

When compared with the systems studied In reference 2, all of which 

employed high-density propellants in the boosters, the size increase 

required for boosters using an all-LOX/LHz propulsion subsystem was 

substantial. However, system performance and weights compared quite 

favorably. This indicates that the use of a single propellant would not 

imply serious performance penalties when used as in this study. Certainly, 

a system using the same engines and propellants for both the orbiter and 

boosters would simplify operations and would result in reduced development, 

Inventory, and manufacturing costs. 

CONCLUSIONS 

The results of this study suggest that: 

(1) A fully reusable, two-stage manned launch system utilizing allLOX/LH2 propellants, SSMEs, and near-term technology would weigh 

approximately 2.9 Mlb at lift off for delivery of 70 klb of 

payload to low-Earth orbit. 

(2) The a1 1-LOXILH2 System yielded the highest payload-to-gross- weight ratio when compared with other launch systems that used 

lower performing solids or RP in the boosters for the same type 

of logistics mission. 
6 

(3) The payload-to-dry-weight ratio of the a1 l-LOXlLH2 3-3-4 system ranked second to that of the launch system utilizing LOXILH2 in 

the core vehicle and RP in the boosters. 

REFERENCES 

1. Stanley, Douglas 0.; and Talay, Theodore A.: 

Fuel Propulsion systems for AMLS Applications. A paper presented at 

the AIAAIASMEISAEIASEE 25th Joint Propulsion Conference, Monterey, 

California, July 10-14, 1989. 

Dual-Fuel versus Single2. MacConochie, Ian 0.; Naftel, J. Chris; and Widman, Frederick W.: 

Performance of a Circular Body Earth to Orbit Transport with Various 

Strap On Boosters. NASA Technical Memorandum 100603, April, 1988. 

. 

3. Powell, R. W. and Freeman, D. C.: Application of a Tip Fin 

Controller to the Shuttle Orbiter for Improved Yaw Control. AIAA 

Paper No. 81-0074, January 1981. 

4. Lepsch, R. A. and MacConochie, I. 0.: Subsonic Aerodynamic 

Characteristics of a Circular Body Earth to Orbit Transport. Paper No. 

AIAA 86-1801. 

Conference, San Diego, California, June 9-11, 1986. 

and Static Aeroelastic Characteristics of an Oblique Wing Research 

Airplane. NASA Technical Paper 2224, October 1984. 

Presented at the AIAA 4th Applied Aerodynamics 

5. Curry, Robert E. and Sim, Alex G.: In Flight Total Forces, Moments, 

6. MacConochie, Ian O., et al: Shuttle 11: Subsystem Weights Program 

Development. SAWE Paper No. 1764. Presented at the 46th Annual 

Conference of the Society of Allied Weight Engineeers, Seattle, 

Washington, May 18-20, 1987. 

7. Brauer, G. L., et al: Capabilities and Applications of the Program 

to Optimize Simulated Trajectories (POST) .- NASA CR-2770, February 

1977. 

8. Oblique Wing Transonic Transport Configuration Development. NASA 

CR-151928, January 1977. 
Table 1. Vehicle Weights Comparison 

P/L-to-dry-wt P/L-to-gross-wt 

.CBV (ref. 2) 

5 LOWLH2 SSMEs 

with Shuttle SRBs 

66.0 

CBV (ref. 2) 

5 LONLH2 SSMEs 

with graphite case 

Shuttle SRBs 

83.0 0.1 749 

0.201 2 

3 0.01 73 

CBV (ref. 2) 1 0.0205 

5 LOWLH2 SSMEs 

with 6 LOWRP engine 

crossfeed boosters 

84.0 

2-3-3 CBV System, 

all-LOWLH2 

3-4-4 CBV System, 

all-LOWLH2 

37.1 

70.4 

0.1361 

0.1856 

~~ 

0.01 91 ** 

2 0.0243 

Vehicle Dry wt, 

kl b 

Gross wi 

kl b 

Payload 

kl b I Ratio Rank* Ratio Rank* 

Shuttle 47.0 61 5.0 4500 5 

649.5 5000 4 

474.5 4800 3 

41 7.5 41 00 2 

272.5 1941 ** 

379.3 2902 

- 

1 

* For rankings, '1' represents the best vehicle for the particular characteristic, while '5' implies 

** Not ranked because of the large difference in payload capability compared with other vehicles 

in the matrix; hence, the secondary effects Of vehicle site makes the data not comparable. 

the worst. 

-7- 
~ 

Table 2. Two-Stage Manned Launch Vehicle Weights Summary 

Component 

1 .o 

2.0 

3.0 

4.0 

5.0 

6.0 

7.0 

8.0 

9.0 

10.0 

12.0 

13.0 

14.0 

15.0 

16.0 

Wing Group 

Tail Group 

Body Group 

Crew Module 

Fore body 

Mid Fuselage 

Aft Fuselage 

Thrust Structure 

Body Flap 

Fuel Tank (LH2) 

Oxidizer Tank (LOX) 

Thermal Protection System 

Landing Gear 

Main Propulsion 

RCS Propulsion 

OMS Propulsion 

Prime Power 

Elec. Conv. and Distr. 

Surface Controls 

Avionics 

Environmental Control 

Personnel Provisions 

Margin 

Inert (Dry) Weight 

Weight, Ib 

2-3-3 Con f ig uta tio n 

Core 

8,852 

106 

44,044 

559 

450 

4,662 

4,767 

1,799 

21 1 

22,542 

9,054 

18,058 

4,536 

19,785 

2,409 

3,475 

847 

2,639 

3,179 

2,678 

1,169 

900 

11,267 

123,944 

2 Boosters 

26,658 

295 

56,343 

0 

4,712 

11,096 

9,235 

4,676 

31 5 

18,772 

7,537 

0 

6,209 

59,434 

0 

0 

1,269 

1,836 

3,172 

4,048 

508 

0 

15,177 

174,949 

3-4-4 Configuration 

Core 

12,197 

146 

63,662 

559 

600 

5,543 

6,569 

2,732 

29 1 

33,783 

13,585 

26,017 

6,904 

30,052 

3,668 

5,279 

1,037 

2,829 

4,095 

3,149 

1,169 

900 

14,110 

175,214 

2 Boosters 

37,086 

41 0 

81,702 

0 

8,870 

13,390 

13,324 

7,000 

438 

27,586 

11,094 

0 

9,058 

76,992 

0 

0 

1,576 

2,142 

4,078 

4,656 

508 

0 

21,822 

240,030 

. 

-8- 
Table 2. Continued 

536 

166,433 

0 

166,433 

Weight, Ib 

709 

175,658 

0 

175,658 

~ 

2-3-3 Configuration 

22,460 

0 

3-4-4 Configuration 

XqLz 

0 

0 

Component 

0 

285,052 

61 1 

5,23 1 

1,305,457 

0 

241,095 

798 

5,330 

1,057,927 

17.0 Personnel 

18.0 Payload Accommodations 

19.0 Payload Returned 

20.0 Residual Fluids 

653 

4,200 

37,100 

653 

2,200 

70,400 

81 5 1,065 

Landed Weight 249,282 241,095 

21 .O OMS and RCS Reserves 0 0 ~ ~ 

Insertion Weight 249,282 241,095 

22.0 RCS Propellant (Entry) 0 

~~ 

1,085 1,652 0 

Descent Weight 

I 

167,518 I 175,658 250,934 I 241,095 

23.0 OMS and RCS Propellant 

24.0 Payload Discharged 

~ 

34,118 0 

Injected Weight 189,978 1 175,658 

25.0 Ascent Reserves & Residual 

26.0 lnflight Losses 

27.0 Ascent Propellant 

402 

3,839 

858,176 

532 

3,553 

709,029 

Gross Lift-off Weight 1,052,395 080,772 1,596,351 I 1,305,150 

TOTAL SYSTEM GLOW 1,941,167 2,901,501 

-9- 
~ 

Table 3. Moments of Inertia and Centers of Gravity 

(a) Core and boosters 

Configuration 

2-3-3 Configuration 

Core I 

IXX 

IYY 

zz 

Booster Ixx 

I 

lYY 

ZW 

zz 

y"Q 

cg 

3-4-4 Configuration 

Core I 

IXX 

IYY 

zz 

Booster Ixx 

I 

IYY 

ZZ 

Moments of Inertia*, slug-ft2, and c.g.**, percent 

Gross 

1,100,000 

26,000,000 

30,000,000 

73.7 

0.0 

-0.1 

190,000 

6,800,000 

7,800,000 

45.0 

0.0 

0.3 

2,000,000 

51,000,000 

60,000,000 

73.3 

0.0 

-0.1 

360,000 

13,000,000 

15,000,000 

43.9 

0.0 

0.3 

Landed 

990,000 

10,000,000 

12,000,000 

69.8 

0.0 

-1 .o 

530,000 

3,900,000 

4,000,000 

72.6 

0.0 

1.8 

1,800,000 

19,000,000 

22,000,000 

69.3 

0.0 

-0.9 

970,000 

7,200,000 

7,500,000 

72.9 

0.0 

1.7 

Inert 

800,000 

5,600,000 

6,800,000 

71.6 

0.0 

-1.5 

170,000 

2,300,000 

2,100,000 

72.8 

0.0 

1.9 

1,500,000 

11,000,000 

13,000,000 

71.5 

0.0 

-1.4 

320,000 . 

4,200,000 

3,900,000 

73.1 

0.0 

1.8 

* The propellants are treated as a viscous fluid when calculating moments of inertia for the 

gross weight condition. Moments of inertia for the boosters for the landed condition are 

I with wings deployed. 

I ** Centers of gravity are given as percentages of body length taken from the nose of thevehicle 

to the base heat shield. Y and Z are zero on the nose centerline. Z is positive upward. 

-10- 
Table 3. Concluded 

(b) System 

con f i g u ia t i o n 

Core with 2 Boosters 

2-3-3 System 

-cg 

X 

-cg 

Z 

3-4-4 System 

-cg 

X 

-cg 

Z 

c.g., percent 

Liftoff 

65.91 

1.63 

65.36 

1.42 

Staging 

74.38 

0.66 

73.99 

0.53 

-11 - 
cn 

c 

cn 

m 

a: 

c/) 

CY 

J 

n 

a 

0 

rrw 

vj 

-12- 
r Crew compartment 

147.7 ft ,pd 

Figure 2. Two-engine orbiter. 

-13- 
r Crew compartment 

I r LOX 

Figure 3. Three-engine orbiter. 

-14- 
,.I' 

I' 

I i' 

I 'I 

'4 

I 

I 

I I 

I 

I I I I 

I 1: 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

It 

I I1 

I -; 

I . :: 

'--d 

LOX 7 f LH2 

0 0 I 

I,- 116.3 ft -4 

Figure 4. Three-engine booster. 

-15- 
I 

I 

I 

;0 

I 

I 

I 

II 

O+O I I ; 

I I 

I 1- 

-u I ILI I I 

I I 

I I 

I I 

I I 

I I 

I I 

I 

I II 

I :: 

'I 

I '4 

\ 4 -90' 

LOX 7 f LH2 

Figure 5. Four-engine booster. 

-16- 
. 

43 

40 

35 

30 

25 

20 

15 

10 

5 

0 

3-engine core vehicle 

2-engine core vehicle 

- 

20 40 

/ 

- 

60 80 

Payload, klb 

- 

100 120 

Figure 6. Plot of main propulsion subsystem weight versus payload. 

-17- 
0 

m 

m 

0 

0 

m 

0 

m 

T3- 

0 

0 * 

0 

m 

m 

0 

0 

m 

0 

m 

N 

0 

0 

cv 

0 

m 

7 

0 

0 

F 

0 

m 

0 

5 .- 

3 

5 

8 

-18- 
0 

Lo 

Lo 

0 

0 

Lo 

0 

Lo 

w 

0 

0 

w 

0 

Lo 

c3 

80 *% 

zi= 

6 

oE 

0 

0 

cv 

0 

0 

F 

3 

3 

r 

3 

n 

- 19- 
Report Documentation Page 

I. Report No. 

NASA TM-101513 

2. Government Accesston No. 

4. Title and Subtitle 

17. Key Words (Suggested by Authorls)) 

Two-Stage Reusable Launch Sys tern Utilizing a Winged 

Core Vehicle and Glideback Boosters 

18. Distribution Statement 

7. AuthorW 

19. Security Classif. [of this report) 20. Security Classif. {of this page) 

Unclassified Unclassified 

Ian 0. MacConochie 

James A. Martin 

21. No. of pages 22. Price 

20 A03 

James S. Wood 

Miles 0. Duquette 

9. Performing Organization Name and Address 

NASA Langley Research Center 

Hampton, VA 23665-5225 

2. Sponsoring Agency Name and Address 

National Aeronautics and Space Adminstration 

Washington, DC 2 054 6-0001 

5. Supplementary Notes 

3. Recipient’s Catalog No. 

5. Rewn Date 

July 1989 

6. Performing Organization Code 

8. Performing Organization Report No. 

10. Work Unit No. 

5 0 6-4 9- 1 1-0 1 

11. Contract or Grant No. 

TType of Report and Period Covered 

Technical Memorandum 

14. Sponsoring Agency Code 

James S. Wood and Miles 0. Duquette are graduate students in the joint NASAGeorge Washington University Graduate Program. 

6. Abstract 

A near-term technology launch system is described in which Space Shuttle 

main engines are used on a manned orbiter and also on twin strap-on unmanned 

boosters. The orbiter has a circular body and clipped delta wings. The twin 

strap-on boosters have a circular body and deployable oblique wings for a 

glideback recovery. The dry and gross weights of the system, capable of 

delivering 70klb of cargo to orbit, are compared with a similar system with 

hydrocarbon-fueled boosters and with the current Shuttle. 

Launch Vehicle 

Two- S t ag e 

Unclassified - Unlimited 

Subject Category 16