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An efficient system for transportation to and from earth orbit
Acta Astronautica Vol. 9, No. II, pp. 671-676. 1982 Printed in Great Britain.
0094-57651821110671-06503.00]0 Pergamon Press Ltd.
AN EFFICIENT SYSTEM FOR TRANSPORTATION TO AND FROM EARTH ORBITt EDWARD LANTZ 19121 Rhodes Way, Gaithersburg,MD 20879, U.S.A. (Received 17 November 1981; revised version received 4 May 1982)
Abstract--The Space Shuttle with its winged Orbiter has paved the way for a further advance toward efficient transportation to space. Since the orbiter and its wings are designedfor the dynamic pressures and heat loads of reentry, they could also be used for lift duringthe accelerationto orbit. If the atmosphereis not only used for lift but also for ramjet engines, an efficientsystem can be designed if the craft is initiallyaccelerated to 0.4 kin/see by a high-energylauncher; so it can be initiallypropelled solely by ramjet engines. With this system over 20% of the gross lift off mass can be placed into orbit and returned to Earth. This relatively large percentage provides a good potential for efficientlytransportinglarge payloads to and from low Earth orbit.
1. INTRODUCTION The Space Shuttle with its winged Orbiter has paved the way for a further advance toward efficient transportation to space. Since the Orbiter and its wings are designed for the dynamic pressures and heat loads of reentry, they could also be used for lift during the acceleration to orbit. By using lift nearly all of the acceleration of a future Space Shuttle could be in the direction needed for Earth orbit, i.e. a horizontal direction. However, when the present Space Shuttle takes off, its rocket engines lift and accelerate about 1,150,000 kg of oxygen. This oxygen has to be stored in the Shuttle prior to every trip to orbit. But about 20% of the Earth's atmosphere is oxygen, which ramjet and supersonic combustion ramjet (scramjet) engines use instead of stored oxygen. This gives them a much higher specific impulse than rocket engines. Ramjet technology has been developed over a long period of time[l]. Scramjet technology has now been developed to the point where it is possible to make some initial predictions of the performance of large scramjet engines [2]. Also, as shown by Sutton in 194913], rocket engines are inefficient at low speeds. When a rocket is setting on a launch pad, at zero speed, all of the generated energy goes into the exhaust gases. At that instant of time no energy is going into the acceleration of the vehicle. Thus the propulsive efficiency of a rocket at zero speed is zero. One method of eliminating this low-speed inefficiency is to use a pneumatic launcher[4]. Since very little onboard fuel is expended during a launch, the specific impulse of a launcher is very much higher than that for a rocket. To take advantage of both the very high specific impulse of a launcher and the oxygen in the air, a future Space Shuttle could be launched at sufficient speed to allow it to be initially propelled solely by ramjet engines. tPaper presented at the XXXIInd Congress of the International Astronautical Federation, Rome, Italy, 7-12 September 1981.
Martin in 1977 found[5] that hydrogen is the best fuel for this application. However, hydrogen is a low density fuel. Large tanks will be needed to contain it. Consideration of the aerodynamic heating on these tanks and the decrease in the specific impulse of subsonic burning ramjets above about 2 km/sec leads to the conclusion that a scramjet propelled orbiter should have two ramjet booster planes, one attached to each wing, which would accelerate it to 2 km/sec and then fly back to their base, while scramjets propel the remaining craft towards orbit. 2. SPACE TRANSPORTATIONSYSTEM
Thus this space transportation system would consist of a high-energy launcher; two ramjet propelled booster planes, which would return and land at the launch site; and a craft, a scramjet orbiter, which would be capable of going into low Earth orbit and returning to and landing on Earth. The major portion of the acceleration would be done by scramjets, but rocket engines would be used for insertion into orbit. 2.1 High-energy launcher A differential pressure of 0.69 MN/m 2 acting on two pistons, 9.0 mdia., will provide a gross driving force of about 90 MN. This low pressure and large diameter will allow dynamic lubrication systems, which will prevent metal to metal contact. These pistons would be inside of 3 km long pressure tubes, which would be on an inclined plane, as described by Lantz in 1980[4]. There are at least two methods for transmitting the pressure force, which is on the pistons inside of the tubes, to the craft, which is outside. This force acting through this distance can provide a launch speed of Mach 1.2 (0.4 km/sec) at an altitude of 300 m. To offset the disadvantages of going through the speed of sound on the launcher, there are at least two advantages. These are: (1) The higher thrust needed in the transonic regime can be supplied at a very high specific impulse. (2) The transonic moving of the center of pressure will 671
672
E. LANrz
take place while the craft is on the launcher. Thus, after the launch the center of pressure will already be aft in its supersonic position; so the craft will not have to be retrimmed for supersonic flight. The 3 km length of this launcher gives a maximum acceleration of 3 g's. This could be reduced by increasing the length of the launcher. At 0.4 km/sec at an altitude of 300 m the nominal dynamic pressure will be 96 kN/m 2. This will be sufficient for reliable ramjet operation. 2.2 Ramjet boosterplanes The amount of thrust needed from the ramjet boosters depends on the overall takeoff mass of the entire craft and vice versa. Thus it is necessary to use an iterative process to determine the characteristics of the booster planes and the scramjet orbiter (SJO). Previous iterations showed that the gross takeoff mass from the launcher will need to be 985,000 kg. From the lift and drag coefficients given by Martin in 197716] it is found that the thrust required to lift a hypersonic aircraft of this mass at a speed of 0.4 km/sec at a 20° angle of attack is 2,123,660newtons. For an acceleration of 0.14g an additional 1,351,420 newtons is required. Thus the total thrust of the two boosters is assumed to be 3,475,000 newtons. An estimate of the mass of the ramjet engines which will produce this thrust can be obtained from data reported by Avery in 195511]. This shows that at 0.4 km/sec a thrust to weight ratio of 10 can be expected. With the present high temperature material and regenerative cooling technologies this ratio is now probably much larger. However, the engines for these boosters will have to operate from 0.4 to 2.0 km/sec; so a variable geometry will be required. Thus it is assumed that a variablegeometry, hydrogen-fueled ramjet, which is capable of operating from 0.4 to 2.0 km/sec with a thrust to weight ratio of 10 can be developed. This gives a ramjet mass of 17,700 kg for each of the two boosters. In order to maintain constant ramjet thrust while climbing, the speed has to increase to compensate for the decreased density of air; i.e. the dynamic pressure has to be kept constant. For a launch speed of 0.4 km/sec at an altitude of 300 m the nominal dynamic pressure will be 96 kN/m 2. From an aerodynamic heating standpoint this will be higher than desired for sustained climbing. It should, however, provide ensured operation immediately after launch to get the craft up to an altitude of about 5 kin. To provide for this extra thrust, the design point for the ramjet engines is taken as that at 0.4 km/sec at 5 km altitude. At this point the dynamic pressure will be 57.5 kN/m 2 for which a required ramjet intake area of 28.3 m2 was calculated from Lapede's ramjet performance equation [7]. The potential specific impulse of hydrogen burning ramjets as a function of flight speed was reported by Lane in 196218], the Space Planners Guide in 196519], Goethert in 1969[10], and Kramer in 1980111]. These are shown plotted in Fig. 1. As shown by the dotted line an average of these was assumed. This gives a peak value of about 3900 sec at Mach 3.5 and an average of 3300 sec over the the range from 0.4 to 2.0 km/sec. At an average
FLight speed, 04
km/sec
I 0 i
5000
I ! I
Assumed /
4000
Goethert 969 [ I0] ' i
~ 3000E "2_ ~ 2000 I
i I I P
I000-
I I
I
'
I
2
I 5
r 4
I
I
5
~;
Fright MACH no
Fig. 1.
acceleration of 0.14 g, a 1.6 km/sec velocity increase will require 19.4 rain. From this it is found that 62.741 kg of hydrogen will be needed for each booster. At the staging point the craft will be about 1400 km down range. Thus hyrogen will be needed to fly the booster planes back to their base. Assuming most of this return is done at the peak specific impulse, this will be about 4000 kg. Allowing for evaporation loss, the total mass of hydrogen per booster is taken as 73,444 kg. This will require a volume of 1050 m 3. It appears the best place to store this hydrogen would be in two cylindrical tanks in the fuselage with one on top of the other. This orientation would increase the rigidity in the vertical direction, and it would also minimize the underbody surface area, which will undergo a large portion of the aerodynamic heating. This one-overthe-other orientation would also increase the fineness ratio of the fuselage, which is desirable for supersonic aircraft. If both of these tanks have an average diameter of 4.3 m, a length of 36.2 m will be required. While the actual diameters will have to vary along their lengths to conform to the area rule of supersonic aerodynamics, the average diameter can be 4.3 m. Hemispherical ends on the tanks could be blended into the rest of the fuselage in an overall length of about 50 m. By assuming the same length to wing span ratio as used by Martin in 1977112] but a smaller maximum wing loading of 1000 kg/m2, the sizes given in Table 1 were derived. Glatt in 1974113] reported a computer program, called WAATS, which NASA had developed to estimate the weights of advanced transportation systems. The equations used in the WAATS program are given in the Appendix of Martin's 1977 paper[14]. The ones that are needed for this concept are reproduced in Appendix A, where a list of symbols is also provided. These equations were used to calculate the masses given in Table 2. As shown in the WATTS report[15] the equation for the thermal protection system would provide sufficient
An efficientsystem for transportation to and from Earth orbit
673
Table 1. Booster component sizes and factors Length of fuselage,1 I 50 m Average height of fuselage,h - I0 m ~Body wetted a r e a ~ s b Average width of fuselage - 5 m iI S b " 1285 m Wing span = 50/1.8 " 27.8 m Wing length I 22.8 m! wing } length I 11:4 m 9 Exposed wing a r e a , s e " 132,000/1000 - ~jz mAverage chord - 132/22.8 - 5.8 m! maximum o h o r d ~ l l . 6 m Theoretical total wing planform area,sth- 132 ÷ 5xli.6-190 m 2 Maximum wing root thickness/structural span of win~,r- 0.15 Ultimate load faetor,n - max. aoc. x^safeZy margin/g=3.6 Maximum dynamic pressure,q = 96 kN/m z Wetted area of booster, sw = 1635 m 2
Table 2. Booster masses 17700 ltg
Ramjet e n g i n e s
Wings 3000 Tail 50 Basic body structure 9300 Thrust structure 2660 Hydrogen tanks 8140 Thermal protection system 12770 Landing gear 1590 Prime power 177~ Electrical conversion & distribution 1550 Landing mass,m I Hydrogen m a s s Gross takeoff m a s s
5853~* 73444
131--~
* Also assumed to be m e for booster.
microquartz insulation to protect an aluminum structure against a 1316°C surface temperature for 1 hr. Since the total flight time of the ramjet booster planes will be less than 40 min, this should be adequate. From this it is seen that each 50 m long booster plane will have a takeoff mass of 132,000 kg. With an average acceleration of 0.14g and a rate of climb from 5 km up which maintains the dynamic pressure at 57.5 kN/m 2, the altitude will be 27 km when the speed gets up to 2 km/sec. At this point, which will be about 1400km down range, the booster planes will be released. After flying back to the base the calculated landing mass is 58,500 kg. 2.3 Scramjet and rocket engine operation Since the gross mass of the craft at launch was 985,000 kg and the gross mass of each of the two boosters was found to be 132,000 kg, the mass of the scramjet orbiter, or SJO, at booster staging will be 721,000 kg. Previous iterations with the WAATS method[13] showed that the largest single component mass is that for the scramjet engines. Thus it is important to minimise the required thrust and hence the drag on the SJO. Martin's data[6] for the drag and lift coefficients of a representative hypersonic craft show that at speeds from 2 to 10 km/sec the minimum drag tO lift ratio of 0.23 occurs at a 5° angle of attack. If the scramjet engines are designed to direct the thrust in the direction of the desired climb angle when the angle of attack is 50, minimum drag should be obtained. The thrust required to lift the SJO in this condition is 1,625,000 N. Since the major portion of the acceleration to orbit will be done by scramjets, it is
desirable to not have this take too long; so the acceleration for this phase was assumed to be increased to 0.2 g. If the small effect of the small climb angle is neglected, the thrust for the 0.2 g acceleration will be 1,413,000N. Thus the total thrust required from the scramjet engines will be 3,038,000 N at a dynamic pressure of 57.5 kN/m 2 at an altitude of 27 km. Jones in 1978 [2] gave a method for estimating the mass of the scramjet engines. The data in this reference show that at a dynamic pressure of 57.5 kN/m 2 it should be possible to get a thrust to weight ratio of 4.6. This results in a mass of 67,400 kg. The required intake area can also be obtained from this data. This is found to be 53 m2. An inlet span of 35 m would make the height about 1.5 m, but it is assumed that this would be split in half so there would be a double row of 0.76 m high scramjet engines supported by a lower structure. The length of the scramjet engine would be 7.2 m and the nozzle length (i.e. wing surface aft of the integrated scram jet) would need to be at least 6.0 m long. Since any horizontal velocity in the Earth's atmosphere is a tangential velocity, it generates a centrifugal force which tends to counteract gravity. A consequence of this is that less and less thrust is needed to support the craft as the horizontal velocity increases. Also, as hydrogen is burned the remaining mass decreases; so the thrust needed for 0.2g acceleration also decreases. By using these reductions with the specific impulses reported by Goethert in 1%9110] and shown in Fig. 2, the hydrogen requirements shown in Table 3 were calculated. Adding these up gives a total of 350,085 kg of hydrogen for the acceleration to 7.8 km/sec. At this point the mass of the SJO is 370,915 kg. With the thrust known as a function of velocity the corresponding altitude can be calculated as shown in Table 4, where the altitude at a velocity of 7.8 km/sec is seen to be 56 km. At this point rocket engines would be started. A velocity increment of 1.41 km/sec is needed to increase the orbit height to 161 km. With a rocket engine specific inpulse of 455 sec, which the Space Shuttle engines have, 14,600 kg of hydrogen and 86,000kg of oxygen will be needed. With some margin for evaporation losses, the mass of the SJO in a 161 km orbit will be about 255,000 kg. The total amount of hydrogen needed in the SJO will be 385,000 kg, and 86,000 kg of oxygen will be needed. The corresponding volumes are 54%m ~ of hydrogen and 72 m 3 of oxygen. If two fuselage tanks, like those for the
674
E. LANTZ Table 3. Scramjet hydrogen requirements Initial Required velocltv thrust* (km/sec) (N)
HydroKen Time Hydrogen flow rate (k~J'see) (see) (kg)
SJO mass
Earth weight
(kg)
.__
(N)
Trajectory
weiR?
721,ooo
2.0
3,o38,0oo
111
663
73,593
647,4o7
6,344,589
5.266,009
3.3
2,48o, loo
11o
357
39,321
6o8,o86
5,959,24o
4,46b.431
4.0
2,219,~17
119
510
60,863
547,223
5,362,789
3.217,673
5.0
1,812,623
132
510
67,448
479,775
4,701,798
2.021,773
6.0
1,405,367
120
510
61,200
418,575
4,102,035
902,448
7.0
1,027,970
117
408
47,660
370,915
3,634,962
127,224
7,8
756,253 35o,o85
*Required thrust
=
0.20 Earth weight
+
0,23 Trajectory weigh~
Table 4. Scramjet trajectory Initial velocity (km/sec)
Final velocity (km/sec)
2.0
3.3
Required thrust (N)
Required dynamic I pressure ° (kN/m ~ )
3,038,000
Air Altitude density2 (kg/m3) (km)
57.5
.0288
27
3.3
4.0
2,480,100
46.9
.0086
34
4.0
5,0
2,219,817
42,0
.0053
37
5.0
6.0
1,812,623
34,3
.oo27
42.5
6.0
7.0
1,4o5,367
26.6
.0015
47
7.0
7.8
1,027,970
19.5
.0008
52.5
756,253
I~.3
.ooo5
56
7.8 1.
Required dynamic pressure
2.
Air density
=
-
Required thrust 3,038,000 • 57.5
2 x Dynamic P~r • I03 ( v • 103) 2
boosters but with a larger diameter of 7 m are used, they will be able to store 38.5 m3 of hydrogen per meter of tank. As noted by Love in 1977116] a McDonnell Douglas design for an Advanced Supersonic Transport has a fuselage length of about 90 m; so this length should also be possible for the SJO. If 22 m of the top part of the fuselage is reserved for the payload, it appears that about 122 m of tanks will fit into the fuselage as shown in Fig. 3. This will provide a volume of 4697 m 3. This will leave a volume of 861 m3 to be stored in the wings. For this craft a length to wing span ratio of 2.0 is assumed. The mass of the rocket engines has to be included. As shown in Table 4, if 0.2 g acceleration is to be maintained after the scramjet engines cut off, the rocket engines will have to provide an initial thrust of 756,253 N. If they have a thrust to weight ratio of 81, as the Space Shuttle
=
,002 • Dyn. Pr. v2
main engines have, their mass will be 950"kg. Since this is about one-third of the thrust developed by a single Space Shuttle main engine, these engines will be considerably smaller. Consequently they may fit into the space between the hemispherical ends at the aft end of the fuselage.
2.4 Scramjet orbiter ( SJO) summary In Table 5 the calculated component and system masses in the SJO are listed. It should be noted that these include only those necessary for going to a space station in orbit and returning to Earth. For this type of operation the mass of the expendables can be assumed to be in the 12,581kg margin. Also, since the total aerodynamic heating time for the SJO will be a little over an hour,
675
An efficientsystem for transportation to and from Earth orbit
5000-
2.0 ==
F L i g h t speed, k m / s e c 3.0 4.0 5.0 6.0 I I I 1
Table 5. SJO masses 7.0 I
8.0 I
i ,, ~ 4ooo-.,,, ,., "
/G°ethert
X
_= 3 o o o
I
~---------~-q-2 8 0 0 sec ~
._E
L ~ -
2 3 0 0 sec , 9 0 0 se: L--- 1400 sec
I000-
5
t
I
I0
I
15 20 FLight MACH rio.
\
'
Wings Tail Basic body structure thrust structure Oxygen tank Hydrogen tanks Thermal protection system Landing gear Auxiliary propulsion Prime power Electrical conversion & distribution Margin
22,25~
303 27,121
@,650
993 30,3~3 38,761 6,9@1 3,506 1,72@
5.@29
12.581
SJO mass to orbit, me . . . .
255,000
Payload to 161 km orblt-
-32.000
I
25
Shuttle there will also be actual data for the thermal protection system. This thermal protection system equation was checked by using it to calculate the mass of all of the thermal protection insulation on the present Space Shuttle Orbiter. This equation gives an answer of 7541 kg where the quoted number is 7028kg[17]. Thus this equation appears to be satisfactory for this estimation.
Fig. 2.
/
6?,@00 kg 950
SJO landing mass, m i ~ 2 2 3 , 0 0 0
-r " -.~-- 9 0 0 sec
I
ScramJet engines Rocket engines
7
'
Fig. 3.
some of the margin[15] may have to be used for additional insulation. Assuming a 32,000 kg payload is left in orbit, the mass landed on Earth would be 223,000 kg. 3. DISCUSSION These calculations show that with a gross takeoff mass of 985,000 kg, 225,000 kg could be placed in Earth orbit. This is well over 20% of the gross takeoff mass. The comparable percentage for the present Space Shuttle is about five. A percentage of about sixteen was reported by Kramer in 1980111|. The development of higherstrength, lighter-weight, and higher-temperature materials and coatings will reduce the mass of the craft which is in this percentage. This means there is considerable potential in this system for increasing the payload without increasing the gross takeoff mass and size. In regard to the validity of this estimation method it should be noted, that except for the thermal protection system, there is full-scale, actual data for generating these equations. With successive flight of the Space
4. CONCLUSIONS (1) With this system the maximum acceleration will occur during the brief launch period, and this can be reduced by making the launcher longer. The use of wings allows the acceleration during the rest of the flight to be between 0.14 to 0.20g. This will provide a comfortable trip to orbit. (2) All components of this system will be fully reusable; so it can be operated like an airline shuttle. (3) The rocket engines will be used only for vacuum operation; so they can be made highly efficient for orbit insertion at a speed of 7.8 km/sec. (4) Over 20% of the gross lift off mass can be placed into orbit and returned to Earth. This is a much higher percentage that can be obtained with any presently proposed system. The development of higher-strength, lighter-weight, and higher-temperature materials and coatings will reduce the mass of the craft, which is in this percentage. This will allow this system to efficiently transport large payloads to and from low Earth orbit. REFERENCES
1. Avery W. H. Twenty-five years of ramjet development. Jet Propulsion 25, 604-614 (1955). 2. Jones R. A. and Huber P. W. Toward scramjet aircraft. Astronautics & Aeronautics 38-48 (1978). 3. Sutton G. P. Rocket Propulsion Elements, pp. 15-18. Wiley, New York (1949). 4. Lantz E. High-energylauncher for commercialtransportation to space. J. Spacecrayt & Rockets 17, 163-164 (1980). 5. Martin J. A. An Evaluation o[ Composite Propulsion /or Single-Stage-to-Orbit Vehicles Designed [or Horizontal Take-Off. NASA Technical Memorandum NASA TM X3554; Washington,D.C.; p. 9 (1977). 6. lbid, p. 16. 7. Lapedes D. N. (Editor), Encyclopedia o[ Energy, p. 601. McGraw Hill, New York (1977).
676
E. LANTZ
8. Lane R. I. Recoverable air-breathing boosters for space vehicles. J. Royal Aeronautical Soc. 66, 371-384 (1962). 9. U.S. Air Force Systems Command; Space Planners Guide, pp. V-25-26. Andrews Air Force Base, Maryland (1965). 10. Goethert B. H. and Shanrokhi F. (Editors) Proc. of Reusable Launch & Reentry Vehicles for Space Flight, Vol. 1, Aerodynamics; p. 18. The University of Tennessee Space Institute, Tullahoma, Tennessee (1969). 11. Kramer P. A. and Biihler R. D. Hybrid rocket/air breathing propulsion for ballistic space transportation. J. Spacecraft & Rockets 17, 334--341 (1980). 12. Martin J. A. An Evaluation of Composite Propulsion for
Single-Stage-to-Orbit VehiclesDesignedfor Horizontal Take Off; NASA Technical Memorandum; NASA TM X-3554; p. 5. Washington, D.C. (1977). 13. Glatt, C. R. WAATS--A Computer Program for Weights Analysis of Advance Transportation Systems. NASA CR2420 (1974). 14. Martin J. A. An Evaluation of Composite Propulsion for
Single-Stage-to-Orbit VehiclesDesignedfor Horizontal Take Off, p. 12. NASA Technical Memorandum NASA TM X3554; Washington, D.C. (1977). 15. Glatt C. R. WAATS--A Computer Program for Weights Analysis of Advanced Transportation Systems, pp. 29, 30. NASA CR-2420 (1974). 16. Love T. U.S. Keeping the door ajar on second generation SST. The Washington Star, pp. B4, BS. Washington, D.C., 6 May (1979). 17. U.S. National Aeronautics & Space Administration. Space Shuttle. NASA SP-407; Washington, D.C., pp. 48, 49 (1976).
The equations used to determine the masses of the components and systems in the ramjet booster planes and the scramjet orbiter are as follows:
Wing Tail Body basic structure
(0.044 kg) ( ~ )
°584
(1.21 kg)(O.15Sth)tl
(0.523kg)(ln)°'Sq°~6Sb TM
0.015 T g
Oxygen tanks
(27.0 kg) (~o°)°s~3
Hydrogen tanks
(32.3 kg) (~ff) °7°~
Thermal protection
(7.81 kg/m 2) Sw
Landing gear Auxiliary propulsion
(0.00676 kg) m i 124 0.01375me
Prime power Electrical conversion and distribution
1774 kg (0.135 kg) me 0"7213~0"36[k'
These were obtained from the Appendix of NASA report TM X-3554 by Martin (1977). The symbols used are defined as follows: Symbol g h l me m~ mo n
q p, p~
Equation
Equation
Thrust structure
mn
APPENDIX A
Component or system
Component or system
r
Sb St
Sth S~ T
Definition Acceleration of Earth's gravity Average height of fuselage Length of fuselage Total mass in Earth orbit Required mass of liquid hydrogen Landing mass Required mass of liquid oxygen Maximum acceleration x safety margin g Maximum dynamic pressure Density of liquid hydrogen Density of liquid oxygen Maximum wing root thickness Structural span of wing Body wetted area Exposed wing area Theoretical total wing planform area Total wetted area Maximum thrust
Unit 9.8 m/sec 2 m m kg kg kg kg N/m: kg/m 3 kg/m ~ m2 m2 m2 m2 N
0094-57651821110671-06503.00]0 Pergamon Press Ltd.
AN EFFICIENT SYSTEM FOR TRANSPORTATION TO AND FROM EARTH ORBITt EDWARD LANTZ 19121 Rhodes Way, Gaithersburg,MD 20879, U.S.A. (Received 17 November 1981; revised version received 4 May 1982)
Abstract--The Space Shuttle with its winged Orbiter has paved the way for a further advance toward efficient transportation to space. Since the orbiter and its wings are designedfor the dynamic pressures and heat loads of reentry, they could also be used for lift duringthe accelerationto orbit. If the atmosphereis not only used for lift but also for ramjet engines, an efficientsystem can be designed if the craft is initiallyaccelerated to 0.4 kin/see by a high-energylauncher; so it can be initiallypropelled solely by ramjet engines. With this system over 20% of the gross lift off mass can be placed into orbit and returned to Earth. This relatively large percentage provides a good potential for efficientlytransportinglarge payloads to and from low Earth orbit.
1. INTRODUCTION The Space Shuttle with its winged Orbiter has paved the way for a further advance toward efficient transportation to space. Since the Orbiter and its wings are designed for the dynamic pressures and heat loads of reentry, they could also be used for lift during the acceleration to orbit. By using lift nearly all of the acceleration of a future Space Shuttle could be in the direction needed for Earth orbit, i.e. a horizontal direction. However, when the present Space Shuttle takes off, its rocket engines lift and accelerate about 1,150,000 kg of oxygen. This oxygen has to be stored in the Shuttle prior to every trip to orbit. But about 20% of the Earth's atmosphere is oxygen, which ramjet and supersonic combustion ramjet (scramjet) engines use instead of stored oxygen. This gives them a much higher specific impulse than rocket engines. Ramjet technology has been developed over a long period of time[l]. Scramjet technology has now been developed to the point where it is possible to make some initial predictions of the performance of large scramjet engines [2]. Also, as shown by Sutton in 194913], rocket engines are inefficient at low speeds. When a rocket is setting on a launch pad, at zero speed, all of the generated energy goes into the exhaust gases. At that instant of time no energy is going into the acceleration of the vehicle. Thus the propulsive efficiency of a rocket at zero speed is zero. One method of eliminating this low-speed inefficiency is to use a pneumatic launcher[4]. Since very little onboard fuel is expended during a launch, the specific impulse of a launcher is very much higher than that for a rocket. To take advantage of both the very high specific impulse of a launcher and the oxygen in the air, a future Space Shuttle could be launched at sufficient speed to allow it to be initially propelled solely by ramjet engines. tPaper presented at the XXXIInd Congress of the International Astronautical Federation, Rome, Italy, 7-12 September 1981.
Martin in 1977 found[5] that hydrogen is the best fuel for this application. However, hydrogen is a low density fuel. Large tanks will be needed to contain it. Consideration of the aerodynamic heating on these tanks and the decrease in the specific impulse of subsonic burning ramjets above about 2 km/sec leads to the conclusion that a scramjet propelled orbiter should have two ramjet booster planes, one attached to each wing, which would accelerate it to 2 km/sec and then fly back to their base, while scramjets propel the remaining craft towards orbit. 2. SPACE TRANSPORTATIONSYSTEM
Thus this space transportation system would consist of a high-energy launcher; two ramjet propelled booster planes, which would return and land at the launch site; and a craft, a scramjet orbiter, which would be capable of going into low Earth orbit and returning to and landing on Earth. The major portion of the acceleration would be done by scramjets, but rocket engines would be used for insertion into orbit. 2.1 High-energy launcher A differential pressure of 0.69 MN/m 2 acting on two pistons, 9.0 mdia., will provide a gross driving force of about 90 MN. This low pressure and large diameter will allow dynamic lubrication systems, which will prevent metal to metal contact. These pistons would be inside of 3 km long pressure tubes, which would be on an inclined plane, as described by Lantz in 1980[4]. There are at least two methods for transmitting the pressure force, which is on the pistons inside of the tubes, to the craft, which is outside. This force acting through this distance can provide a launch speed of Mach 1.2 (0.4 km/sec) at an altitude of 300 m. To offset the disadvantages of going through the speed of sound on the launcher, there are at least two advantages. These are: (1) The higher thrust needed in the transonic regime can be supplied at a very high specific impulse. (2) The transonic moving of the center of pressure will 671
672
E. LANrz
take place while the craft is on the launcher. Thus, after the launch the center of pressure will already be aft in its supersonic position; so the craft will not have to be retrimmed for supersonic flight. The 3 km length of this launcher gives a maximum acceleration of 3 g's. This could be reduced by increasing the length of the launcher. At 0.4 km/sec at an altitude of 300 m the nominal dynamic pressure will be 96 kN/m 2. This will be sufficient for reliable ramjet operation. 2.2 Ramjet boosterplanes The amount of thrust needed from the ramjet boosters depends on the overall takeoff mass of the entire craft and vice versa. Thus it is necessary to use an iterative process to determine the characteristics of the booster planes and the scramjet orbiter (SJO). Previous iterations showed that the gross takeoff mass from the launcher will need to be 985,000 kg. From the lift and drag coefficients given by Martin in 197716] it is found that the thrust required to lift a hypersonic aircraft of this mass at a speed of 0.4 km/sec at a 20° angle of attack is 2,123,660newtons. For an acceleration of 0.14g an additional 1,351,420 newtons is required. Thus the total thrust of the two boosters is assumed to be 3,475,000 newtons. An estimate of the mass of the ramjet engines which will produce this thrust can be obtained from data reported by Avery in 195511]. This shows that at 0.4 km/sec a thrust to weight ratio of 10 can be expected. With the present high temperature material and regenerative cooling technologies this ratio is now probably much larger. However, the engines for these boosters will have to operate from 0.4 to 2.0 km/sec; so a variable geometry will be required. Thus it is assumed that a variablegeometry, hydrogen-fueled ramjet, which is capable of operating from 0.4 to 2.0 km/sec with a thrust to weight ratio of 10 can be developed. This gives a ramjet mass of 17,700 kg for each of the two boosters. In order to maintain constant ramjet thrust while climbing, the speed has to increase to compensate for the decreased density of air; i.e. the dynamic pressure has to be kept constant. For a launch speed of 0.4 km/sec at an altitude of 300 m the nominal dynamic pressure will be 96 kN/m 2. From an aerodynamic heating standpoint this will be higher than desired for sustained climbing. It should, however, provide ensured operation immediately after launch to get the craft up to an altitude of about 5 kin. To provide for this extra thrust, the design point for the ramjet engines is taken as that at 0.4 km/sec at 5 km altitude. At this point the dynamic pressure will be 57.5 kN/m 2 for which a required ramjet intake area of 28.3 m2 was calculated from Lapede's ramjet performance equation [7]. The potential specific impulse of hydrogen burning ramjets as a function of flight speed was reported by Lane in 196218], the Space Planners Guide in 196519], Goethert in 1969[10], and Kramer in 1980111]. These are shown plotted in Fig. 1. As shown by the dotted line an average of these was assumed. This gives a peak value of about 3900 sec at Mach 3.5 and an average of 3300 sec over the the range from 0.4 to 2.0 km/sec. At an average
FLight speed, 04
km/sec
I 0 i
5000
I ! I
Assumed /
4000
Goethert 969 [ I0] ' i
~ 3000E "2_ ~ 2000 I
i I I P
I000-
I I
I
'
I
2
I 5
r 4
I
I
5
~;
Fright MACH no
Fig. 1.
acceleration of 0.14 g, a 1.6 km/sec velocity increase will require 19.4 rain. From this it is found that 62.741 kg of hydrogen will be needed for each booster. At the staging point the craft will be about 1400 km down range. Thus hyrogen will be needed to fly the booster planes back to their base. Assuming most of this return is done at the peak specific impulse, this will be about 4000 kg. Allowing for evaporation loss, the total mass of hydrogen per booster is taken as 73,444 kg. This will require a volume of 1050 m 3. It appears the best place to store this hydrogen would be in two cylindrical tanks in the fuselage with one on top of the other. This orientation would increase the rigidity in the vertical direction, and it would also minimize the underbody surface area, which will undergo a large portion of the aerodynamic heating. This one-overthe-other orientation would also increase the fineness ratio of the fuselage, which is desirable for supersonic aircraft. If both of these tanks have an average diameter of 4.3 m, a length of 36.2 m will be required. While the actual diameters will have to vary along their lengths to conform to the area rule of supersonic aerodynamics, the average diameter can be 4.3 m. Hemispherical ends on the tanks could be blended into the rest of the fuselage in an overall length of about 50 m. By assuming the same length to wing span ratio as used by Martin in 1977112] but a smaller maximum wing loading of 1000 kg/m2, the sizes given in Table 1 were derived. Glatt in 1974113] reported a computer program, called WAATS, which NASA had developed to estimate the weights of advanced transportation systems. The equations used in the WAATS program are given in the Appendix of Martin's 1977 paper[14]. The ones that are needed for this concept are reproduced in Appendix A, where a list of symbols is also provided. These equations were used to calculate the masses given in Table 2. As shown in the WATTS report[15] the equation for the thermal protection system would provide sufficient
An efficientsystem for transportation to and from Earth orbit
673
Table 1. Booster component sizes and factors Length of fuselage,1 I 50 m Average height of fuselage,h - I0 m ~Body wetted a r e a ~ s b Average width of fuselage - 5 m iI S b " 1285 m Wing span = 50/1.8 " 27.8 m Wing length I 22.8 m! wing } length I 11:4 m 9 Exposed wing a r e a , s e " 132,000/1000 - ~jz mAverage chord - 132/22.8 - 5.8 m! maximum o h o r d ~ l l . 6 m Theoretical total wing planform area,sth- 132 ÷ 5xli.6-190 m 2 Maximum wing root thickness/structural span of win~,r- 0.15 Ultimate load faetor,n - max. aoc. x^safeZy margin/g=3.6 Maximum dynamic pressure,q = 96 kN/m z Wetted area of booster, sw = 1635 m 2
Table 2. Booster masses 17700 ltg
Ramjet e n g i n e s
Wings 3000 Tail 50 Basic body structure 9300 Thrust structure 2660 Hydrogen tanks 8140 Thermal protection system 12770 Landing gear 1590 Prime power 177~ Electrical conversion & distribution 1550 Landing mass,m I Hydrogen m a s s Gross takeoff m a s s
5853~* 73444
131--~
* Also assumed to be m e for booster.
microquartz insulation to protect an aluminum structure against a 1316°C surface temperature for 1 hr. Since the total flight time of the ramjet booster planes will be less than 40 min, this should be adequate. From this it is seen that each 50 m long booster plane will have a takeoff mass of 132,000 kg. With an average acceleration of 0.14g and a rate of climb from 5 km up which maintains the dynamic pressure at 57.5 kN/m 2, the altitude will be 27 km when the speed gets up to 2 km/sec. At this point, which will be about 1400km down range, the booster planes will be released. After flying back to the base the calculated landing mass is 58,500 kg. 2.3 Scramjet and rocket engine operation Since the gross mass of the craft at launch was 985,000 kg and the gross mass of each of the two boosters was found to be 132,000 kg, the mass of the scramjet orbiter, or SJO, at booster staging will be 721,000 kg. Previous iterations with the WAATS method[13] showed that the largest single component mass is that for the scramjet engines. Thus it is important to minimise the required thrust and hence the drag on the SJO. Martin's data[6] for the drag and lift coefficients of a representative hypersonic craft show that at speeds from 2 to 10 km/sec the minimum drag tO lift ratio of 0.23 occurs at a 5° angle of attack. If the scramjet engines are designed to direct the thrust in the direction of the desired climb angle when the angle of attack is 50, minimum drag should be obtained. The thrust required to lift the SJO in this condition is 1,625,000 N. Since the major portion of the acceleration to orbit will be done by scramjets, it is
desirable to not have this take too long; so the acceleration for this phase was assumed to be increased to 0.2 g. If the small effect of the small climb angle is neglected, the thrust for the 0.2 g acceleration will be 1,413,000N. Thus the total thrust required from the scramjet engines will be 3,038,000 N at a dynamic pressure of 57.5 kN/m 2 at an altitude of 27 km. Jones in 1978 [2] gave a method for estimating the mass of the scramjet engines. The data in this reference show that at a dynamic pressure of 57.5 kN/m 2 it should be possible to get a thrust to weight ratio of 4.6. This results in a mass of 67,400 kg. The required intake area can also be obtained from this data. This is found to be 53 m2. An inlet span of 35 m would make the height about 1.5 m, but it is assumed that this would be split in half so there would be a double row of 0.76 m high scramjet engines supported by a lower structure. The length of the scramjet engine would be 7.2 m and the nozzle length (i.e. wing surface aft of the integrated scram jet) would need to be at least 6.0 m long. Since any horizontal velocity in the Earth's atmosphere is a tangential velocity, it generates a centrifugal force which tends to counteract gravity. A consequence of this is that less and less thrust is needed to support the craft as the horizontal velocity increases. Also, as hydrogen is burned the remaining mass decreases; so the thrust needed for 0.2g acceleration also decreases. By using these reductions with the specific impulses reported by Goethert in 1%9110] and shown in Fig. 2, the hydrogen requirements shown in Table 3 were calculated. Adding these up gives a total of 350,085 kg of hydrogen for the acceleration to 7.8 km/sec. At this point the mass of the SJO is 370,915 kg. With the thrust known as a function of velocity the corresponding altitude can be calculated as shown in Table 4, where the altitude at a velocity of 7.8 km/sec is seen to be 56 km. At this point rocket engines would be started. A velocity increment of 1.41 km/sec is needed to increase the orbit height to 161 km. With a rocket engine specific inpulse of 455 sec, which the Space Shuttle engines have, 14,600 kg of hydrogen and 86,000kg of oxygen will be needed. With some margin for evaporation losses, the mass of the SJO in a 161 km orbit will be about 255,000 kg. The total amount of hydrogen needed in the SJO will be 385,000 kg, and 86,000 kg of oxygen will be needed. The corresponding volumes are 54%m ~ of hydrogen and 72 m 3 of oxygen. If two fuselage tanks, like those for the
674
E. LANTZ Table 3. Scramjet hydrogen requirements Initial Required velocltv thrust* (km/sec) (N)
HydroKen Time Hydrogen flow rate (k~J'see) (see) (kg)
SJO mass
Earth weight
(kg)
.__
(N)
Trajectory
weiR?
721,ooo
2.0
3,o38,0oo
111
663
73,593
647,4o7
6,344,589
5.266,009
3.3
2,48o, loo
11o
357
39,321
6o8,o86
5,959,24o
4,46b.431
4.0
2,219,~17
119
510
60,863
547,223
5,362,789
3.217,673
5.0
1,812,623
132
510
67,448
479,775
4,701,798
2.021,773
6.0
1,405,367
120
510
61,200
418,575
4,102,035
902,448
7.0
1,027,970
117
408
47,660
370,915
3,634,962
127,224
7,8
756,253 35o,o85
*Required thrust
=
0.20 Earth weight
+
0,23 Trajectory weigh~
Table 4. Scramjet trajectory Initial velocity (km/sec)
Final velocity (km/sec)
2.0
3.3
Required thrust (N)
Required dynamic I pressure ° (kN/m ~ )
3,038,000
Air Altitude density2 (kg/m3) (km)
57.5
.0288
27
3.3
4.0
2,480,100
46.9
.0086
34
4.0
5,0
2,219,817
42,0
.0053
37
5.0
6.0
1,812,623
34,3
.oo27
42.5
6.0
7.0
1,4o5,367
26.6
.0015
47
7.0
7.8
1,027,970
19.5
.0008
52.5
756,253
I~.3
.ooo5
56
7.8 1.
Required dynamic pressure
2.
Air density
=
-
Required thrust 3,038,000 • 57.5
2 x Dynamic P~r • I03 ( v • 103) 2
boosters but with a larger diameter of 7 m are used, they will be able to store 38.5 m3 of hydrogen per meter of tank. As noted by Love in 1977116] a McDonnell Douglas design for an Advanced Supersonic Transport has a fuselage length of about 90 m; so this length should also be possible for the SJO. If 22 m of the top part of the fuselage is reserved for the payload, it appears that about 122 m of tanks will fit into the fuselage as shown in Fig. 3. This will provide a volume of 4697 m 3. This will leave a volume of 861 m3 to be stored in the wings. For this craft a length to wing span ratio of 2.0 is assumed. The mass of the rocket engines has to be included. As shown in Table 4, if 0.2 g acceleration is to be maintained after the scramjet engines cut off, the rocket engines will have to provide an initial thrust of 756,253 N. If they have a thrust to weight ratio of 81, as the Space Shuttle
=
,002 • Dyn. Pr. v2
main engines have, their mass will be 950"kg. Since this is about one-third of the thrust developed by a single Space Shuttle main engine, these engines will be considerably smaller. Consequently they may fit into the space between the hemispherical ends at the aft end of the fuselage.
2.4 Scramjet orbiter ( SJO) summary In Table 5 the calculated component and system masses in the SJO are listed. It should be noted that these include only those necessary for going to a space station in orbit and returning to Earth. For this type of operation the mass of the expendables can be assumed to be in the 12,581kg margin. Also, since the total aerodynamic heating time for the SJO will be a little over an hour,
675
An efficientsystem for transportation to and from Earth orbit
5000-
2.0 ==
F L i g h t speed, k m / s e c 3.0 4.0 5.0 6.0 I I I 1
Table 5. SJO masses 7.0 I
8.0 I
i ,, ~ 4ooo-.,,, ,., "
/G°ethert
X
_= 3 o o o
I
~---------~-q-2 8 0 0 sec ~
._E
L ~ -
2 3 0 0 sec , 9 0 0 se: L--- 1400 sec
I000-
5
t
I
I0
I
15 20 FLight MACH rio.
\
'
Wings Tail Basic body structure thrust structure Oxygen tank Hydrogen tanks Thermal protection system Landing gear Auxiliary propulsion Prime power Electrical conversion & distribution Margin
22,25~
303 27,121
@,650
993 30,3~3 38,761 6,9@1 3,506 1,72@
5.@29
12.581
SJO mass to orbit, me . . . .
255,000
Payload to 161 km orblt-
-32.000
I
25
Shuttle there will also be actual data for the thermal protection system. This thermal protection system equation was checked by using it to calculate the mass of all of the thermal protection insulation on the present Space Shuttle Orbiter. This equation gives an answer of 7541 kg where the quoted number is 7028kg[17]. Thus this equation appears to be satisfactory for this estimation.
Fig. 2.
/
6?,@00 kg 950
SJO landing mass, m i ~ 2 2 3 , 0 0 0
-r " -.~-- 9 0 0 sec
I
ScramJet engines Rocket engines
7
'
Fig. 3.
some of the margin[15] may have to be used for additional insulation. Assuming a 32,000 kg payload is left in orbit, the mass landed on Earth would be 223,000 kg. 3. DISCUSSION These calculations show that with a gross takeoff mass of 985,000 kg, 225,000 kg could be placed in Earth orbit. This is well over 20% of the gross takeoff mass. The comparable percentage for the present Space Shuttle is about five. A percentage of about sixteen was reported by Kramer in 1980111|. The development of higherstrength, lighter-weight, and higher-temperature materials and coatings will reduce the mass of the craft which is in this percentage. This means there is considerable potential in this system for increasing the payload without increasing the gross takeoff mass and size. In regard to the validity of this estimation method it should be noted, that except for the thermal protection system, there is full-scale, actual data for generating these equations. With successive flight of the Space
4. CONCLUSIONS (1) With this system the maximum acceleration will occur during the brief launch period, and this can be reduced by making the launcher longer. The use of wings allows the acceleration during the rest of the flight to be between 0.14 to 0.20g. This will provide a comfortable trip to orbit. (2) All components of this system will be fully reusable; so it can be operated like an airline shuttle. (3) The rocket engines will be used only for vacuum operation; so they can be made highly efficient for orbit insertion at a speed of 7.8 km/sec. (4) Over 20% of the gross lift off mass can be placed into orbit and returned to Earth. This is a much higher percentage that can be obtained with any presently proposed system. The development of higher-strength, lighter-weight, and higher-temperature materials and coatings will reduce the mass of the craft, which is in this percentage. This will allow this system to efficiently transport large payloads to and from low Earth orbit. REFERENCES
1. Avery W. H. Twenty-five years of ramjet development. Jet Propulsion 25, 604-614 (1955). 2. Jones R. A. and Huber P. W. Toward scramjet aircraft. Astronautics & Aeronautics 38-48 (1978). 3. Sutton G. P. Rocket Propulsion Elements, pp. 15-18. Wiley, New York (1949). 4. Lantz E. High-energylauncher for commercialtransportation to space. J. Spacecrayt & Rockets 17, 163-164 (1980). 5. Martin J. A. An Evaluation o[ Composite Propulsion /or Single-Stage-to-Orbit Vehicles Designed [or Horizontal Take-Off. NASA Technical Memorandum NASA TM X3554; Washington,D.C.; p. 9 (1977). 6. lbid, p. 16. 7. Lapedes D. N. (Editor), Encyclopedia o[ Energy, p. 601. McGraw Hill, New York (1977).
676
E. LANTZ
8. Lane R. I. Recoverable air-breathing boosters for space vehicles. J. Royal Aeronautical Soc. 66, 371-384 (1962). 9. U.S. Air Force Systems Command; Space Planners Guide, pp. V-25-26. Andrews Air Force Base, Maryland (1965). 10. Goethert B. H. and Shanrokhi F. (Editors) Proc. of Reusable Launch & Reentry Vehicles for Space Flight, Vol. 1, Aerodynamics; p. 18. The University of Tennessee Space Institute, Tullahoma, Tennessee (1969). 11. Kramer P. A. and Biihler R. D. Hybrid rocket/air breathing propulsion for ballistic space transportation. J. Spacecraft & Rockets 17, 334--341 (1980). 12. Martin J. A. An Evaluation of Composite Propulsion for
Single-Stage-to-Orbit VehiclesDesignedfor Horizontal Take Off; NASA Technical Memorandum; NASA TM X-3554; p. 5. Washington, D.C. (1977). 13. Glatt, C. R. WAATS--A Computer Program for Weights Analysis of Advance Transportation Systems. NASA CR2420 (1974). 14. Martin J. A. An Evaluation of Composite Propulsion for
Single-Stage-to-Orbit VehiclesDesignedfor Horizontal Take Off, p. 12. NASA Technical Memorandum NASA TM X3554; Washington, D.C. (1977). 15. Glatt C. R. WAATS--A Computer Program for Weights Analysis of Advanced Transportation Systems, pp. 29, 30. NASA CR-2420 (1974). 16. Love T. U.S. Keeping the door ajar on second generation SST. The Washington Star, pp. B4, BS. Washington, D.C., 6 May (1979). 17. U.S. National Aeronautics & Space Administration. Space Shuttle. NASA SP-407; Washington, D.C., pp. 48, 49 (1976).
The equations used to determine the masses of the components and systems in the ramjet booster planes and the scramjet orbiter are as follows:
Wing Tail Body basic structure
(0.044 kg) ( ~ )
°584
(1.21 kg)(O.15Sth)tl
(0.523kg)(ln)°'Sq°~6Sb TM
0.015 T g
Oxygen tanks
(27.0 kg) (~o°)°s~3
Hydrogen tanks
(32.3 kg) (~ff) °7°~
Thermal protection
(7.81 kg/m 2) Sw
Landing gear Auxiliary propulsion
(0.00676 kg) m i 124 0.01375me
Prime power Electrical conversion and distribution
1774 kg (0.135 kg) me 0"7213~0"36[k'
These were obtained from the Appendix of NASA report TM X-3554 by Martin (1977). The symbols used are defined as follows: Symbol g h l me m~ mo n
q p, p~
Equation
Equation
Thrust structure
mn
APPENDIX A
Component or system
Component or system
r
Sb St
Sth S~ T
Definition Acceleration of Earth's gravity Average height of fuselage Length of fuselage Total mass in Earth orbit Required mass of liquid hydrogen Landing mass Required mass of liquid oxygen Maximum acceleration x safety margin g Maximum dynamic pressure Density of liquid hydrogen Density of liquid oxygen Maximum wing root thickness Structural span of wing Body wetted area Exposed wing area Theoretical total wing planform area Total wetted area Maximum thrust
Unit 9.8 m/sec 2 m m kg kg kg kg N/m: kg/m 3 kg/m ~ m2 m2 m2 m2 N