Space freighters for the 21st century

Acta AstronauticaVol. 29, No. 2, pp. 121-137, 1993 Printed in Great Britain. All rights reserved

0094-5765/93 $6.00+ 0.00 Copyright © 1993 Pergamon Press Ltd

SPACE FREIGHTERS FOR THE 21ST CENTURY~" H. H. KOELLE Technical University of Berlin, Aerospace Institute, Marchstr. 12, D-1000 Berlin 10, Germany

(Received 26 March 1991; receivedfor publication 15 October 1992) Almract--The useful life cycle of space launchers derived from the military long range missiles, developed during the fifties in East and West, will come to an end with the turn of the century. They have served the space faring nations of this planet well in the past and are a good basis to start from to develop space freighters for the 21st century. The market potential expected in a changing geopolitical environment and mission requirements give a fairly good frame of reference in which future launch vehicle developments can be discussed. The near future will have to live with existing space transportation systems such as Space Shuttle, Proton, Titan 4 and Energia when larger payloads look for transportation. The unmanned Shuttle C concept is marginal in performance and offers minor improvements in economy only. Thus it is not the space launcher we are looking for, at best a stop gap solution. In the present geopolitical situation it might become acceptable to use the Energia for large civilian payloads by all space faring nations. This would be an economical solution for medium-sized space freighters and serve the market for the next 20 years. The concept of an "Advanced Launch Vehicle" as favoured by the U.S. Air Force for their conceived requirements is not a popular project anymore, since even the latest concepts in connection with the space defense initiative (SDI) are no longer based on such a space launcher. Considering the fact that a Moon project and more so a Mars expedition appears likely for the first half of the 21st century it is easy to see that present launch vehicles are not suitable for such a program. The answer would be a heavy "Space Freighter" in the Post-Saturn class as already analysed in the sixties for those missions. Since these early studies the rule "What you can do on Earth you should do there"! is now more apparent than ever. This will reduce the number and complexities of expensive operations in space! Or in other words: The larger the launch vehicle, the cheaper and faster the space program! A representative space freighter which can transport 300 metric tons to low Earth orbit or about 100 metric tons to the geostationary and lunar orbits is presented with respect to performance, size, operations, schedule and cost to indicate that the technology is not the limiting factor. Most of it is available now. The problem is the market and the determination to provide for the future in time--with the decline in military expenditures during the next decade it appears probable that a major new launch vehicle development is feasible from the financial viewpoint and desirable from the political viewpoint if seen in the international context.

1. INTRODUCTION During the first half o f the 20th century it became clear that fairly large rockets would be needed for the transportation of people and cargo into space. Military rockets paved the way to get the first unmanned satellites and the first people into an orbit about the Earth in the middle o f this century. It was in 1952 that the first engineering studies were made by Wernher yon Braun and his associates to prove that conventional chemical rockets would suffice to send people to Mars. His orbital carrier vehicle had a launch mass o f 6400 M T and a payload capability o f 25 M T with a partially recoverable concept (growth ratio 6400/25 = 256). In the same year Kratft Ehricke proposed a launch vehicle with a launch mass of 784 M T and a payload of 5 M T ( M = 157) and a ferry system to support a manned space station in the 2 h orbit. H. H. Koelle designed a carrier vehicle in connection with his master thesis at the University Stuttgart with a launch mass of

tPaper IAF-90-195 presented at the 41st Congress of the International Astronautical Federation, Dresden, Germany, 8-12 October 1990.

871 M T and a payload of 3.5 M T in a winged return vehicle ( M = 250). These three proposals had two things in common: conventional propellants and partial recovery. After joining the von Braum team at Huntsville, the author initiated the design of a 1,000,000 lb launcher in April 1957, convinced that the time would come when such a big bird would be required. This led to a contract with the Advanced Research Projects Agency ( A R P A ) in July 1958 for a booster called J U N O V at that time. This became the Saturn I launcher when taken over by N A S A in 1959. At about the same time, a new big engine with a thrust of 1,000,000 Ib was contracted with the Rocketdyne C o m p a n y inspiring Milton Rosen and Carl Sehwenk to their proposal for a large M o o n rocket, called N O V A , which was presented at the lOth International Astronautical Congress in L o n d o n in September 1959. This launch vehicle had a take-off mass o f 3015 M T and the capability to carry a lunar lander of 16.2 M T to the Moon. This was certainly one of the inputs leading to the Apollo lunar landing program initiated by President Kennedy, in M a y 1961. This in turn led to an uprating o f the Saturn launch vehicle program and vehicle size, culminating 121

122

H.H. KOELLE

in the selection of the Saturn V configuration with a launch mass of 2900 MT and an escape capability of 45 MT towards the Moon. This vehicle had a low Earth orbit (LEO) capability of 120MT, with a growth factor M = 24. Due to lack of mission assignments and funds, the production of this largest launcher to date was discontinued in 1970. The last launch of a Saturn V was in 1973 to orbit the space-laboratory Skylab. In the early sixties, however, the future of spaceflight looked bright. Thus, studies were initiated by the Future Projects Office of the NASA Marshall Space Flight Center on the next generation of launch vehicles as soon as the planning concerning the Saturn V and the Apollo program was completed. There development activities began following two lines of thinking: (1) A reusable winged rocket plane for the transport of 12 people or 3 metric tons of cargo for the logistic support of a space station went through the analysis and preliminary design cycle leading eventually to the Space Shuttle concept in 1972 (not reported here), and (2) a heavy lift cargo vehicle to triple the Saturn V payload capability, was studied in detail. The primary contributors to these contracted studies in those years were the Martin Company, General Dynamics-Space Division and the Douglas Company. The name NOVA became unpopular with Congress and these investigations were renamed "Post Saturn" launch vehicle studies and lasted through 1965. The results of these preliminary design studies were widely published by Bono and Kalitinsky (1963) and summarized in two reports (Huber, 1964; Sanders, 1965). A typical Class I design used a cluster of 18 F-1 engines in the first stage and 3 M-1 engines in the second stage; it had a launch mass of 10,350 metric tons and a LEO payload of 350 metric tons, resulting in a growth factor of 30. Class II vehicles had a reusable first stage, Class III vehicles had new high pressure engines and were fully recoverable, while Class IV vehicles had even a nuclear upper stage including gas-core and nuclear pulse engines. This series of studies was complemented by an analysis of solid boosters and their potential for heavy lift vehicles by Boeing in 1968. Thus, all options were on the table in 1968, but the increasing commitment in Vietnam requiring all available resources practically killed all dreams to use the mid-80 window for a manned Mars mission, and thus all planning for heavy lift launch vehicles in the U.S.A. ceased. During the late seventies, the interest in big launch vehicles was revived by feasibility studies of space solar power system (SSPS) as proposed by Peter Glaser already in 1968. The looming energy crisis caused the responsible Agencies of the U.S. Government to sponsor several studies on the feasibility of the SSPS in 1978-1979 spending about 25 million

dollars. The Boeing company (among others) made detailed design studies on suitable launchers reported by Nansen (1978) and Henley (1978). By that time the Space Shuttle Main Engine was under development as well as a suitable propulsion system for heavy space freighters, permitting even the design of single stage-to-orbit (SSTO) launch vehicles. Payloads were in the 500 metric ton class and the market volume connected with the SSPS project clearly favoured fully reusable vehicles. This renewed interest in heavy space freighters prompted the restudy of such vehicles, this time based on the SSME. A Ph.D. thesis was used as a vehicle for a system analysis in depth (Ress, 1979) leading to a modular, fully reusable two-stage concept with a take-off mass of 10,000 MT

(A)

.

.

~

Sta|e 2

$ta$o 3

Fig. 1. (A) Caption opposite.

S p a c e freighters

and a 100 MT GEO payload capability. This study was followed by several others culminating in three reports (Arend, 1987; Altmann and Kerinnis, 1989) which resulted in a fairly detailed assessment of the potential of the selected concept of a heavy space freighter with an optimized launch mass of 6000 MT. It is this concept (see Fig. 1) which is presented in this report in summary fashion and can serve as a reference for future analysis and program development.

123 2. VEHICLE DESIGN APPROACH

2.1. System philosophy Space transportation systems must be designed observing the criteria --reliability --availability ----economy --growth potential.

A u U p p e r structure

(B)

B - B o t t o m structure C = Plaps D =* E n g i n e c o v e r s E ,= P r o p e l l a n t tanks F :

T h r u s t structure

G = P r i m a r y structure H = Landing gear

I = Heat shield K = Ensines

Stase 3

Stage 2

-

A

C

~k------- O

m p

m K

D

Stage I

D

I

K

E

Fig. I. (B) Vehicle cross-sections. AA 29/2--E

B

124

H.H. KOELLE

One of the greatest contributions towards fulfillment of those requirements is simplicity of design and operations. Economy is achieved with payload and mission flexibility leading to a long operational life and reasonable production rates. The criteria of reliability and availability are enhanced by a design based on available subsystems and a modular concept. Growth potential can be provided by choosing a geometry with a high volume payload compartment and permitting the incorporation of technological improvements as they become available. Unfortunately, these desirable attributes of a launch vehicle cannot be optimized as individual parameters, but within the constraints of the system and its environment. Thus, trade-offs cannot be avoided. It is well known that the expected market size and vehicle life time determine more than anything else the vehicle size and design approach. Reusability is of interest only if the launch rate is greater than one per month and the life cycle longer than two decades. The space freighter concept described in this report is based on the assumption that this condition will be met in the first half of the next century. Another basic assumption underlying this vehicle concept is the fact that the technology of space propulsion has achieved a high degree of maturity. High-pressure, turbopump hydrogen/oxygen engines of the type used by the Space Shuttle have proven to be reliable and to offer high performance. They are also desirable from the viewpoint of environmental conditions. Competitors to this class of engines are not in sight for many decades to come; therefore, no reason exists to wait for better propulsion systems. However, there is room for improvement with respect to engine life, maintainability and production cost. These will certainly be emphasized as soon as the demand for these engines increase. With respect to availability of such a vehicle system several influential parameters will have to be observed. The situation the decision maker finds himself in is a classical one: the market will develop only if the vehicle is ready to be used, but the initiation of the development of this vehicle can be justified only if the market is clearly visible. Past experience has shown that in many instances the launch vehicles developed were too small for the payloads then looking for transportation. Thus, the conclusion can only be that the next generation space freighter should have plenty of margin with respect to payload mass and volume if a long life cycle is expected. It is also important to use available technology and hardware, and a great deal of commonality in a multi-vehicle fleet. All of these aspects are part of the design philosophy used.

2.2. Design concept The basic assumption for this vehicle concept is that the primary markets are the geostationary and lunar orbits. These will require fairly high launch rates if and when a lunar base and space solar power

plants are developed, constructed and operated. This excludes practically single-stage-to-orbit vehicle concepts. They are also not desirable from the environmental point of view, since they tend to be larger, produce more noise and require more propellants. A fully reusable, three stage vehicle concept was selected with stage separation of the third stage in LEO (about 150 km) to be able to return the second stage to the launch site after one revolution. This requires also a moderate amount of lift and controlability of the landing area which can be accomplished by the flaps of a ballistic vehicle. A ballistic reentry requires a low ballistic factor to keep the heat loads down, thus, a fairly large diameter was chosen for the vehicle in spite of the somewhat higher drag. However, this is more than compensated for by a higher specific impulse and the advantage for a high volume payload compartment. A sectionalized, modular concept was chosen to keep the production costs low, using flat surfaces on the outer structure of a twelve-sided pyramid, a geometry as shown in Fig. 1. Within this geometry, hydrogen and oxygen tanks of a similar cylindrical shape and diameter, but different length and insulation are arranged in such a way, that the stage volume is minimized. Only derivatives of the SSME engine are used in all stages. They will have shorter nozzles in the first stage and longer nozzles in the second and third stage. The third stage features a toroidel hydrogen tank to keep the stage length down in addition to a set of spherical oxygen tanks. The heat shields of all stages employ individual concepts and materials due to different heat loads, emphasizing low operational cost. The concept selected permits a very sizeable payload compartment allowing a variety of payload shapes, sizes, and densities, which is very important because liquid hydrogen will be one of the more frequent cargoes. With this general approach all desirable criteria are observed and an optimum size vehicle with respect the market envisaged can be designed. The technical details of such a design are presented in the following section. 3. VEHICLE DESCRIPTION

The size of the vehicle was selected on the basis of the market scenario envisioned for the first half of the 21st century. Also it should not duplicate existing or past payload capabilities. The goal of the Post-Saturn studies was a vehicle with at least three times the payload capability of the Saturn V to be operated at specific transportation cost of less than one-third of the Saturn V. Thus a payload target of 100 metric tons for GEO and lunar orbit missions was selected, which would allow payloads of about 350 MT to LEO. Payloads of this size are of interest in connection with manned missions to other planets. A study on the influence of vehicle size on operating cost was

Space freighters Table I. Basic dimensions of the space freighter Neptune 2000+

Width (m) Height (m) Cross-section (m 2) Top-structure (m 2) (with shroud) Bottom-structure (m 2) Volume (m 3)

Stage I

Stage 2

Stage 3

41.054 38.062 1.355 1.985

29,000 21.207 676 800

3.317 40.327

1.467 11.977

22.376 33.400 403 400 (I.165) 427 4.742

made (Arend, 1987) indicating that the selected size with a take-off mass of 6000 MT was an optimum size for the market model used. With this payload and vehicle mass targets, the preliminary design led to a vehicle with a diameter of 41 m and a total height of 72 m with the standard payload compartment. The stage dimensions are given in Table 1. While the dimensions of the vehicle will not be changed during its life cycle, the materials of the subsystems will be upgraded as the advancing technology will permit within the product improvement program, if economically justified. The masses estimated and presented in this mass model were calculated at the subsystem level on the basis of a preliminary design. They should be representative, but certainly could use some improvement. The general mass characteristics, however, should not change very much after a detailed design has been completed. A considerable contribution to the stage mass is resulting from the type and number of engines selected. A SSME type engine, reengineered for this application, with two different expansion ratios in an original configuration, and an improved version for the later vehicles has been selected. A summary of the engine data is given in Table 2. A reusable vehicle requires adequate heat shields for the reentry and recovery of the vehicle stages. The first stage will be recovered down range on a mobile platform in case an island is not within reach. The second stage returns to a landing site near the launch site as does the third stage. The materials selected for the individual stages depend on the heat loads, the

125

geometric dimensions, the operational concept and cost factors. Another major contributor to the vehicle mass are the propellant tanks for liquid hydrogen and liquid oxygen. These are very similar to those of the Space Shuttle tanks, but are optimized with respect to the diameter in terms of systems costs. The vehicle equipment for this vehicle is classified by type and comprises guidance and control equipment, hydraulics as required for operating the flaps, telemetry and measuring equipment and the electrical systems including power supply, distribution and cable harnesses. These have not been considered in detail but their masses have been estimated with the help of estimating relationships in analogy to other launch vehicles. Since these masses are not very large if compared to those of the other subsystems, some margin of error can be tolerated. A summary of the mass characteristics of the Neptune 2000+ is presented in Table 3. The advanced version of this vehicle uses the most costeffective materials in the individual subsystems. The prototype vehicle is the heavier one, since it is conservatively designed and based on proven aluminium alloys. The advanced vehicles use more costly, but lighter materials in many, but not all subsystems. However, the higher costs are more than compensated for by the higher payload resulting in lower specific transportation cost, but not in smaller launch cost per flight! The absolute mass values do not give the complete picture of the mass characteristics of a launch vehicle. Mass ratios are much more illustrative and more useful if comparing launch vehicles. Therefore, Table 4 has been prepared, presenting the most interesting mass ratios: ----(performance) effective mass ratio --propellant fraction (of the individual stage mass) --propellant ratio (which includes the payload) --payload ratio (payload: take-off mass) --growth ratio (take-off mass:net payload) --net growth ratio (dry mass:net payload).

Table 2. Propulsion system data Parameters Cross-section ascent Cross-section descent Reference nose radius Sea level thrust p. eng. Vacuum thrust p. eng. Number of engines Nozzle area ratio Specific sea level impulse Specific vacuum impulse Engine mass flow Single engine mass (first generation) Total propulsion system (first generation)

Singie engine mass (advanced engine)

Dimensions

Stage I

Stage 2

Stage 3

m2 m kN kN --s s kg/s

1345 1558 29 1845 2082 40 20 400 451 470

676 878 18 1802 2167 9 120 388 469 470

402 800 14 1802 1045 I 120 388 469 470

ks

2500

3500

2500

MT

100

31.5

2.5

k8

2000

2800

2000

MT

80

25.2

2

m2

Total propulsion system

(advanced engine)

126

H.H. KOELLE

Table 3. Summary mass model in metric tons of the prototype space freighter for transportation to LEO (1), GEO (2) and LUO (3) Prototype (metric tons) Stage 1 Stage 2 Stage3 Take-off mass 1 6000.00 1592.05 344.22 2 6000.00 1592.05 344.22 3 5990.00 1582.05 334.22 Net payload 1 1592.05 344.22 285.02 2 1592.05 344.22 93.99 3 1582.05 334.22 71.55 Equipment 8.90 3.50 1.76 Structure 457.62 119.93 24.41 Engines 100.00 31.50 2.50 Recovery equipment 1 43.37 21.44 14.89 2 43.37 21.44 16.96 3 43.37 21.44 18.66 Residuals 37.50 10.45 2.11 Propellant reserves 55.68 15.68 2.11 Propellant consumption 1 3705.42 1045.33 13.41 2 3705.42 1045.33 200.48 3 3705.42 1045.33 211.13 Cut-off mass 1 2294.58 546.72 330.82 2 2294.58 546.72 143.74 3 2294.58 546.72 123.10 Stage net mass 1 702.52 202.50 45.79 2 702.52 202.50 49.75 3 702.52 202.50 51.55

The mass ratios for both vehicle versions (prototype and advanced) are integrated in Table 4 to make the comparison easier. These ratios appear quite plausible if compared with vehicles of similar size and concept. But if one considers that the propellant mass fraction of the expendable 2nd stage of the Saturn V vehicle, also using L O X / L H 2 was in terms of hardware close to 0.92 then the above figures look quite realistic. The result of these calculations indicate that growth ratios of about 20 can be achieved for a ballistic 3-stage space freighter with regard to L E O missions, increasing to about 60 for G E O missions. These are good and representative target figures for future mission planning! A vehicle of this type and size in a developing market will experience complex operational conditions leading to extensive use of resources. Therefore, a detailed model for the operational phase is required to get a good hold on operational cost determining the system's economy to a large extent. The next section describes the flight operational concept and performance of the vehicle under consideration.

Ratios

4. VEHICLE PERFORMANCE

4.1. Flight operations concept A vehicle of this type and size will probably use more than one launch site. A typical scenario would be an off-shore launch site near Cape Canaveral for launches due East and for logistic supply o f a lunar base, and in addition a near equatorial launch site for support of space solar power plants in the geostationary orbit, if this develops into a major market. The empty vehicle will be prepared for launch on a mobile catamaran-like launch platform in a protected area of a harbour a few miles inland of the coastline. It will then be towed to a launch position some 20 km out to sea for reasons of safety and noise reduction in inhabited areas. There it will be fueled by a tanker, which is highly automated. The launch control center is located close to the launch preparation and assembly facilities. The standard ascent trajectory resembles the ascent trajectory of the Saturn and similar launch vehicles. If payload size and mission profile permit, the engines of the first stage are throttled to about 90% soon after lift-off. This is desirable for several reasons: (1) differential throttling for attitude control, (2) engine out capability and (3) engine lifetime. The tilting program will lead to a cut-off of the second stage in about 150 km altitude and to a return to the launch site within one revolution, with a combination of modular drag and lift control. This maneuver is supported by the thrusters of the attitude control system at minimum expense of propellants. The second stage is aerodynamically stable. Three of the nine main engines are reignited for the final landing at a dry landing site near the refurbishment facility. The first stage will come down some 700 km downrange by drag modulated ballistic descent. The four centre engines are reignited for the final landing approach on the mobile landing platform. This platform is a flat 100 x 100 m structure supported by two refurbished and adapted tankers. The platform has its own diesel engines and can be moved in certain limits during the landing approach to support the booster guidance and control system. It is also conceivable that islands located downrange of the launch site might be employed as a landing site, even if a doglegging maneuver is required by the booster. The

Table 4. Mass ratios of the prototype and advanced versions for the LEO cargo mission Stage 1 Stage 2 Stage 3 Total vehicles

Effective mass ratio = take-off/cut-off mass

Propellant fraction = used prop./stage mass

pro adv pro

adv Propellant ratio used prop./take-off mass

Payload ratio = net payload/take-off mass

pro adv pro

adv Growth factor = take-off mass/net payload Net growth factor dry mass/net payload

pro

2.614 2.680

2.912 2.828

0.863 0.889

0.861 0.894

1.040

7.916

1.036 0.272 0.360

7.852 0.857 0.886

0.6t7

0.656

0.038

0.794

0.626 0.266 0.277 3.769

0.646 0.218 0.258 4.600

0.038 0.833 0.878 1.207

0.808 0.048 0.062 21.05

adv

3.619

3.892

1.323

18.61

pro adv

0.383

0.512 0.313

0.166 0.096

2.925 1.760

0.293

Space freighters first stage will be cleaned after landing and transferred to a transport ship which brings it back to the refurbishment site within 2 days. The third stage of this space freighter will provide the velocity increment required from orbital velocity to the transfer velocity to G E O or to LUO, and after arrival the additional velocity for adaption to local orbital conditions. After unloading, the stage will be refueled as necessary and provide the velocity increment required for a direct return to the launch site. The landing will be accomplished in similar fashion as the second stage with the assistance of modular lift and drag as well as an auxiliary propulsion system during the final landing approach near the refurbishment facility. A vehicle of this size and type to be employed for a number of different missions over a long life cycle requires a great amount of ground facilities and support equipment. These have been identified for the purpose of taking their cost into consideration. The ground support system is in line with past experience. A detailed study of the ground facilities and equipment was not considered to be essential at this time since many studies of the past (see Introduction) have led to a rather thorough understanding of the problems involved. It is quite clear that the operational procedures for preparing such a vehicle for launch and turning it around are not simple ones. The Space Shuttle is a point in case. It has taught us many lessons, particularly when a flight crew is involved. The Heavy Space Freighter will be a cargo carrier, unmanned and will be employed with increasing launch rates for logistic missions. It resembles in a way the operation of a fleet of Jumbo jets in a freight market. Some of the operational procedures observed today in the operation of air-freighters can be adapted for use in the space operation of the future. Refurbishment operations, particularly of engines and hot structures, have been studied in some depth during the past decades. This know-how will certainly be used when detailing this operational process. Here is not the room and the time to get into these details other than to make cost estimates possible. Bono (1963), Eisner (1970) and Nansen (1978) have been used as a source for relevant information.

Table

Ascent Initial acceleration Max. acceleration Altitude Burning time Velocity Max. drag Max. prep. Altitude M = I Time M = I Pressure M = 1

127

4.2. Standard trajectories Standard trajectories for ascent and descent for individual stages and missions have been calculated but they leave room for further optimization and trade-offs with vehicle parameters. Thus, in general one can say that the resulting performance figures a r e on the conservative side. One particular feature of the vehicle concept proposed is the cluster of 40 hinged engines of the first stage. This allows an improvement of the thrust coefficient because the expansion ratio will be adapted with decreasing ambient air pressure. The control during ascent is effected by differential thrust of the outer engines. Since the lever arm is about 20 m this is very effective considering that the present SSME can be operated up to a 109% thrust. Some development effort is certainly required to realize such a concept, but there is no doubt that this can be done. The flaps of the individual stages could also be used to support attitude control. Other alternatives not yet considered might be even more advantageous. Table 5 summarizes the most important parameters of a standard ascent trajectory; they are satisfactory for planning purposes. The descent trajectories are typical for ballistic vehicles. They make use of a high drag coefficient and a low cross sectional load. This places the point of maximum deceleration and heat load in a fairly high altitude and thus reduces the mass requirements for heat protection. This design also uses extendable flaps to increase the cross section, also allowing drag modulation by a suitable program. Unsymmetrical flap extension allows to produce a certain amount of lift and thus a fairly good control of the landing point. If this is not sufficient, some thrust must be applied. The standard trajectory presented here assumes that the flaps are extended half of the maximum angle at the average. The use of parachutes is not envisaged in this application other than for control of emergency purposes. The final velocity remaining after deceleration by drag which is in the order of 100 m/s or less will be reduced to zero at touchdown by igniting and throttling some of the stage engines alternatively by a separate propulsion system.

5. Characteristicdata points for the ascent trajectory Dimensions Stage 1 Stage 2 m/s 12.3 12.3 m s-2 36.3 35.4 km 75.445 152.201 s 197 246 m s- t 2893 7809 kN N m-2 km s N m -2

12,652 19,102 I0 78 9,4

------

Stage 3 6.3 12.6 LUO 449 Depart. trajectory ------

128

H.H. KOELLE Table 6. Characteristic d a t a points of descent trajectories Descent Ballistic factor* Max. drag

LID Pressure on flaps Stagnation pressure Max. q Qcum Flight time Distance L E O / G E O Distance L U O Max. deceleration Brake propellants Terminal velocity

Dimensions

Stage 1

Stage 2

Stage 3

kg m -2 kN -kN m -2 k N m -2 kW m -2 kWh m - 2 s km km m s -2 MT m s- ]

265 52430 0.1 21.18 46.08 26.4 0.281 615 688 692.0 80.6 19 74

153 4739 0.3 3.81 6.77 147.0 12.4 Optional Launch site Launch site 24.8 4.5 53

42 2316 0.3 4.16 5.44 1330.0 27.5 Optional Launch site Launch site 45.1 2 27

*Flaps fully extended.

The resulting reentry trajectories as calculated produced the more important parameters and data points of these typical reentry trajectories as summarized in Table 6. Each trajectory, particularly all reentry trajectories, produce a heat load on the structure of the individual stages. This heat load has a maximum at the leading edges and reaches a peak at a particular combination of velocity and altitude. The heat loads are smaller along the longitudinal dimension of a stage. Thus, it is important for the calculation of the heat shield and its mass to know the variation of the heat rate as a function of time for the distinguished locations as well as the cumulative heat load for any particular point. This information is required to select a particular concept and material for each of the stages. This heavy space freighter would probably use a metallic heat shield in the first stage, a ceramic tile heat shield in the second stage and will probably have an exchangeable ablative heat shield in the third stage.

4.3. Mission capabilities Using the mass models and the results of the trajectory calculation, the mission capabilities can be determined. They are summarized in Table 7. Since fairly high launch rates are expected to support the lunar base logistically, it is unlikely that the optimum launch window can be used all the time which is limited to once per month. In this case the lunar orbit payload capability would be equal to the payload capability for GEO missions. But for this model a 10% reduction for the average flight was assumed to be a realistic payload for lunar missions. Furthermore, a gradual improvement of the payload capability of the Neptune is expected during its Table 7. Payload capabilities for reference missions (metric tons) Payload capabilities Prototype Stage 1 Stage 2 Stage 3 A d v a n c e d vehicle Stage I Stage 2

Stage 3

LEO

GEO

LUO

1592 344 285

1592 344 94

1582 344 71

life cycle of 50 years due to product improvement. This results in a payload growth with time as shown in Fig. 2. This will contribute to the improvement of the cost-effectiveness of this transportation system.

4.4. Operational system performance The payload capabilities given above are those of single flights with 100% reliability. An actual flight program, however, will require different annual flight numbers for the reference missions and probably with an increasing trend due to the expected market growth. Since it is quite difficult to estimate the markets over a 50 year time period, another approach was selected for this model. The production rates have been used as independent input variables at three different levels (programs A, B and C). These production rates begin with one vehicle per year in the first year and increase in certain time intervals thereafter. The highest production rate envisioned is five vehicles per annum in the 34th year of program C. The smallest program (A) keeps the production rate to 2 p.a. as shown in Fig. 3. With these production rates a total of 75, 125 and 200 vehicles will be produced during the 50 year life cycle of this vehicle. Following this line of thinking, assumptions have to be made for the loss rate of vehicles due to catastrophic failures and concerning the design lifetime of a particular vehicle. In this model it was

1658

1648

426 322

426 132

416 106

-

350

-

300

-

250

-

200

-

150

-

100

-

O-'--"

g. O

1658

400 gg O

eL

50

Geostationary o r b i t

l

L

I

L

L

I

10

20

30

40

50

60

No. of operational

year

Fig. 2. Payload growth vs time.

Space freighters 5

-

Production 4 -

rate

p.a.

129 800

of vehicles

700

IC

-

C

-

B

6O0 ,~ 500 m

3

.~ 400

300

=

A

2OO

I

100 10

20

30

40

50

Fig. 5. Annual number of launches vs time. 10

20

30

40

50

Fig. 3. Production rates for vehicle stages vs time for three selected programs.

assumed that a particular vehicle will be taken off flight status after 10 years of service. This will result in a particular number of vehicles in the inventory on fright status. Figure 4 shows the number of vehicles in the inventory. If, in addition, assumptions are made for the turnaround time, which certainly will decrease as function of time with experience and product improvement activities, the annual number of launches of the available vehicle inventory can be calculated. This growth of launch rate p.a. vs time is given in Fig. 5. In case the distribution of missions within a program is known (see Fig. 6 for this model) the individual mission performance or annual payload volume can be calculated. This approach allows to balance various operational parameters to result in plausible production rates, flight rates, inventories, turnaround-time and vehicle life times. Figure 7 is the result of the operational assumptions made with respect to the turnaround times and individual vehicle flight numbers vs time for the three selected programs. At first sight at least they look reasonable, although for some people of this generation it might stress their imagination to envision space programs requiring hundreds of flights of this space freighter towards the middle of the next century. But time will show whether this was an optimistic or pessimistic view. Anyhow, this model is open for any variation also in direction of smaller programs. But it has been stated before that a minimum of one flight per month is required to justify

a reusable space freighter of this type and size, which amounts to about 1000 metric tons p.a. to GEO and LUO. On the basis of this design concept producing a mass model and performance model for a heavy space freighter, an operational concept was developed leading to production rates, flight rates and other operational parameters required for estimating the cost for such a transportation system. The costing procedure used and the results obtained are summarized in the following section. 5. VEHICLE COST MODEL

5.1. Model philosophy Any launch vehicle concept has to pass an economy test of a fairly detailed cost model. Past experience accumulated during the last 30 years has been assembled in cost estimating relationships which can be applied to new launch vehicle concepts, however with a great deal of discretion. The Apollo project had one negative influence on space transportation cost, namely the developers were to some extent spoiled by plenty of money, due to the deadlines set. This resulted in overstaffing and a relatively low manpower efficiency. This trend must not last forever. It depends very strongly on the political and organizational environment on how costly space transportation will be in the future. Top management and its underlying motivation has a stronger influence on cost than technical details. It is the know-how at the top and the capability to control contractors 100

GEO m

50 "~ ~ 40

o

-

(%) 50

C

20

A o o 10

20

30

40

Fig. 4. Vehicle inventories vs time.

50

25 Years

Fig. 6. Distribution of missions vs time (in %).

50

130

H.H. KOELLE 200

~, 150 --

tq

160 '~ "~ 90 -- / ~ ~ 60 -- ~ ~ 3o

"o

,//"f"~ Program A ./,/" - - - ProgramB - - - - Program C

120 ~ ' ~ -

Z

80

~

40 .,-, ~ ,~,

0 25 50 Years Fig. 7. Turnaround-times and number of flights per vehicle vs time during a 10-year operational period. ~

o

0

which determines the cost of a project. These facts limit all efforts to make cost projections. Using past experience and derived cost estimating relationships thus has the elements of an art. There is more to it than applying statistics. It depends on the experience and the judgement of the estimator to interpret the available data and adapt it to the situation he expects to be realistic, perhaps only desirable. In this sense the cost estimates presented here are rather "should" cost than projected cost. It is important, however, that these costs are representative, plausible and suitable for the comparison of alternative concepts and their optimization. The life-cycle-cost model used for analysing the Neptune concept is of the structure shown in Fig. 8. 5.2. Development cost

Development costs were estimated with the help of CERs on the basis of man-years converted to 1990 $ values. One man-year represents an amount of $180,000 and includes all overhead charges. Table 8

is a summary of the results and gives the estimated cost for the prototype vehicle. If one would want to go out to develop the advanced vehicle from the outset, the development cost would be approx. 8% higher. These figures might seem rather low. It should be kept in mind that no new propulsion systems are developed, only adapted, to this application, and furthermore that this is a sectionalized vehicle where most of the testing can be done with a 1/12 section of the structure and tanks. The same effect can be realized when producing the test hardware. If first unit cost and original facility cost are added, the total grows to $20 x 10 9. 5.3. Production cost

Production costs are estimated in terms of a first unit cost for reference purposes, and a mass production factor due to learning. The cost for producing the prototype is not included in the development cost, but in this line item. The production cost is heavily impacted by the choice of the sectionalized modular concept and the fact that the structural surfaces are plane surfaces relatively easy to manufacture. The cost estimates arrived at for the three stages including effects of commonality and preproduction of already available hardware is summarized in Table 9. Also of interest is the reduction of vehicle unit cost vs time when all learning effects are taken into consideration. This trend is shown in Fig. 9. The engine unit cost vs time are contributing heavily to the vehicle cost and therefore they are presented separately permitting a check of the plausibility of these estimates. They are presented in Fig. 10.

Life Cycle Cost LCC I

~

_ _1O p e r a t i ° n pellants Launch preparation Launch and mission

Production

Spares

Fig. 8. Life-cycle-cost model--overview.

Space freighters

131

Table 8. Development¢o6t for the prototype vehicle(excludingfirst unit cost and ground facilities)--million1990$ Subsystem Stage I Stage 2 Stage 3 Total Structure TPS Tanks Equipment Engine modifications

3 i 25 271 902 1625 341

1059 408 656 942 93

386 69 427 638 107

4570 748 1984 3205 541

Sub-total

6264

3157

1626

11048

Management, integration and tests

3396

Total

14443

Table 9. First unit cost of prototype vehicle(million1990$) (1 MY = 170,000 1990$) Stage I Stage 2 Stage 3 Total Structure TPS Tanks Equipment Engines Sub-total Management, integration checkout

670.7 239.9 304.3 ! 46.7 832.2

234,4 195.7 132.2 78.9 228.6

67.8 13.9 55.4 50.4 55.0

973.0 449.5 491.9 276.0 I I 15.8

2193.8

869.9

242.5

3306.2

--

--

--

202.3

Total

3508.5

5.4. Operation cost The operation costs cover all expenditures except the development and production cost. They do include spare parts, replacements and interest. If the non-recurring and the recurring costs are prorated over the life cycle "specific transportation cost" for individual missions within a multi-mission program can be calculated. The acquisition costs for the ground facilities including the ground equipment are estimated on the basis of first unit cost for each of the facilities plus additional units to be purchased in time before the capacity of these facilities is exceeded. Thus, the acquisition cost is directly related to the number of launches from a particular launch site within the program constraints. The turnaround times resulting

1 p.a.

_

from the assumptions on vehicles and their operation made in previous chapters are the dominating input parameter for this cost element. The manpower requirements for launch preparation, launch and mission control, refurbishment are quite sizeable. The man-years estimated by this method are converted into 1990 dollars by a multiplier of $0.16 million/ man-year for operation. Spare parts are another important contributor to the operations cost. They are assumed to be a function of launch rate. The spare part rates are estimates at the subsystem level and their costs are derived from the production cost of that particular subsystem in the specific operational year. In case heat shields have to be replaced for every flight, as in the third stage for GEO and LUO missions, the full unit cost is charged to the program operations cost. The product improvement costs leading from the prototype vehicle to the advanced vehicle in due course are in the order of the original development costs, but are prorated as a burden over the entire operational life cycle.

"N33 3p.a. "~ O

24

.--, 22 es g$

20 18

1

E' 16

-

O~

~

age SSME rst stage SSME ---O

¢'~

.o

12

I I 5

I 10

I 15

I 20

I 25

Fig. 9. Vehicle unit production cost vs operational life time

for program B.

I

I

I

I

10 20 30 40 50 No. of operational year

t

60

Fig. 10. Unit production cost of 1st and 2nd stage engines vs operational life time for program B.

132

H.H. KOELLE

~ ' 120

[] Development [] Ground-facilities [] Production

6 90

.~

60

L~

Operation

30 0

L)

0 25

50

Years Fig. 11. Cost per flight for LEO missions vs time.

5.5. Cost-effectiveness The most important parameter of a space transportation system is the "specific operating cost", based on specific missions within a multi-mission program. All non-recurring and recurring costs are integrated into the systems cost and prorated over the missions flown in a particular year. This generates the cost trend which can be presented as a function of program size or transportation volume respectively. Life cycle average cost as well as annual cost data are of interest. The first curves presented in Fig. 11 give the cost per flight as function of time. They are in the order of about $115 million per flight during the first years of the flight operation gradually diminishing to about $42 million per flight. The cost figures given are valid for program B, they are slightly higher for program A and slightly lower for program C. The production and operations cost dominate the trend; the development cost and facility costs are relatively small in these fairly large programs lasting fifty years! Some irregularities observed in the above trends are the result of step functions with respect to launch rates and purchase of facilities occurring at low numbers. Figure 12 presents the overall trend of the total specific transportation cost as a function of the life cycle transport volume which is an indicator of program size. The upper limit of the ranges given for the three missions represents a vehicle close to the prototype configuration, which is considered to be conservative, the lower limit of the indicated range represents the advanced configuration to be introduced from the beginning of the program leading to optimistic trends. 6. PROJECT PROGRAMMATICS

responsible decision makers that a heavy space freighter will be needed for the development of extraterrestrial resources. This situation will then demand a fairly aggressive schedule which can be realized only by using technology at hand. In such a scenario the existing problems and risks to be taken are pretty clear. The workpackages to be accomplished are fairly easy to define; some fifty areas requiring major development activities have been identified so far. There is no space in this report to describe them in detail. Additional insight can be obtained by presenting the distribution of the individual cost shares over subsystems, which is shown in Fig. 13 for program B. The other programs have very similar trends. This general picture of cost trends is supplemented by a graphical presentation of the major contributors to these costs indicating again the domination of the systems cost by the production and operations cost for systems of this size (Fig. 14). It is considered possible that the actual development of a vehicle of this size could be initiated by about 1995 if the political scenario will gradually deemphasize weapon system development in favour of international global development projects. This might seem optimistic today, but it is a chance to be pointed out to the responsible de~sion makers. In this case a development schedule as shown in Fig. 15 can be considered as typical for a heavy space freighter of this size and concept. With these assumptions a first flight in the year 2001 does not seem inconceivable, but is admittedly optimistic from today's viewpoint. But not more

600

500 LUO

400

GEO ~].~

300

t-----J 200

100

LEO

6.1. Schedule The vehicle concept presented in this report was based on the assumption that scarce resources will not permit a radical departure from available or near term technology. Development funds are also not expected to be available before the assembly operation of Space Station Freedom gets underway. On the other hand by this time it will be apparent to the

I 500

I 1000

I 1500 LC cumulative payload (103 MT)

Fig. 12. Average life cycle specific transportation cost for LEO, GEO and LUO missions vs mission cumulative payload volume within the multimission space program of 50 years.

Space freighters

133 Spares Tanks 0.7% /~-~T

Average operation cost

IIIIII

~Structure E76g.i;%s ~ E q u~ipm~n7t%

Lau48e' g~~eS~:ystemmanagement • h~re~~A-~I ~,~/Refurbishment 1.0%

PS 19.3%

0.7%

5.3%

Propellants 5.2% ~ l n t e g r a t i o n 3 2% ~

'

~

Recovery 5.7%

Refurbishment Tanks 3.4%

TPS 9.7%

Engines~Structure 30.7% ~ 12•7%

General sup. Missions control 10.5% 1.5% Equipment 43.5% Fig. 13. Distribution of total operational costs over the individual cost elements for program B. optimistic than the first manned lunar landing in 1969 as seen from the viewpoint of the year 1959!

6.2. Financing The development of a space freighter of this size, based on existing or near term technology, was estimated to require funds in the order of 20 billion 1990 dollars including the fabrication of a prototype vehicle and a first set of ground facilities. These funds are required in a 7 year time span averaging approx. $3 billion a year, assuming that these funds will come from public funds, since all previous launch vehicles developed were in the national interest. Typical distribution of resource requirements during the development phase of a launch vehicle system over 10 years' planning and development cycle is shown in Fig. 16. This is used to calculate the rate of expenditure for a cash flow analysis. Production and operation costs are distributed over time in those years when they originate• The ~'q Operation

[ ] Development

[ ] Production

•

Product improvement

]Ground-facilities 100 /

75

\ \ _

x

~

/ /

% 50

/\

/

. /

~ /

\

/

i ,

I

\

/ \

/

_

/ ~

\

x

~

/ /

-

/

/

/

' /

\

~

I

~

same is true for additional facilities• Product improvement costs are charged as a burden on an annual basis assuming that there has to be a minimum crew of experts for trouble-shooting all the time during systems operation. If not required for troubleshooting these experts will spend their time on product improvement activities resulting in a more efficient vehicle as represented in this report by the advanced version of the Neptune. The production cost of newly manufactured vehicles is spread over the 2 years preceding delivery. In this way all expenditures can be added up as function of time for any size of a program. The result of such a calculation is shown for program B on the negative scale of Fig. 17. The income side of this cash flow scenario is determined by the prices which can be obtained in a competitive or non-competitive market and the interest to be paid for funds required to finance production and operation to the extent it cannot be covered by the income. Figure 17 is one example of such a calculation. It is no more than an illustrative scenario depending on many unknown factors at this time. If this calculation is typical, a positive cash flow can be expected after some 20 years of flight operation. This is too long a time to attract private capital. Thus, it is obvious that development decisions of this type will have to be made by the interested government(s). Figure 18 is the result of a sensitivity analysis with respect to program size for programs A, B and C indicating that the cash flow might even take 30 years before turning positive for smaller programs.

6.3. Organization and management

0

I[1111111111111111

....

i111

i ii

i l l l l l

ii

i iiii11

R

1000 2500 Cumulative transport v o l u m e [10 3 MT] Fig. 14. Distribution of major cost elements vs cumulative transportation volume.

The organization and management approach taken will be influenced strongly by the degree in which such a project is internationalized or influenced by commercial interests. In case the project of a space solar power plant system (SSPS) should become an interesting commercial proposition, it is conceivable that the project

134

H.H. KOELLE Major project milestones

ATP PRR SRR V T T

PDR V

Mockups

D~i!~'~'~ Manufactu"ring

CDR T

CSR V

I

Evaluatio~

System design Structure

Subsystem-

TPS

development

Engines Equipment

Qualification model

Engineering model 1 [~.Manfact-" c

Engineering model 2

1Testl ,

,. , , Jmte~tion

First unit

+Check-out Production and test facilities

'

[~Construction + C h e c k ~

I I

~//////~

Operational flight facilities

Calendar year

1995

I

I 1

V esi gn / / / / / / / ,

1996

1997

]

1998

1999

2000

2001

2002

Fig. 15. Typical milestones for Neptune subsystems development. organization might be set up and financed by those international companies investing in the long range supply of energy resources to the tune of $100 billion a year at this time. It is unlikely, however, that this will happen in the next decade. The more likely alternative seems to be that a determined national government recognizes that a heavy space freighter is essential for pre-eminence in space and that the present launch vehicles will not suffice for large scale space operations in the next

100

-

-

90 •

80 --

/

o /

60-51o" / Cumulative rO 40 --

/

expenditures

/

20 ~

16 --/

1

' ~ ] 7 I 8 I 9 110

l

.

Years

4~s'---~Pro)eet l[Ist flight~---'4~" Fig. 16. Financial resources required typically vs development time.

century. The U.S.A. is one of the nations to be able to afford the development of a heavy space freighter around the turn of the century in light of the present trends of a reduced burden of defense spending. But also the former Soviet Union has the capability and resources to enter such a project on or about the year 2000 when the Energia launch vehicle has completed its development cycle. Finally, Europe and Japan should have the resources and perhaps the determination to participate in such a development or even join forces if they want to develop an independent capability for political reasons as in the case of a manned space station. The most desirable scenario, however, would be an international consortium which jointly enters such a development in an organization such as INTELSAT or MARISAT. The exploration of extraterrestrial resources and their development for the benefit of humankind is a global and not a national risk. In the long run national efforts will have to be coordinated and concentrated to achieve such goals as to provide safe and clean energy for a growing population on this planet without destroying the ecology on Earth. Furthermore, the settlement of the Moon and perhaps Mars are tasks which can be solved successfully only if the industrial nations of this planet work together. The resulting organizational set-up will be a compromise between the goals, visions, interests,

Space freighters

~

135

20

50

10

25

10

25

2o

50

Devel~ r . . . . Fig. 17. Expenditures, income and cash flow vs time for program B. contributions and capabilities of those countries deciding to join the project. Many alternatives are available and it is premature to draft a detailed organizational structure. These things can only be discussed seriously as soon as one or more national governments have indicated their intent to get involved in such an enterprise. The announcement of President Bush to return to the Moon and to eventually send a manned expedition to Mars early in the next century are hopeful signs that this may be the case in the coming decade. 7. SUMMARY OF THE REPORT

Space freighters for large payloads have been studied since 1952. The NASA budget proposal of 1962 already contained a line item for the NOVA launcher to the amount of $48.5 million to initiate the

development of an 8x F-1 engine expendable space freighter for the lunar program (see Statement of General Don R. Ostrander, Director of NASA Launch Vehicle Programs, to the Committee of Aeronautical and Space Sciences U.S. Senate, 8 June, 1961). These funds were used, however, for the Saturn V launch vehicle. NOVA and Post-Saturn class vehicles were studied in the U.S.A. to 1967 and then discontinued when it became clear that the space market would not possess a priority high enough to support such a development after completion of the Apollo program. Studies of heavy space freighters were sporadically carried out thereafter in the U.S.A. and Europe and were intensified in the late seventies in connection with feasibility studies of space solar power systems (SSPS). They were not continued in the eighties because the energy crisis was brought under control and SSPS were put on the back burner.

30

o

10 ¢

5 o

10

-10 -

15

////20//

/

25.... Operational""3035 ...." years

40

45

r~ . . . . . . . . . Program A ..

...-'" /

~

Program B

.....

Program C

-50 Fig. 18. Profit trends for programs A, B and C (6% interest rate).

136

H.H. KOELLE

President Bush's announcement that the U.S. would return to the Moon, and eventually conduct a manned expedition to Mars, again brings up the question of suitable space transportation. Present launch vehicles are too small to carry the required payloads and propellants economically for these missions. A Shuttle-C, unmanned for cargo missions to low Earth orbit, is one of the options to raise payload capabilities quickly, but this is a stop gap solution at best, not leading to an economical space transportation system for the 21st century. The Energia launch vehicle of the Soviet Union duplicates the payload capabilities of the former Saturn V vehicle used successfully during the Apollo program. However, it does not appear acceptable from the political viewpoint to rent this launch vehicle for the logistic support of an initial lunar outpost at this time. A long range solution to the problem is a heavy reusable space freighter employing available technology which can take on all missions planned for the first half of the 21st century. Various vehicle concepts have been proposed during the last three decades. One of these resulted from extensive studies at the Aerospace Institute of the Technical University Berlin under the direction of the author, who was also in charge of the preliminary design of the Saturn vehicle family from 1957 to 1962 at NASA/MSFC under the leadership of Dr von Braun. The development schedule of this state-of-the-art vehicle should be less than 7 years, since no new engines are to be developed. The vehicle proposed is sectionalized and modular in design, reducing development and production costs. The development cost including a prototype vehicle and a first set of ground facilities should be close to $20 billion (1990), if managed professionally without too many political perturbations. First unit costs have been estimated close to $3.5 billion. These figures look reasonable and feasible if compared with present budgets and projected costs. With the present emphasis on global cooperation and reduction in defense outlays to be expected in the foreseeable future, development investigations supported by some experiments resulted in the preliminary design of a vehicle called "Neptune 2000 Plus" presented in this summary report. It is a fully reusable, ballistic vehicle with three stages, using derivatives of the Space Shuttle Main Engine (SSME). It has a payload capability of about 350 MT to LEO and of about 100MT for missions to the geostationary or lunar orbits. Its take-off mass is 6000 MT and the dimensions are 42 m maximum cross-section and a length of 71 m. This report contains detailed mass models for a prototype vehicle using standard aluminium alloys as the primary structural material, and for an advanced version optimized with respect to cost, using

advanced materials in those subsystems where it is economically justified. The initial launch site is expected to be some 20 km off the Florida coast not far away from Cape Canaveral. The launch pedestal is mounted on a catamaran type ship. The launch preparation is carried out in a harbor, the fueling at the launch position out at sea. This launch site will be used during the development phase and for logistic supply of a lunar base as well as for manned planetary expeditions. A second launch site near the equator will be needed in case a space solar power system is to be supported logistically by this space freighter. This appears fairly realistic if a life cycle of 50 years is considered to be a likely proposition. The report presents standard trajectories for the ascent and descent of all three stages including mechanical and heat loads. The first stage recovery is effected either on a down range island or on a floating platform. The second stage returns to the launch site after one revolution from an altitude of about 150 km using modulated drag and lift ending in a vertical landing on dry land near the refurbishment site with the help of some of its engines. The third stage returning from GEO or a flight from the lunar orbit goes through the same operational procedure, but has to withstand greater heat loads. Operational behaviour of this vehicle has been studied as well, e.g. individual vehicle life will be 10 years in the flight line which results in a few dozen flights in the early operational period increasing to about 200 flights per vehicle towards the end of the life cycle. Turnaround-times will last several weeks in the beginning of the 50 year system life cycle and funds for this project might become available in the late nineties, leading to a first flight in the first year of the next century. This is in agreement with the goal to establish a lunar base in the first decade of the 21st century. The specific transportation costs of such a space transportation system are estimated to be in the order of $200 (1990)/kg payload for LEO missions which might come down to $100/kg in the case of very large launch rates. GEO and LUO missions will begin with costs in the order of $120 million per flight, resulting in specific payload costs of about $1000/kg. In a high launch rate market these costs may come down to about $500/kg. After several decades of operation the costs per launch have been estimated to be somewhat above $40 million (1990) per flight. Production rates envisioned are from one to five per annum. Some selected operational parameters as they develop over time are presented in this report including detailed cost figures. After an analysis period of nearly four decades of this particular subject and taking present technological and political trends into consideration, the following conclusions can be drawn with respect to

Space freighters the possible acquisition of a heavy space freighter system: (1) Existing launch vehicles are neither large nor economical enough to satisfy the expected space markets of the next century. (2) The next generation space freighter should have a sizeable performance margin to provide growth potential for a 50 year life cycle. (3) Propulsion systems other than presently available chemical propulsion systems of the SSME variety are not required to design and operate launch vehicle economically taking off from the surfaces of the Earth or the Moon. (4) Space transportation systems should avoid as far as possible cumbersome and time-consuming assembly and fueling operations in space requiring manned support, because manned orbital operations are expensive. This leads to rather big launch vehicles, concentrating the activities on the Earth surface, where the manpower is cheapest. (5) A payload capability of about 100 MT to GEO and LUO orbits, or 300 MT to LEO representing a quantum jump over the Saturn V and Energia capabilities, appears to be a desirable goal for the next generation of a big space freighter. (6) Such a space freighter can become available early in the first decade of the 21st century to be complemented by a second generation Space Shuttle for passenger transport to LEO to provide the essential space transportation for the development of extraterrestrial resources. (7) The global resources presently spent on space system development and applications are in the order of $50 billion per year. The total global expenditures in the area of military activities are near $1000 billion per annum, with the expectation that these can be reduced during the next decades. This leaves room for additional programs, hopefully also in space. Therefore, it does not seem unrealistic that funds for the development of a heavy space freighter could be made available shortly after the turn of the century. (8) For the reasons given above, detailed design and system analysis and corresponding planning activities should be initiated for a heavy space freighter as soon as possible.

137 REFERENCES

Other than the historical publications listed in the introduction, the following contributions are pertinent to the subject and have been sources of the information summarized in this report. 1. A. Heyser and H. Riedel, Aerodynamische Beiwerte der wiederverwendbaren Tr~gerrakete NEPTUN. DVL (1968). 2. A. Heyser, H. Riedel and H. Edmunds, EinfluB yon Bremsklappen auf die aerodynamischen Beiwerte der ersten Wiedereintrittsstufe der wiederverwendbaren Tnigerrakete NEPTUN. DFVLR, Institut fiir angewandte Gasdynamik, Porz-Wahn (1968). 3. H. K. Petrick, Berechnung yon Ablations- und Isolationsw/irmeschutzystemen ffr grol3e wiederverwendbare Tr/igersysteme. Bericht TUB-IR, No. 4 (1969). 4. E. Elsner, Vergleichende Analyse yon Systemen gro~r Nutzlastkapazit~it fiir den direkten Transport ErdeMars unter besonderer Berfcksichtigung yon Wirtschaftlichkeit und friiber Verfiigbarkeit. Dissertation, TU Berlin (1970). 5. E. Eisner, Projekt NEPTUN--Entwurfskriterien ffir groBe wiederverwendbar¢ Tr/igersysteme. Bericht BMBW W 70.38, 29 S. (1970). 6. U. Stark, Untersuchung fiber die Erh6hung des Hcckdruckes bci wiederverwendbaren Triijersystemen mit Pilzdfsen unter besonderer Berfcksichtigung yon Systemeinflfissen. Dissertation, TU Berlin (1971). 7. H. K. Petrick, Belastungen yon Platten durch Wechseldruckfelder und akkustische Messungen yon Pilzdiisenstr6mungen hinsichtlich dynamischer Beanspruchung yon Hautstrukturen wiederverwendbarer Tr/igerraketen. Dissertation, TU Berlin (1972). 8. G. R. Woodcock, The low cost heavy lift vehicle. NASA CR-141856. The Boeing Co. (1975). 9. R. W. Jaeger, Ein Beitrag zur Analyse der Fertigungskosten yon Raketenzellen. Dissertation, TU Berlin (1976). 10. H. V. Davis, Space solar power--the transportation challenge in space manufacturing facilities---space colonies. AIAA (1977). 11. E. Birkholz, RETURN--analyse kostenoptimaler Wiedereintauchbahnen und Bergungsverfahren. ILRMitt. 44 (1977). 12. D.E. Koelle, Transcost--cost model for space transportation systems development and operations. MBBReport No. URV-180 (1988). 13. Z. Matijevic, Analyse und Entwurf einer universellen Raumf/ihre f/Jr groSe Frachtraketen. ILR-Mitt. 199 (1988). 14. I. Bekey, NASA's plans for manned missions to the Moon and Mars. Spaceflight 31, 297-302 (1989).