Economy of modular low-technology launch vehicles

Acta Astronautica Vol.6, pp. 1727-1743 PergamonPress Ltd., 1979. Printedin Great Fritain

Economy of modular low-technology launch vehiclest

H. O. R U P P E , R. H. S C H M U C K E R , W. D I T T B E R N E R , D. H A Y N AND M. B R A I T I N G E R Lehrstuhl f/ir Raumfahrttechnik, Technische Universit~it M/inchen, Richard-Wagner-Strasse 18/!11. 8000 Mtinchen 2, F.R.G.

Ahstract--The paper comprises an assessment of the design and the economics of so-called "lowcost" simple modular launch vehicles. It is shown that the performance is very marginal and that the cost per launch cannot compete with technically more advanced fully reusable vehicles. Especially a private-funded development cannot be amortized economically in case of an expendable launch vehicle.

Introduction space transportation has been realized during the last two decades, permitting nearly all conceivable missions within our solar system either with hardware available, or with hardware which could be developed without doubt. This apparently very satisfactory situation is unsatisfactory nevertheless, con-, sidering the high costs involved. Therefore, the main forward-looking activity in this field has concentrated upon cost reduction. Philosophically, two diametrically opposed approaches are possible: --Highly refined launch vehicles, which will be very expensive, with the capability of recovery and reuses are developed; a large fleet (making development cost rather unimportant), a large number of uses per vehicle (rendering first cost of smaller impact) a n d a high flight rate (reducing operational cost) are envisioned to make this course attractive. --Very simple vehicles (hopefully, cheap) are used in an expendable mode; lower performance is made up for by increased size and/or more stages. Again, flight rates should not be too low to keep per-flight cost down. Any reliability decrease is unacceptable, because of the high payload cost. For the second case two different systems will be discussed: (a) Different simple designs for each stage ("Big Dumb Booster"), and (b) Each stage built up from a different number of the same basic module. RELIABLE

tPaper presented at the XXIXth Congress of the International Astronautical Federation, Dubrovnik, Yugoslavia, 1-8 October 1978. 1727

1728

H.O. Ruppe

et al.

Expendable launch vehicle design alternatives There is s o m e hope that present-day expendable launch vehicles are not at the end of their development potential. Cost reductions with such designs can be expected from two development philosophies:

Different simple designs for each stage--Big Dumb Booster In this class, the "Big Dumb B o o s t e r " has been studied in some depth; even experimental hardware has been developed. This choice has been abandoned in competition with the Shuttle: it did not offer more economic promise, and its "future promise" was even more limited. For this study, we proceed as follows: for two missions, (a) L o w Earth Orbit, due East from KSC (h = 200km, vld = 9200 [m/sec]) and (b) Injection to Escape-Speed (rid = 11700 [m/sec]) we investigated 9 different possibilities of staging and combining L o x / K e r o - and H2/O2-Steps (see Table 1, first column).

Mathematical background With the following definitions of the structure-ratio e, assuming 5% residuals, "wet" cutoff stage mass (without payload!), ewi = stage-fuelled mass M/

eu

dry cutoff stage mass = ewi - 0.005 stage-fuelled mass M~

we get the propellant mass for each stage: Pi = M i ( l

(1)

l~Di)

--

the launching weight for an n-stager is defined as n

(2)

Mo = Me + ~ Mi

where Mp is the payload-mass (which includes G & c ) . For performance opti-

Table 1. Stage

Vld

Coebinltlon

9ZO0

I1700

v*rsion of Optl~tzltton

Structure-Ratio

E, 0,087 0,087 0,058 0,0763

0,087 O,~87 O,OS8 0,064Z

~2

E3

J

S~:. ImpulSe Lsec] (Sp) [SpZ [SO3 440 440 46O 440 4,10 440 460 460

GrowthFactor G 28,5 I 28,5

paylo4d I ~ormltzed with ~ost • cheal:~lst Version L S/kg] of each Mission 2 340,4 Z 340,4

13,0

O~

Z1,6

Z 100,9

i

291,0 Z91,0 ~0

261,2 i

~o solutien I

SS,Z

i e6,0

o:, 2a 8'6.a t ';'S8,'

Economy o[ modular low-technology launch vehicles

1729

Table 1. (Contd.) Stage Combination

Vt~ Vernon ~ [Cm/sec;I of Oott- [ t

1

~

.ct~c~ure-~tto

$oec,:rMoulsesec ..,

I 8 , 0 8 7 ~ 0 o. 78 I

't.2

~..1

I 44o I ~o I

~-rovth- plyloid I ,',omali ze~ ~ th I ~ost . , cr,eaoest Version of e~cnM~ssion

I Factor

I u.s

1 051,9 ] I 10Sl,9

130.8 130,8 124,7

2 362,6 I ~44,Z I

~

137,6

437,8 178,8 1437,8c0~178,8 $ 800,8 354,1 ~9 800,85 354,! 12~:.i ?23,3 112,2 109,? 140,0

~I

9200

2

I

o

4

0,088 0.087

I

0,0992

55,3 50,4 29,2 45,~

0,087 0.088 0.0gO2

136,7 O~ 129,6 186,4 185,2 CO 169,5 266,4 260,6 ~ CO 3 636,0 0Z!.9 989,7 123,1 9:,7 123,1 CO 117,0 I 1Zl,3 12~.3 CO 116,8 i 149,5 ;49,5 GO 137,7 Z07,4

207,4 180.0

440

sc'qlT~-Stliq~r LOX/kero- H2,0~ ~6,e I ~8.o ] 1oo.o 9,7 CO QO

440

z~ Two-$tagii- LOX/Kero * NZ.,02

10o

] ;Ox,'KE,OI

35,7 19,I

37,1 18,4 0Q,2 Z8,7

48,0 $0,4

08,4 55,2

1 566,Z 101,7 ~

CO

I 638,3 , 180 ~ 121,0 ~3,6 107,4 ,Z " ;03,8 I 98~,4 I,:I,2 l ,; 117,~ CO 114,1

173(}

H.O. Ruppe et al.

mization we have to minimize M0, i.e. d e t e r m i n e P~, so that M0 is a minimum, fulfilling the side-condition for the required ideal velocity.

Via = ~

ci In ri

(3)

i=1

w h e r e c~ is the e x h a u s t - v e l o c i t y of each stage and initial m a s s of stage i r~ --- cutoff mass o f stage i' To find the transport cost for each kg of p a y l o a d , we i n t r o d u c e the following cost models: 1. P r o p e l l a n t - c o s t :

CpH,/O, = 0.5 $/kg,

cp,,,,,Ko~,' = 0.2 $]kg

C~H~/o; = 1000 $/kg,

c,~.~,K,,~,,= 500 $/kg

2. S t r u c t u r e - c o s t :

3. F i x e d c h a r g e s : (a) f o r each first stage:

(b) for each additional stage:

c~t = 1.5"

1075

Cfix2 = 0.5 • 1075.

With this. we are able to optimize p a y l o a d cost in two w a y s , holding E c o n s t a n t : V e r s i o n l: p e r f o r m a n c e optimization (see eqn 2), multiplying the results a f t e r w a r d s with the a b o v e f a c t o r s (see eqn 4), V e r s i o n 2: minimizing

Co = ~ (EDi " Mi " c~, + Pi " cp, + c~i).

(4)

il

Next, let us refine this model in a w a y that the cost of changing specific impulse and c h a n g i n g structural fractions will be taken into consideration. This can easily be d o n e by introducing E as a restricted variable and a c o s t - c u r v e on Lr (see Fig. 1) with the following e-values (a) " n o r m a l " Ew~,( -- r e f e r e n c e - v a l u e ) E,.H~/O, = 0.087 and E,.j ,,x..K,.,,, = 0.068, (b) "limit" Ewet ( = best, but at e x t r e m e cost) 2 EWlim ~ 3 Ew'

(c) variable e~,., = E.,,

Economy of modular low-technology launch vehicles

1731

[c 2 I

290

300

310 320 330

430

440

450 4 6 0

Isp Isp Fig. I. Cost factor assumption for increasing performance. we get for stage-cost (the exponent 2/3 results from quite subtle considerations): Ci =/Vii • ~.Di .

.

.

.

[ci " Csi + Pi " Cpi + Cfixi

.

(5)

~D/ - - (~Wlimi/'

so we have to minimize Co

Ci.

(6)

i=1

For comparison, we compute performance-optimization with best E~ = Elim and best Isp: this will be the most lightweight structure ( = V e r s i o n 3, Column 3, in Table 1). V e r s i o n 4 (see Column 3, Table 1) is the result of eqn (6). All four versions are computed for a constant payload mass of 200 t. Those who do not believe in our scaling, should multiply the last two columns by their own factor (e.g. Saturn V: k = 2). Conclusions ---The worse the system, the more it can be improved by optimization --Optimization requires better structure-ratios with increasing missiondemands - - U p p e r stages have to be " b e t t e r " than lower stages - - F o r both missions the Saturn-V-like three-stager is the cheapest (Combination No. 8) - - F o r high-energy missions, L O X / K e r o as last stage doubles transport-cost in the case of the two-stager and is 2/3 times higher in the case of a three-stager. - - T h e r e is no way to a reasonable single-stager - - M o s t surprising: the cost-optimized versions are not at all "big dumb boosters", but highly refined vehicles. (Just first stages seem to show a trend towards "big and d u m b " ) This cost model would reflect that the historic trend towards launch vehicle refinement is justifiable on economic grounds. Modular clustered vehicle (OTRAG, 1978) Consider a relatively small tank/structure/engine system, which we will call a module. From a number of modules a stage and from several such stages the space transport rocket is assembled.

1732

H.O. Ruppe et al.

Module design The basic principles of one module are:

- - L i q u i d propellants: Kerosine (Diesel Off) as fuel Nitric acid as o x y d i z e r

price: 0.3 $/kg

--Tanks : 4 clustered steel extruded tabular tanks of 0.3 m $. The tanks are segmented into 3 m long pieces. End plates are flat pieces.

--Feedsystem : Nitrogen b l o w d o w n f r o m ullage space, a b o v e liquids.

--Engine: Simple, tubular f r o m the same steel tubing as above. Ablation-cooled.

--Ignition: Hypergolic p r e b u r n - - s i n g l e start capability (Furfurvlalcohol). Thrust shall be 3 tsl, at Pc = 20 bar L o w pressure limit: pj = 10 bar (p~ = pressure at tank bottom) Thrust levels are 0%, 50% and 100% ( u n s y m m e t r i c throttling is utilized for direction control). Figures 2-5 show the basic idea of tank and engine configuration of a single module and a possible a r r a n g e m e n t for a launch vehicle.

Specific impulse Figure 6 shows the v a c u u m specific impulse, assuming equilibrium flow conditions. The calculations show an o p t i m u m mixture ratio r = 5. With an area ratio E = 6 the engine shall have the following theoretical p e r f o r m a n c e : Isp, the, vac, O D E = 272.5 sec c*, vac, the, O D E = 1566.6 m/sec Isp*t = 159.69 sec (go = 981 m/sec2). The various impulse-losses are estimated in the following table; e - - 6 for all cases.

F r o m a p e r f o r m a n c e point of view, those are very optimistic figures: 247 nt~p -- 27--3= 0.904 C o m p a r i s o n with the J 2-engine (H2/O2-Saturn V) 424.8

rlt,o = 448.2 = 0.947 */~p* is scaled for c*

Economy of modular Iow-tet~nology taunck vehicles ,

1733

Fig. 2. Overall view of a four-tank OTRAG module, length = 6 m (OTRAG, 1977).

P c = 54 bar, • = 27.5, bell-nozzle regeneratively cooled etc! So we may conclude that the simple Kerosine/nitric acid-engine may well have losses even larger than we assumed. Flight p e r f o r m a n c e and vehicle data Speed requirement: ( - ) Earth rotation v* 200 km circular orbit g-loss Drag-loss Nozzle-end pressure

8030 m/sec 1300 m/sec 180 mlsec 120 m/sec 9630 m/sec

Reserves L o w limit Upper limit

- 460 m/sec 9170 m/sec 90 m/sec 9260 m/sec 9110 m/sec 9460 m/sec

1734

H.O. Ruppe et al.

"Correct arrangement" (hquid volume: gas volume : equal for both tanks h.

Solution of equal len(Jhts Move ,5/2 over to the other s t d e Gas manifold

_..t__-7

:i

~" ~'

Se parahon bulkhead

I L i i

-4

__t Fig. 3. Tank module design features.

Engine

Die

"

Oxydizer /

~

~

-

"

'

~

:

/

:

~

'

-

~

0

Feed lines from tank to engine

--

Bolted with one central bolt to tank

~

"

~

OaXnYdizer

Diesel Oll

Fig. 4. Engine/tank chester arrangement of one module.

Economy of modular low-technology launch vehicles

I st s t a g e : 1 2 0 m o d u l e s

1735

*

2nO stage : 3 4 m o O u l e s 3rd stage : 12 modules 4th stage : 3 modules • 48

o f this are 3 m longer

Fig. 5. Parallel clustering concept for a large launch vehicle with 120 modules as first stage, 34 modules serving as second stage, 12 modules as third stage and 3 modules as fourth stage. 350

~

~

~

-

~

t -I000 t00

300

o

~

~"

~

250

200 35

6o

5.0

~

1

40

k 4 5

4.0

50

55

6 0

rO/F Fig.

6. Specific

Impulse

vs

mixture ratio with expansion (Kerosene/Nitric acid).

ratio as parameter

1736

H.O. Ruppe et al.

Table 2, Impulse-losses Type

Loss in %

(a) c*/c related c*-reduction, combustion efficiency pressure drop boundary layer r-shifting ablation contamination (c*~ c*ol>F)" 1/4

-

2.6

-1.5

-0.1 -0.5 -0.5 -- 0.9 6.1

(b) c only boundary layer ablation contamination

- 1.5 -0.15 -0.55 2.2

(c) During flight operation blow-down-principle throttling of engine throat expansion (ablation) base pressure gain

-0.5 -0.1 -1.7

+1.I 1.2 9.5%

-

Total l~p, vac, average: 247 s = 2423 m/s.

Per[ormance summary Using the above results the investigation for a 3 × 3 = 9 module cluster with three 12 m long stages shows that the system has no orbital capability. According to the complex pressure dependance of time and acceleration the relevant equations were solved by a computer-program. Results can be seen in Fig. 7. The large OTRAG carrier has 13 × 13 modules, with 24 m tanks. The outer ring of the modules (i.e. 48 together) have their tank lengths increased by 3 m. This longer tanks form a part of the payload shroud. Besides, it increases performance (by additional propellant loading) and insures control (together with pressure gas exhaust) during coast phase between 1st and 2nd stage, and during staging. Some general observations: ----4 stages to orbit give optimum growth factor; still results in sufficient launch acceleration. --Propellant performance is low (I~p.... = 247 s), partly because of the low engine expansion ratio (e = 6.17) and the simple conical nozzle.

_

Economy of modular low-technology launch vehicles

1737

Table 3. Module mass for different tank lengths L=6m P0' P0 H h l (2L-l)

31.5383 32.7076 4.8636 1.8909 3.3112 8.6887

L=24m 27.5207 31.9041 18.2316 7.0884 12.5821 35.4179

bar bar m m m m

Loading

Fuel Oxydator Gas(N2), Fuel Gas (Ng, Oxydator

103.32 516.61 3.58

387.31 1936.55 12.09

kg kg kg

10.00

43.84

kg

Total loading 2

633.51 1267.02

2379.79 4759.58

kg kg

Dry mass

465.0 1732.02

685.0 5444.58

kg kg

1227.46

4601.24

kg

27.00 477.56

54.00 789.34

kg kg

Launch mass NIL

99% propellants Engine use (ablative) Cutoff-mass Mc

rMoaule= ~ML

3.6268

6.8976

--

-I = ~ r

0.2757

0.1449

--

---The b l o w d o w n f e e d s y s t e m leads to a significant f u r t h e r d e c r e a s e in a v e r a g e specific impulse: 232 s. - - E n g i n e lifetime (ablation-cooled) m a y not suffice f o r the long burn d u r a t i o n (120 sec and more) - - N o z z l e t h r u s t e r o s i o n m a y lead to f u r t h e r r e d u c t i o n of engine p e r f o r m a n c e . - - T h e r e is an o p t i m u m g a s / p r o p e l l a n t loading, d e p e n d i n g significantly u p o n the p o l y t r o p i c constant. W o r s t case is slightly larger t h a n adiabatic, b e c a u s e of J o u l e - T h o m s o n effect; b e s t case w o u l d be i s o t h e r m i c (impossible). - - I n i t i a l c o n d i t i o n s prior to l a u n c h (gas p r e s s u r e and t e m p e r a t u r e ; liquid m a s s and t e m p e r a t u r e ) h a v e to be precisely equal in c o r r e s p o n d i n g tanks. Besides " n o r m a l " p r o b l e m s o f m e a s u r i n g , difficulties m a y be p o s e d b y climatic c o n d i t i o n s (wind etc.) and sunshine, acting u n s y m m e t r i c a l l y u p o n the vehicle. - - S t a g i n g and inflight ignition w o u l d require m o r e detailed studies.

H, O. Ruppe et al.

1738

?

m/s

~ m/s

p

..~" i. . . . . . . .

T

..... i • . . . .

i

',

I

!

~

i

bor

s

i

,

j

,,o,.oo ,o,.ok

f~p

'

T .......

/,

~

7--~

~

o

/

!

300

i l i

-

rn

r~ . . . . . . . . . . . . . . .

v

./.

.,o

:

kg/'s

i

,.oo.......,oi

,oo., 24

PO00

2300

soo

2e5o

20

220

is

zlo

i

i

2 3 '~o

i 0

2200

I0 ~200

0 D

20

40

60

80

I00

120

2 2

I~0

s

Fig. 7. Performance characteristics as 1,;,, velocity, mass flow, tank bottom pressure, exhaust velocity and thrust acceleration vs burn time for a three-stage 9 module cluster (M,, = 6 655 kg).

- - S i n c e structural factors of our idealized design are very favourable, no further gain can be expected in this field (with the exception of reduction of engine n u m b e r in final stages). - - T h e vehicle appears to be difficult to be controlled by differential thrust throttling; fins m a y relieve this situation, if the problem of too much stability can be avoided. Flight control remains a very critical area both within the first stage and for the long and slender last stage, outside the atmosphere. - - V e h i c l e reliability remains a critical area, in our opinion. No sensible analytical approach will lead to a significant solution of this complex problem. - - T h e extended ring of modules adds 30.4 t to the launch mass and provides a speed increment of about 55 m/s. Subtracting that we get for the most optimistic estimate of speed requirement: 9050m/s, and the most pessimistic case: 9420 m/s.

Capability Assuming a G & C weight of 400 kg, we get Optimistic limits: I,~ = 260 sec (certainly way out! O T R A G claim) Vr~q = 9050 m/s Residuals: 0.5% Payload: 8.7 t

Economy o[ modular low-technology launch vehicles

1739

Pessimistic limit: hp = 232 s Vreq= 9420 m/s Residuals: 1% Payload: 2.9 t Launch mass: 957 t (approx) Growth factor: 110-330 Probable payload 3.5 t; G = 273 (Corresponds to 232, 9210, 1%) Sensitivity: - 47 kg per promille residuals _+ 144 kg per sec/so - 2.85 kg per m/s velocity requirement.

Cost analysis The assessment of the overall transportation cost of a modular launch vehicle must include, besides the analysis of production, vehicle assembly and operation cost, the development cost with interest and gain and the role of propellant cost versus performance. Hardware and operation cost Hardware production cost. Normally, the prediction of the launch vehicle production cost is based on historical data. For conventional launchers with non-modular design, this approach results in reasonable accuracy. The nonconventional modular launch vehicle should therefore be different in cost compared to the above-mentioned type. Since no historical data for such an approach are available, two limits are used for cost prediction: The upper cost limit is characterized by typical space vehicle cost data, while the lower limit is based on the mass production data of the normal industry, especially such of car industry. In order to come up with an estimate for specific structure cost, the basic construction material for such a pressure-fed vehicle has to be taken into account. For structure mass and therefore vehicle size reduction, a high-strength material is required. Using flow-turned steel may result in a tensile strength up to 150 - 200 N/mm2; the cost for the material range from 5 to 50 $/kg (e.g. Ultrafort ®), a somewhat arbitrary estimate is 10 $/kg. On the basis of the previously mentioned facts, the hardware production cost can be assessed. Three different launch rates per year are used for estimation of the production numbers, 3, 10 and, as an upper limit, 50. The associated numbers of engines, tanks and modules are summarized in Table 4. To perform these large production numbers, several production and assembly lines are required. In Table 5 some of these numbers are listed. The resulting hardware production cost including guidance and control are summarized in Table 6. In addition, some estimates for production investment, including offices, machine shops, social buildings and a small airport for transportation to launch site or suitable port are also listed.

1740

H.O. Ruppe et al.

Table 4. Component production numbers of a modular launch vehicle Launches/year Component (unit)

Units/ launch vehicle

Modules Engines Tank elements Tank domes Propellant valves Tanking quick disconnects Tank dome screws

169 679 5408 10,816 1352 676

3

Units/year 10

50

507 2028 16,224 32,448 4056 2028

1609 6760 54,080 108,160 13,520 6760

8450 33,800 270,400 540,800 67,600 33,800

850,000 2.5

• 10 6

8.5 • l0s 42.5

• 10 6

Table 5. Production and assembly line numbers for tanks and engines

Launches/year

Number of production and assembly lines 3 10 50

Shifts/Day

1

2

3

Tank elements Tank domes Tank assemblyt Engine assemblyt

8 8 12 5

15 15 21 8

50 50 70 25

tlncluding production and quality control and leakage testing. Table 6. Hardware production cost and total production investment Launches/year

3

10 50

Production cost ($/kg) 60 45 35 Vehicle hardware cost (excluding launch site assembly and check-out)t 7 5 4 1065 Investment for production (middle european production site) 1065 100 150 300 tWithout ROI.

T h e d a t a a r e t h a t f o r a m i n i m u m i n v e s t m e n t p r o g r a m ; it is h a r d t o f o r e s e e t h a t t h e s e l o w c o s t f i g u r e s c a n b e a c h i e v e d . B u t n e v e r t h e l e s s , t h e d a t a l i s t e d will be used for the assessment of the low limit launch cost.

Economy o/modular low-technology launch vehicles

1741

Propellant cost T h e propellant cost of pure nitric acid (approx. 100%) and gas oil is about 0.3 $/kg. For transportation to launch site roughly another 0.3 $/kg have to be added. Besides propellant transportation, production near the launch facility is possible. This requires the establishment of a chemical and a p o w e r plant; the i n v e s t m e n t should be one to several 100 Mio$ Hardware transportation to launch site and assembly/check-out cost Each engine and tank subunit, including two d o m e s with quick connection devices, are assembled and partially checked in Europe. After transportation to the launch site, the c o m p o n e n t s are finally assembled and c h e c k e d again. Hardware transportation cost Due to the sensitivity of the launch vehicle hardware to handling and transportation overloads, an air transportation seems to be the o p t i m u m choice. On basis of a B 747 cargo carrier, approx. 3 flights per launch vehicle with the hardware and payload are required, since the low overall density of the engines and tanks results in a volume rather than a mass payload limitation of the plane. Thus, roughly 0.4Mio $ for transportation have to be added to the overall vehicle cost. Vehicle assembly and check-out Vehicle a s s e m b l y can be p e r f o r m e d in two ways: vertical a s s e m b l y and horizontal assembly. In Table 7, the required s u b a s s e m b l y facilities, launch pads, launcher a s s e m b l y and operation crew for vertical a s s e m b l y are listed. In case of horizontal a s s e m b l y , the figures change s o m e w h a t to larger numbers. Table 7. Important data for vertical assembly of modular launch vehicle (3 shifts a day) Launches/year

3

Number of 4 engine/ 4 tank modules 4/4 Module subassembly facilities Subassembly time/launcher (days) Launch pad preparation (days) Launch vehicle assembly time, including check-out (days) Payload integration, Tanking, final launch preparations and Launch (days) Number of launch pads Launch site crew

507

I0

50

1690 8450

10

30

150

80

28

6

15

15

15

20

20

20

l0

l0

10

2 400

3 12 800 3500

1742

H.O. Ruppe et

al.

Table 8. Transportation cost (Mio S/launch) Launches/year

3

10

51)

Development cost investment --Production --Launch Site

500

500

600

100 10(l

150 200

300 500

Overall investment

70(1

850

1400

7 0.4 0.5 6.6

5 0.4 0.5 4.0

4 0.4 0.5 3.5

14.5

9.9

8.4

16 23 23

6 8 8

2 2.5 2.5

76.5

32

15.4

l,aunch vehicle hardware l,aunch site transport. Propellants Assembly. . . . . . launch

Interest ROI Gain

Transportation cost The transportation cost, derived from production, operation and investments are listed in Table 8. The result of this transportation cost analysis is, in its overall aspect, very similar to that of conventional launch vehicles. The launch or transportation cost, not including amortization etc. covers a range, comparable to that of conventional launchers, but in spite of the low specific hardware production cost, the overall cost exceed that of conventional rockets by a factor of roughly 1.5. With increasing launch rate, the transportation cost drop sharply, similar to normal rockets. If development cost and all investments are included, this picture worsens dramatically. Costs are significantly increased up to a level of 5000-20,000 $/kg which completely eliminates the private launcher development. Is there a possibility to reduce transportation cost below that of STS or ESA'S ARIANE in this way? By the choice of the propellant combination, the overall transportation costs are fixed, though the propellant cost can normally be neglected. The proper selection will result in a specific transportation cost reduction of a factor of 1.5-1.7 compared to HNO3/RP1. The following specific costs are computed for a payload of about 3000 kg: Launches/year

3

10

50

Specific transportation cost (S/kg) 25,000 I 1,000 5000

Economy of modular low-technology launch vehicles

1743

It is obviously to see that a private c o m p a n y cannot compete with a non-profit (government) organization launcher development, since amortization, interest rate etc. represent a significant fraction of the overall cost, roughly a factor of 2, when high launch rates are considered. This picture worsens for lower launch rates so that the resulting specific transportation cost for L E O are as high as those occurring for GEO. Therefore, it can be concluded that, from a cost standpoint, a private development, especially that on basis of "low cost propellants" with "simplified high modular structure principle", does not represent a promising solution at all. No method for cost reduction seems to be within sight, to come down with transportation cost to new highly refined systems.

References Kayser L. T., (1977) Personal communication. OTRAG (1977) Press releases. OTRAG (1978) Press releases. Rockwell Internatl (1976) Space Shuttle-System Summary. Ruppe H. O. (1966) Introduction to Astronautics, Vols. I and II. Academic Press, New York and London. Ruppe H. O. (1975) Kaysers (T)Raumfahrt Umschau, Vol. 23.