The transcost-model for launch vehicle cost estimation and its application to future systems analysis

Acta Astronautica Vol. 11, No. 12, pp. 803-817, 1984

0094-5765/84 $3.00 + .00 Pergamon Press Ltd.

Printed in the U.S.A.

THE TRANSCOST-MODEL FOR LAUNCH VEHICLE COST ESTIMATION AND ITS APPLICATION TO FUTURE SYSTEMS ANALYSISt DIETRICH E. KOELLE* Messerschmitt-Bflkow-Blohm GmbH (MBB), Space Division, EO.B. 801169, D-8000 Miinchen, ER.G.

(Received 20 March 1984) Abstract--Tbe first part of the paper describes the structure of the analytical cost estimation model (1982 edition) for launch vehicle development, fabrication and launch operations cost. Especially the new approach for a cost assessment of operations cost including refurbishment (in case of reusable vehicles), direct and indirect operations is presented for discussion and subsequent improvements by introduction of more reference values. The model uses the Man-Year (MY) as cost unit which is independent from inflation and currency exchange rate changes. The second part of the paper deals with its application to future systems analysis and cost comparison with the example of a potential future European launcher (Post-Ariane-4) with 15 tons LEO payload capability: ten different two-stage launch vehicle concepts (expendable, semi-reusable and fully reusable) with storable and cryogenic propellants are analysed with respect to development cost and cost per launch. The key problem for a future European launch vehicle is the optimum solution between the (limited) development effort and the desired minimum launch cost. More advanced (partially) reusable systems could provide an essential reduction in cost per launch, require, however, a higher development effort. In such a case an analytical cost model based on realistic reference data can provide important data for the vehicle concept selection process.

I. B A C K G R O U N D AND M O D E L S C H E M A T I C

I. I Survey and background The TRANSCOST (82) space transportation cost model is based to a certain extent on the relevant parts of a cost analysis performed in the period 1965 to 1970 of some space projects[l]. It was published in a summarized form as paper and preprint at the IAF Congress Vienna, 197212] and subsequently in the Journal RAUMFAHRTFORSCHUNG[3]. ESA translated and published the complete cost model as Report[4]. The work was also translated and published in a Russian Journal[5]. Another effort more specifically dedicated to space transportation has been made in the 1974 to 77 period by a joint analysis of MBB and the University of Bedim The results have been published in Acta Astronautica[6] under the title "Future low-cost space transportation system analysis" by D. E. Koelle and H. H. Koelle. Here the first attempt was made to define cost model elements for operations cost, the most important cost factor for space transportation if it comes to reusable vehicles. This effort was continued in the IAF-paper 78A2717], which has also been used as an input for the new TRANSCOST Model. Its first version has been prepared in 1980 as part of the ESA Study on "Future Space Transportation Systems." ?Paper presented at the 12th Symposium of the International Academy of Astronautics (IAA) on "Space Economics and Benefits," 33rd Congress of the International Astronautical Federation OAF), Paris, France, 27 Seplember-20 October 1982. *Director, Advanced Space Systems and Technology, MBB Space Division.

This paper reports about the updated 1982 Edition (Rev. 1) and its application for the analysis of a Future European Launcher after the ARIANE 4 family.

1.2 Cost model schematic The TRANSCOST Model is organized in three major parts which are interconnected (Fig. 1). The advantage of this subdivision is the possibility of making cost comparisons in all three areas separately or, by example, include or exclude the development cost (the question of development cost amortization, which has been disregarded in the past, but can be considered in case of fully reusable systems for commercial application). The subdivision in the three areas shown in Fig. 1 seems also to be useful for flexibility in application since there are different programs scopes to be analyzed analytically, by example --development program with flight test vehicles and launch cost (add-on of submodels 1, 2 and 3) --flight program cost with the vehicle hardware required and the related operations cost (add-on of submodels 2 and 3). The flight operations cost (3) also include the constant indirect operations cost of the launch site which have to be shared by the number of launches performed. The cost model is structured such that the following data can be derived: • specific transportation cost (with or without vehicle and development amortization) • development cost for a transportation system (with or without a given number of test flights) 803

804

D. E. KOELLE SUBMOOEL

1

Development Cost (Elements + System)

I Test Flight

Hardware

SUBMOOEL

Vehicle Cost

2

I SUeMOOEL 3

Flight Operations Cost

t Indirect Operations Cost

Fig. 1. Organization of the TRANSCOST-Model in three major submodels. • fabrication cost per vehicle or for a given fleet size • total cost per launch with or without vehicle fabrication (share) and development cost amortization • the total transportation cost for a given payload scenario • the optimum vehicle s&e (payload capability) for a given annual cargo transportation requirement. The SPECIFIC TRANSPORTATION COST as the key value for comparison of transportation systems is defined as follows:

C. = Flight Ops. Cost + Vehicle Cost (share*)** [MYgY] Launch Vehicle Payload The effort in all cases is expressed in Man-Years (MY) as a constant figure, independent from hour rate increases and currency exchange rate changes. For conversion to the actual value of the effort, the cost of I MY is inserted for the relevant year (or average for a time period). The value of 1 MY includes not only the direct labor cost but all other costs too (including material, components, computer cost, travel cost, indirect burden, fee, etc.). The MY value is an industrial average of total cost divided by productive hours per man and year (about 1 820 h in the USA, and 1 700 h in Europe). The following values have been used for the past years: EUROPE 48 000 AU 52 300 AU 57 600 AU 66 000 AU 72 500 AU 80 000 AU 88 000 AU 95 000 AU

1977 1978 1979 1980 1981 1982 1983 1984

USA 72 000 Dollars 80 000 Dollars 85 000 Dollars 92 500 Dollars 103 000 Dollars 111 000 Dollars 116 000 Dollars 120 000 Dollars

The basic CER's (Cost Estimation Relationship) are based on statistical data and are conceived for similar technical systems, i.e. *(for re-usable vehicles) **Addition of development cost share is optional.

--liquid propellant rocket engines --expendable launch vehicle stages - - m a n n e d winged space vehicles - - u n m a n n e d ballistic reusable stages, etc. It is important to convert all reference project cost values in the MY figures, and to define appropriate degression factors for development CER's (in order to take into account the technological status of the project dealt with), and to apply learning factors in case of production cost values. The new submodel of OPERATIONS COST is presented here as a basis for discussion. It is the first time these complex cost factors have been assessed in a model, and everybody who could contribute to this problem is invited to address the author in order to achieve a more mature model in the future. List of symbols is given in Appendix I at end of paper. 2. DEVELOPMENT COST OF SPACE TRANSPORTATION

ELEMENTSANDSYSTEMS(Submodel1)

2.1 Fundamentals The development cost model is based on[ 1] analytical results for launch vehicle stages, rocket engines and manned space vehicles. The scope of the development cost submodel includes the main vehicle elements, system engineering, integration and testing and the ground installations, but no flight hardware and flight test vehicle operations. Those have to be derived from the other two submodels if complete program cost are required. The total industrial development cost of the complete launch vehicle have been defined as the sum of the single elements cost plus 10% for system engineering and testing, i.e. Cov,h = 1.1

Hs +

HE MY

with Hs being the development effort for stage systems and Hr the effort for rocket-engines (as far as applicable, i.e. if an existing engine is being used, then HE becomes zero).

The TRANSCOST-Model for launch vehicle cost estimation and its application to future systems analysis It has been found sufficient to restrict the submodel elements to their two major cost items, and to include the other smaller ones. By example, the payload shroud is considered to be part of the first stage (in case of expendable vehicles) and the instrumentation compartment belongs to the final stage. The basic development cost elements have the form of

I H = a'MX'f,'f2"f3

I MY

with H

--- development effort in MY (Man-Years)

a [ = specific values for each type of equipment or xj system (defining slope and level of the reference curve) M = the reference system mass (kg). The three factors ft, f2, f3 represent the major impacts of overall system development status, the technology status applied and company/team experience: / l = correction factor for the overall technical development status: --first generation system: fi = 1.25 --technology already proven by similar systems elsewhere: f~ = 0.8 to 1.0

IAR LANDER 36 oo0 MY)

M__~Y

I

I

I

- - s a m e system as already built (only modifications or change in size): fj = 0.4 to 0.8 fz = technical quality factor, different and specific for each type of system f3 = team experience factor - - n e w team/company: f 3 = 1.1 to 1.3 ---company/industry team with some related experience f3 = 0.9 to 1.1 ---company/industry team with previous relevant experience: f3 = 0.6to0.9. 2.2 Stage~vehicle systems d e v e l o p m e n t cost 2.2.1 Definitions. This group of space systems comprises stage/vehicle systems, excluding the engines, but including all other equipment as well as stage adaptors or shrouds, as applicable. The basic development cost effort a • M x is derived from realistic reference values, introducing a suitable technical quality factor (f2) which is important to achieve the desired accuracy. Figure 2 shows original reference points for the launch vehicle system CERs. There is evidently a clear trend of specific cost reduction with growing system mass, however, not proportional: the double vehicle mass in this case leads to a development cost increase of about 25% (or a decrease in specific cost of about 37%). Using the stage net mass as reference leads automat-

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.

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Fig. 2. Specific development cost of different types of space transportation vehicles.

10 7 kg

8015

D. E. KOELLE

ically to the problem that a more lightweight, advanced structure would theoretically reduce the cost, in reality, however, application of advanced technology is a cost driver. This problem has been overcome by an appropriate definition of the quality factor f2 which is related to the stage K-factor MN/Me. A degression factor analysis led to the optimum value of

( Kr,f~ 1.8

with K,,f being the average K-factor. It can be derived by using Fig. 3 which shows the present range of Kvalues vs. vehicle propellant mass for the different groups of stages. 2.2.2 Expendable stages system development cost. The development test analysis of expendable stage systems led to the relationship

The resulting definition of the basic development effort is HM = 6500 M °21. f~" f2" f3 [MY] indicating more than twice the development cost of expendable stages. The model curve defined is in good agreement with the cost estimates published for the Rockwell Booster (1971) and the Boeing Two-Stage SPS launcher with 420 Mg payload in LEO (1979). 2.2.4 Reusable ballistic vehicles (stages) dev. cost. A third type of launch vehicle systems are reusable ballistic stages, however, no reference project cost data exist yet. Therefore, a provisional assumption has been made as shown in Fig. 2 with HR = 4080 • M °2' • f , • S2 " f3 [MY].

Hs = 3 1 4 0 " M °2'. f , . f 2 - f 3 [ M Y ] with M being the stage net mass excluding the rocket engine(s). 2.2.3 Winged (manned) systems development cost. While in case of launcher stages and rocket engines a number of reference projects exists, this is not the case for winged manned space vehicles. Only the Shuttle Orbiter development cost are available with 6380 Mio. Dollars or 89 000 MY (without SSME, ET and SRMs). Based on the reference curve for the specific development cost of stages (Fig. 2) a parallel curve has been assumed for manned vehicles using the Shuttle Orbiter reference point with fl = 1.25 since there is no reason that these curves should cross each other at any size.

This assumes a 30% higher effort than for expendable stages, independent from the anyhow higher development cost due to the increased net mass. 2.3 Rocket engine development cost The available reference points of rocket engine development cost have been assembled in Figs. 2-4. The CER derived is based on the fact that the RL-10 and SSME projects should be above average as first engines of its type, while the conventional F1 engine should range below the average effort. The slope of the curve or "scale factor," in addition, is identical to the cost vs. thrust level relationship presented by Pratt and Whitney.

0.28 0.26 0.24

0.22 0.20 0.18 0.10 0.14 0.12 STO 76)

0.10 0.08

rT)

0.06 0,04 0.02 10

20

30

50

100

200

300 500

1000

2000 5000 10000 PROPELLANT MASS Mg

Fig. 3. Stage mass fraction values vs vehicle propellant mass for five types of launch vehicle stages.

The TRANSCOST-Modelfor launch vehicle cost estimation and its application to future systems analysis

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ENGINE DEVELOPMENT EFFORT o,

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Fig. 4. Liquid propellant engine development effort. The resulting CER for turbopump engines is He = 1 6 2 . M ° ' S s . f z . f 2 . f 3 M Y and for pressure-fed engines:

of a complete launch vehicle system are defined by the sum of the single elements cost plus a factor 1.02 '~ for management, integration and checkout:

Cev,h = 1"02~

F, +

Fe MY )

He = 9 2 . M ° 5 ~ . f ~ - f 2 " L M Y The quality factor f2 is mainly determined by the operational reliabilityto be demonstrated by extensive ground test f~ings. Figures 2 - 5 shows the reference value f2 = 1,0 for a statistic reliability value of about 0.995 and the cost/reliability relationship, taken from [7]. U.S. Rocket Engines have a much greater qualification program than European engines as shown by the examples in Figs. 2-5, because of their use for manned missions. The Viking Engine qualification initially comprised only 180 firing tests (before the fast launch). After the engine failure of flight 02 some 140 additional tests have been made.

3. VEHICLE COST SUBMODEL (Fabrication, Integration and Test of Flight Hardware)

3.1 Cost model structure

The vehicle cost submodel has the same basic structure as the development cost submodel, i.e. it is based on the relevant hardware mass of similar systems. The total industrial fabrication, assembly and test cost

(N = number of different stages and rocket engines). The basic cost elements have the structure F = n'a'M~'f4.

Indirect cost like appreciation of buildings and facilities are not included. The fabrication rigs and tools cost are included in the development effort. 3.2 The "learning factor" The so called "learning factor" was defined by T, P. Wright in 1936 and is being used since that time. The factor p = 80% or 0.80 defined that doubling the production number will reduce the unit cost to 80%, i.e. the second unit costs only 80% of the fast, the fourth unit only 80% of the 2nd and the 8th unit only 80% of the fourth, etc. The learning factor p depends on the type and size of vehicle as well as on the production rate (units per year). For large systems and small production rates the learning factor becomes practically nil, for quantity production of smaller units, however, it is a major cost factor. Reference [6] provides charts for the cost reduction

808

D. E. KOELLE

TURBOPUMP ROCKEr ENO,NES

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Fig. 5. Quality factor f2 as function of the number of engine test firings and the demonstrated engine reliability.

factor f4 vs. production number with the learning factor as parameter, as well as a learning factor estimation diagram.

3.3 Cost relationship f o r rocket motors The costing relations have been derived by statistical evaluation and regression of actual project cost data. It is important in this case to relate all cost values to the first production unit by use of a learning factor and the actual total production number. In case of SOLID PROPELLANT ROCKET MOTORS the following CER has been established: FR = n • 3.9 • M °31 • f4 [MY] with M being the motor case and nozzle mass (in kg). For LIQUID PROPELLANT ENGINES the following cost relationships have been found: (a) for relatively simple, single-use engines with storable propellants FE = n ' 2 . 5 ' M

°~-f4[MY]

(b) for standard-type single-use cryogenic engines (HJO2) Fe = n • 4.0 • M °'46 • f4 [MY]

(c) for complex, multi-use cryogenic engines Fe = n • 5.0 • M °'~ • f4 [MY]

3.4 Cost relationship (CERs) f o r launch vehicle stages 3.4.1 Ballistic stages~vehicles. A large variety of expendable stages has been built in different numbers which can serve as reference values. It is required, however, to re-calculate the unit no. 1 cost taking into account the learning influence depending on vehicle size and number of units built. The result of this exercise is shown in Figs. 3 - 5 , resulting in a satisfactory cost model vs. mass (with the exception of the AR-L. 140 and the S-II stages). There is no evident cost difference between stages with storable or cryogenic propellants, however, cryo-stages are larger and have higher net weights, so that higher cost occur automatically. The same should apply for reusable stages. The stage net mass does not include the engine(s), to be calculated by another specific CER (Section 3.3), however, the interstage adapter and/or the payload shroud, if applicable, as well as the guidance equipment. The CER derived from the reference points of Figs. 3 - 5 for ballistic stages is Fs = n • 5.0 • M ° ~ • f , [MY].

The "Scale factor" resulting from the CER shows that a vehicle of 10 times larger mass requires only 3 times higher cost, or, doubling the vehicle mass increases the cost only by some 40%. Manned ballistic vehicles follow the same law as shown in Figs. 3 - 5 , the relevant CER is as follows: Fs = n • 16.6 • M °'46 • f4 [MY].

The TRANSCOST-Modelfor launch vehicle cost estimationand its applicationto future systems analysis

809

MY

FIRST UNIT PRODUCTION COST

103

z

~

104

z

5

10n

2

5

108

2

i

107

Fig. 6. Launch vehicle stages unit cost vs stage net mass (without engines)--First flight unit. 3.4.2 W i n g e d v e h i c l e s ( r e u s a b l e ) . For winged space vehicles only few reference points exist: the Space Shuttle Orbiter production cost and the cost projected by Boeing for its Heavy-lift SPS launchers[5]. Due to the more complex structure and the expensive equipment of the crew compartment and associated safety measures, the specific and total cost of winged vehicles are considerably higher. A preliminary assessment as indicated in Figs. 3-5 leads to the following CER for manned winged Orbiter vehicles without tankage, using the Space Shuttle Orbiter as reference point: Fs = n " 54 . M

°'~.f4[MY].

For winged second stages with integral tankage the overall specific cost must be lower due to the larger share of (lower) tank fabrication cost. A preliminary assessment is Fs = n •

27 • M °-46 • f4 [MY].

For u n m a n n e d first stage fly-back vehicles lower specific cost can be assumed due to the lack of crew and cabin provisions and the relaxed thermal protection: Fs = n •

10 • M °'46 • f4 [MY].

This is twice as expensive as ballistic stages and considered to be realistic due to the higher structural complexity, the required landing gear and guidance equip-

ment, as well as the fly-back engines plus associated equipment and mechanisms.

4. FLIGHT OPERATIONS COST (Submodel 3)

4.1 Cost m o d e l structure

The direct operations cost, including launch management, prelaunch operations (assembly, checkout), launch and mission control, propellants, recovery and transportation represent some 5 to 10% of the total cost per launch in case of expendable vehicles (ARIANE 1 : 7%). For reusable vehicles the refurbishment cost have to be added, however, the hardware cost become negligible. Further the indirect operations cost (IOC) of the launch site (administration, safety, engineering support, facilities maintenance) have to be taken into account. These cost have been 1980 some 18 MAU for Kourou. For the NASA KSC (ETR) the equivalent in the 72-74 period was 90 Mio. $. The specific problem of a cost model for operations cost is the lack of a data base. In the area of development and production cost both the cost and the product are well defined by the industrial contract. In the operations area nothing comparable exists--one of the reasons that cost models of this type have never before really been established. Therefore, this model can only be considered as a preliminary first effort and it needs verification and improvement in the future. Figure 7 shows the scope of the flight operations submodel. The vehicle cost would come from submodel 2 if required, while the development cost amortization as

810

D. E. KOELLE Flight Operations Co~ Model

I

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J-[

Direct Operations Cost (DOC)

I

No, of Re-uses 1

I I ,o.o, [ Vehicle

t

Prelaunch Operations

Cm

Cc

Launches per year I

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Indirect I Operation I Cost (lOC) J

Refurbish- } ment ] Cost (RFC) I

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1 .eco.e

Mission Control

Propellant Cost

and Transport

Ce

Cp

Cb

Veh¢le q Cost iAmorti zationiI

I I

Charge

I d

Fig. 7. Scheme of flight operations cost submodel.

another possible cost factor is not shown here. It can however be introduced by submodel 1. The total operations cost model is defined as follows:

Co = DOC + Refurb.Cost + IOC + Add.Charges Co = (C, + Cc + Ce + Cp + C,) + Cp.F + Cxcc + Ca [MY per launch] (LpA) = Launches per Year). 4.2 Direct operations cost (DOC) As indicated in Fig. 4 these cost have been subdivided in 5 groups which are discussed in the following subsections. 4.2.1 Technical system management. It is assumed that these cost are a fixed amount per launch, independent from vehicle size, however, influenced by the number and type of stages. Assuming a certain basic value plus a figure specific for each stage (taking into account stage complexity and interface management) the following model can be established: C , = (5 + at + a2 + a3 + • . • a , ) L -°35(MY) at,z,3

=

dedicated values per stage ---expendable stages --reusable stages - - m a n n e d systems

a = 3 a = 4 a = 5

(MY) (MY) (MY).

This relationship means that the number of man-hours per launch is reduced to 50% for each increase of the launch rate by one order of magnitude.

Application of this model to Ariane 1 results in 8.6 MY per launch (4 LpA), equivalent to 0.51 MAU (79) or 10% of the launch operations cost. In case of a large-scale SPS launch operation of 1 500 flights per year with a winged two-stage vehicle to LEO we get 1.16 MY per flight which is equivalent to 104 K$ (79). Boeing uses a figure of 114 KS or 6.1% of the total launch operations cost (excluding hardware). 4.2.2 Pre-launch operations (assembly, checkout). These activities are very comparable in their characteristics to the previous category. A reference v#ue is 0.7 MS (74) for the DELTA launcher ( = 13 MY). The following cost estimation relationship (CER) has been established: C~ = (16 + bl + b~ + . , . b , ) L -°35 dedicated value for each type of stage - - s o l i d motor kick stage b = 6 (MY) ----expendable stages b = 12 (MY) --expendable LH2 stages b = 15 (MY) --reusable stages b = 20 (MY) - - m a n n e d vehicles b = 25 (MY). 4.2.3 Launch and mission control operations and software. The effort for these activities depends on the vehicle complexity, on the mission profile, and to a smaller extent on the annual launch rate L. The following CER has been defined: bl.2, 3 =

C, = (4 + d~ + d2 + d , ) ' L -°~5 (MY) dE2.. = dedicated values for each type of vehicle: ---expendable stages d t = 1 (MY)

The TRANSCOST-Modelfor launch vehiclecost estimationand its applicationto future systems analysis --cryogenic orbital stages --reusable first stages --unmanned reusable orbital stages --manned orbital systems

2 d~ = 2 dl = 4 d~ = 6 d I =

A certain verification of these CER's was feasible by the data provided in [3] for the DELTA launch vehicle. 4.2.4 Propellant cost (Fig. 8). The propellant cost per launch depend to a certain extent on the vehicle size (in contrast to the previous cost factors) and also on the annual launch rate. Both determine the required quantity per year and that is--according to the ground rules of the chemical industry--a major influence factor on the specific cost. The propellant delivery cost in the USA for quantities required in the Space Shuttle Program have been quoted as follows: 1967: 550 $/Mg LH2 and 27 $/Mg[12] 1977: 1 530 $/Mg LH2 and 44 $/Mg LOX (Boeing) 1979: 1 912 $/Mg LH2 and 54 $/Mg LOX[12] 1982: 3 238 $/Mg LH2[13] incl. 25% transportationcost. These values in MY plus some additional data are depicted in Figs. 4 - 5 showing the trend of cost reduction with increasing production rate. It is also shown that the LH2 production cost in Europe (Air Liquide and Linde) are much higher than in the USA. (Because of this fact the LH2 required at Kourou is shipped from the US.) Based on these data and assuming a mixture ratio of r = 6.0 a propellant CER can be established for LH2/ LOX:

Ce = 0.016. M e ' a ( M e "

L) 0~6[ MY ] - L~_]

with Me = total propellant mass of the launch vehicle (all stages) in Mg L = number of launches per year, a = boil-off rate The boil-off rate of LHz between propellant delivery and the actual launch are about 35% as indicated in [13] for the Space Shuttle. For LOX the boil-off rate is lower, perhaps 15 to 20%. The overall boil-offlosses, however, are mainly determined by the LH2 losses, due to the cost ratio between LH2 and LOX of about 25-35, reduced to 4 - 6 by the mixture ratio. Figure 8 shows the specific propellant cost per launch, related to 103 Mg propellants. If more than one launch vehicle is being supplied than the higher annual propellant mass required will reduce the cost of one program. 4.2.5 Recovery cost. This cost item refers (only) to reusable ballistic first stages and/or boosters which fall into the ocean and have to be recovered and towed back to the refurbishment or launch site. A study performed by the Company Harms in Hamburg for ESA/CNES in relation to the recovery of Ariane first stages resulted in cost of 1 Million DM (82) per retrieval for a frequency of 6 retrievals per year. This

811

would be equivalent to about 5 MY according to aerospace cost levels. The specific cost will be strongly influenced by the annual launch rate with a step function for an additional set of equipment and crew. As a first approach the following CER has been set up: C, = 6 + 7.0 L°'7 [MY]. This results in about 5 MY for a launch rate of 6 per year, and decreases to 3.8 MY for 12 LpA.

4.3 Refurbishment cost 4.3.1 Definition. The refurbishment cost are a special category related both to fabrication cost (of spare units) and operations cost. The refurbishment cost are only relevant for reusable vehicles. No real reference data exist yet about this type of costs for space vehicles. The number of flights per vehicle (or reuses) certainly will have an influence on the total refurbishment cost, i.e. the amount of refurbishment required may limit the vehicle lifetime. The possible and most economic number of flights for a reusable vehicle is still an open issue: the present range of assumptions is between 20 and 300, for rocket engines between 10 and 100.

4.3.2 Cost model. The following cost estimation relationships is structured such that the major contributions to the total refurbishment cost are separated. The numerical effort in each case can only be taken as a preliminary assumption, until more experience exists in this area. However, it is anticipated that at least a representative comparison for the relative refurbishment effort of different launch vehicles is feasible with this CER: C~ = N" (Rs + RG) + n RE + 2.5" lO-SrsFs + 5" 10-5 re FE 4,

(1!0 for winged fhst stage vehicles) N

=

number of stages

number of engines (total vehicle) Rs = stage system refurbishment effort MY R~ = guidance, control, checkout refurbishment effort Rr = rocket engine refurbishment effort F = fabrication effort (see section 3). The last two terms represent the material share of the refurbishment, i.e. the parts to be replaced, which will increase with the number of reflights. The values indicated assume exchange parts cost of 0.25% of the vehicle and engine fabrication cost per flight, with 100 reflights for the system and 50 per engine. Higher reflight numbers increase the percentage value and a lower number of reflights decreases the parts cost share per flight. n

=

812

D. E. KOELLE MY 103 Mg

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PRODUCTION RATE rMg/DAY~

Fig. 8. Propellant cost (LH2 and LOX) vs production rate.

The other definitions are:

Rs = 3 MY per stage for inspection, partial deassem-

R~ =

Re = The (10%)

bly/assembly and test operations of unmanned ballistic systems = 4 MY for manned winged systems 2 MY per stage for inspection, checkout and partial replacement of guidance, control and telemetry equipment for unmanned ballistic systems = 2.5 MY for manned winged systems 0.2 MY for inspection, control and refurbishment of each main engine. values include the required management activities and quality assurance (15%).

4.4 Indirect operations cost (IOC) 4.4.1 Definition. The indirect operations are those cost related to the launch base and flight range. In principle, these are fixed cost for - - s i t e management and administration --engineering support --facilities maintenance and amortization --spares storage and supply vehicles. However, there are secondary influences by the annual launch rate due to the number of facilities required for an increasing number of launches. 4.4.2 Cost assessment. Only few reference values are available, also the assessment is difficult because the IOC fee is a value fixed administratively. For a Delta launch 1974 an amount of 2.61 Mio. Dollar (or 48 MY) was charged, in 1972 a sum of 3.1 Mio. $ for an ATLAS-CENTAUR launch (70 MY).

Another way to assess these cost are the overall cost: 90 to 92 Mio. $ have been budgeted in FYs 72-74 for NASA-KSC (ETR), where 15 to 20 launches have been performed annually, resulting in an average figure of 4.5 to 6 Mio. Dollars per launch (90 to 120 MY). Regarding CSG Kourou the fixed cost for ARIANE are about 2 MAU (79) according to ESA (telex dated 11.6.80) for a launch rate of 3 per year. This equals 34 MY per launch and is lower than comparable US values. 4.4.3 Cost model. The actual IOC cost per launch depend on the size and type of launch site with the launch rate as the major factor. The number of launch pads and/ or vehicle preparation facilities required to achieve a certain launch rate is another factor to be taken into account. Since different approaches are taken for launch preparation and launch operations this total launch capability is taken as "launch site capacity" = L*, indicating the maximum number of launches which could be performed with the existing equipment and facilities. The staffing of the launch site needs to be adapted to the amount of launch activities, therefore, the IOC cannot simply be defined as a certain number divided by the launch rate. Figure 9 shows the model established with few reference points. The mathematical cost relationship is as follows: C.+lOC- -

4 0 . (L*) °'34 L055 MY per launch

L* = launch site capacity (LpA) L = actual launch rate (LpA).

The TRANSCOST-Model for launch vehicle cost estimation and its application to future systems analysis

7-17ll

MY/Launch~--

813

C 40 (L*) 0.34 IO = L 0.55

100~ 50

--

30

R

,5i~ 0 0 0 ' 0 0 ~

20 10

r~

--

1

3

5

10

20 30

50

100

200 300 500 1000 2000 LAUNCH RATE L (LpA)

Fig. 9. Indirect launch operations cost model.

4.5 Additional charges Launch cost can include a charge for vehicle fabrication cost amortization in case of reusable systems, or even an amortization charge for the development cost. A third type of charge would be a purely administrative charge of the responsible organization or company.

5. APPLICATIONTO A FUTURE EUROPEAN LAUNCHER (FEL or Post-ARIANE 4 Vehicle)

material (space processing) and it would mean an essential cost reduction to retrieve complete satellite platforms for a later reuse. While a semi-reusable vehicle is considered to be the first step of a new development programme, the final goal should be the fully reusable system, which would provide the transportation economy we need for the future commercial activities in space. As shown later there is a potential to reduce specific transportation cost/MAU/ Mg) to some 15% of present ARIANE 1 cost.

5.1 Vehicle requirements The expected growth in communication satellites as well as the economic competition will require a more capable and more efficient launch vehicle in Europe by the year 1995 at the latest. This means that the development should be initiated by 1985/86. Present preliminary understanding of the requirements is - - G E O payload capability 4 to 4.5 Mg - - L E O payload capability 15 to 18 Mg --Payload shroud diameter 5.5 to 6.5 m --Specific launch cost <50% of AR-44 --Application of the new HM.60 engine (900 kN vacuum thrust, 710 kN S.L.thrust). The basic concept is a two-stage launcher to LEO with one reusable stage. Reusability of the first stage would cut the launch cost to 50-60% compared to an expendable vehicle, while reusability of the second stage would reduce the launch cost to a minor extent, but allow •

retrieval of payloads from LEO,

a highly desirable capability. It is required for return of

5.2 Concept Candidates A variety of launch vehicle concepts can be conceived which fulfil the basic requirements. Ten different vehicle concepts have been analysed which can be grouped as follows: CONCEPT 1: Ariane 4 first stage + cryogenic second stage (fully expendable) CONCEPTS 2, 3, 4: New cryogenic core vehicle with boosters (P.8, L.40, P.150) CONCEPTS 5, 6, 7: Two-stage cryogenic ballistic vehicles CONCEPTS 8, 9, 10: Winged fly-back first stage plus different cryogenic second stages. Figure 10 presents a survey and Table 1 the stage characteristics and propellant mass values of the 10 concepts. All vehicles have been sized for 15 Mg LEO payload and 9 200 m/s total delta V capability (200 km orbit with launch from Kourou). CONCEPT 1 is the most conventional solution of an expendable launch vehicle. It is using an AR-44 L In'st stage with 5 Viking engines, 4 L.40 Liquid Boosters and a new cryogenic second stage (H.60). The launch mass

814

D.E. KOELLE CANDIDATE LAUNCH VEHICLE CONFIGURATIONS

FIG. 10

5

1

6

8

3

/

i

i

!

•

J

I

J

¥

Fig. 10. Candidate launch vehicle configurations. is 478 Mg, the payload volume and growth capability are marginal. CONCEPT 2 as an initial expendable version consists of a relatively large cryogenic core vehicle with 385 Mg propellants and 7.6 m diameter. In addition, 4 solid motor boosters (P 8) as used in the AR-4 program are required to achieve the necessary lift-off thrust level. The core vehicle net mass is estimated to 34 Mg including 4 HM.60 engines (cf. Fig. 3). CONCEPT 3 uses the same size core vehicle as described above, however, in a reusable configuration. This increases the net mass to some 50 Mg including 7 HM.60 engines. In this case 8 liquid propellant L.40 boosters are required to achieve LEO with the desired payload. The total lift-off mass of 810 Mg is relatively high, however, the booster concept allows the coverage of a wide payload range by using 4 to 12 booster units (3 to 20 Mg LEO payload). CONCEPT 4 is characterized by two large 250 Mg solid boosters, and a smaller core stage with 115 Mg propellants. It is under study by CNES as an expendable vehicle with potential recovery and reuse of the solid boosters. CONCEPT 5 is a two-stage vehicle with reusable first stage and an expendable second stage with H. 10 tankage (40 Mg propellants). This concept with 357 Mg launch

mass is not optimized in size in order to allow a followdevelopment to a fully reusable concept without change of the first stage (concept 5). A mass-optimized concept of this type would only have a lift-off mass of 275 Mg with 165 Mg propellants in the first and 65 Mg in the second stage. CONCEPT 6 is the fully reusable two-stage system with a 65 Mg propellant second stage (RH.65). This stage can be used for payload retrieval and it can land near or at Kourou. The fast stage is the same as in concept 4 and has 7 HM.60 engines. CONCEPT 7 is a variation of concept 6 with nonoptimum staging, assuming two equal-sized stages with 180 Mg propellants. In this case the two-step development scheme of concept 5/6 is abandoned in favour of a straight-forward approach towards a fully reusable vehicle with reduced development effort. The non-optimum staging leads to an increase in launch mass from 383 to 435 Mg. CONCEPT 8 is a winged fly-back first stage equipped with turbo engines for return to Kourou. The propellant load required for ascent is some 220 Mg, while the expendable second stage has some 60 Mg propellants (H.60). The total launch mass would be 367 Mg with some 280 Mg for the first stage. 7 HM.60 engines are required. The problem of this concept is the fact that it cannot

Table 1. Candidate future European launch vehicle data VEHICLE CONCEPT TYPE Booster Stage 1 Stage 2 Launch Mass (Mg)

1

2

4 × L.40 L.200 H.60

4 x P8 H.385 --

478

477

3

4

5

6

7

8

9

10

CRYO-CORE TWO-STAGEBALLISTIC WINGED 1st STAGE . . . . . . . . 8 × L.40 2 x P.150 RH.250 RH.250 RH.180 WH.220 WH.310 WH.550 RH.385 H.ll5 H.40 RH.65 RH.180 H.60 RH.90 WH.60 810

495

L = Liquid storable propellant stage H = Cryogenic propellant stage

357

383

435

367

RH = Reusable cryogenic stage WH = Winged cryogenic stage

506

780

The TRANSCOST-Modelfor launch vehicle cost estimation and its application to future systems analysis

815

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" i

8' ~

\

\ ~

TOTAL ~

COST OF

",,.

7.

~s

COST

,

,

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10

20

30

40

3,

I

50

60

10 MAU

VEHICLE COSTPER LAUNCH Fig. 11. Total cost of vehicle development and 150 launches for different vehicle concepts. be extended to a fully reusable system. The first stage would be too small, and a new vehicle development required for an improved follow-on system. CONCEPT 9 is a fully reusable approach with a flyback ftrst stage and a ballistic second stage, as presently studied by SNIAS for ESA. The fast stage has some 310 Mg propellants and a total mass of 380 Mg. The second stage has 90 Mg propellants and sits atop the flyback fn'st stage, requiring an expendable adapter. CONCEPT 10 is a vehicle with two winged stages, comparable to the initial NASA Shuttle studies of 1970/ 71. The relatively heavy second stage requires a large first stage propellant mass of 550 Mg, so that the launch mass reaches a value of some 780 Mg. 5.3 Vehicle cost The TRANSCOST model allows to perform an assessment of both development cost and cost per flight. Since the operations cost are similar for the two-stage concepts with the same payload, only vehicle cost have been compared, however, including refurbishment in ease of reusable stages. The first important criteria are development cost since Europe cannot afford, or would not be willing to spend any amount for a future launch vehicle (3 000 MAU might be the upper limit). The second important value are the launch cost, and the general trend is that a decrease in launch cost can be realized by increased development cost.

Figure 11 illustrates the development cost and vehicle cost (per launch) of the 10 concepts investigated (Fig. 10). The exact values are shown in Table 2. The stage cost have been calculated with the CER's of Section 3 and the following unit prices: 1 Viking engine 1.2 MAU (80) 1 VEGA engine (HM.60) 4.5 MAU (80) 1 Solid Booster (P8) 1.5 MAU (80) 1 Liquid Booster (L.40) 3.5 MAU (80). Only 10 uses have been assumed for ballistic stages, 50 flights for a winged fast stage and 25 flights for a winged second stage. The refurbishment rate per flight is assumed to be 3.5% of production cost and only 2% for the winged fly-back vehicle. These assumptions are relative pessimistic compared to US studies, and may lead later to lower cost than those shown in Table 2 for the reusable versions. The production learning factor has been taken into account with p = 0.92. For expendable vehicles a total quantity of 40 units has been considered (f4 = 0.70), for ballistic reusable vehicles = 6 units (f4 = 0.86) and for winged vehicles = 3 units (f4 = 0.91). 5.4 Conclusions The results, depicted in Fig. 11 show the expected trend of decreasing launch cost with increasing development cost. In order to arrive at a comparison of total programme cost of development investment plus the cost of 150 launches parametric curves have been introduced in Fig.

~With Booster r e c o v e r y and reuse.

TOTAL (MAU '80)

(1)

(5)

(4)

1.8

8.6 4.5

47.3

__

--

14"0 12.4 6.0

1400

TOTAL (MAU '80)

VEHICLE COST PER FLIGHT --booster --stage 1 --engines --refurbishment --stage 2 --engine(s) --refurbishment --management, integration

575 70 630 125

1 AR4 • 1st STAGE

DEVELOPMENT COST --engme(HM.60) --smge 1 --smge 2 --AIT

VEHICLE CONCEPT

(4)

(4)

---53.2

1.5

6"0 27.7 18.0

1875

575 1130 -170

38.4/10 31.5/10 2.2 2.3

28.0

39.5 (25-30) 1

(7)

(8)/

+ 1020

+50 + 870 -+ 100

-18.0 4.5 -1.7

10.0

34.2 (28-30) I

/

1900

575 320 835 170

2 3 4 CRYO-CORE VEHICLES

(1)

(7)

29.1

. 34.5/10 31.5/10 2.1 14.1 4.5 -1.8

3330

625 1880 524 301

(3)

(7)

.

15.3

. 34.5/10 31.5/10 2.1 22.2/10 12.5/10 1.1 2.0

+ 1010

--920 90

(2)

(8)

.

14.2

25.1/10 36.0/10 1.9 26.3/10 9.0/10 1.1 1.6

2700

625 1725 100 250

5 6 7 TWO-STAGE BALLISTIC VEHICLES

.

(1)

(7)

19.3

80.4/50 31.5/50 2.2 8.6 4.5 -1.8

6140

625 4330 630 555

(3)

12.1

92.2/50 45.0/50 2.7 23.1/10 9.0/10 1.5 2.0

7520

625 4610 1600 685

(3)

(12)

m

20.1

117.6/50 54.O/50 3.3 142.0/25 9.O/25 5.1 2.1

9620

625 5150 2970 875

8 9 10 WINGED FIRST STAGE CONCEPTS

T a b l e 2. D e v e l o p m e n t cost and v e h i c l e cost p e r flight for the c a n d i d a t e l a u n c h e r c o n c e p t s

rrl

oo

The TRANSCOST-MOdel for launch vehicle cost estimation and its application to future systems analysis 11. The resulting most economic solution with minimum overall cost is concept 7, a two-stage reusable ballistic vehicle with two equal-sized stages. Concept 4 is the second best concept from the total cost standpoint; however, it does not provide payload return capability. The winged fly-back candidates especially concept 9 show attractive launch cost values; however, the required development effort seems to be prohibitive high for Europe. The fact that a step-wise development from a semireusable vehicle to a fully reusable vehicle is not feasible since the first stage must grow in size substantially and results in a new vehicle, contributes to the fact that launchers of this type cannot be considered as a realistic alternative for a Post-Ariane 4 vehicle. The expendable concept 2 on the other side exhibits the lowest development effort, (comparable to 1 and 4) but has the potential of a later improvement to a reusable vehicle with payload return capability. The analysis has clearly shown the economic situation and the cost differences for a number of launch vehicle concepts, This is intended as a first preliminary overview. Much more work is required until a final decision for a launch vehicle concept can be made, taking also into account political conditions. It has been shown, however, that the investment into a new launch vehicle development is worthwhile: several billion A U can finally be saved and the market for space activities increased (lower launch cost will attract more customers). In addition, the value of retrieved spacecraft and material from orbit must be taken into account, which is feasible with a fully reusable vehicle (without additional investment into a separate recovery system). Payload return capability is essential for future industrial utilization of space. Europe should create this capability in combination with its next generation of launch vehicles. REFERENCES

!. D. E. Koelle, Statistisch-analytische Kostenmodelle fOr die Entwickburg und Fertigung von Raumfahrtger~it, Technical University, Munich, July 1971. 2. D. E. Koelle, Cost prediction of space projects. Proceedings 23rd IAF Congress, Vienna (1972). 3. D. E. Koelle, Raumfahrtforschung 17, 167 (1973). 4. D. E. Koelle, ESA Report No. TT-6, Dec. 1973. 5. D. E. Koelle, BPT 12/72. 6. D. E. Koelle and H. H. Koelle, Future low-cost space transportation system analysis, ActaAstronautica 6, 16351667 (1979). 7. D. E. Koelle, Performance and Cost Analysis for an

8.

9. 10. 11. 12. 13.

817

SSTO + OTV Heavy Cargo Transportation System to Geosync Orbit, IAF-Paper 78-A 27, 29th IAF Congress, Dubrovnik, Sept. 1978. E.W. Bormett, McDonnell-Douglas, A Cost History of the Thor-Delta Launch Vehicle Family, 25th IAF-Congress, Amsterdam, Oct. 1974. MBB Space Tug Study Report, Phase B, 1972. R. H. Nansen and H. DiRamio, Heavy Lift Freighters-A Transportation System of the Future, Boeing Co. AIAA Paper 78-316 (1978). D.E. Koelle, Cost Model for Space Transportation Systems and Operations (TRANSCOST), MBB-TN-RX1-328/B, 1983 Edition, Sept. 1983. R. Hess, H. H. Koelle and T. N. Berlin. The heavy space freighter, paper 79-50, Int. Astronaut. Cong., Munich (1979). Aviation Week, May 2, 1983, 86.

APPENDIX 1

List of Symbols C = total cost in MY C* = specific transportation cost (MY/Mg) F = fabrication/integration/test effort for a specific element in MY f~ = correction factor for the technical development standard (0.5 to 1.25) f2 = technical quality factor, specific for each type of system f3 = team experience factor f4 = cost reduction factor for series production H = development effort for a specific element in MY L = launches per year, launch rate (LpA) L* = launch site capacity (LpA) M = basic reference system mass in kg N = number of (different) stages per vehicle n = number of (equal) engines per stage or vehicle P = learning factor r = number of flight per vehicle or engine (life cycle) R = refurbishment effort for a specific element (MY) x = cost formula value indication the mass sensitivity (scale factor)

Subscripts: A c D e E F G 1 M m N O p P R S t

= = = = = = = = = = = = = = = = =

administrative and/or amortization charges prelaunch operations development launch and mission operations rocket engines (liquid props) fabrication guidance, control and recovery equipment launch and mission control activities manned vehicle stages management and mission software net mass operations propellant mass propellants reusable ballistic stages stage system transportation