Impact of mission requirements and constraints on conceptual launch vehicle design

Aerospace Science and Technology, 1999, no. 6,391-401

Impact of mission requirements and constraints on conceptual launch vehicle design

M. Rahn a*, U.M. Schijttle b, E. Messerschmid b

a MAN Technologie AG, Aerospace Division, RS, P.O. Box 1347, 85751 Karlsfeld, Germany b Space Systems Institute, Stuttgart University, 70550 Stuttgart, Germany

(Received 26 February 1999, revised and accepted 25 May 1999)

Rahn M., Schijttle U.M., Messerschmid E., Aerospace Science and Technology, 1999, no. 6,391401

Abstract The objective of this paper is to analyse the impact of mission requirements and constraints on both the optimum vehicle design and the effects on flight path selection for two types of reusable two-stage-to-orbit launch vehicles. The first vehicle type considered provides horizontal take-off and landing capabilities and is intended to be propelled by an airbreathing propulsion system during stage 1 flight. The second vehicle type assumes a vertical launch and is accelerated by a rocket propulsion system during the booster stage ascent flight. The analysis employs a design tool for simultaneous system and mission optimization. It consists of a CAD-based preliminary vehicle design tool, aerodynamic and aerothermodynamic calculation software, flight simulation programs, and a two-level decomposition optimization algorithm enabling simultaneous system and flight optimization. The results to be presented show that the cruise flight requirement for an European launched mission of the airbreathing vehicle results in a loss of 60 % payload mass as compared to a mere accelerated ascent for a near equatorial mission into the same target orbit assuming constant take- off mass. The strong dependencies of mission requirements on both the optimal vehicle design and the ascent performance are determined for the rocket-powered vehicle type by varying the inclination and altitude of the target orbit. 0 1999 Editions scientifiques et medicales Elsevier SAS. All rights reserved

system and mission optimization/conceptual vehicle design / airbreathmg vehicle / rocket vehicle / trajectory optimization I nonlinear programming

Zusammenfassung Einfhd3 von MissionsanforderPlngen und -bedingungen auf den konzeptionellen Entwurf von Raumtransportsystemen. In der vorliegenden Veriiffentlichung wird der EinfluD von Missionsanforderungen und -bedingungen sowohl auf den optimalen Fahrzeugentwurf als such auf die optimale Flugbahn fur zwei voll wiederverwendbare zweistufige Tmgersysteme analysiert. Zum einen wird ein horizontal startendes und landendes Fahrzeug betrachtet, dessen erste Stufe durch ein luftatrnendes Antriebssystem beschleunigt wird, zum anderen wird ein vertikal startendes Tragersystem mit Raketenantrieb in der ersten Stufe betrachtet. Die Analyse verwendet ein Programmsystem zur simultanen Missions- und Systemoptimierung. Letzteres bein- haltet Module fur den konzeptionellen Fahrzeugentwurf mit CAD, fur die Aero- und Aerothermodynamik, Modelle ftlr die Flugbahnsimulation und ein zweistufiges Gptimierungsverfahren fur die simultane Flugbahn- und Systemoptimierung. Die Ergebnisse der Analyse verdeutlichen, daft die Marschfluganfor- derung des luftatmenden Tragersystems fur eine Europamission eine Verminderung der Nutzlastkapazit8t urn 60 % zur Folge hat, verglichen mit einer reinen Beschleunigungsmission fllr einen Startplatz in der N&e des Equators bei ansonst identischen Randbedingungen. Ftir das zweite, raketengetriebene Fahrzeug wird die Abhlingigkeit des Fahrzeugentwurfs und der Nutzlastkapazitit von den ,Missionsbedingungen in Form von variierten Bahninklinationen und Zielorbithohen dargestellt. 0 1999 Editions scientifiques et medicales Elsevier SAS. All rights reserved

System- und Missionsoptimierung / konzeptioneller Fahneugentwurf / luftatnwndes l%igersystem / raketengetriebenes ‘Migersystem / F’lugbahnoptimierung / nichtlhware Programmterung

* Correspondence and reprints.

Aerospace Science and Technology, 1270-9638,99/O6/ 0 1999 Jbtions scientifiques et m~dicales Elsevier SAS. All rights reserved

392 M. Rahn, U.M. SchWIe, E. Messerscknid

Nomenclature

A, B coefficients for the mass model a acceleration [m/s 2] F thrust [N] h flight altitude [ml i inclination [deg] 1 length [m] m mass [kg] M Mach number SC scaling factor t time [s] 21 velocity [m/s ]

Symbols

d thrust level factor F mass fraction c7 mass split

Subscripts

0, E initial, end value 1, u lower, upper bound opt optimum per perigee Pl payload pr propellant ram ramjet engine ref reference rk rocket engine S structure st staging turb turbojet engine

1. Introduction

In the world-wide search for more cost efficient access to space, reusability of launch vehicles has been identified as the most promising approach [ 1, II]. Conceptual studies on future space transportation systems (STS) conducted over the past decades have emphasized winged vehicles for both single-stage (SSTO) and two-stage to orbit (TSTO) missions employing rocket and airbreathing propulsion. Depending on mission requirements with respect to operational versatility, the systems are designed for vertical or horizontal take off and horizontal landing (VII-IL, HTHL respectively). All concepts envisaged so far require advances in technology beyond the present state of the art. Currently, efforts are undertaken to identify, develop and demonstrate the enabling technologies [9, 10, 191.

The optimal design of launch vehicles and ascent trajectories is heavily dependent on numerical simulation and optimization techniques. Conflicting goals regarding mission requirements and constraints, flight path control, engine performance and weight, vehicle design and flight loads have to be matched by an appropriate system design and flight optimization strategy. Therefore, accu- rate modeling of vehicle properties and sophisticated software tools are required to predict the feasibility and capabilities of a launch system considered because of the small performance margins typically involved. For this purpose, an iterative analysis and design procedure for performance optimization of STS has been developed previously to support trade-off and sensitivity studies. The technique described in references [ 14, 151 simulta- neously defines system design variables and flight controls, minimizing the performance index while satis- fying preset constraints.

The objective of this paper is to apply the design optimization approach previously developed for two-stage launch vehicle concepts described in the literature [7, 81, in order to evaluate the effects of different mission requirements. Results omitting economic aspects will be presented for two space transportation systems that are designed for horizontal take off in the first case and for vertical take off in the second example. The former mission considers cruise capabilities of the launch configuration which the latter does not require. Both concepts however, impose a high cross-range requirement on the re- entering orbiter stages. With the exception of the carrier stage of the first example, which is powered by a hydro- gen fed turboramjet propulsion system, all vehicle stages use advanced LOX/LH2 rocket engines. Along the ascent and re-entry or flyback paths, respectively, typical inequality constraints on dynamic pressure, acceleration, aerodynamic load factor and heating rates are considered. Numerical simulation results extending earlier fin- dings exhibit a high sensitivity of mission constraints on vehicle design and performance. Finally, to put the results into perspective, a comparison with predictions of FESTIP-studies is provided.

2. Analysis method

The sequential multistep optimization scheme denoted STSOPT (Space Transportation System Optimization) has been applied in this study. It has been developed previously for space transportation system optimization. The general approach of STSOFT is depicted infigure I and is described in detail in references [13, 151, and hence will only briefly be summarized here.

The scheme comprises iteratively called vehicle design sequences, aero- and aerothermodynamic caicula- tions, vehicle design verification, and an optimization tool to account for interdisciplinary dependencies of an STS and its mission. The optimization tool sequentially

Impact of mission requirements and constraints on conceptual launch vehicle design Einfu$I von Missionsanforderungen und -bedingungen auf a!en konzeptionellen Entwqf von Raumtransportsystemen 393

cation and data- 7. Graphical

v&&n + interpretation

z.~~~~~~a a -on fCrnj i

t 3.Vehicle weight

estimation

sys@+ +nd%&t !r&&swion

t Geophysical & dynamic model

Figure 1. Scheme of the multidisciplinary design and optimization tool STSOPT.

performs numerical flight simulation, simplified vehicle scaling, and simultaneous system and flight optimization in an automated procedure. All of the seven program blocks shown (figure Z) are separate tools controlled by the central data-manager of STSOPT. This manager also handles the data transfer required for interprocess com- munication between the various program tools which is based on a file-system database. According to figure 1 the approach can be stated as follows:

1. Determine the mission of the STS to be analysed and select propulsion system properties based on the available database. The database also contains aerodynamic characteristics of reference configurations and weight data for a large variety of launch vehicles.

2. The vehicle design is accomplished by means of the CAD software denoted Pro/ENGINEER [12]. This CAD-system is a 3D volume modeler and designs com- ponents as solid volumes and surface models, respectively, unlike traditional 2D CAD-systems (e.g. ACAD) which create drawings of a model. The main advantage of this approach is that the designed models are fully parametric, i.e. they are described mathematically and dimensions can be assigned for later manipulations. In addition, the Pro/DEVELOP extension of Pro/ENGI- NEER allows the software developer to extend the origi- nal Pro/ENGINEER functionality by arbitrary routines written in the computer language C. Also, the parametric database of a designed part can be accessed, existing menus be extended, and new ones be included into the Pro/Engineer working area.

Within the frame of the STSOPT program-tool the CAD has to accomplish three vehicle design tasks: Firstly, the parametric design of a new vehicle model plus determination of the main lineal, area, and volume dimensions of tanks, cabin, and vehicle contour; Secondly, the modification task to accommodate an exis- ing vehicle model to propellant and wing area requirements; Thirdly, to create a structured surface grid serving as input for the aerodynamic module.

3. The weights module uses statistical data from existing and projected launch vehicles which are fitted using the method of least squares, to define empirical weight equations. The user can take into account different levels of technology assumptions by specifying technology parameters employed in the mass models. These use statistical mass correlations of the form

(1) i=l i=l

where rni is the subsystem mass of a component i, Ai, Bi are the empirical coefficients defining the technology to be employed, and xi represents the function of physi- cal properties influencing the component mass. In total, weight relationships for more than two dozen subsys- tems of a launch vehicle are used.

4. For the estimation of the aerodynamic vehicle properties, simplified engineering tools are used. From sub- sonic to low supersonic speeds (up to Mach 3 - 4) the panel method is employed. In the hypersonic flight regi- me (Mach 14 - 5) the modified Newtonian and Prandtl- Meyer expansion flow theories are used along a paneli- zed vehicle surface model to calculate the global aerodynamic coefficients [5].

The aerodynamic heating analysis follows the classi- cal path in that the vehicle considered is represented by sphere, cone, cylinder, and flat plate shape elements for which semi-empirical solutions are available. For the purpose of trajectory shaping during flight performance optimization the heating environment is monitored by a stagnation point heat flux model.

5. The simultaneous mission and design optimization tool is based on a parallelised two-level decomposition algorithm described in detail in reference [14]. It employs a point-mass approximation for flight simulation which is described by a set of differential equations defined in a relative velocity coordinate system to define the motion in three degrees of freedom. The optimization tasks of the decomposition algorithm are formulated as nonlinear programming problems. A sequential quadra- tic programming method NLPQL [17] that is supported by a gradient projection method (GPA) [ 181 is employed to solve the optimization tasks.

6. During the course of optimization an automatic vehicle design scaling is required to accommodate varying propellant and payload volumes with respect to the vehicle reference design previously defined in step 2. The sizing task is simplified, assuming the vehicle to shrink or to expand uniformly, i.e. photographic scaling. It requires the utilization of two scaling factors for lineal, area, and volume dimensions scaling of a winged launch system. These are defined as follows:

- the volume factor (SC,~) relates the propellant and payload volume requirements of subsequent iteration steps by

1999, no. 6

394 M. Rahn, U.M. Schiittle, E. Messerschmid

fuselage volume of the i-th iteration volume of the reference configuration (2)

- the aerodynamic area factor (Scar) aims to keep the specific wing loadings constant:

SC,, = d- aerodyn. area of the i-th iteration

aerodyn. area of the reference configuration

(3) The geometrical properties of the iterated vehicle are

modified by the corresponding scaling factor depending on their affiliation to fuselage and wing features of the geometric vehicle model, respectively.

In the vehicle design verification process, the user has to decide whether or not to continue the iterative steps 2, 4, and 5. This decision depends on the accommodated vehicle scalings after the optimization step. Thus, the difference between both the iterated and the reference vehicle serves as decision criteria for the expert user. Ideally, both scaling factors are 1.0 for the converged solution. If not, a vehicle redesign is required in the CAD-based vehicle design module according to actual payload, propellant, and aerodynamic area requirements. Then, a new geometric reference dataset and panel model of the vehicle is generated for further processing. Also, the re-evaluation of the aerodynamic coefficients (held constant in the optimization part) is performed, and the optimization tool must be restarted.

7. The programs required for graphical representation of calculated results are supplied by STSOPT in the visualization tool box.

3. System and mission optimization of the airbreathing HTHL vehicle concept

In this first application section the airbreathing launch vehicle is examined firstly for payload optimization and secondly for minimization of the take-off mass at the specified payload capability, for the same target orbit. The mission example selected assumes an accelerated ascent from the near equatorial launch site in Kourou, the launch site of the European Ariane launchers, for payload delivery into a low earth orbit (463 km) at 28.5 deg inclination. In comparison with the results of a study [ 151 that has been conducted earlier for the same vehicle concept and target orbit, but a different launch site, Istres in Europe, these results will point out the influence of mission requirements on vehicle design.

3.1. Vehicle concept

A comprehensive description of the reference vehicle concept that is required for the vehicle design approach

of the STSOPT software, is given in references [2, 81, and so will only briefly be summarized.

The horizontal take-off and horizontal landing TSTO reference concept, depicted infigure 2, is of Sanger type. It consists of a manned hypersonic stage that is powered by five LH2 fuelled turboramjet engines. The second winged stage considered, attached to the booster stage until staging, will continue the ascent flight into the orbit after staging. Its main propulsion system consists of an advanced high pressure LOX/LH2 rocket engine. The total launch mass mg = 366 Mg is split up into mass fractions of rno1 = 251 Mg for the booster stage and rn02 = 115 Mg for the orbiter stage.

Figure 2. The airbreathing launcher of type Stinger.

3.2. Optimization results

Table I compares characteristic results of the simultaneous system and flight optimization of the airbreathing launch system. The results listed for both Istres missions are taken from a study that has been described in reference [ 151. A payload improvement of roughly 30 % from mpl,ref = 2.2 Mg to 2.9 Mg could be found for the Istres mission, which is mainly caused by an extension of the airbreathing flight phase of the booster vehicle until staging conditions are reached at Mach number Mst = 7.4 in h,t = 31.5 km altitude. This flight envelope extension requires enhanced ramjet engines indicated by the increase of the installed engine thrust ratio &ar,, by the maximum allowed value of + 30 %. The engine scaling restriction has been chosen to avoid inadmis- sible results due to unreasonable engine design while assuming a linear scaling of the propulsion system performance datasets.

A change of the launch site from Europe to the near equatorial Kourou site allows a payload mass increase of 62 % to mpl = 4.7 Mg if the payload mass is optimized for constant take-off mass, and a take-off mass reduction of 19 % to mu = 296.4 Mg if the lift-off mass is minimized, respectively.

Two remarks are in order regarding the optimization results. First the mass split (cr = m02/mo) of the launch

Aemspucr Science and Technology

Impact of mission requirements and constraints on conceptual launch vehicle design EinjluJ von Missionsanforderungen und -bedingungen auf den konzeptionellen Entwurf von Raumtramportsystemen 395

Table I. Ontimization results for the airbreathing TSTO vehicle con&p t.

Launch site

mQl Mzl

Kt F-1 hs, [km1

4 rurb L-1

rprm L-1 @rk L-1 MS, L-1

mprl [Mgl mpr2 Msl 0 L-1

m02 [WA m0 Pbl

-SO -45

longitude [deg]

Figure 3. Optimal trajectories of the airbreathing vehicle concept for a launch in Kourou.

1 m. = constant; * mpl = constant i. 1”’ LL.,.. fuel n oxidizer m payload

system for both Kourou missions has been found optimal at (T = 0.41 in contrast to CT = 0.31 for the Istres designed vehicle. This results in an orbiter mass increase of 32 % from mu2 = 114.3 Mg for the Istres vehicle design to mn2 = 150.8 Mg for the Kourou one. Second, the optimal staging conditions for the Kourou examples decrease to A4,t = 6.6 in hst = 30.7 km and I& = 6.4, in h,t = 29.6 km, respectively, although the ramjet- engine design parameter thrust level has been increased to q&m = 1.3.

The absent cross-range condition of the booster stage for the Kourou vehicle design in conjunction with its flyback to the launch site requirement are responsible for both effects. In the ideal case, the vehicle potential and kinetic energy should be sufficient high at the staging conditions to let the booster stage return to the launch site by a gliding flight, without any propulsion support. Since the shutting down of the airbreathing engines is not recommended due to security reasons, the engine throttle settings should be reduced to a minimum setting in-stead. Therefore, the lower staging Machnumber cau- sing the increased mass split of the vehicles, is the ans- wer of the optimization process to define the optimal booster flyback conditions. The same remark applies to the location of the stage separation as shown injigure 3 depicting the optimal flight path for the payload optimized vehicle.

The optimization process determines two possible flight controls providing the same payload mass at the same mission and system parameters (M,t, hst, q5, 0). The two solutions basically differ in reaching orbit via a 1999, no. 6

case 1: design for the lstres mission lift-off mass m,, = 366.OMg, m,, = 2.9 Mg

case 2: design for the Kourou mission lift-off mass m. = 366.0Mg, md = 4.7Mg

ml

case 3: design for the Kourou mission , = 2.2Mg, m. = 296.4Mg

1 [ml 1 [ml

Figure 4. Optimal vehicle design for the three mission examples considered.

396 M. Fbhn, U-M. S&We, E. Messerschmid

flight over the northern or southern Earth hemisphere. According to the statements above, it is not astonishing that the airbreathing winged vehicles do not perform a planar ascent. Unlike conventional rocket launchers the airbreathing vehicle is heading almost westwards at take- off, which requires a more or less distinct turn prior to staging. Subsequent to separation the booster stage turns back to the launch site, while the upper stage continues its ascent to perigee of a 90 x 463 km transfer orbit into the desired 463 km target orbit.

For all three system optimization examples considered in table Z, the turbojet and the orbiter stage rocket propulsion system design of the reference STS has been confirmed. Both propulsion scaling factors are held almost constant according to their initial values $turb = f#+k = 1.0, providing an initial acceleration capability of a = 0.6 g for the booster and a = 1.3 g for the orbiter stage. The same applies to the transition Machnumber IV&, which determines the event of swit- ching from turbojet to ramjet engine mode.

3.3. Optimal vehicle design

Different missions require different optimal vehicle designs of the same vehicle concept since the calculated propellant and payload masses of the optimal trajectories are the main parameters influencing the vehicle design step. As demonstrated in figure 4, the mission requirement of a launch in Kourou results in a smaller fuselage design since the propellant requirements are less for the booster stage compared to the Istres vehicle design (table r). For the payload optimized variant, a fuselage length of 1 = 72.6 m has been found optimal, that is, 11.5 % shorter than the fuselage of the optimal Istres vehicle launched at Istres ( 1 = 82.0 m). A further length reduction to 1 = 66.4 m has been determined for the example minimizing lift-off mass. The wing design has been adjusted in the course of system optimization in order to accomplish the requirement for constant wing loads in both payload optimized and lift-off mass minimized design alternatives. The payload maximized upper stage design has been enlarged by about 10 % to accommodate the increased payload bay and propellant tanks while the lift-off mass minimized orbiter stage remains almost unchanged compared to the Istres design.

Detailed information on the individual subsystem weights is provided by figure 5 for the booster stage and

figure 6 for the orbiter of the three optimum vehicle configurations described. The influence of the modified vehicle dimensions, wing areas, propellant, and payload bay volumes on the three mass distributions can clearly be identified. It should be remembered that for the booster stage the payload mpl comprises both propellants, structure and payload of stage 2.

launch mass [Mg] dry mass [Mg] m,,, = 366.0 msr = 161.6

Figure 5. Comparison of vehicle mass predictions of the airbreathing booster stage for the missions considered (indices l-3 refer to system variants considered).

launch mass [Mg] dry mass [Mg]

6 payload mass [Mg]

Figure 6. Comparison of vehicle mass predictions of the orbiter stage for the missions considered.

4. System and mission optimization of the VT&IL racket vehkle concept

In the second application example a fully reusable TSTO rocket vehicle of given GLOW is considered for payload optimization. All missions specify vertical launch in Kourou but have different target orbits in terms of altitude and inclination, in order to evaluate their impact on both vehicle design and ascent performance. The present study comprises a sensitivity study and, hence, extends the previous work described in reference

Aerospace Science and Technology

Impact of mission requirements and constraints on conceptual launch vehicle design Einjluj’ van Missionsanforderungen und -bedingungen auf den lconzeptionellen Entwu$von Raumtransportsystemen 397

[6] defining a reusable launch vehicle for a mission to a circular target orbit at h = 463 km and i = 28.5 deg inclination.

4.1. Vehicle concept

The reusable winged rocket vehicle examined is part of an Ariane-X launcher family sketched in figure 7. It has been designed by an evolutionary strategy, aiming for the development of a fully reusable winged launch system based on an Ariane-V extended expendable launcher, see reference [7]. The vehicle considered (figure 7, right hand side) consists of a winged booster and orbiter stage. The first stage is able to return to the launch site after staging by the use of two wing mounted turbojet engines. A payload ratio of ~~1 = 1.29 % has been pre- dicted for the vehicle of mg = 368.2 Mg GLOW and a stage ratio of D = m02/mo = 0.28.

- 70 m

- 60 m

- 50 m

-4Om

- 30 m

- 20 m

- 10m

- Om

Figure 7. The Ariane-X launcher family of reference [7].

4.2. Optimization results

The results of the study in reference [6] have shown that the evolutionary design strategy results in a payload loss of 28 % compared to an optimally redesigned vehicle for the reference orbit. These previous calcula- tions considered the booster stage fly-back only by the inclusion of a constant propellant mass in the vehicle design step. In contrast, the present study fully considers the return of the first stage to the launch site in the simultaneous system and flight optimization process. Now a slightly lower payload mass of mpl = 6.0 Mg has been calculated for the reference orbit compared to mpt = 6.11 Mg of reference [6]. This is due to the slightly increased propellant requirement for the flyback of the booster stage of mpr = 5.9 Mg compared to the assumed mpr = 5.8 Mg in [6]. However, our results confirm the fuel assumption selected in [6] for the flyback segment of the booster stage.

1999, no. 6

Figure 8 shows the sensitivity of the payload, express- ed as specific payload ratio pSpeC = mpl/mpl,ref, as a function of the orbit altitude and inclination for the At&e-X vehicle. The optimized vehicle design of the reference mission serves as baseline and its payload rnpl,ref = 6.0 Mg indicates a ratio pLspeC = 100 %. Rvo bunches of curves are pictured. The solid curves repre- sent the payload ratio for the vehicle designs that have been optimized for the specific missions considered, while the dashed lines indicate the payload ratio calculated at constant reference vehicle design for optimal ascent trajectories. Both bunches show the expected behaviour, in that the payload decreases with increasing orbit altitude and inclination. Also, they illustrate the importance of proper vehicle design point selection since payload losses are increasing the greater the deviation from the specified design point.

Optimal ascent trajectories into a 463 km target orbit for the three inclinations 5.3 deg, 28.5 deg, and 97.5 deg are depicted in figure 9. The typical steep ascents of

100 500 1000 2000 orbit altitude [km]

Figure 8. Specific payload ratio of the Ariane-X concept as a function of orbit altitude and inclination.

-50 -45 -40

longitude [deg]

Figure 9. Optimal trajectories of the Ariane-X launcher in the 463 km orbit for three inclinations 5.3, 28.5, and 97.5 deg.

398 M. Rah, U.M. Sch(ittle, E. Messerschmid

rocket launchers are illustrated, as well as the overshooting transfer behaviour of the optimal orbiter flight. In addition the flyback segment of the booster stage is illustrated for the mission of 28.5 deg inclined orbit showing a pronounced climb after staging following a steep des- cent and the powered turning flight with final heading to the launch site.

The impact of the target orbit altitude on optimal vehicle design is shown in figure 10 for missions of constant inclination i = 28.5 deg. It depicts five missions and system parameters optimally determined in the course of optimization, including the resulting mass split 0 = m&mo. These system parameters essentially affect the stages propellant consumption and thus the optimal mass split of the vehicle. The parameters com- prise the staging velocity uSt and height hSt, the engine scaling factors f&t and f&2 affecting the main engines thrust to weight ratio of both stages and the perigee altitude hper of the transfer orbit to the circular target orbit.

The system reacts to higher target altitude requirements by an extension of the first stage flight envelope indicated by the increase of the staging velocity and altitude from uSt = 2.92 km/s in hst = 61.7 km for the 200 km target orbit to ust = 3.29 km/s in hst = 65.2 km for the 2000 km orbit. As a result, the optimal vehicle mass split decreases and finally amounts to (T = 0.198. The two engine scaling factors 4 show an almost

1.50

!I 1.25

1 .oo

0.75

0.50 I

I , 0.24

- b-j 0.20

1 0.18

li,,i,] , 0.16

1 500 1000 2000

100 1 1 4.2

80 - - 3.4 *--_

-He 70 - VSf ___--- _--- --

- 3.0 h ,\I -I

60 100

_^ ..,.. x,1 I 2.6

500 1000 2000 orbit altitude [km]

Figure 10. Influence of orbit altitude on optimal system design for the Ariane-X vehicle.

constant trend with a slightly increasing tendency for the booster stage engine factor C&l and a slightly decreasing curve for the upper stage propulsion scaling factor &k2. Its mean value for the booster stage $rkl = 1.26 causes a strong thrust to weight ratio of 1.63 g at lift-off since a ratio of f&k1 = 1.0 indicates a ratio of 1.30 g. The engine scaling ratio &a = 0.88 effects a moderate thrust to weight ratio of 0.93 g for the winged upper stage after staging. This relatively low acceleration capability of the orbiter in conjunction with the low altitude of the transfer orbit perigee causes the overshooting trajectories observed infigure 9. Thus, it seems to be optimal to have a downscaled and therefore lighter propulsion system for the upper stage at the cost of a longer time required for the remaining ascent flight after stage separation.

4.3. Vehicle design

Table II shows representative results of the vehicle design optimization for the reference mission considered (h = 463 km and i = 28.5 deg) and the optimal vehicle design for a i = 97.5 deg, 700 km circular polar orbit in comparison to the initial vehicle design of the evolutionary Ariane-X concept of reference [6]. The lift-off mass of the optimized orbiter stage decreases and is rno2 = 82.1 Mg for the reference mission considered, compared to rn02 = 103.1 Mg for the evolutionary design due to an extended booster stage flight envelope [6]. Thus, smaller fuel and oxidizer tanks are required and

Table II. Comparison of A&me-X orbiter stage vehicle designs.

T

Fuel tank [m3]

Oxidizer tank [m3]

Payload bay [m”]

Wing pl. area [m2]

Wings mass [Mg]

Fuselage mass [Mg]

TPS 1Mgl Landing gear [Mg]

Prop. system [Mg]

Total net mass [Mg]

Payload mass [Mg]

Lift-off mass [Mg]

Ariane-X Orbiter Stage

Reference Optimal Optimal design 1 design 1 design2

158 114 104

56 40.3 29.3

68 84.5 59.2

167 157 116

2.7 2.4 1.8

4.8 3.8 3.5

5.9 5.3 4.4

1.1 0.97 0.74

2.6 2.1 1.9

22.5 19.6 16.7

4.7 6.0 3.9

103.1 82.1 72.1

‘mission: 463 km x 28.5 deg; ’ mission: 700 km x 97.5 deg.

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Impact of mission requirements and constraints on conceptual launch vehicle design Einjlu. von Missionsanfonlerzutgen und -bedingungen auf den konzeptionellen En&w-f von Raumtransportsystemen 399

result in a smaller and therefore lighter orbiter design, although the payload bay ca

!i acity has been increased by

about 25 % to VP1 = 84.5 m . Compared to the fuselage, the wing redesign does not effect its weight to the same extend since they are modeled according to constant landing wing loads. In addition, payload delivery back to earth is assumed in the design. Thus, the sum of the total net mass and payload mass exercises a dominating influence on the wing design.

The flight and system optimization results for the 700 km sunsynchronous orbit at 97.5 deg inclination exemplifies the influence of the mission selected on the optimal Ariane-X vehicle design. According tofigure 10, the booster flight is further extended compared to the optimal flight envelope of the reference mission, resulting in an orbiter stage lift-off mass rno2 = 72.1 Mg. The lower initial mass together with lower payload capabilities mpl = 3.9 Mg due to the higher target orbit explains the lower net mass m, = 16.7 Mg of the upper stage. The ascent performance result for the reference vehicle design into the 700 km polar orbit amounts mpl = 2.1 Mg (figure 8) and hence, 46 % less than for the mission optimized vehicle.

5. Comparison of results with concepts of the F’ESTIP study

In order to put the present results into perspective with studies within FESTIP (Future European Space Tran- sportation Investigation Programme), a comparison is provided. Table III summarizes major characteristic data

for the two vehicle concepts evaluated in the present study, and two concepts denoted FSS-9 and FSS- 12 examined in the FESTIP-study [9]. The TSTO FSS-9 vehicle concept consists of a reusable rocket state that carries a reusable rocket second stage in a parallel mode. Similar to the Ariane-X concept, it is designed for vertical launch and horizontal landing (VTHL). However, the FSS-9 concept differs in that it allows all engines run- ning at lift-off supplied by only the first stage propellant tanks (cross-feed-ing). The FESTIP FSS-12 design is an airbreathing TSTO concept comparable to the Stiger type evaluated in the present study. It differs in the propulsion concept considered, assuming turbojet and rocket engines for the first vehicle stage.

Comparing the payload ratio ~~1 = m&m0 of the two VTHL concepts it is noted that, especially for the 250 km polar orbit, considerable differences are evaluated, ~~1 = 0.95 % for Ariane-X and ~~1 = 0.38 % for FSS-9. The difference decreases for the near equatorial orbit at 5 deg inclination (1.76 % for Ariane-X and 1.61 % for FSS-9) but still shows a better payload ratio for the tandem staged Ariane-X concept. However, seve- ral remarks are in order regarding these results. The lift- off acceleration amounts to 1.4 g for FSS-9 compared to 1.6 g for Ariane-X resulting in lower gravity losses. In addition, the parallel staging concept FSS-9 experiences increasing drag losses during ascent compared to the tandem staging Ariane-X concept. Both items yield increasing propellant requirements for the FSS-9 concept. Another important point concerns the depth of modeling of the vehicles considered. Here, the accuracy of the FSS-9 vehicle model exceeds the Ariane-X model in

Table III. Comparison of major characteristic data for the FESTIP concepts FSS-9 [16] and FSS-12 [3, 161 and concepts Ariane-X and S&nger-Type.

mpl (250 km circ. x 98”) bfgl q(250 km circ. x 5”) [Mgl mpl (463 km circ. x 28.5”: WI m0 WI M0.i PM Mdry,i PM M pr0pellant.i Pkl M tank,! Pkl Mengines.i [Mid F .T”gl”eS WI

F-1

VTHL Rocket concept HTHL Airbreathing / Rocket concept

t

FFSTIP FSS-9 Ariane-X booster orbiter booster orbiter

521.3 394.4 53.4 329.8 7.0 15.0

4 x 1716.7

9.03w 9.44(3

2.0 8.4

126.9 32.5 86.0 2.0 4.9

1 x 1787

368.2 268.1 45.1 241.1 7.7 13.0

3 x 2454

9.4

3.5 6.5 6.0

82.1 19.6 56.6 2.1 2.1

1 x 754

FESTIP FSS-12 Sgnger-Type(‘) 1 st stage 2nd stage 1 st stage 2nd stage

3.3 7.0

4.7 389.8 269.8 142.4 127.4

120.0 23.8

2 x 825

150.8 31.8 113.4 1.9 3.8

x 1920

6.6

366.0 215.2 152.1 63.1 6.7

44.2 5 x 420c4) 5 x 350(5)

6.6

1

(I) Kourou mission; (*I polar orbit; c3) equatorial orbit; c4) turbojet engines; 0) ramjet engines.

1999, no. 6

400

terms of aerodynamics, mass, propulsion, and aerother- modynamics. Thus, the mass model of the Ariane-X upper stage (a = 0.24) may be too optimistic compared to the more conservative model of the FSS-9 orbiter stage (a = 0.26). Finally, the present FSS-9 configuration has not yet achieved a fully converged design status [4]. The staging conditions exemplify this fact, since they are held fixed for FSS-9 on the basis of an engineering assessment of the ascent performance, in contrast to the system and mission optimized Ariane-X vehicle design presented in this study.

The ascent performance results and vehicle design data published in reference [3] for the FSS-12 concept allow only limited comparison with the results of the simultaneous flight and system optimization of the present study. They both indicate similar payload ratios of approximately ,+,I = 1.4 - 1.8 % for low inclined orbits at different staging mass splits. An optimal value of m&mo = 0.41 has been found for the S&ringer-type vehicle in contrast to moz/rno = 0.31 for the FSS-12 concept. This is mainly caused by the additional propellant mass required for the partly rocket thrusted booster stage of the FSS-12 vehicle although both concepts as- sume staging at the same Machnumber M,t = 6.6.

6. Conclusions

The performance optimization of two types of reusable two-stage to orbit launch vehicle has been conducted in the present study to evaluate the impact of mission constraints on conceptual launch vehicle design. For this purpose, the sequential multistep optimization scheme STSOPT comprising CAD-based conceptual vehicle design tools, aerodynamic and aerothermodynamic calculation software, flight simulation programs and a two- level decomposition optimization algorithm enabling simultaneous system and flight optimization has been applied.

The results for the airbreatbing TSTO vehicle concept presented identify the strong influence of the missions selected on both the optimal flight path and vehicle design. It has been found that a change of the launch site from Europe to the near equatorial Kourou site allows a payload mass increase of 62 % to mpl = 4.7 Mg if the payload mass is optimized for constant take-off mass, and a take-off mass reduction of 19 % to mg = 296.4 Mg if the lift-off mass is minimized.

The sensitivity analysis of the TSTO rocket propelled Ariane-X vehicle design with respect to target orbit altitude and inclination demonstrates the importance of proper vehicle design point selection, since payload losses increase the greater the deviation from the specified design point. This means a vehicle design point of approximately i = 20 deg for missions in the range of inclinations 0 5 i 5 40 deg and i = 60 deg in the range 40 < i I 80 deg is recommended.

M. Rahn, U.M. SchWle, E. Messmdnid

The final comparison with results derived from the FESTIP study confirms the fact that a proper vehicle performance comparison depends strongly on similar simulation model assumptions since the mathematical description models significantly affects the optimization results.

Acknowledgment

The support of this work by the Deutsche Forschungsgemeinschaft as part of the Sonderforschungs- bereich 259 “High Temperature Problems of Reusable Space Transportation Systems” at Stuttgart University is gratefully acknowledged.

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1999, no. 6

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Impact of mission requirements and constraints on conceptual launch vehicle design