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As part of phase 2 of the X-33 Program, NASA selected an integrated lifting body/aerospike engine con guration as the study vehicle for the conceptual analysis of a single-stage-to-orbit reusable launch vehicle. A team at NASA Langley Research Center participated in the screening and evaluation of a number of proposed vehicle con gurations in the early phases of the conceptual design process. The performance analyses that supported these studies were conducted to assess the effect of the vehicle's lifting capability, linear aerospike engine, and metallic thermal protection system on the weight and performance of the vehicle. These performance studies were conducted in a multidisciplinary fashion that indirectly linked the trajectory optimization with weight estimation and aerothermal analysis tools. This approach was necessary to develop optimized ascent and entry trajectories that met all vehicle design constraints. Signi cant improvements in ascent performance were achieved when the vehicle ew a lifting trajectory and varied the engine mixture ratio during ight. Also, a considerable reduction in empty weight was possible by adjusting the total oxidizer-to-fuel and liftoff thrust-to-weight ratios. However, the optimal ascent ight pro le had to be altered to ensure that the vehicle could be trimmed in pitch using only the ow diverting capability of the aerospike engine. Likewise, the optimal entry trajectory had to be tailored to meet thermal protection system heating rate and transition constraints while satisfying a crossrange requirement. NomenclatureC L = lift coef cient I sp = speci c impulse, s M e = edge Mach number M 1 = freestream Mach number O=F = total oxidizer-to-fuel ratio q = dynamic pressure, psf q ¢ ® = dynamic pressure times angle-of-attack,psf ¢ deg Re µ = momentum thickness Reynolds number S = aerodynamic reference area, ft 2 T =W = thrust-to-weightratio .T =W / eng = engine thrust-to-weightratio W = entry weight, lb W empty = empty weight, lb W ins = inserted weight, lb X=L = body position over vehicle length ® = angle of attack, deg 1 payload = change in payload from

.jectories that met all vehicle design constraints. Significant improvements in ascent performancewere achieved when the vehicle flew a lifting trajectory and varied the engine mixture ratio during flight. Also, a considerable reduction in empty weight was possible by adjusting the total oxidizer-to-fuel and lifloff thrust-toweight ratios. However, the optimal ascent flight profile had to be altered to ensure that the vehicle could be trimmed in pitch using only the flow diverting capability of the aerospike engine. Likewise, the optimal entry trajectory had to be tailored to meet TPS heating rate and transition constraints while satisfying a crossrange requirement.

Design of the thermal protection system for any hypersonic ight vehicle requires determination of both the peak temperatures over the surface and the heating-rate history along the ight pro le. The process used to generate the time-dependent aerothermal environments for design of the X-34 thermal protection system is described. The process evolved from a relatively simplistic approach based solely on engineering methods applied to critical areas to an integrated approach using ground-test data with detailed computational and engineering methods to predict the aeroheating over the entire vehicle. A brief description of the trajectory development leading to the selection of the thermal protection system design trajectory is included. Comparisons of engineering heating predictions with wind-tunnel test data and with results obtained using a Navier-Stokes ow eld code and an inviscid/boundarylayer method are shown. Good agreement is demonstrated among all of these methods for both the ground-test condition and the peak heating ight condition. Nomenclatureh = heat transfer coef cient, Btu-s/ft 4 or lbm/ft 2 -s L = reference length, in. M = Mach number P q = heat transfer rate, Btu/ft 2 -s Re = Reynolds number x = axial location, in. y = spanwise location, in. ® = angle of attack, deg ± = control surface de ection, deg " = emissivity Subscripts BF = body ap cl = centerline ( ) el = elevon FR = Fay-Riddell min = MINIVER ref = reference value 1 = freestream conditions

Bowing of metallic thermal protection systems for reentry of a previously proposed single-stage-to-orbit reusable launch vehicle was studied. The outer layer of current metallic thermal protection system concepts typically consists of a honeycomb panel made of a hightemperature nickel alloy. During portions of reentry when the thermal protection system is exposed to rapidly varying heating rates, a significant temperature gradient develops across the honeycomb panel thickness, resulting in bowing of the honeycomb panel. The deformations of the honeycomb panel increase the roughness of the outer mold line of the vehicle, which could possibly result in premature boundary layer transition, resulting in significantly higher downstream heating rates. The aerothermal loads and parameters for three locations on the centerline of the windward side of this vehicle were calculated using an engineering code. The transient temperature distributions through a metallic thermal protection system were obtained using 1-D finite volume thermal analysis, and the resulting displacements of the thermal protection system were calculated. The maximum deflection of the thermal protection system throughout the reentry trajectory was 6.4 mm. The maximum ratio of deflection to boundary layer thickness was 0.032. Based on previously developed distributed roughness correlations, it was concluded that these defections will not result in tripping the hypersonic boundary layer.