Christopher J. Riley

Engineering Aerothermal Analysis for X-34 Thermal Protection System Design

Design of the thermal protection system for any hypersonic flight vehicle requires determination of both the peak temperatures over the surface and the heating-rate history along the flight profile. In this paper, the process used to generate the aerothermal environments required for the X-34 Testbed Technology Demonstrator thermal protection system design is described as it has evolved from a relatively simplistic approach based on engineering methods applied to critical areas to one of detailed analyses 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 flowfield code and an inviscid/boundary layer method are shown. Good agreement is demonstrated among all these methods for both the ground-test condition and the peak heating flight condition. Finally, the detailed analysis using engineering methods to interpolate the surface-heating-rate results from the inviscid/boundary layer method to predict the required thermal environments is described and results presented.

Aeroheating Predictions for X-34 Using an Inviscid Boundary-Layer Method

Radiative equilibrium surface temperatures and surface heating rates from a combined inviscid-boundary layer method are presented for the X-34 Reusable Launch V ehicle for several points along the hypersonic descent portion of its trajectory. I n viscid, perfect-gas solutions are generated with the Langley Aerothermodynamic Upwind Relaxation Algorithm LAURA and the Data-Parallel Lower-Upper Relaxation DPLUR code. Surface temperatures and heating rates are then computed using the Langley Approximate Three-Dimensional Convective Heating LATCH engineering code employing both laminar and turbulent ow models. The combined inviscid-boundary layer method provides accurate predictions of surface temperatures over most of the vehicle and requires much less computational eeort than a Navier-Stokes code. This enables the generation of a more thorough aerothermal database which is necessary to design the thermal protection system and specify the vehicles ight limits.

Engineering calculations of three-dimensional inviscid hypersonic flowfields

An approximate solution technique has been developed for three-dimensional, inviscid, hypersonic flows. The method uses Maslens explicit pressure equation and the assumption of approximate stream surfaces in the shock layer. This approximation represents a simplification of Maslens asymmetric method. The solution procedure involves iteratively changing the shock shape in the subsonic-transonic region until the correct body shape is obtained. Beyond this region, the shock surface is determined by using a marching procedure. Results are presented herein for a paraboloid and elliptic cone at angle of attack. Calculated surface pressure distributions, shock shapes, and property profiles are compared with experimental data and finite-difference solutions of the Euler equations. Comparisons of the results of the present method with experimental data and detailed predictions are very good. Since the present method provides a very rapid computational procedure, it can be used for parametric or preliminary design applications. One useful application would be to incorporate a heating procedure for aerothermal studies.

Surface pressure and streamline effects on laminar heating calculations

The effect of streamline geometry and pressure distributions on surface heating rates is examined for slender, spherically blunted cones. The resulting modifications to an approximate aeroheating code include a curve fit of pressures computed by an Euler solution over a range of Mach numbers and small cone angles. An existing correlation based on pressures computed by the method of characteristi cs is used for larger cone angles. The streamline geometry is computed using the surface pressures and inviscid surface properties. Streamline calculations based on inviscid surface conditions rather than boundary-layer edge properties are demonstrated to yield improved heating analyses. However, the heating rates are calculated using inviscid properties at the boundarylayer edge. Resulting heating rates compare favorably with solutions from the viscous-shock-layer equations.

Effects of surface pressures and streamline metrics on the calculation of laminar heating rates

The effect of streamline geometry and pressure distributions on surface heating rates is examined for slender, spherically blunted cones. The modifications to the approximate aeroheating code include a curve fit of pressures computed by an Euler solution over a range of Mach numbers and cone angles. The streamline geometry is then found using the surface pressures and inviscid surface properties. Previously, streamlines were determined using the inviscid properties at the edge of the boundary layer when accounting for the effects of entropy-layer swallowing. Streamline calculations are now based on inviscid surface conditions rather than boundary-layer edge properties. However, the heating rates are calculated using inviscid properties at the edge of the boundary layer. Resulting heating rates compare favorably with solutions from the viscous-shock-layer equations.