Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

A new method for calculating the three-dimensional inviscid surface streamlines and streamline metrics using Cartesian coordinates and time as the independent variable of integration has been developed. The technique calculates the streamline from a specified point on the body to a point near the stagnation point by using a prescribed pressure distribution in the Euler equations. The differential equations, which are singular at the stagnation point, are of the two-point boundary-value-problem type. Laminar heating rates are calculated using the axisymmetric analog concept for three-dimensional boundary layers and approximate solutions to the axisymmetric boundary-layer equations. Results for elliptic conic forebody geometries indicate that the location of the point of maximum heating depends on the angle of attack of the conic and the type of conic in the plane of symmetry. This location is in general different from the stagnation point. The new method was found to give smooth predictions of heat transfer in the nose region where previous methods gave oscillatory results.

* Senior Research Scientist, Reacting Flow Environments Branch. Senior Member AIAA † Senior Research Scientist. Senior Member AIAA ‡ Senior Research Scientist. Associate Fellow AIAA § Senior Research Scientist. Senior Member AIAA ** Research Scientist, Aerothermodynamics Branch. Senior Member AIAA †† Research Scientist, Aerothermodynamics Branch. Fellow AIAA ‡‡ Research Scientist, Aerothermodynamics Branch. Fellow AIAA Copyright © 2003 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17, U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for Governmental purposes Abstract Details of the radiative heating analysis for the forebody of a candidate Titan aerocapture orbiter are presented. The radiative heating rates are obtained through a posteriori analysis of high-fidelity thermochemical nonequilibrium flow fields computed using modern techniques of computational fluid dynamics. Results from axisymmetric and three-dimensional analysis are presented at several points on candidate aerocapture trajectories in various model atmospheres of Titan. The radiative heating rates are found to be up to five times the peak convective heating rates, indicating that an accurate knowledge of the uncertainty of the radiative heating predictions is critical for a Titan aerocapture mission. The results also show that – (1) the radiative heating rates are dominated by the violet band system of CN, and (2) the gas mixture is optically thin. The predicted radiative heating is found to be very sensitive to the dissociation rate of molecular nitrogen – a factor of two increase in the rate, which is within the experimental uncertainty, results in a 25% decrease in the radiative heating.

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

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.

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

A higher-order viscous-shock-layer method has been developed and is used to obtain physically consistent results under varying degrees of low-density conditions for perfect-gas and nonequilibrium flows past slender bodies. The method of solution is a spatial-marching, implicit finite-difference technique, which employs a Vigneron pressure condition in the subsonic nose region. Higher-order body and shock slip conditions are employed with the method to obtain solutions for the low-density flows. Detailed comparisons with the direct simulation Monte Carlo method and Navier-Stokes calculations clearly show that the higher-order terms included in the HVSL equations are required to predict comparable values of surface pressure and heat transfer rate at higher altitude. The deficiency of the standard viscous shock-layer method in predicting low-density flows can not be corrected by the slip conditions alone as considered by earlier researchers.

Results are obtained for cylindrical leading edges of proposed transatmospheric vehicles by employing a two-dimensional viscous shock-layer code for nonequilibrium gas flows. The accuracy and efficiency of the planar code is verified through detailed comparisons with other predictions. This study includes results for 6-deg half-angle bodies with nose radii ranging from 0.01 to 2.0 ft for both cylindrically blunted wedges and spherically blunted cones (included for comparison). Some results are presented as a ratio of the noncatalytic to the corresponding fully catalytic heating value to illustrate the maximum potential for a heating reduction in dissociated nonequilibrium flows. Generally, this ratio and the individual heating rates are smaller for cylindrically blunted wedges with small nose radii as compared to the spherically blunted cones (for the same nose radius). Therefore, a larger potential exists for heating reduction in cylindrically blunted as compared with the spherically blunted surfaces. However, the results presented at higher altitudes (where the slip effects become important) show that the spherically blunted nose gives lower stagnation-point heating due to stronger merged shock-layer effects as compared with a cylindrically blunted nose.

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Results are obtained for cylindrical leading edges of proposed transatmospheric vehicles by employing a 2D viscous shock-layer code for nonequilibrium and perfect gas flows. The accuracy and efficiency of the planar code is verified through detailed comparisons with other predictions. It is found to be as accurate and robust as its axisymmetric counterpart. This study includes results for nose radii ranging from 0.01 to 2.0 ft and half-angles of 5 and 6 deg for both cylindrically-blunted wedges and spherically blunted cones (included for comparison). Some results are presented as a ratio of the noncatalytic to the corresponding fully catalytic heating value to illustrate the maximum potential for a heating reduction in dissociated nonequilibrium flows. This ratio and the individual heating rates are smaller for cylindrically-blunted wedges with small nose radii, relative to the spherically-blunted cones (at the same radius). Therefore, a larger potential exists for heating reduction in cylindrically-blunted as compared with the spherically-blunted surfaces for finite-rate chemistry.

A solution procedure is presented that considerably improves the computational efficiency of the viscous-shock-layer technique, especially for long slender bodies. The ''predictor-corrector'' procedure suggested for obtaining the shock shape beyond the nose region requires only a single global pass. The accuracy of the present method is demonstrated by comparison with globally iterated results over the entire body and with ground- and flight-test data. The new procedure results in computer run times one-third to one-half of the times required for the full-body global iteration procedure. Furthermore, the algebraic expressions used to specify the initial shock shape eliminate the need for a shock shape generated by external means and permit immediate introduction of the full viscous-shock-layer equations.

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Preliminary results from the STS-35 and STS-40 flight of the Shuttle Infrared Leeside Temperature Sensing (SILTS) experiment aboard the Shuttle Orbiter Columbia are presented. Infrared images are shown in false-color indicating the level and distribution of surface temperature over the vehicle's leeside fuselage during entry. Features evident in the imagery are related to their causative aerodynamic flow phenomena. Quantitative comparisons of the infrared image data with in situ temperature measurements obtained with thermocouples located at the aerodynamic surface of the thermal protection materials are presented.

Flight- and ground-test heat-transfer data, detailed predictions, and engineering solutions have been compared. The impact of several parameters on heat transfer and the capability of three engineering codes to predict these results have been demonstrated. Results have shown that fairly good agreement with data and detailed solutions can be obtained, but good engineering judgment is required in choosing the options in the codes. In particular, comparison of the results of the engineering codes, AEROHEAT, INCHES, and MINIVER, with Reentry F flight data and ground-test heat-transfer data for a range of cone angles, and with the predictions obtained using the detailed VSL3D code, has shown very good agreement in the regions of applicability of the engineering codes. The impact of several flowfield and vehicle parameters, including entropy, pressure gradient, nose bluntness, gas chemistry, and angle of attack on heating levels has been shwon to be important. Particular care must be exercised when using engineering codes since comparisons have demonstrated that the parameters of this study can significantly influence the actual heating levels and the prediction capability of a code. The engineering codes provide the user with relatively simple techniques to define the aerothermal environment for parametric or preliminary design studies.

A new method for calculating the three-dimensional inviscid surface streamlines and streamline metrics using Cartesian coordinates and time as the independent variable of integration has been developed. The technique calculates the streamline from a specified point on the body to a point near the stagnation point by using a prescribed pressure distribution in the Euler equations. The differential equations, which are singular at the stagnation point, are of the two point boundary value problem type. Laminar heating rates are calculated using the axisymmetric analog concept for three-dimensional boundary layers and approximate solutions to the axisymmetric boundary layer equations. Results for elliptic conic forebody geometries show that location of the point of maximum heating depends on the type of conic in the plane of symmetry and the angle of attack, and that this location is in general different from the stagnation point. The new method was found to give smooth predictions of heat transfer in the nose region where previous methods gave oscillatory results.

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

A numerical study was performed to assess the applicability of some current techniques which can be used for aerothermal predictions over slender spherically blunted cones. Predictions using a viscous-shock-layer method and several engineering approaches were compared with experimental results from flight and ground-based tests, with each other, and with other detailed results. Good agreement was obtained in comparisons with laminar and turbulent heating data from the Reentry F flight vehicle and with the wind-tunnel data. In particular, the viscous-shock-layer method was shown to yield excellent comparisons and should be useful in providing detailed flowfield and surface values for slender blunted cones. Additional predictions were obtained with these methods for two 5-deg half-angle cones with different nose radii to illustrate the effects of nose bluntness and angle of attack on drag and heat transfer. These results demonstrate the benefit of nose blunting with respect to heating and drag for laminar and transitional flow at zero-lift conditions and the benefit of heating reduction at angle of attack. Detailed comparison of the engineering code predictions with the viscous-shock-layer results for these additional cases generally showed good agreement except for the laminar prodictions at angle of attack on the forward cone surface. Nonequilibrium calculations at 0-deg angle of attack showed that substantial benefits of low surface catalysis existed only in the nose-dominated region.