Kathryn E. Wurster

Development of Advanced Metallic-Thermal-Protection System Prototype Hardware

A new metallic thermal-protection-system concept has been designed, analyzed, and fabricated. A specific location on a slngle-stage-to-orbit reusable launch vehicle was selected to develop loads and requirements needed todesign prototype panels. The design loads include ascent and entry heating rates, pressures, acoustics, and accelerations. Additional design issues were identified and discussed. An iterative sizing procedure was used to size the thermal protection system panels for thermal and structural loads as part of an integrated wall construction that included the thermal protection system and cryogenic tank structure. The panels were sized to maintain acceptable temperatures on the underlying structure and to operate under the design structural loading. Detailed creep analyses were also performed on critical components of the panels. Four 18-in.-square metallic thermal-protection-system panels were fabricated. A lightweight, thermally compliant support system to connect the thermal protection system to the cryogenic tank structure was designed and fabricated.

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.

Boundary Layer Transition on X-43A

The successful Mach 7 and 10 flights of the first fully integrated scramjet propulsion systems by the Hyper-X (X-43A) program have provided the means with which to verify the original design methodologies and assumptions. As part of Hyper-Xs propulsion-airframe integration, the forebody was designed to include a spanwise array of vortex generators to promote boundary layer transition ahead of the engine. Turbulence at the inlet is thought to provide the most reliable engine design and allows direct scaling of flight results to ground- based data. Pre-flight estimations of boundary layer transition, for both Mach 7 and 10 flight conditions, suggested that forebody boundary layer trips were required to ensure fully turbulent conditions upstream of the inlet. This paper presents the results of an analysis of the thermocouple measurements used to infer the dynamics of the transition process during the trajectories for both flights, on both the lower surface (to assess trip performance) and the upper surface (to assess natural transition). The approach used in the analysis of the thermocouple data is outlined, along with a discussion of the calculated local flow properties that correspond to the transition events as identified in the flight data. The present analysis has confirmed that the boundary layer trips performed as expected for both flights, providing turbulent flow ahead of the inlet during critical portions of the trajectory, while the upper surface was laminar as predicted by the pre-flight analysis.

Metallic Thermal Protection System Requirements, Environments, and Integrated Concepts

Achieving the ultimate goal of an economically viable reusable launch vehicle will eventually require developing Federal Aviation Regulation-type performance-based requirements and certification by the Federal Aviation Administration, as is currently done for commercial transports. Because the necessary requirements do not currently exist, there is no verifiable and traceable link between thermal protection system design implementation and resulting performance, safety, and cost. An initial attempt has been made to outline a set of performance-based thermal protection system design requirements. Critical requirements that will have a profound effect on the economic viability of a reusable launch vehicle, such as those for ground hail strike, lightning strike, bird strike, rain/rain erosion, and on-orbit dehris/micrometeoroid hypervelocity impact have been proposed. In addition to design requirements, the importance of both compiling a comprehensive loads envelope and deriving time- and location-consistent loads for thermal protection system design and sizing is addressed. Including ascent abort trajectories as limit-load cases and on-orbit debris/micrometeoroid hypervelocity impact as one of the discrete-source-damage cases is Imperative because of their significant impact on thermal protection system design and resulting performance, reliability, and operability. General features of a suite of integrated airframe concepts is summarized, and the specific details of a metallic thermal protection system concept having design flexibility that enables weight and operability to be traded and balanced is described.

Multidisciplinary Analysis of a Lifting Body Launch Vehicle

As part of phase 2 of the X-33 Program, NASA selected an integrated lifting body/aerospike engine configuration 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 configurations in the early phases of the conceptual design process. The performance analyses that supported these studies were conducted to assess the effect of the vehicles 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. Significant improvements in ascent performance were 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 liftoff thrust-to-weight 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 thermal protection system heating rate and transition constraints while satisfying a crossrange requirement.

Electric vs Chemical Propulsion for a Large-Cargo Orbit Transfer Vehicle

Techniques for sizing electrically or chemically propelled orbit transfer vehicles and analyzing fleet requirements are used in a comparative analysis of the two concepts for various levels of traffic to geosynchronous orbit. The vehicle masses, fuel requirements, and fleet sizes are determined and translated into launch vehicle payload requirements. Technology projections beyond normal growth are made and their effect on the comparative advantages of the concepts is determined. A preliminary cost analysis indicates that electric propulsion greatly reduces launch vehicle requirements and would be competitive with chemical propulsion if the technology of power generation systems advances to where reusability can be achieved at low cost.

Flowfield and vehicle parameter influence on results of engineering aerothermal methods

Flightand 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 shown 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.

Effect of Computational Method on Discrete Roughness Correlations for Shuttle Orbiter

A reanalysis of discrete roughness boundary-layer transition data using a consistent computational method for comparison to other published results has been completed. The primary objective of this effort was to investigate the influence of the computational approach on the resulting transition correlation. The experimental results were previously obtained on Space Shuttle Orbiter models in the NASA Langley Research Center 20-Inch Mach 6 Air Tunnel. The phosphor thermography system was used to monitor the status of the boundary layer via global heat- transfer images of the orbiter windward surface. The existing roughness transition database included a variation in the size and location of discrete roughness trips along the centerline of 0.0075-scale models at an angle of attack of 40 deg. Various correlative approaches were attempted, with the roughness transition correlations based on edge properties providing the most reliable results. When a consistent computational method is used to compute edge conditions, transition data sets for different moderately blunt configurations at several angles of attack are shown to collapse to a well-behaved correlation. The shuttle experimental dataset presented herein, therefore, can be used to calibrate the preferred computational method of the end user for use in the future designs of the next-generation space access vehicles.

Metallic Thermal Protection System Technology Development: Concepts, Requirements And Assessment Overview

A technology development program was conducted to evolve an earlier metallic thermal protection system (TPS) panel design, with the goals of: improving operations features, increasing adaptability (ease of attaching to a variety of tank shapes and structural concepts), and reducing weight. The resulting Adaptable Robust Metallic Operable Reusable (ARMOR) TPS system incorporates a high degree of design flexibility (allowing weight and operability to be traded and balanced) and can also be easily integrated with a large variety of tank shapes, airframe structural arrangements and airframe structure/material concepts. An initial attempt has been made to establish a set of performance based TPS design requirements. A set of general (FARtype) requirements have been proposed, focusing on defining categories that must be included for a comprehensive design. Load cases required for TPS design must reflect the full flight envelope, including a comprehensive set of limit loads, However, including additional loads. such as ascent abort trajectories, as ultimate load cases, and on-orbit debris/micro-meteoroid hypervelocity impact, as one of the discrete -source -damage load cases, will have a significant impact on system design and resulting performance, reliability and operability. Although these load cases have not been established, they are of paramount importance for reusable vehicles, and until properly included, all sizing results and assessments of reliability and operability must be considered optimistic at a minimum.

Conceptual design of a fully reusable manned launch system

The conceptual design of a rocket-powered, two-stage fully reusable launch vehicle has been performed as a part of the advanced manned launch system (AMLS) study by NASA. The main goals of the AMLS study are to provide routine, low-cost manned access to space. Technologies and system approaches have been studied that would contribute to significant reductions in operating time and manpower relative to current systems. System and operational characteristics of the two-stage fully reusable vehicle are presented, and the various tools and methods used in the design process are summarized. The results of a series of trade studies performed to examine the effect of varying major vehicle parameters on the reference two-stage fully reusable vehicle are also summarized.