The lifting body legacy - X-33

Aero-thermal analysis of lifting body configurations in hypersonic flow

The preliminary design phase of any program is the key to its eventual successful development. The more advanced a design, the more this tends to be true. For this reason, the preliminary design phase is particularly important in the design of aerospace systems. Errors in preliminary design tend to be fundamental and tend to cause programs to be abandoned, or to be changed fundamentally, and at great cost later in the design development. In order to facilitate the identification of attractive design regimes, the NASA Ames Research Center initiated a project, supporting the 2/sup nd/ generation reusable launch vehicle (RLV) Program, to develop an advanced suite of conceptual design tools. The tool, called risk oriented program optimization tool (ROPOT) was developed for evaluation of reusable space transportation system (RSTS) conceptual analysis. The ROPOT tool is a modular lego-block approach that is built on risk drivers from the space shuttle. This paper illustrates how this tool was adapted and applied to the decision-making about future space transportation system architectures, including new orbital space plane designs that are to be launched on the new evolved expendable launch vehicles. ROPOT has been further developed to provide the capability to explore crew survival capabilities of the vehicles throughout the mission, and develop insights and concerning the key trades involving performance, cost, crew safety and developmental risk. The tool demonstrated the ability to capture the key trades necessary to make informed decisions concerning alternative vehicle concepts, clearly identify strengths and weaknesses of design alternatives, and identify the most fruitful areas for achieving overall benefits. The results of this effort demonstrate the ability of ROPOT to be easily adapted to evaluate a broad range of vehicle types ranging from fully-reusable advanced concepts of the 2/sup nd/ generation program, to near term expendable vehicles.

The development and application of the differential evolution (DE) optimisation algorithm to the problem of worst-case analysis of nonlinear control laws for hypersonic re-entry vehicles is described. The algorithm is applied to the problem of evaluating a proposed nonlinear handling qualities clearance criterion for a detailed simulation model of a hypersonic re-entry vehicle (also known as a reusable launch vehicle (RLV)) having a full-authority nonlinear dynamic inversion (NDI) flight control law. A hybrid version of the differential evolution algorithm, incorporating local gradient-based optimisation, is also developed and evaluated. Comparisons of computational complexity and global convergence properties reveal the significant benefits which may be obtained through hybridisation of the standard differential evolution algorithm. The proposed optimisation-based approach to worst-case analysis is shown to have significant potential for improving both the reliability and efficiency of the flight clearance process for next generation RLV’s.

Our aerospace customers are demanding that we drastically reduce the cost of operating and supporting our products. Our space customer in particular is looking for the next generation of reusable launch vehicle systems to support more aircraft like operation. To achieve this goal requires more than an evolution in materials, processes and systems, what is required is a paradigm shift in the design of the launch vehicles and the processing systems that support the launch vehicles. This paper describes the Automated Informed Maintenance System (AIM) we are developing for NASA's Space Launch Initiative (SLI) Second Generation Reusable Launch Vehicle (RLV). Our system includes an Integrated Health Management (IHM) system for the launch vehicles and ground support systems, which features model based diagnostics and prognostics. Health Management data is used by our AIM decision support and process aids to automatically plan maintenance, generate work orders and schedule maintenance activities along with the resources required to execute these processes. Our system will automate the ground processing for a spaceport handling multiple RLVs executing multiple missions. To accomplish this task we are applying the latest web based distributed computing technologies and application development techniques.

A major objective of next generation reusable launch vehicle (RLV) programs includes significant improvements in vehicle safety, reliability, and operational costs. In this paper, trajectory generation on approach/ landing (A&L) for RLVs using motion primitives (MPs) and neighboring optimal control (NOC) is discussed. The proposed trajectory generation approach at A&L phase is based on an MP scheme which consists of trims and maneuvers. From an initial point to a given touchdown point, all feasible trajectories that satisfy certain constraints are generated and saved into a trajectory database. An optimal trajectory can then be found off-line by using Dijkstra's algorithm. After perturbations are imposed on the initial states of the off-line optimal trajectory, it is reshaped into a neighboring feasible trajectory on-line by using NOC approach. At this point, a neighboring feasible trajectory existence theorem (NFTET) is investigated and its proof is provided as well. The results show that the vehicle with stuck effectors can be recovered from failures in real time. Finally, robustness issues on NOC approach are briefly discussed.

NASA's Integrated Space Transportation Plan — 3 rd generation reusable launch vehicle technology update

The aim of the present thesis is the study of some problems of hypersonic rarefied Aerodynamics. More specifically the subjects here considered are: a) Development of a new parameter to detect non-equilibrium region. As already pointed out, the identification of non-equilibrium regions is important for an accurate solution of a flow field. More specifically an hybrid code needs a parameter to determine what method (DSMC or CFD), has to be used in the solution of the flow field. In this thesis, a new parameter to detect the non equilibrium region is proposed. This parameter is based on the Crocco theorem. The assumption on which the new parameter relies is that a theoretical relation, based on the hypotheses of equilibrium, as the one of the Crocco theorem, is not verified in non-equilibrium. The new parameter has been computed as the difference of the terms forming the Crocco theorem equation. Thus, one can expect that the higher is non-equilibrium, the larger is the mismatch between the terms and therefore the higher is the parameter. b) Improvement of approximate methods (bridging formulae) for the evaluation of aerodynamic coefficients of a re-entry vehicle in high altitude flight. At the first stage of a design of a re-entry vehicle it could be important to determine in a fast way the aerodynamic forces coefficient. The achieve this goal the well known “bridging formulae” are used. In this thesis a “new” methodology (here called “new” bridging formula) has been developed. The “new” bridging formula, has been successfully tuned to sphere and it has been also tested on two current capsules: EXPERT and ORION. c) Analysis and comparison of several chemical models: 1) peculiar of a DSMC approach such as quantic models (classic and new), Fan-Shen and Bird, 2) “classic” models such as the Gupta and Park models with and without ionization reactions. It is well known that one of the most important problems in the design of a capsule is the evaluation of heat flux during the re-entry. This evaluation has to provide information about the design of the Thermal Protection System (TPS). To this regard it is important to underline that due to the endothermic characteristic of the reactions, the chemical model affects the computation of the heat flux. A very deep analysis about the difference in the computation of heat flux between a direct simulation Monte Carlo code (DS2V) and a computational fluid dynamics code (H3NS) has been carried out. To this purpose a method to implement the Park model in a DSMC code has been developed. d) Application of DSMC codes to evaluate the aerodynamic coefficients of a current capsule (EXPERT) and a future aerospace vehicle (FAST20XX). For the EXPERT capsule computer tests have been carried out in the altitude flight with particular attention to the longitudinal stability of the capsule. As for as FAST20XX, the demonstration and validation of the numerical tools able to predict aero-thermal loads on a space re-entry vehicle at high altitude conditions has been carried out. This goal is fulfilled by the characterization of the DLR low density wind tunnel V2G and by an aerodynamic analysis of the available model of a lifting body.

Computational fluid dynamic (CFD) simulations in support of the SHARP-B2 (Slender Hypervelocity Aerothermodynamic Research Probes) program are presented. Pre-flight simulations of the aerothermal flight environment have been made to aid in the design of the flight hardware and instrumentation. Simulations of a critical flight qualification ground test performed in the 20 MW Panel Test Facility (PTF) arc jet, which is equipped with a semi-elliptical nozzle, have also been performed. The arc jet simulations were performed with the same CFD code that is used in the flight calculations. For both the flight and the arc jet simulations, the CFD results are used to provide surface heat transfer coefficients that serve as boundary conditions for an in-depth conduction thermal response code. By using the same CFD tool to compute the aerothermal environments for both the ground test and the flight test environment, useful comparisons are made regarding the traceability of the ground test environment to flight. Introduction The NASA Ames Research Center has been developing new Ultra High Temperature Ceramics (UHTC's) with the goal that these materials will enable sharp leading edges to be used for future space vehicles. Because of their unique thermo-structural properties, these materials are capable of operation at temperatures near 5100 °F without ablation. The materials have been developed and tested in ground-based arc jet facilities, and a flight test program designated SHARP (Slender Hypervelocity Aerothermodynamic Research Probes) has been initiated. The first flight demonstration, SHARP-B1, incorporated a UHTC nosetip mounted on a U.S. Air Force Mkl2A entry vehicle in collaboration with Sandia National Laboratory (SNL), and was successfully flown in May, 1997. The second flight test, SHARP-B2, which incorporated four instrumented UHTC strakes mounted on the side of the entry vehicle, was flown in September 2000. The goal of these flight tests is to assess the performance of the materials under entry conditions. The use of Computational Fluid Dynamic (CFD) simulations for the design of Thermal Protection Systems (TPS) has evolved to an extent that they have been integrated into the preliminary design process of Reusable Launch Vehicle (RLV) programs such as the X-33. Real-gas CFD simulations with finite-rate chemistry and proper boundary conditions for coupling with material thermal response codes are now routinely performed on high-end laboratory workstations. Much of the ground based testing of advanced TPS components is done in arc-heated test facilities such as those located in the Arc Jet Complex at the NASA Ames Research Center. These facilities are capable of simulating the high enthalpy flow environment experienced by entry vehicles for a broad range of test conditions and configurations. Semi-elliptical nozzles are available for the testing of flat panels in a high enthalpy boundary layer environment. CFD is • Aerospace Engineer, Reacting Flow Environments Branch, Member AIAA. t Senior Research Scientist, ELORET, Member AIAA. Copyright © 2001 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. All other rights are reserved by the copyright owner. (c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

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.

Reusable launch vehicle technology program

In response to military requirements for responsive and low-cost space access, this study provides an objective empty weight analysis of potential reusable launch vehicle (RLV) configurations. Many studies have been performed with hydrogen-fueled RLVs, but there has been limited research into RLVs using hydrocarbon scramjet engines, which have the potential to be more practical and operable for responsive military operations. To fill this void, this study used hydrocarbon dual-mode scramjet (DMSJ) engine performance data from the U.S. Air Force HyTech program in the propulsion systems of the rocket-basedcombined-cycle (RBCC) and turbine-based-combined-cycle (TBCC) booster vehicles. Each two-stage-to-orbit (TSTO) RLV has a fixed payload requirement of 20,000 lbf to low Earth orbit. The booster propulsion systems considered in the baseline study include rocket, turbine, RBCC, and TBCC. Empty weight sensitivity analyses considered the effects of the orbiter fuel selection and the effects of the turbine installed thrust to weight ratios for the RLVs using afterburning turbine engines. Based on the assumptions in this study, the vertical-takeoff-horizontal-landing (VTHL) RLVs are lighter than the horizontal-takeoffhorizontal-landing (HTHL) RLVs. Additionally, the change in orbiter propellant has either negligible or no empty weight savings for the VTHL RLVs, while it leads to substantial empty weight savings for the HTHL RLVs. For the HTHL RLVs, increasing the turbine installed thrust to weight ratio causes a significant decrease in RLV empty weight.

Two studies were conducted in order to investigate the technical communication practices of Russian and US aerospace engineers and scientists. Both studies had the same five objectives: to solicit the opinions of aerospace engineers and scientists regarding the importance of technical communication to their professions; to determine the use and production of technical communication by aerospace engineers and scientists; to seek their views about the appropriate content of the undergraduate course in technical communication; to determine aerospace engineers' and scientists' use of libraries, technical information centers, and online databases; and to determine the use and importance of computer and information technology to them. Responses to a self-administered questionnaire that was distributed to Russian aerospace engineers and scientists at the Central Aero-Hydrodynamic Institute (TsAGI) and to their US counterparts at the NASA Ames Research Center and the NASA Langley Research Center are presented.< <ETX xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">&gt;</ETX>

For the reusable launch vehicle(RLV), in the cruise stage without power and the initial load phase, due to the aerodynamic pressure is small, and the aerodynamic surface control efficiency is relatively low, it can not achieve the attitude adjustment of the RLV. This design is based on the continuous system, and applies to the model reference adaptive control law algorithms, to achieve the design of the reaction control system(RCS). According to the control torque which is required by the RLV attitude adjustment, the RCS distribution method of mixed-integer programming is used, and the control parameters of the reaction control system is adjusted in real time,thus to achieve the adjustment of the RLV's attitude. Through the software-digital simulation analysis, this method has a good controllability and robustness.At the same time, the stability of the model reference adaptive RCS is analyzed theoretically, and getting a certain amount of stability region.

At the end of the NASA High Speed Research (HSR) Program, NASA Langley Research Center (LaRC) began a program to screen the high-temperature Polymeric Composite Materials (PMCs) characterized by the HSR Durability Program for possible use in Reusable Launch Vehicles (RLVs) operating under extreme temperature conditions. The HSR Program focused on developing material-related technologies to enable a High Speed Civil Transport (HSCT) capable of operating temperatures ranging from 54 C (-65 F) to 177 C (350 F). A high-temperature polymeric resin, PETI-5 was used in the HSR Program to satisfy the requirements for performance and durability for a PMC. For RLVs, it was anticipated that this high temperature material would contribute to reducing the overall weight of a vehicle by eliminating or reducing the thermal protection required to protect the internal structural elements of the vehicle and increasing the structural strain limits. The tests were performed to determine temperature-dependent mechanical and physical proper-ties of IM7/PETI-5 composite over a temperature range from cryogenic temperature -253 C (-423F) to the material's maximum use temperature of 230 C (450 F). This paper presents results from the test program for the temperature-dependent mechanical and physical properties of IM7/PETI-5 composite in the temperature range from -253 C (-423 F) to 27 C (80 F).

Many parameter identification techniques have been used at the NASA Ames Research Center, Dryden Flight Research Facility at Edwards Air Force Base to determine the aeroelastic stability of new and modified research vehicles in flight. This paper presents a summary of each technique used with emphasis on fast Fourier transform methods. Experiences gained from application of these techniques to various flight test programs are discussed. Also presented are data-smoothing techniques used for test data distorted by noise. Data are presented for various aircraft to demonstrate the accuracy of each parameter identification technique discussed.