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Aerosciences, Aero-Propulsion and Flight Mechanics Technology Development for NASA's Next Generation Launch Technology Program
... Moreover, hypersonic wind tunnels typically does not allow full scale testing. The realization that CFD can heavily be relied upon in design of hyperplanes is exemplified by Cockrell [17], and with the increased (cheap) availability of computational resources, has been widely adopted in various national and international programs in the past two decades as discussed in Section 1. ...
This paper presents an overview of the use of CFD to help in the design of various aspects of the Destinus hyperplanes. CFD is not only relied upon in the aerodynamic shape definition but also in engine related design tasks including intakes, compressors, turbines, nozzles as well as feeding components. The different tasks are complementary to rapid engineering design tools and aim at improving general in-house understanding, semi-empirical modeling capability as well as build surrogate models. Furthermore, the affordability of CFD can be taken advantage of at much earlier stages in design and development tasks. The versatility of such an endeavour requires a robust CFD solver capable of tackling a diverse range of conditions, from subsonic to hypersonic, with associated physics such as chemical reactions for ideal and real gases, multi-phase flows and conjugate heat transfer. To this end, the commercial software Simcenter STAR-CCM+ is intensively used at Destinus.
... The X-43D/RCCFD is an ongoing part of the Hyper-X program. The end goal of the program is to create an operational reusable combined-cycle flight demonstrator (RCCFD) [52] [50]. ...
Resurgence of interest in hypersonic and space-access vehicles increases the demand for an educated workforce. A valuable approach to immediately educate a new generation of engineers is simply to remember the past. The objective of this paper is to present the historical development of high-speed airbreathing and rocket propulsion systems. These systems include the ramjet, scramjet, rocket, and combined cycle engines. It is observed that a readily available comprehensive accounting of said systems in a singular location is absent. Additionally, due to the nature of these high-speed systems, information is often classified or restricted; as such, this paper solely focuses on information publicly available. A study of this type is approached from two avenues. First, a chronological accounting of both U.S. and international high-speed programs. Second, a data driven analytical accounting of said programs to arrive at propulsion and vehicle design regressions. This paper presents the first step- the chronological accounting. A comprehensive database of ramjet, scramjet, rockets, and combined cycle systems has been assembled and analyzed for technology evolution. The findings of the analysis are visualized in a chronological infographic identifying significant milestones.
... The physical difficulty of designing entry vehicles originates from the large degree of coupling between the various disciplines involved in the design [1-3]. The disciplines which can be accounted for and integrated during the design are: trajectory optimization [4-6], guidance, navigation, and control (GN&C) technology [7,8], aerodynamics and aerothermodynamics [9][10][11], thermal-structural analysis [12][13][14], and thermal protection system (TPS) development [15][16][17][18][19]. Efforts have been made in developing a collaborative or a multidisciplinary optimization process that considers some of the disciplines of interest during an integrated design [20][21][22][23]. However, none of these efforts considers how uncertainty in the atmospheric conditions, in the entry parameters of the vehicle, in the condition of the vehicle during entry, and in the performance of the TPS will influence the design and provide a risk assessment. ...
This paper presents an approach for performing Multi-disciplinary Design Optimization under Uncertainty (MDO-U) and an application for minimizing the thickness of the Thermal Protection System (TPS) for an entry vehicle configuration. Uncertainties due to material variability and the operating environment are considered. The value of considering uncertainty during the optimization process is demonstrated by comparing the performance of two optimal configurations, one identified through a deterministic approach and the other by considering uncertainty. Been able to account for uncertainty in engineering simulations and in product design is both important and challenging. In particular, during design optimization the performance of the optimized design typically lays closely to the boundaries of some of the constraints. Since variability alters the actual performance of a system, the optimal design can exhibit performance in the infeasible domain when the impact of the uncertainty has not been accounted properly during the optimization.
This chapter presents an overview of flow phenomena related specifically to scramjet propulsion as it integrates with its airframe. It addresses some characteristics of the hypersonic flowfield, paying attention to phenomena such as boundary‐layer transition, inlet starting, inlet spillage, thermal choking, and combustor‐inlet interaction. Aerothermodynamic analysis requires accurate modeling of shock wave‐boundary‐layer interaction, and shock‐shock vehicle interactions to assess how those interactions affect the vehicle boundary layers, since they can lead to regions of enhanced aerothermodynamic loading. At hypersonic speeds, the engine inlet compresses the ingested atmospheric airflow by the deceleration of the ram flow, which converts the freestream kinetic energy into pressure energy, using a supersonic diffuser. In a scramjet inlet, the flow is not diffused to sonic condition as the ramjet inlet does. Shockwave interference heating was seen in a flight test of NASA X‐15 hypersonic research aircraft.
An integration environment has been developed for conducting multidiscipline design optimization analysis under uncertainty. It facilitates solution of multiple optimization problems in parallel with multiple sets of objectives and constraints originating from different design disciplines while simultaneously accounting for uncertainty during the optimization process. A seamless general purpose integration capability facilitates exchanging data between the optimization processes and the solvers which are used for evaluating the objective functions and the constraints. Metamodels can be developed and used instead of the actual solvers during the highly iterative optimization process in order to expedite the computations. Uncertainties are introduced in the optimization by considering the constraints which depend on any random variables and any random parameters as probabilistic. Satisfying the probabilistic constraints in the presence of uncertainty introduces sufficient conservatism in the solution and eliminates the need for further application of safety factors. The work presented in this paper considers trajectory, aerothermal, aerodynamic, thermal, and structural computations when performing the design optimization for the Thermal Protection System (TPS) and for the structure of a TSTO upper stage vehicle. Sixteen different sections are considered on the vehicle when designing the TPS. The trajectory bank angle schedule, the angle of attack schedule, the thickness of the sixteen different TPS sections, and twenty seven thicknesses associated with the structure are considered when reducing the overall weight of the vehicle while satisfying the imposed constraints. Uncertainties are considered in three control angles of the trajectory, in the material strength, the thrust load and the 2.5G loads. The results from the multi-discipline optimization without and with uncertainty are discussed, and a comparison between the deterministic and the probabilistic optimum is made. © 2010 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
An overview of some of the activities in hypersonic airbreathing aerodynamics and propulsion airframe integration is presented for the Space Vehicle Technology Institute. The Institute is a multi-university joint NASA-DoD program that was created as a center for research and education in future launch vehicle technologies. Perhaps more than any other type of flight vehicle, next generation space launchers will have to be analyzed as completely integrated aerodynamic-propulsion systems. Optimal aerodynamics will be vital to the development of efficient, engine-integrated launch vehicle forms, especially using airbreathing propulsion. To this end, inverse design approaches, design tradeoffs, and an understanding of relevant basic flow physics are all part of the Space Vehicle Technology Institute program. The relevance of these efforts to NASA activities is also described.
A collaborative software development approach is described. The software product is an adaptation of proven computational capabilities combined with new ca-pabilities to form the Agency's next generation aerothermodynamic and aerodynamic analysis and design tools. To efficiently produce a cohesive, robust, and extensible software suite, the approach uses agile software development techniques; specifically, project retrospectives, the Scrum status meeting format, and a subset of Extreme Programming's coding practices are employed. Examples are provided which demon-strate the substantial benefits derived from employing these practices. Also included is a discussion of issues encountered when porting legacy Fortran 77 code to For-tran 95 and a Fortran 95 coding standard.
Computational fluid dynamics tools have been used extensively in the analysis and development of the X-43A Hyper-X Research Vehicle. A significant element of this analysis is the prediction of integrated vehicle aeropropulsive performance, which includes an integration of aerodynamic and propulsion flowfields. The development of the Mach 7 X-43A required a preflight assessment of longitudinal and lateral-directional aeropropulsive characteristics near the target flight-test condition. The development of this preflight database was accomplished through extensive aerodynamic wind-tunnel testing and a combination of three-dimensional inviscid airframe calculations and cowl-to-tail scramjet cycle analyses to generate longitudinal performance increments between mission sequences. These increments were measured directly and validated through tests of the Hyper-X flight engine and vehicle flowpath simulator in the NASA Langley Research Center 8-Foot High Temperature Tunnel. Predictions were refined with tip-to-tail Navier-Stokes calculations, which also provided information on scramjet exhaust plume expansion in the aftbody region. A qualitative assessment of lateral-directional stability characteristics was made through a series of tip-to-tail inviscid calculations, including a simulation of the powered scramjet flight-test condition. Additional comparisons with wind-tunnel force and moment data as well as surface pressure measurements from the Hyper-X flight engine and vehicle flowpath simulator model and wind-tunnel testing were made to assess solution accuracy.
A comprehensive treatment to the optimal atmospheric ascent problem of launch vehicles subject to path constraints and final condition constraints is presented. The development is particularly tailored for the purpose of the eventual realization of closed-loop endoatmospheric ascent guidance. It is demonstrated that the classical finite difference method for two-point boundary-value problems is well suited for solving the optimal ascent problem onboard. The performance and execution of the guidance algorithm are assessed with a series of open-loop and closed-loop tests using the vehicle data of a reusable launch vehicle. The test results render strong supporting evidence to the belief that closed-loop optimal endoatmospheric ascent guidance is now feasible.
An integrated tool developed for the construction of unstructured numerical grids about complex configurations is presented. The tool is highly integrated and employs the native underlying geometry modeling kernel used to define the target domain for providing topological and geometric access used by the grid generation procedures. This access greatly reduces the overall process time required to generate grids for complex models. In addition, the tool is based on an underlying framework that enables the integration of new grid generation technology as it becomes available. The GridEx package presented herein is under development at the NASA Langley Research Center as part of the Fast Adaptive Aerospace Tools initiative.
For reusable launch vehicles, it is common to separate the guidance and flight control problem into an outer loop and an inner loop. The inner loop uses aerodynamic and propulsive controls to achieve a commanded attitude. The attitude commands are generated by an outer loop. This outer loop uses inner loop commands and control variables to achieve desired positions/velocities. This paper describes improvements and evaluation results for combined inner-outer loop adaptive flight control architecture for reusable launch vehicles. This architecture includes the online training of a neural network to correct for force and moment model errors. The result is a system that can respond to large model errors (significant failure scenarios) in both inner loop attitude control as well as in trajectory tracking. The resulting system can be used to improve performance and/or to increase guidance, navigation, and control system tolerance to model error, failures, and the environment. Simulation results are presented for failure scenarios.
Recently performed linear stability analyses suggested that transition could be delayed in hypersonic boundary layers by using an ultrasonically absorptive surface to damp the second mode (Mack mode). Boundary-layer transition experiments were performed on a sharp 5.06-deg half-angle round cone at zero angle of attack in the T5 Hypervelocity Shock Tunnel to test this concept. The cone was constructed with a smooth surface around half the cone circumference (to serve as a control) and an acoustically absorptive porous surface on the other half. Test gases investigated included nitrogen and carbon dioxide at M∞ ≃ 5 with specific reservoir enthalpy ranging from 1.3 to 13.0 MJ/kg and reservoir pressure ranging from 9.0 to 50.0 MPa. Comparisons were performed to ensure that previous results obtained in similar experiments (on a regular smooth surface) were reproduced, and the results were extended to examine the effects of the porous surface. These experiments indicated that the porous surface was highly effective in delaying transition provided that the pore size was significantly smaller than the viscous length scale.
In a recent flight experiment to study hypersonic crossflow transition, boundary layer characteristics were documented. A smooth steel glove was mounted on the first stage delta wing of Orbital Sciences Corporation's Pegasus (R) launch vehicle and was flown at speeds of up to Mach 8 and altitudes of up to 250,000 ft. The wing-glove experiment was flown as a secondary payload off the coast of Florida in October 1998. This paper describes the measurement system developed. Samples of the results obtained for different parts of the trajectory are included to show the characteristics and quality of the data. Thermocouples and pressure sensors (including Preston tubes, Stanton tubes, and a "probeless" pressure rake showing boundary layer profiles) measured the time-averaged flow. Surface hot-films and high-frequency pressure transducers measured flow dynamics. Because the vehicle was not recoverable, it was necessary to design a system for real-time onboard processing and transmission. Onboard processing included spectral averaging. The quality and consistency of data obtained was good and met the experiment requirements.
A methodology for on-board planning of sub-orbital entry trajectories is developed. The algorithm is able to generate in a time frame consistent with on-board environment a three-degree-of-freedom (3DOF) feasible entry trajectory, given the boundary conditions and vehicle modeling. This trajectory is then tracked by feedback guidance laws which issue guidance commands. The current trajectory planning algorithm complements the recently developed method for on-board 3DOF entry trajectory generation for orbital missions, and provides full-envelope autonomous adaptive entry guidance capability. The algorithm is validated and verified by extensive high fidelity simulations using a sub-orbital reusable launch vehicle model and difficult mission scenarios including failures and aborts.
Boundary layer trip devices for the Hper-X forebody have been experimentally examined in several wind tunnels. Five different trip configurations were compared in three hypersonic facilities, the LaRC 20-Inch Mach 6 Air Tunnel, the LaRC 31 -Inch Mach 10 Air Tunnel, and in the HYPULSE Reflected Shock Tunnel at GASL. Heat transfer distributions, utilizing the phosphor thermography and thin-film techniques, shock system details, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. Parametric variations include angles-of-attack of 0-deg, 2-deg, and 4-deg; Reynolds numbers based on model length of 1.2 to 15.4 million: and inlet cowl door simulated in both open and closed positions. Comparisons of transition due to discrete roughness elements have led to the selection of a trip configuration for the Hyper-X Mach 7 flight vehicle.
A multi-grid, flux-difference-split, finite-volume code, VULCAN, is presented for solving the elliptic and parabolized form of the equations governing three-dimensional, turbulent, calorically perfect and non-equilibrium chemically reacting flows. The space marching algorithms developed to improve convergence rate and or reduce computational cost are emphasized. The algorithms presented are extensions to the class of implicit pseudo-time iterative, upwind space-marching schemes. A full approximate storage, full multi-grid scheme is also described which is used to accelerate the convergence of a Gauss-Seidel relaxation method. The multi-grid algorithm is shown to significantly improve convergence on high aspect ratio grids.
In this paper, a new nonlinear control synthesis technique (θ - D approximation) is presented. This approach achieves suboptimal solutions to nonlinear optimal control problems in the sense that it solves the Hamilton-Jacobi-Bellman (HJB) equation approximately by adding perturbations to the cost function. By manipulating the perturbation terms both semi-globally asymptotic stability and suboptimality properties can be obtained. The convergence and stability proofs are given. This method overcomes the large control for large initial states problem that occurs in some other Taylor expansion based methods. It does not need time-consuming online computations like the state dependent Riccati equation (SDRE) technique. A vector problem is investigated to demonstrate the effectiveness of this new technique.
A novel flight control system for a reusable launch vehicle during Terminal Area Energy Management (TAEM), approach and landing phases of flight is presented. In order to provide accurate tracking and bring the vehicle safely to the landing site, the controller must be robust to external unknown disturbances, unmodeled dynamics and plant uncertainties and also be able to perform well under a wide range of operating conditions. This is accomplished using a new approach in Sliding Mode Control which has multiple loop structure with different time scales, and is driven by sliding mode observers combined with gain adaptation. Comparisons with a classically designed controller are made using a well defined, reusable launch vehicle model and considerable improvements in performance are observed. Implications are that this type of design can improve vehicle safety and reliability by improved robustness, and decrease development and operational cost by reducing analysis time.
Researchers at the NASA Langley Research Center are engaged in numerous activities to optimize the development and integration of active flow control structures. There are two different types of measurements that are required for the active flow control arena. First, there are the traditional wind tunnel measurements which are designed to help the researcher develop a fundamental understanding of the aerodynamic cause and effect phenomena associated with various active flow control experiments. Second, there are the measurements that provide the global, high-bandwidth sensor information that enables the integration of active flow control systems to flight vehicles. This paper defines what is meant by active flow control, discusses some of the traditional as well as some of the newest techniques that are currently used in the wind tunnel environment, and outlines the challenges of active flow control measurements for flight. © 2002 by the American Institute of Aeronautics and Astronautics, Inc.
In this paper, we present a direct fault tolerant control (DFTC) technique for the attitude control of a reusable launch vehicle (RLV), where by "direct" we mean that no explicit fault identification is used. Any partial or complete failure of control actuators and effectors in a high-gain (integral feedback) control loop will be inferred from saturation of one or more commanded control signals generated by the controller. The saturation causes a reduction in the effective gain, or bandwidth of the feedback loop, which can be modeled as an increase in singular perturbation in the loop. In order to maintain stability, the bandwidth of the nominal (reduced-order) system will be reduced proportionally according to the singular perturbation theory. The presented DFTC technique automatically handles momentary saturations and integrator windup caused by excessive disturbances, guidance command or dispersions under normal vehicle conditions. For multi-input, multi-output (MIMO) systems with redundant control effectors, such as the RLV attitude control system, algorithms for determining the direction of the (vector-valued) bandwidth reduction are discussed and illustrated by an example. The stability for variable bandwidth adaptation is ensured by linear time-varying PD-eigenvalues or by receding horizon optimal control method. An illustrative example and dispersion test results of a high-fidelity reusable launch vehicle model are presented to demonstrate the effectiveness of the DFTC technique.
The direct simulation Monte Carlo method is applied in this paper to simulate the time-dependent development of a three-dimensional model flow. The flows are initially chaotic and are driven by a centrifugal forcing that mimics that of the centrifugal force in the Taylor flows. The computations have been performed in parallel computer clusters. The results presented include those for two domain sizes with various levels of forcing. The results show that the long-time-averaged flow structures resemble that of the Tayor-Gortler vortices. The evolution of the dominant Fourier modes is examined. Single point perturbation signals are oscillatory. Pathline flow visualization shows that the perturbation describes coherent vortical flow structures. © 2003 by the authors. Published by the American Institute of Aeronautics and Astronautics, Inc.
Advanced guidance and control (AG&C) technologies are critical for meeting safety/reliability and cost requirements for the next generation of reusable launch vehicle (RLV). This becomes clear upon examining the number of expendable launch vehicle failures in the recent past where AG&C technologies would have saved a RLV with the same failure mode, the additional vehicle problems where this technology applies, and the costs associated with mission design with or without all these failure issues. The state-of-the-art in guidance and control technology, as well as in computing technology, is at the point where we can look to the possibility of being able to safely return a RLV in any situation where it can physically be recovered. This paper outlines reasons for AG&C, current technology efforts, and the additional work needed for making this goal a reality.
We investigate a new type of flow-tagging velocimetry technique for hypersonic flows. The technique involves exciting a thin line of nitric oxide molecules with a laser beam and then, after some delay, acquiring an image of the displaced line. One component of velocity is determined from the time of flight. This method is applied to measure the velocity profile in a Mach 8.5 laminar, hypersonic boundary layer in the Australian National Universities T2 free-piston shock tunnel. The velocity is measured with an uncertainty of approximately 2%. Comparison with a CFD simulation of the flow shows reasonable agreement.
Detailed aeroheating information is critical to the successful design of a thermal protection system (TPS) for an aerospace vehicle. NASA Langley Research Center's (LaRC) phosphor thermography method is described. Development of theory is provided for a new weighted two color relative-intensity fluorescence theory for quantitatively determining surface temperatures on hypersonic wind-tunnel models and an improved application of the one-dimensional conduction theory for use in determining global heating mappings. The phosphor methodology at LaRC is presented including descriptions of phosphor model fabrication, test facilities, and phosphor video acquisition systems. A discussion of the calibration procedures, data reduction, and data analysis is given. Estimates of the total uncertainties (with a 95% confidence level) associated with the phosphor technique are shown to be approximately 7-10% in LaRC's 31-Inch Mach 10 Tunnel and 8-10% in the 20-Inch Mach 6 Tunnel. A comparison with thin-film measurements using 5.08-cm-radius hemispheres shows the phosphor data to be within 7% of thin-film measurements and to agree even better with predictions via a LATCH computational fluid dynamics (CFD) solution. Good agreement between phosphor data and LAURA CFD computations on the forebody of a vertical takeoff/vertical lander configuration at four angles of attack is also shown. In addition, a comparison is given between Mach 6 phosphor data and laminar and turbulent solutions generated using the LAURA, GASP, and LATCH CFD codes on the X-34 configuration. The phosphor process outlined is believed to provide the aerothermodynamic community with a valuable capability for rapidly obtaining (three to four weeks) detailed heating information needed in TPS design.
Tests of the Hyper-X scramjet engine flowpath have been conducted in the HYPULSE shock tunnel at conditions duplicating the stagnation enthalpy at flight Mach 7, 10, and 15. For the tests at Mach 7 and 10 HYPULSE was operated as a reflected-shock tunnel; at the Mach 15 condition, HYPULSE was operated as a shock-expansion tunnel. The test conditions matched the stagnation enthalpy of a scramjet engine on an aerospace vehicle accelerating through the atmosphere along a 1000 psf dynamic pressure trajectory. Test parameter variation included fuel equivalence ratios from lean (0.8) to rich (1.5+); fuel composition from pure hydrogen to mixtures of 2% and 5% silane in hydrogen by volume; and inflow pressure and Mach number made by changing the scramjet model mounting angle in the HYPULSE test chamber. Data sources were wall pressures and heat flux distributions and schlieren and fuel plume imaging in the combustor/nozzle sections. Data are presented for calibration of the facility nozzles and the scramjet engine model. Comparisons of pressure distributions and flowpath streamtube performance estimates are made for the three Mach numbers tested.
An implicit, Navier-Stokes solution algorithm is presented for the computation of turbulent flow on unstructured grids. The inviscid fluxes are computed using an upwind algorithm and the solution is advanced in time using a backward-Euler time-stepping scheme. At each time step, the linear system of equations is approximately solved with a point-implicit relaxation scheme. This methodology provides a viable and robust algorithm for computing turbulent flows on unstructured meshes. Results are shown for subsonic flow over a NACA 0012 airfoil and for transonic flow over an RAE 2822 airfoil exhibiting a strong upper-surface shock. In addition, results are shown for 3- and 4-element airfoil configurations. For the calculations, two 1-equation turbulence models are utilized. For the NACA 0012 airfoil a pressure distribution and force data are compared with other computational results as well as with the experiment. Comparisons of computed pressure distributions and velocity profiles with experimental data are shown for the RAE airfoil and for the 3-element configuration. For the 4-element case, comparisons of surface pressure distributions with the experiment are made. In general, the agreement between the computations and the experiment is good.
This paper provides an overview of NASA's Hyper-X Program; a focused hypersonic technology effort designed to move hypersonic, airbreathing vehicle technology from the laboratory environment to the flight environment. This paper presents an overview of the flight test program, research objectives, approach, schedule and status. Substantial experimental database and concept validation have been completed. The program is currently concentrating on the first, Mach 7, vehicle development, verification and validation in preparation for wind-tunnel testing in 1998 and flight testing in 1999. Parallel to this effort the Mach 5 and 10 vehicle designs are being finalized. Detailed analytical and experimental evaluation of the Mach 7 vehicle at the flight conditions is nearing completion, and will provide a database for validation of design methods once flight test data are available.
Two optical systems capable of measuring model attitude and deformation were compared to inertial devices employed to acquire wind tunnel model angle of attack measurements during the sting mounted full span 30% geometric scale flexible configuration of the Northrop Grumman Unmanned Combat Air Vehicle (UCAV) installed in the NASA Langley Transonic Dynamics Tunnel (TDT). The overall purpose of the test at TDT was to evaluate smart materials and structures adaptive wing technology. The optical techniques that were compared to inertial devices employed to measure angle of attack for this test were: (1) an Optotrak (registered) system, an optical system consisting of two sensors, each containing a pair of orthogonally oriented linear arrays to compute spatial positions of a set of active markers; and (2) Video Model Deformation (VMD) system, providing a single view of passive targets using a constrained photogrammetric solution whose primary function was to measure wing and control surface deformations. The Optotrak system was installed for this test for the first time at TDT in order to assess the usefulness of the system for future static and dynamic deformation measurements.
The scope and significance of propulsion airframe integration (PAI) for hypersonic airbreathing vehicles is presented through a discussion of the PAI test techniques utilized at NASA Langley Research Center. Four primary types of PAI model tests utilized at NASA Langley for hypersonic airbreathing vehicles are discussed. The four types of PAI test models examined are the forebody/inlet test model, the partial-width/truncated propulsion flowpath test model, the powered exhaust simulation test model, and the full-length/width propulsion flowpath test model. The test technique for each of these four types of PAI test models is described, and the relevant PAI issues addressed by each test technique are illustrated through the presentation of recent PAI test data.
This paper presents the status of the airbreathing hypersonic airplane and space-access vision-operational-vehicle design matrix, with emphasis on horizontal takeoff and landing systems being, studied at Langley, it reflects the synergies and issues, and indicates the thrust of the effort to resolve the design matrix including Mach 5 to 10 airplanes with global-reach potential, pop-up and dual-role transatmospheric vehicles and airbreathing launch systems. The convergence of several critical systems/technologies across the vehicle matrix is indicated. This is particularly true for the low speed propulsion system for large unassisted horizontal takeoff vehicles which favor turbines and/or perhaps pulse detonation engines that do not require LOX which imposes loading concerns and mission Flexibility restraints.
An independent twelve degree-of-freedom simulation of the X-43A separation trajectory was created with the Program to Optimize Simulated trajectories (POST II). This simulation modeled the multi-body dynamics of the X-43A and its booster and included the effect of two pyrotechnically actuated pistons used to push the vehicles apart as well as aerodynamic interaction forces and moments between the two vehicles. The simulation was developed to validate trajectory studies conducted with a 14 degree-of-freedom simulation created early in the program using the Automatic Dynamic Analysis of Mechanics Systems (ADAMS) simulation software. The POST simulation was less detailed than the official ADAMS-based simulation used by the Project, but was simpler, more concise and ran faster, while providing similar results. The increase in speed provided by the POST simulation provided the Project with an alternate analysis tool. This tool was ideal for performing separation control logic trade studies that required the running of numerous Monte Carlo trajectories.
Airframe-integrated scramjet engine tests have been completed at Mach 7 in the NASA Langley 8-Foot High Temperature Tunnel under the Hyper-X program. These tests provided critical engine data as well as design and database verification for the Mach 7 flight tests of the Hyper-X research vehicle (X-43), which will provide the first-ever airframe-integrated scramjet flight data. The first model tested was the Hyper-X Engine Model (HXEM), and the second was the Hyper-X Flight Engine (HXFE). The HXEM, a partial-width, full-height engine that is mounted on an airframe structure to simulate the forebody features of the X-43, was tested to provide data linking flowpath development databases to the complete airframe-integrated three-dimensional flight configuration, and to isolate effects of ground testing conditions and techniques. The HXFE, an exact geometric representation of the X-43 scramjet engine mounted on an airframe structure that duplicates the entire three-dimensional propulsion flowpath from the vehicle leading edge to the vehicle base, was tested to verify the complete design as it will be flight tested. This paper presents an overview of these two tests, their importance to the Hyper-X program, and the significance of their contribution to scramjet database development.
This paper provides an overview of the activities associated with the aerodynamic database which is being developed in support of NASA's Hyper-X scramjet flight experiments. Three flight tests are planned as part of the Hyper-X program. Each will utilize a small, nonrecoverable research vehicle with an airframe integrated scramjet propulsion engine. The research vehicles will be individually rocket boosted to the scramjet engine test points at Mach 7 and Mach 10. The research vehicles will then separate from the first stage booster vehicle and the scramjet engine test will be conducted prior to the terminal decent phase of the flight. An overview is provided of the activities associated with the development of the Hyper-X aerodynamic database, including wind tunnel test activities and parallel CFD analysis efforts for all phases of the Hyper-X flight tests. A brief summary of the Hyper-X research vehicle aerodynamic characteristics is provided, including the direct and indirect effects of the airframe integrated scramjet propulsion system operation on the basic airframe stability and control characteristics. Brief comments on the planned post flight data analysis efforts are also included.
Airbreathing launch vehicles continue to be a subject of great interest in the space access community. In particular, horizontal takeoff and horizontal landing vehicles are attractive with their airplane-like benefits and flexibility for future space launch requirements. The most promising of these concepts involve airframe integrated propulsion systems, in which the external undersurface of the vehicle forms part of the propulsion flowpath. Combining of airframe and engine functions in this manner involves all of the design disciplines interacting at once. Design and optimization of these configurations is a most difficult activity, requiring a multi-discipline process to analytically resolve the numerous interactions among the design variables. This paper describes the design and optimization of one configuration in this vehicle class, a lifting body with turbine-based low-speed propulsion. The integration of propulsion and airframe, both from an aero-propulsive and mechanical perspe...
The NASA X-43 "Hyper-X" hypersonic research vehicle will be boosted to a Mach 7 flight test condition mounted on the nose of an Orbital Sciences Pegasus launch vehicle. The separation of the research vehicle from the Pegasus presents some unique aerodynamic problems, for which computational fluid dynamics has played a role in the analysis. This paper describes the use of several CFD methods for investigating the aerodynamics of the research and launch vehicles in close proximity. Specifically addressed are unsteady effects, aerodynamic database extrapolation, and differences between wind tunnel and flight environments. Introduction The Hyper-X research program was initiated in 1996 to demonstrate in-flight hypersonic scramjet propulsion. To get to the flight test conditions, the 12 ft long research vehicle (HXRV) is mounted on the nose of the first stage of an Orbital Sciences Corporation Pegasus booster. Given the non-axisymmetric shape of the HXRV, it is mounted onto an adapter tha...
An implicit code for computing inviscid and viscous incompressible flows on unstructured grids is described. The foundation of the code is a backward Euler time discretization for which the linear system is approximately solved at each time step with either a point implicit method or a preconditioned generalized minimal residual (GMRES) technique. For the GMRES calculations, several techniques are investigated for forming the matrix-vector product. Convergence acceleration is achieved through a multigrid scheme that uses nonnested coarse grids that are generated using a technique described in the present paper. Convergence characteristics are investigated and results are compared with an exact solution for the inviscid flow over a four-element airfoil. Viscous results, which are compared with experimental data, include the turbulent flow over a NACA 4412 airfoil, a three-element airfoil for which Mach number effects are investigated, and three-dimensional flow over a wing with a partial-span flap.
Prediction of Hyper-X Stage Separation Aerodynamics Using CFD AIAA Paper 2000-4009, Presented at the 18 Calculations from Infrared Thermographic Data
- Dilley
- D Arthur
- Pao
- Jenn
Dilley, Arthur D. and Pao, Jenn L. " Prediction of Hyper-X Stage Separation Aerodynamics Using CFD, " AIAA Paper 2000-4009, Presented at the 18 Calculations from Infrared Thermographic Data, " AIAA Paper 2003-3634, Presented at the 36th AIAA Thermophysics Conference, June 2003.
Hypersonic Technology (HyTech) Program Overview AIAA Paper 98-1566, Presented at the AIAA International Space Planes and Hypersonic Systems and Technologies Conference
- Robert A Mercier
- T M Ronald
Mercier, Robert A.; Ronald, T.M.F. " Hypersonic Technology (HyTech) Program Overview, " AIAA Paper 98-1566, Presented at the AIAA International Space Planes and Hypersonic Systems and Technologies Conference, April 1998.
Scramjet Performance
- Griffin Y Anderson
- Charles R Mcclinton
- Weidner
- John
Anderson, Griffin Y.; McClinton, Charles R. and Weidner, John P. " Scramjet Performance ", Scramjet Propulsion, AIAA Progress in Aeronautics and Astronautics, vol. 189, 2000.
The NASA Hyper-X Program
- Delma C Freeman
- David E Reubush
- Charles R Mcclinton
- Rausch
- L Vincent
- J Crawford
- Larry
Hypervelocity Capability of the HYPULSE Sock-Expansion Tube Tunnel for Scramjet Testing
- R O Foelsche
- R C Rogers
- C Y Tsai
- R J Bakos
- A Shih
AIAA Paper 98-1566, Presented at theAIAA International Space Planes and Hypersonic Systems and Technologies Conference
- Robert A Mercier
- T M Ronald
AIAA Paper 2002-4558, Presented at the AIAA Guidance, Navigation and Control Conference
- Ping Lu
- Bruce Tsai