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Sample records for airbreathing hypersonic vehicles

  1. Flight testing of airbreathing hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Hicks, John W.

    1993-01-01

    Using the scramjet engine as the prime example of a hypersonic airbreathing concept, this paper reviews the history of and addresses the need for hypersonic flight tests. It also describes how such tests can contribute to the development of airbreathing technology. Aspects of captive-carry and free-flight concepts are compared. An incremental flight envelope expansion technique for manned flight vehicles is also described. Such critical issues as required instrumentation technology and proper scaling of experimental devices are addressed. Lastly, examples of international flight test approaches, existing programs, or concepts currently under study, development, or both, are given.

  2. Airbreathing hypersonic vehicle design and analysis methods

    NASA Technical Reports Server (NTRS)

    Lockwood, Mary Kae; Petley, Dennis H.; Hunt, James L.; Martin, John G.

    1996-01-01

    The design, analysis, and optimization of airbreathing hypersonic vehicles requires analyses involving many highly coupled disciplines at levels of accuracy exceeding those traditionally considered in a conceptual or preliminary-level design. Discipline analysis methods including propulsion, structures, thermal management, geometry, aerodynamics, performance, synthesis, sizing, closure, and cost are discussed. Also, the on-going integration of these methods into a working environment, known as HOLIST, is described.

  3. Airbreathing Hypersonic Technology Vision Vehicles and Development Dreams

    NASA Technical Reports Server (NTRS)

    McClinton, C. R.; Hunt, J. L.; Ricketts, R. H.; Reukauf, P.; Peddie, C. L.

    1999-01-01

    Significant advancements in hypersonic airbreathing vehicle technology have been made in the country's research centers and industry over the past 40 years. Some of that technology is being validated with the X-43 flight tests. This paper presents an overview of hypersonic airbreathing technology status within the US, and a hypersonic technology development plan. This plan builds on the nation's large investment in hypersonics. This affordable, incremental plan focuses technology development on hypersonic systems, which could be operating by the 2020's.

  4. Prospects for future hypersonic air-breathing vehicles

    NASA Technical Reports Server (NTRS)

    Beach, H. L., Jr.; Blankson, Isaiah M.

    1991-01-01

    An overview of the technical progress achieved in key areas of hypersonic airbreathing vehicle development is presented. The context for hypersonic applications is discussed with emphasis placed on technology issues and requirements, particularly for propulsion and technology integration. Attention is given to CFD technology which allows the consideration of configurations and extrapolations to flight conditions that cannot be simulated on the ground.

  5. Prospects for future hypersonic air-breathing vehicles

    NASA Technical Reports Server (NTRS)

    Beach, H. L., Jr.; Blankson, Isaiah M.

    1991-01-01

    The age of hypersonics is (almost) here. This is evident from the amount of activity in the United States, Europe, the USSR and Japan; this activity is a reflection of technical progress in key areas which will enable new vehicle systems, as well as renewed interest in the utilization of these systems. The current situation, at least in the United States, is the product of an interesting history which is briefly reviewed here. The context for hypersonic applications is discussed, but the emphasis is on hypersonic technology issues and needs, particularly for propulsion and technology integration. The paper concludes with prospects for accomplishing the objective of air-breathing hypersonic vehicle systems.

  6. Airbreathing Hypersonic Vision-Operational-Vehicles Design Matrix

    NASA Technical Reports Server (NTRS)

    Hunt, James L.; Pegg, Robert J.; Petley, Dennis H.

    1999-01-01

    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.

  7. Airbreathing Hypersonic Vision-Operational-Vehicles Design Matrix

    NASA Technical Reports Server (NTRS)

    Hunt, James L.; Pegg, Robert J.; Petley, Dennis H.

    1999-01-01

    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.

  8. Boundary Layer Control for Hypersonic Airbreathing Vehicles

    NASA Technical Reports Server (NTRS)

    Berry, Scott A.; Nowak, Robert J.; Horvath, Thomas J.

    2004-01-01

    Active and passive methods for tripping hypersonic boundary layers have been examined in NASA Langley Research Center wind tunnels using a Hyper-X model. This investigation assessed several concepts for forcing transition, including passive discrete roughness elements and active mass addition (or blowing), in the 20-Inch Mach 6 Air and the 31-Inch Mach 10 Air Tunnels. Heat transfer distributions obtained via phosphor thermography, shock system details, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. The comparisons between the active and passive methods for boundary layer control were conducted at test conditions that nearly match the Hyper-X nominal Mach 7 flight test-point of an angle-of-attack of 2-deg and length Reynolds number of 5.6 million. For passive roughness, the primary parametric variation was a range of trip heights within the calculated boundary layer thickness for several trip concepts. The passive roughness study resulted in a swept ramp configuration, scaled to be roughly 0.6 of the calculated boundary layer thickness, being selected for the Mach 7 flight vehicle. For the active blowing study, the manifold pressure was systematically varied (while monitoring the mass flow) for each configuration to determine the jet penetration height, with schlieren, and transition movement, with the phosphor system, for comparison to the passive results. All the blowing concepts tested, which included various rows of sonic orifices (holes), two- and three-dimensional slots, and random porosity, provided transition onset near the trip location with manifold stagnation pressures on the order of 40 times the model surface static pressure, which is adequate to ensure sonic jets. The present results indicate that the jet penetration height for blowing was roughly half the height required with passive roughness elements for an equivalent amount of transition movement.

  9. Airframe Research and Technology for Hypersonic Airbreathing Vehicles

    NASA Technical Reports Server (NTRS)

    Glass, David E.; Merski, N. Ronald; Glass, Christopher E.

    2002-01-01

    The Hypersonics Investment Area (HIA) within NASA's Advanced Space Transportation Program (ASTP) has the responsibility to develop hypersonic airbreathing vehicles for access to space. The Airframe Research and Technology (AR and T) Project, as one of six projects in the HIA, will push the state-of-the-art in airframe and vehicle systems for low-cost, reliable, and safe space transportation. The individual technologies within the project are focused on advanced, breakthrough technologies in airframe and vehicle systems and cross-cutting activities that are the basis for improvements in these disciplines. Both low and medium technology readiness level (TRL) activities are being pursued. The key technical areas that will be addressed by the project include analysis and design tools, integrated vehicle health management (IVHM), composite (polymer, metal, and ceramic matrix) materials development, thermal/structural wall concepts, thermal protection systems, seals, leading edges, aerothermodynamics, and airframe/propulsion flowpath technology. Each of the technical areas or sub-projects within the Airframe R and T Project is described in this paper.

  10. Composite predictive flight control for airbreathing hypersonic vehicles

    NASA Astrophysics Data System (ADS)

    Yang, Jun; Zhao, Zhenhua; Li, Shihua; Zheng, Wei Xing

    2014-09-01

    The robust optimised tracking control problem for a generic airbreathing hypersonic vehicle (AHV) subject to nonvanishing mismatched disturbances/uncertainties is investigated in this paper. A baseline nonlinear model predictive control (MPC) method is firstly introduced for optimised tracking control of the nominal dynamics. A nonlinear-disturbance-observer-based control law is then developed for robustness enhancement in the presence of both external disturbances and uncertainties. Compared with the existing robust tracking control methods for AHVs, the proposed composite nonlinear MPC method obtains not only promising robustness and disturbance rejection performance but also optimised nominal tracking control performance. The merits of the proposed method are validated by implementing simulation studies on the AHV system.

  11. Survey of Aerothermodynamics Facilities Useful for the Design of Hypersonic Vehicles Using Air-Breathing Propulsion

    NASA Technical Reports Server (NTRS)

    Arnold, James O.; Deiwert, George S.

    1997-01-01

    This paper surveys the use of aerothermodynamic facilities which have been useful in the study of external flows and propulsion aspects of hypersonic, air-breathing vehicles. While the paper is not a survey of all facilities, it covers the utility of shock tunnels and conventional hypersonic blow-down facilities which have been used for hypersonic air-breather studies. The problems confronting researchers in the field of aerothermodynamics are outlined. Results from the T5 GALCIT tunnel for the shock-on lip problem are outlined. Experiments on combustors and short expansion nozzles using the semi-free jet method have been conducted in large shock tunnels. An example which employed the NASA Ames 16-Inch shock tunnel is outlined, and the philosophy of the test technique is described. Conventional blow-down hypersonic wind tunnels are quite useful in hypersonic air-breathing studies. Results from an expansion ramp experiment, simulating the nozzle on a hypersonic air-breather from the NASA Ames 3.5 Foot Hypersonic wind tunnel are summarized. Similar work on expansion nozzles conducted in the NASA Langley hypersonic wind tunnel complex is cited. Free-jet air-frame propulsion integration and configuration stability experiments conducted at Langley in the hypersonic wind tunnel complex on a small generic model are also summarized.

  12. Impact of aeroelasticity on propulsion and longitudinal flight dynamics of an air-breathing hypersonic vehicle

    NASA Technical Reports Server (NTRS)

    Raney, David L.; Mcminn, John D.; Pototzky, Anthony S.; Wooley, Christine L.

    1993-01-01

    Many air-breathing hypersonic aerospacecraft design concepts incorporate an elongated fuselage forebody acting as the aerodynamic compression surface for a hypersonic combustion module, or scram jet. This highly integrated design approach creates the potential for an unprecedented form of aero-propulsive-elastic interaction in which deflections of the vehicle fuselage give rise to propulsion transients, producing force and moment variations that may adversely impact the rigid body flight dynamics and/or further excite the fuselage bending modes. To investigate the potential for such interactions, a math model was developed which included the longitudinal flight dynamics, propulsion system, and first seven elastic modes of a hypersonic air-breathing vehicle. Perturbation time histories from a simulation incorporating this math model are presented that quantify the propulsive force and moment variations resulting from aeroelastic vehicle deflections. Root locus plots are presented to illustrate the effect of feeding the propulsive perturbations back into the aeroelastic model. A concluding section summarizes the implications of the observed effects for highly integrated hypersonic air-breathing vehicle concepts.

  13. Impact of aeroelasticity on propulsion and longitudinal flight dynamics of an air-breathing hypersonic vehicle

    NASA Astrophysics Data System (ADS)

    Raney, David L.; McMinn, John D.; Pototzky, Anthony S.; Wooley, Christine L.

    1993-04-01

    Many air-breathing hypersonic aerospacecraft design concepts incorporate an elongated fuselage forebody acting as the aerodynamic compression surface for a hypersonic combustion module, or scram jet. This highly integrated design approach creates the potential for an unprecedented form of aero-propulsive-elastic interaction in which deflections of the vehicle fuselage give rise to propulsion transients, producing force and moment variations that may adversely impact the rigid body flight dynamics and/or further excite the fuselage bending modes. To investigate the potential for such interactions, a math model was developed which included the longitudinal flight dynamics, propulsion system, and first seven elastic modes of a hypersonic air-breathing vehicle. Perturbation time histories from a simulation incorporating this math model are presented that quantify the propulsive force and moment variations resulting from aeroelastic vehicle deflections. Root locus plots are presented to illustrate the effect of feeding the propulsive perturbations back into the aeroelastic model. A concluding section summarizes the implications of the observed effects for highly integrated hypersonic air-breathing vehicle concepts.

  14. A fast ascent trajectory optimization method for hypersonic air-breathing vehicles

    NASA Astrophysics Data System (ADS)

    Murillo, Oscar J., Jr.

    The objective of this dissertation is to investigate a fast and reliable method to generate three-dimensional optimal ascent trajectories for hypersonic air-breathing vehicles. The problem is notoriously difficult because of the strong nonlinear coupling amongst aerodynamics, propulsion, vehicle attitude and trajectory state. As such an algorithm matures, the ultimate goal is to realize optimal closed-loop ascent guidance for hypersonic air-breathing vehicles. The problem is formulated as a fuel-optimal control problem. The corresponding necessary conditions are given. It is shown how the original problem of search for the optimal control commands can be reduced to a univariate root-finding problem at each point along the trajectory. A finite difference scheme is used to numerically solve the associated two-point-boundary-value problem. Evaluation of the approach is done through open-loop solutions and closed-loop simulations. The results show promising potential of the proposed approach as a rapid trajectory optimization tool for the class of hypersonic air-breathing vehicles.

  15. Geometry Modeling and Adaptive Control of Air-Breathing Hypersonic Vehicles

    NASA Astrophysics Data System (ADS)

    Vick, Tyler Joseph

    Air-breathing hypersonic vehicles have the potential to provide global reach and affordable access to space. Recent technological advancements have made scramjet-powered flight achievable, as evidenced by the successes of the X-43A and X-51A flight test programs over the last decade. Air-breathing hypersonic vehicles present unique modeling and control challenges in large part due to the fact that scramjet propulsion systems are highly integrated into the airframe, resulting in strongly coupled and often unstable dynamics. Additionally, the extreme flight conditions and inability to test fully integrated vehicle systems larger than X-51 before flight leads to inherent uncertainty in hypersonic flight. This thesis presents a means to design vehicle geometries, simulate vehicle dynamics, and develop and analyze control systems for hypersonic vehicles. First, a software tool for generating three-dimensional watertight vehicle surface meshes from simple design parameters is developed. These surface meshes are compatible with existing vehicle analysis tools, with which databases of aerodynamic and propulsive forces and moments can be constructed. A six-degree-of-freedom nonlinear dynamics simulation model which incorporates this data is presented. Inner-loop longitudinal and lateral control systems are designed and analyzed utilizing the simulation model. The first is an output feedback proportional-integral linear controller designed using linear quadratic regulator techniques. The second is a model reference adaptive controller (MRAC) which augments this baseline linear controller with an adaptive element. The performance and robustness of each controller are analyzed through simulated time responses to angle-of-attack and bank angle commands, while various uncertainties are introduced. The MRAC architecture enables the controller to adapt in a nonlinear fashion to deviations from the desired response, allowing for improved tracking performance, stability, and

  16. Air-Breathing Hypersonic Vehicle Tracking Control Based on Adaptive Dynamic Programming.

    PubMed

    Mu, Chaoxu; Ni, Zhen; Sun, Changyin; He, Haibo

    2017-03-01

    In this paper, we propose a data-driven supplementary control approach with adaptive learning capability for air-breathing hypersonic vehicle tracking control based on action-dependent heuristic dynamic programming (ADHDP). The control action is generated by the combination of sliding mode control (SMC) and the ADHDP controller to track the desired velocity and the desired altitude. In particular, the ADHDP controller observes the differences between the actual velocity/altitude and the desired velocity/altitude, and then provides a supplementary control action accordingly. The ADHDP controller does not rely on the accurate mathematical model function and is data driven. Meanwhile, it is capable to adjust its parameters online over time under various working conditions, which is very suitable for hypersonic vehicle system with parameter uncertainties and disturbances. We verify the adaptive supplementary control approach versus the traditional SMC in the cruising flight, and provide three simulation studies to illustrate the improved performance with the proposed approach.

  17. Multi-Disciplinary Design Optimization of Hypersonic Air-Breathing Vehicle

    NASA Astrophysics Data System (ADS)

    Wu, Peng; Tang, Zhili; Sheng, Jianda

    2016-06-01

    A 2D hypersonic vehicle shape with an idealized scramjet is designed at a cruise regime: Mach number (Ma) = 8.0, Angle of attack (AOA) = 0 deg and altitude (H) = 30kms. Then a multi-objective design optimization of the 2D vehicle is carried out by using a Pareto Non-dominated Sorting Genetic Algorithm II (NSGA-II). In the optimization process, the flow around the air-breathing vehicle is simulated by inviscid Euler equations using FLUENT software and the combustion in the combustor is modeled by a methodology based on the well known combination effects of area-varying pipe flow and heat transfer pipe flow. Optimization results reveal tradeoffs among total pressure recovery coefficient of forebody, lift to drag ratio of vehicle, specific impulse of scramjet engine and the maximum temperature on the surface of vehicle.

  18. Propulsion integration of hypersonic air-breathing vehicles utilizing a top-down design methodology

    NASA Astrophysics Data System (ADS)

    Kirkpatrick, Brad Kenneth

    In recent years, a focus of aerospace engineering design has been the development of advanced design methodologies and frameworks to account for increasingly complex and integrated vehicles. Techniques such as parametric modeling, global vehicle analyses, and interdisciplinary data sharing have been employed in an attempt to improve the design process. The purpose of this study is to introduce a new approach to integrated vehicle design known as the top-down design methodology. In the top-down design methodology, the main idea is to relate design changes on the vehicle system and sub-system level to a set of over-arching performance and customer requirements. Rather than focusing on the performance of an individual system, the system is analyzed in terms of the net effect it has on the overall vehicle and other vehicle systems. This detailed level of analysis can only be accomplished through the use of high fidelity computational tools such as Computational Fluid Dynamics (CFD) or Finite Element Analysis (FEA). The utility of the top-down design methodology is investigated through its application to the conceptual and preliminary design of a long-range hypersonic air-breathing vehicle for a hypothetical next generation hypersonic vehicle (NHRV) program. System-level design is demonstrated through the development of the nozzle section of the propulsion system. From this demonstration of the methodology, conclusions are made about the benefits, drawbacks, and cost of using the methodology.

  19. Survey of Aerothermodynamics Facilities Useful for the Design of Hypersonic Vehicles Using Air-Breathing Propulsion

    NASA Technical Reports Server (NTRS)

    Arnold, James O.; Deiwert, G. S.

    1997-01-01

    The dream of producing an air-breathing, hydrogen fueled, hypervelocity aircraft has been before the aerospace community for decades. However, such a craft has not yet been realized, even in an experimental form. Despite the simplicity and beauty of the concept, many formidable problems must be overcome to make this dream a reality. This paper summarizes the aero/aerothermodynamic issues that must be addressed to make the dream a reality and discusses how aerothermodynamics facilities and their modem companion, real-gas computational fluid dynamics (CFD), can help solve the problems blocking the way to realizing the dream. The approach of the paper is first to outline the concept of an air-breathing hypersonic vehicle and then discuss the nose-to-tail aerothermodynamics issues and special aerodynamic problems that arise with such a craft. Then the utility of aerothermodynamic facilities and companion CFD analysis is illustrated by reviewing results from recent United States publications wherein these problems have been addressed. Papers selected for the discussion have k e n chosen such that the review will serve to survey important U.S. aero/aerothermodynamic real gas and conventional wind tunnel facilities that are useful in the study of hypersonic, hydrogen propelled hypervelocity vehicles.

  20. Hypersonic Airbreathing Missile

    NASA Technical Reports Server (NTRS)

    Hunt, J. L.; Lawing, P. L.; Marcum, D. C., Jr. (Inventor)

    1978-01-01

    A hypersonic airbreathing missile using dual mode scramjet engines for propulsion is described. The fuselage is constructed of a material with a high heat sink capacity and is covered with a thermal protective shield and lined with an internal insulating blanket. The engine airframe integration uses the flat lower portion of the lower fuselage to precompress the air entering the scramjet engines. The precompression of air entering the scramjet inlets increases as the angles of attack. This feature results in a highly maneuverable missile which can accelerate as it banks into a turn.

  1. Robust tracking control for an air-breathing hypersonic vehicle with input constraints

    NASA Astrophysics Data System (ADS)

    Gao, Gang; Wang, Jinzhi; Wang, Xianghua

    2014-12-01

    The focus of this paper is on the design and simulation of robust tracking control for an air-breathing hypersonic vehicle (AHV), which is affected by high nonlinearity, uncertain parameters and input constraints. The linearisation method is employed for the longitudinal AHV model about a specific trim condition, and then considering the additive uncertainties of three parameters, the linearised model is just in the form of affine parameter dependence. From this point, the linear parameter-varying method is applied to design the desired controller. The poles for the closed-loop system of the linearised model are placed into a desired vertical strip, and the quadratic stability of the closed-loop system is guaranteed. Input constraints of the AHV are addressed by additional linear matrix inequalities. Finally, the designed controller is evaluated on the nonlinear AHV model and simulation results demonstrate excellent tracking performance with good robustness.

  2. Continuous high order sliding mode controller design for a flexible air-breathing hypersonic vehicle.

    PubMed

    Wang, Jie; Zong, Qun; Su, Rui; Tian, Bailing

    2014-05-01

    This paper investigates the problem of tracking control with uncertainties for a flexible air-breathing hypersonic vehicle (FAHV). In order to overcome the analytical intractability of this model, an Input-Output linearization model is constructed for the purpose of feedback control design. Then, the continuous finite time convergence high order sliding mode controller is designed for the Input-Output linearization model without uncertainties. In addition, a nonlinear disturbance observer is applied to estimate the uncertainties in order to compensate the controller and disturbance suppression, where disturbance observer and controller synthesis design is obtained. Finally, the synthesis of controller and disturbance observer is used to achieve the tracking for the velocity and altitude of the FAHV and simulations are presented to illustrate the effectiveness of the control strategies.

  3. Aerodynamic experimentation with ducted models as applied to hypersonic air-breathing vehicles

    NASA Astrophysics Data System (ADS)

    Goon'ko, Yu. P.

    A methodology of experimentation in high supersonic wind tunnels for studying aerodynamic characteristics of hypersonic flying vehicles powered by air-breathing engines is discussed. Investigations of such total aerodynamic forces as drag, lift and pitching moment at testing the models are implicit when the air flow through the model ducts is accomplished so that to provide the simulation of the external flow around the airplane and flow over the inlets, but the operating engines and, hence, the exhaust jets are not modeled. The methods used for testing such models are based on the measurement of duct stream parameters alongside with the balance measurement of aerodynamic forces acting on the models. In the tests, aerometric tools are used such as narrow metering nozzles (plugs), pitot and static pressure probes, stagnation temperature probes and pressure orifices in walls of the model duct. The aerometric data serve to determine the flow rate and momentum of the stream at the duct exit. The internal non-simulated forces of the model ducts are also determined using the conservation equations for energy, mass flow and momentum, and these forces are eliminated from the aerodynamic test results. The techniques of the said model testing have been well developed as applied to supersonic aircraft, however their application for hypersonic vehicles whose models are tested at high supersonic speeds, Mach number M∞>4, implies some specific features. In the present paper, the results of experimental and theoretical study of these features are discussed. Some experimental data on aerodynamics of hypersonic aircraft models received in methodological tests are also presented. The tunnel experiments have been carried out in the Mach number range M∞=2-6.

  4. Uncertainty analysis and robust trajectory linearization control of a flexible air-breathing hypersonic vehicle

    NASA Astrophysics Data System (ADS)

    Pu, Zhiqiang; Tan, Xiangmin; Fan, Guoliang; Yi, Jianqiang

    2014-08-01

    Flexible air-breathing hypersonic vehicles feature significant uncertainties which pose huge challenges to robust controller designs. In this paper, four major categories of uncertainties are analyzed, that is, uncertainties associated with flexible effects, aerodynamic parameter variations, external environmental disturbances, and control-oriented modeling errors. A uniform nonlinear uncertainty model is explored for the first three uncertainties which lumps all uncertainties together and consequently is beneficial for controller synthesis. The fourth uncertainty is additionally considered in stability analysis. Based on these analyses, the starting point of the control design is to decompose the vehicle dynamics into five functional subsystems. Then a robust trajectory linearization control (TLC) scheme consisting of five robust subsystem controllers is proposed. In each subsystem controller, TLC is combined with the extended state observer (ESO) technique for uncertainty compensation. The stability of the overall closed-loop system with the four aforementioned uncertainties and additional singular perturbations is analyzed. Particularly, the stability of nonlinear ESO is also discussed from a Liénard system perspective. At last, simulations demonstrate the great control performance and the uncertainty rejection ability of the robust scheme.

  5. Evaluation of some significant issues affecting trajectory and control management for air-breathing hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Hattis, Philip D.; Malchow, Harvey L.

    1992-01-01

    Horizontal takeoff airbreathing-propulsion launch vehicles require near-optimal guidance and control which takes into account performance sensitivities to atmospheric characteristics while satisfying physically-derived operational constraints. A generic trajectory/control analysis tool that deepens insight into these considerations has been applied to two versions of a winged-cone vehicle model. Information that is critical to the design and trajectory of these vehicles is derived, and several unusual characteristics of the airbreathing propulsion model are shown to have potentially substantial effects on vehicle dynamics.

  6. Air-breathing hypersonic vehicle guidance and control studies: An integrated trajectory/control analysis methodology, phase 2

    NASA Technical Reports Server (NTRS)

    Hattis, Philip D.; Malchow, Harvey L.

    1992-01-01

    An integrated trajectory/control analysis algorithm has been used to generate trajectories and desired control strategies for two different hypersonic air-breathing vehicle models and orbit targets. Both models used cubic spline curve fit tabulated winged-cone accelerator vehicle representations. Near-fuel-optimal, horizontal takeoff trajectories, imposing a dynamic pressure limit of 1000 psf, were developed. The first model analysis case involved a polar orbit and included the dynamic effects of using elevons to maintain longitudinal trim. Analysis results indicated problems with the adequacy of the propulsion model and highlighted dynamic pressure/altitude instabilities when using vehicle angle of attack as a control variable. Also, the magnitude of computed elevon deflections to maintain trim suggested a need for alternative pitch moment management strategies. The second analysis case was reformulated to use vehicle pitch attitude relative to the local vertical as the control variable. A new, more realistic, air-breathing propulsion model was incorporated. Pitch trim calculations were dropped and an equatorial orbit was specified. Changes in flight characteristics due to the new propulsion model have been identified. Flight regimes demanding rapid attitude changes have been noted. Also, some issues that would affect design of closed-loop controllers were ascertained.

  7. Novel prescribed performance neural control of a flexible air-breathing hypersonic vehicle with unknown initial errors.

    PubMed

    Bu, Xiangwei; Wu, Xiaoyan; Zhu, Fujing; Huang, Jiaqi; Ma, Zhen; Zhang, Rui

    2015-11-01

    A novel prescribed performance neural controller with unknown initial errors is addressed for the longitudinal dynamic model of a flexible air-breathing hypersonic vehicle (FAHV) subject to parametric uncertainties. Different from traditional prescribed performance control (PPC) requiring that the initial errors have to be known accurately, this paper investigates the tracking control without accurate initial errors via exploiting a new performance function. A combined neural back-stepping and minimal learning parameter (MLP) technology is employed for exploring a prescribed performance controller that provides robust tracking of velocity and altitude reference trajectories. The highlight is that the transient performance of velocity and altitude tracking errors is satisfactory and the computational load of neural approximation is low. Finally, numerical simulation results from a nonlinear FAHV model demonstrate the efficacy of the proposed strategy.

  8. Air-breathing hypersonic vehicle guidance and control studies; An integrated trajectory/control analysis methodology: Phase 1

    NASA Technical Reports Server (NTRS)

    Hattis, Philip D.; Malchow, Harvey L.

    1991-01-01

    A tool which generates optimal trajectory/control histories in an integrated manner is generically adapted to the treatment of single-stage-to-orbit air-breathing hypersonic vehicles. The methodology is implemented as a two point boundary value problem solution technique. Its use permits an assessment of an entire near-minimum-fuel trajectory and desired control strategy from takeoff to orbit while satisfying physically derived inequality constraints and while achieving efficient propulsive mode phasing. A simpler analysis strategy that partitions the trajectory into several boundary condition matched segments is also included to construct preliminary trajectory and control history representations with less computational burden than is required for the overall flight profile assessment. A demonstration was accomplished using a tabulated example (winged-cone accelerator) vehicle model that is combined with a newly developed multidimensional cubic spline data smoothing routine. A constrained near-fuel-optimal trajectory, imposing a dynamic pressure limit of 1000 psf, was developed from horizontal takeoff to 20,000 ft/sec relative air speed while aiming for a polar orbit. Previously unspecified propulsive discontinuities were located. Flight regimes demanding rapid attitude changes were identified, dictating control effector and closed-loop controller authority was ascertained after evaluating effector use for vehicle trim. Also, inadequacies in vehicle model representations and specific subsystem models with insufficient fidelity were determined based on unusual control characteristics and/or excessive sensitivity to uncertainty.

  9. Novel adaptive neural control design for a constrained flexible air-breathing hypersonic vehicle based on actuator compensation

    NASA Astrophysics Data System (ADS)

    Bu, Xiangwei; Wu, Xiaoyan; He, Guangjun; Huang, Jiaqi

    2016-03-01

    This paper investigates the design of a novel adaptive neural controller for the longitudinal dynamics of a flexible air-breathing hypersonic vehicle with control input constraints. To reduce the complexity of controller design, the vehicle dynamics is decomposed into the velocity subsystem and the altitude subsystem, respectively. For each subsystem, only one neural network is utilized to approach the lumped unknown function. By employing a minimal-learning parameter method to estimate the norm of ideal weight vectors rather than their elements, there are only two adaptive parameters required for neural approximation. Thus, the computational burden is lower than the ones derived from neural back-stepping schemes. Specially, to deal with the control input constraints, additional systems are exploited to compensate the actuators. Lyapunov synthesis proves that all the closed-loop signals involved are uniformly ultimately bounded. Finally, simulation results show that the adopted compensation scheme can tackle actuator constraint effectively and moreover velocity and altitude can stably track their reference trajectories even when the physical limitations on control inputs are in effect.

  10. A new integration method based on the coupling of mutistage osculating cones waverider and Busemann inlet for hypersonic airbreathing vehicles

    NASA Astrophysics Data System (ADS)

    Wang, Xu-dong; Wang, Jiang-feng; Lyu, Zhen-jun

    2016-09-01

    This paper proposes a new design method employed for a practical integration of multistage compression osculating cones waverider and Busemann inlet. The singularity based on Taylor-Maccoll equations for Busemann inlet is vertified to result in the failure to directly integrate mutistage waverider and Busemann inlet. In order to avoid this limitation, a revised transition portion is incorporated into the integration, which achieves the objective to integrate multistage compression waverider with Busemann inlet in geometrical configuration. Aerodynamic characteristics of the integration and a compared three-stage compression osculating cones waverider are compared and analyzed in the same design conditions with the same leading edge of the forebody. The research results show that the integration with a Busemann inlet as its third-stage compression surface has greater lift-to-drag ratio, more uniform airbreathing inflow at the inlet entrance, and greater total pressure recovery coefficient with the same mass flowrate. And the performance can satisfy the requirements of the forebody-inlet compression for hypersonic vehicles.

  11. Hypersonic airbreathing propulsion/airframe integration

    NASA Technical Reports Server (NTRS)

    Weidner, John P.

    1992-01-01

    Recent interest in airbreathing hypersonic flight has centered around the need to develop advanced space launch systems which can reduce the cost of inserting payloads in orbit and make space more accessible. An effect of the thermal environment is to require the vehicle to operate at high altitudes, in very thin air, to maintain aircraft structural load limits. The high altitudes at which the hypersonic vehicle must operate give rise to the concept of an airframe integrated propulsion system to provide a much larger inlet and nozzle to process the required volume of air at low density, atmospheric conditions. In the integrated system, the forward portion of the vehicle compresses the air flow and serves as the external portion of the inlet; the aftbody completes the expansion process for the nozzle. In addition, the engine, which is contained between the body and the forebody shock wave, lends itself to a modular integration of a number of separate engines. In this manner, a relatively small engine can be defined to allow engine development in existing ground facilities.

  12. Water cooling system for an air-breathing hypersonic test vehicle

    NASA Technical Reports Server (NTRS)

    Petley, Dennis H.; Dziedzic, William M.

    1993-01-01

    This study provides concepts for hypersonic experimental scramjet test vehicles which have low cost and low risk. Cryogenic hydrogen is used as the fuel and coolant. Secondary water cooling systems were designed. Three concepts are shown: an all hydrogen cooling system, a secondary open loop water cooled system, and a secondary closed loop water cooled system. The open loop concept uses high pressure helium (15,000 psi) to drive water through the cooling system while maintaining the pressure in the water tank. The water flows through the turbine side of the turbopump to pump hydrogen fuel. The water is then allowed to vent. In the closed loop concept high pressure, room temperature, compressed liquid water is circulated. In flight water pressure is limited to 6000 psi by venting some of the water. Water is circulated through cooling channels via an ejector which uses high pressure gas to drive a water jet. The cooling systems are presented along with finite difference steady-state and transient analysis results. The results from this study indicate that water used as a secondary coolant can be designed to increase experimental test time, produce minimum venting of fluid and reduce overall development cost.

  13. Airbreathing Hypersonic Systems Focus at NASA Langley Research Center

    NASA Technical Reports Server (NTRS)

    Hunt, James L.; Rausch, Vincent L.

    1998-01-01

    This paper presents the status of the airbreathing hypersonic airplane and space-access vehicle design matrix, reflects on the synergies and issues, and indicates the thrust of the effort to resolve the design matrix and to focus/advance systems technology maturation. Priority is given to the design of the vision operational vehicles followed by flow-down requirements to flight demonstrator vehicles and their design for eventual consideration in the Future-X Program.

  14. LPV H-infinity Control for the Longitudinal Dynamics of a Flexible Air-Breathing Hypersonic Vehicle

    NASA Astrophysics Data System (ADS)

    Hughes, Hunter Douglas

    This dissertation establishes the method needed to synthesize and simulate an Hinfinity Linear Parameter-Varying (LPV) controller for a flexible air-breathing hypersonic vehicle model. A study was conducted to gain the understanding of the elastic effects on the open loop system. It was determined that three modes of vibration would be suitable for the hypersonic vehicle model. It was also discovered from the open loop study that there is strong coupling in the hypersonic vehicle states, especially between the angle of attack, pitch rate, pitch attitude, and the exible modes of the vehicle. This dissertation outlines the procedure for synthesizing a full state feedback Hinfinity LPV controller for the hypersonic vehicle. The full state feedback study looked at both velocity and altitude tracking for the exible vehicle. A parametric study was conducted on each of these controllers to see the effects of changing the number of gridding points in the parameter space and changing the parameter variation rate limits in the system on the robust performance of the controller. As a result of the parametric study, a 7 x 7 grid ranging from Mach 7 to Mach 9 in velocity and from 70,000 feet to 90,000 feet in altitude, and a parameter variation rate limit of [.5 200]T was used for both the velocity tracking and altitude tracking cases. The resulting Hinfinity robust performances were gamma = 2.2224 for the velocity tracking case and = 1:7582 for the altitude tracking case. A linear analysis was then conducted on five different selected trim points from the Hinfinity LPV controller. This was conducted for the velocity tracking and altitude tracking cases. The results of linear analysis show that there is a slight difference in the response of the Hinfinity LPV controller and the fixed point H infinity controller. For the tracking task, the Hinfinity controller responds more quickly, and has a lower Hinfinity performance value. Next, the H infinity LPV controller was simulated

  15. Research in Hypersonic Airbreathing Propulsion at the NASA Langley Research Center

    NASA Technical Reports Server (NTRS)

    Kumar, Ajay; Drummond, J. Philip; McClinton, Charles R.; Hunt, James L.

    2001-01-01

    The NASA Langley Research Center has been conducting research for over four decades to develop technology for an airbreathing-propelled vehicle. Several other organizations within the United States have also been involved in this endeavor. Even though significant progress has been made over this period, a hypersonic airbreathing vehicle has not yet been realized due to low technology maturity. One of the major reasons for the slow progress in technology development has been the low level and cyclic nature of funding. The paper provides a brief historical overview of research in hypersonic airbreathing technology and then discusses current efforts at NASA Langley to develop various analytical, computational, and experimental design tools and their application in the development of future hypersonic airbreathing vehicles. The main focus of this paper is on the hypersonic airbreathing propulsion technology.

  16. Systems Challenges for Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Hunt, James L.; Laruelle, Gerard; Wagner, Alain

    1997-01-01

    This paper examines the system challenges posed by fully reusable hypersonic cruise airplanes and access to space vehicles. Hydrocarbon and hydrogen fueled airplanes are considered with cruise speeds of Mach 5 and 10, respectively. The access to space matrix is examined. Airbreathing and rocket powered, single- and two-stage vehicles are considered. Reference vehicle architectures are presented. Major systems/subsystems challenges are described. Advanced, enhancing systems concepts as well as common system technologies are discussed.

  17. Computational analysis of hypersonic airbreathing aircraft flow fields

    NASA Technical Reports Server (NTRS)

    Dwoyer, Douglas L.; Kumar, Ajay

    1987-01-01

    The general problem of calculating the flow fields associated with hypersonic airbreathing aircraft is presented. Unique aspects of hypersonic aircraft aerodynamics are introduced and their demands on computational fluid dynamics are outlined. Example calculations associated with inlet/forebody integration and hypersonic nozzle design are presented to illustrate the nature of the problems considered.

  18. Tracking control of a class of non-linear systems with applications to cruise control of air-breathing hypersonic vehicles

    NASA Astrophysics Data System (ADS)

    Sun, Hongfei; Yang, Zhiling; Meng, Bin

    2015-05-01

    A new tracking-control method for general non-linear systems is proposed. A virtual controller and some command references are introduced to asymptotically stabilise the system of the tracking error dynamics. Then, the actual controller and command references are derived by solving a system of linear algebraic equations. Compared with other tracking-control methods in the literature, the tracking-controller design in this paper is simple because it needs only to solve a system of linear algebraic equations. The boundedness of the tracking controller and command references is guaranteed by the solvability of the terminal value problem (TVP) of an ordinary differential equation. For non-linear systems with minimum-phase properties, the TVP is automatically solvable. A numerical example shows that the tracking-control method is still available for some systems with non-minimum-phase properties. To enhance the robustness of the tracking controller, a non-linear disturbance observer (NDO) is introduced to estimate the disturbance. The combination of the tracking controller and the NDO is applied to the tracking control of an air-breathing hypersonic vehicle.

  19. X-43 Hypersonic Vehicle Technology Development

    NASA Technical Reports Server (NTRS)

    Voland, Randall T.; Huebner, Lawrence D.; McClinton, Charles R.

    2005-01-01

    NASA recently completed two major programs in Hypersonics: Hyper-X, with the record-breaking flights of the X-43A, and the Next Generation Launch Technology (NGLT) Program. The X-43A flights, the culmination of the Hyper-X Program, were the first-ever examples of a scramjet engine propelling a hypersonic vehicle and provided unique, convincing, detailed flight data required to validate the design tools needed for design and development of future operational hypersonic airbreathing vehicles. Concurrent with Hyper-X, NASA's NGLT Program focused on technologies needed for future revolutionary launch vehicles. The NGLT was "competed" by NASA in response to the President s redirection of the agency to space exploration, after making significant progress towards maturing technologies required to enable airbreathing hypersonic launch vehicles. NGLT quantified the benefits, identified technology needs, developed airframe and propulsion technology, chartered a broad University base, and developed detailed plans to mature and validate hypersonic airbreathing technology for space access. NASA is currently in the process of defining plans for a new Hypersonic Technology Program. Details of that plan are not currently available. This paper highlights results from the successful Mach 7 and 10 flights of the X-43A, and the current state of hypersonic technology.

  20. Propulsion Airframe Integration Test Techniques for Hypersonic Airbreathing Configurations at NASA Langley Research Center

    NASA Technical Reports Server (NTRS)

    Witte, David W.; Huebner, Lawrence D.; Trexler, Carl A.; Cabell, Karen F.; Andrews, Earl H., Jr.

    2003-01-01

    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.

  1. Air-breathing hypersonic cruise - Prospects for Mach 4-7 waverider aircraft

    NASA Technical Reports Server (NTRS)

    Blankson, Isaiah M.

    1992-01-01

    In the Mach 4-7 range, waverider aircraft are considered as candidates for both short- and long-range cruise missions, as hypersonic missiles, and as high L/D highly maneuverable craft. The potential for near- and far-term application of airbreathing engines to the waverider vehicle missions and concepts is presented. Attention is focused on the cruise mission and attempts are made to compare and contrast it with the accelerator mission.

  2. Laser-driven hypersonic air-breathing propulsion simulator

    NASA Technical Reports Server (NTRS)

    Joshi, Prakash B.; Lo, Edmond Y.; Pugh, Evan R.

    1992-01-01

    A feasibility study is presented of simulating airbreathing propulsion on small scale hypersonic models using laser energy. The laser heat addition scheme allows simultaneous inlet and exhaust flows during wind tunnel testing of models with scramjet models. The proposed propulsion simulation concept has extended the Kantrowitz (1974) idea to propulsive wind tunnel models of hypersonic aircraft. Critical issues in aeropropulsive testing of models based on a ramjet power plant are addressed which include transfer of the correct amount of energy to the flowing gas, efficient absorption of laser energy into the gas, and test performance under tunnel reservoir conditions and at reasonable Reynolds numbers.

  3. Thermodynamic Cycle Analysis of Magnetohydrodynamic-Bypass Hypersonic Airbreathing Engines

    NASA Technical Reports Server (NTRS)

    Litchford, R. J.; Cole, J. W.; Bityurin, V. A.; Lineberry, J. T.

    2000-01-01

    The prospects for realizing a magnetohydrodynamic (MHD) bypass hypersonic airbreathing engine are examined from the standpoint of fundamental thermodynamic feasibility. The MHD-bypass engine, first proposed as part of the Russian AJAX vehicle concept, is based on the idea of redistributing energy between various stages of the propulsion system flow train. The system uses an MHD generator to extract a portion of the aerodynamic heating energy from the inlet and an MHD accelerator to reintroduce this power as kinetic energy in the exhaust stream. In this way, the combustor entrance Mach number can be limited to a specified value even as the flight Mach number increases. Thus, the fuel and air can be efficiently mixed and burned within a practical combustor length, and the flight Mach number operating envelope can be extended. In this paper, we quantitatively assess the performance potential and scientific feasibility of MHD-bypass engines using a simplified thermodynamic analysis. This cycle analysis, based on a thermally and calorically perfect gas, incorporates a coupled MHD generator-accelerator system and accounts for aerodynamic losses and thermodynamic process efficiencies in the various engin components. It is found that the flight Mach number range can be significantly extended; however, overall performance is hampered by non-isentropic losses in the MHD devices.

  4. Instrumentation requirements from the user's view. [For airbreathing hypersonic engines

    SciTech Connect

    Harsha, P.T.

    1988-01-01

    The use of combustor diagnostics is considered from the point of view of demonstration of performance of an airbreathing hypersonic engine. The basic need is seen to be that of providing the data necessary to verify performance predictions for the engine as installed in the airplane. This necessitates the use of a diagnostics capability that can provide the inputs required by the computational analyses that will be used to assess this performance. Because of the cost of ground test facilities, a premium is placed on measurement technique reliability and redundancy of instrumentation. A mix of nonintrusive optical techniques and probe-based measurements is seen to be the best approach using current diagnostics capability; one such instrument mix is outlined for a ramjet/scramjet test program. 11 references.

  5. Developing Conceptual Hypersonic Airbreathing Engines Using Design of Experiments Methods

    NASA Technical Reports Server (NTRS)

    Ferlemann, Shelly M.; Robinson, Jeffrey S.; Martin, John G.; Leonard, Charles P.; Taylor, Lawrence W.; Kamhawi, Hilmi

    2000-01-01

    Designing a hypersonic vehicle is a complicated process due to the multi-disciplinary synergy that is required. The greatest challenge involves propulsion-airframe integration. In the past, a two-dimensional flowpath was generated based on the engine performance required for a proposed mission. A three-dimensional CAD geometry was produced from the two-dimensional flowpath for aerodynamic analysis, structural design, and packaging. The aerodynamics, engine performance, and mass properties arc inputs to the vehicle performance tool to determine if the mission goals were met. If the mission goals were not met, then a flowpath and vehicle redesign would begin. This design process might have to be performed several times to produce a "closed" vehicle. This paper will describe an attempt to design a hypersonic cruise vehicle propulsion flowpath using a Design of' Experiments method to reduce the resources necessary to produce a conceptual design with fewer iterations of the design cycle. These methods also allow for more flexible mission analysis and incorporation of additional design constraints at any point. A design system was developed using an object-based software package that would quickly generate each flowpath in the study given the values of the geometric independent variables. These flowpath geometries were put into a hypersonic propulsion code and the engine performance was generated. The propulsion results were loaded into statistical software to produce regression equations that were combined with an aerodynamic database to optimize the flowpath at the vehicle performance level. For this example, the design process was executed twice. The first pass was a cursory look at the independent variables selected to determine which variables are the most important and to test all of the inputs to the optimization process. The second cycle is a more in-depth study with more cases and higher order equations representing the design space.

  6. Thermodynamic Cycle Analysis of Magnetohydrodynamic-Bypass Airbreathing Hypersonic Engines

    NASA Technical Reports Server (NTRS)

    Litchford, Ron J.; Bityurin, Valentine A.; Lineberry, John T.

    1999-01-01

    Established analyses of conventional ramjet/scramjet performance characteristics indicate that a considerable decrease in efficiency can be expected at off-design flight conditions. This can be explained, in large part, by the deterioration of intake mass flow and limited inlet compression at low flight speeds and by the onset of thrust degradation effects associated with increased burner entry temperature at high flight speeds. In combination, these effects tend to impose lower and upper Mach number limits for practical flight. It has been noted, however, that Magnetohydrodynamic (MHD) energy management techniques represent a possible means for extending the flight Mach number envelope of conventional engines. By transferring enthalpy between different stages of the engine cycle, it appears that the onset of thrust degradation may be delayed to higher flight speeds. Obviously, the introduction of additional process inefficiencies is inevitable with this approach, but it is believed that these losses are more than compensated through optimization of the combustion process. The fundamental idea is to use MHD energy conversion processes to extract and bypass a portion of the intake kinetic energy around the burner. We refer to this general class of propulsion system as an MHD-bypass engine. In this paper, we quantitatively assess the performance potential and scientific feasibility of MHD-bypass airbreathing hypersonic engines using ideal gasdynamics and fundamental thermodynamic principles.

  7. Tip-to-tail numerical simulation of a hypersonic air-breathing engine with ethylene fuel

    NASA Astrophysics Data System (ADS)

    Dharavath, Malsur; Manna, P.; Chakraborty, Debasis

    2016-11-01

    End to end CFD simulations of external and internal flow paths of an ethylene fueled hypersonic airbreathing vehicle with including forebody, horizontal fins, vertical fins, intake, combustor, single expansion ramp nozzle are carried out. The performance of the scramjet combustor and vehicle net thrust-drag is calculated for hypersonic cruise condition. Three-dimensional Navier-Stokes equations are solved along with SST-k-ω turbulence model using the commercial CFD software CFX-14. Single step chemical reaction based on fast chemistry assumption is used for combustion of gaseous ethylene fuel. Simulations captured complex shock structures including the shocks generated from the vehicle nose and compression ramps, impingement of cowl-shock on vehicle undersurface and its reflection in the intake and combustor etc. Various thermochemical parameters are analyzed and performance parameters are evaluated for nonreacting and reacting cases. Very good mixing ( 98%) of fuel with incoming air stream is observed. Positive thrust-drag margins are obtained for fuel equivalence ratio of 0.6 and computed combustion efficiency is observed to be 94 %. Effect of equivalence ratio on the vehicle performance is studied parametrically. Though the combustion efficiency has come down by 8% for fuel equivalence ratio of 0.8, net vehicle thrust is increased by 44%. Heat flux distribution on the various walls of the whole vehicle including combustor is estimated for the isothermal wall condition of 1000 K in reacting flow. Higher local heat flux values are observed at all the leading edges of the vehicle (i.e., nose, wing, fin and cowl leading edges) and strut regions of the combustor.

  8. Thermo-acoustic fatigue design for hypersonic vehicle skin panels

    NASA Astrophysics Data System (ADS)

    Wentz, Kenneth R.; Blevins, Robert D.; Holehouse, Ian

    1994-09-01

    Thermo-vibro-acoustic analysis and test of skin panels for airbreathing hypersonic vehicles is made for a generic vehicle and trajectory. Aerothermal analysis shows that impingement of the bow shock wave on the vehicle and engine noise produce high fluctuating pressures and local heat fluxes. Maximum temperatures will exceed 2700 F (1480 C) at the top of the ascent trajectory and engine sound levels will exceed 170 dB at takeoff. As a result, loads due to engine acoustics and shock impingement dominate the design of many transatmospheric vehicle skin panels.

  9. Air-Breathing Launch Vehicle Technology Being Developed

    NASA Technical Reports Server (NTRS)

    Trefny, Charles J.

    2003-01-01

    Of the technical factors that would contribute to lowering the cost of space access, reusability has high potential. The primary objective of the GTX program is to determine whether or not air-breathing propulsion can enable reusable single-stage-to-orbit (SSTO) operations. The approach is based on maturation of a reference vehicle design with focus on the integration and flight-weight construction of its air-breathing rocket-based combined-cycle (RBCC) propulsion system.

  10. Fully reusable launch vehicle with airbreathing booster

    NASA Astrophysics Data System (ADS)

    Huang, Z.

    1983-10-01

    The performance of a two-stage, fully reusable scramjet-Shuttle launch vehicle is examined analytically. The first stage of the hybrid vehicle would be a hydrocarbon/hydrogen fueled air breathing hypersonic aircraft. The winged booster, in one configuration, would be equipped with both turbojets and scramjets, and the Orbiter would be a scaled-up version of the present STS Orbiter. The nominal mission would involve placing a 65,000 lb payload into a 92.5 x 185 km orbit. The size of the manned first stage is dictated by the ventricle gross weight, density, and contents, with the fuel carried in integrated tanks. A delta planform is presented, with an elliptical cone forebody, an elliptical cross-section afterbody, and a smooth transition surface from the end of the forebody to the straightline leading edge. Formulations are defined for the wing loading, gross density, fatness ratio, and breadpoint ratio, as well as for the propulsion, aerodynamics, and trajectories.

  11. Enhanced Control Effector Designs for Airbreathing Transatmospheric Vehicles

    NASA Technical Reports Server (NTRS)

    Cockrell, Charles E.; McMinn, John D.

    1997-01-01

    A study was conducted to evaluate the potential effectiveness of a moveable cowl-trailing-edge design for airbreathing hypersonic single-stage-to-orbit (SSTO) configurations, which can be extended or deflected from the nominal position in order to provide additional pitching moment capability. This additional pitching moment capability may reduce the necessary deflection angle from conventional control surfaces and the associated trim drag penalty. Calculations for a generic SSTO configuration with baseline and modified cowl trailing edge geometries were performed at Mach 6 and 10 initially with an engineering analysis code in order to examine several design parametrics. In order to more accurately model geometries and flow physics, 2-D viscous computational fluid dynamics (CFD) predictions were obtained. FInally, a limited set of 3D CFD predictions were obtained at Mach 6 in order to show the effects of modeling 3D flow fields as well as the full 3D vehicle geometry. Comparisons of aftbody surface pressures and force and moment predictions show differences between initial predictions and 2-D CFD solutions due to geometry modeling and calculation method differences. The 3-D CFD predictions confirm the trends observed in the 2-D solutions and provide additional information on 3D effects. These analyses shows that cowl-trailing-edge extensions were effective in providing additional (nose-down) pitching moment increments as well as increased thrust compared to the baseline geometry. These effects reduce the control surface deflection angle required to trim the vehicle and the associated trim drag. Cowl-trailingedge deflections were not as effective, generating more nose-up pitching moment and decreased thrust compared to non-deflected cases.

  12. Hypersonic Vehicle Trajectory Optimization

    DTIC Science & Technology

    1981-05-01

    number) The purpose of this task is to assemble a computer code for solving the maximum croserange and maximum plane change problems for hypersonic...nodal points and linear interpolation. An existing code for solving the nonlinear programming problem with the augmented-Lagrangian method is used to...Date EnleveEd) geCUmiTY CLAWPICATION OF Tw1ig PAGIi(naem DNA hata9* 4 20.\\ Abstract (cont’d.) numerically by central differences. The code works well in

  13. Ceramic Matrix Composite (CMC) Thermal Protection Systems (TPS) and Hot Structures for Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Glass, David E.

    2008-01-01

    Thermal protection systems (TPS) and hot structures are required for a range of hypersonic vehicles ranging from ballistic reentry to hypersonic cruise vehicles, both within Earth's atmosphere and non-Earth atmospheres. The focus of this paper is on air breathing hypersonic vehicles in the Earth's atmosphere. This includes single-stage to orbit (SSTO), two-stage to orbit (TSTO) accelerators, access to space vehicles, and hypersonic cruise vehicles. This paper will start out with a brief discussion of aerodynamic heating and thermal management techniques to address the high heating, followed by an overview of TPS for rocket-launched and air-breathing vehicles. The argument is presented that as we move from rocket-based vehicles to air-breathing vehicles, we need to move away from the insulated airplane approach used on the Space Shuttle Orbiter to a wide range of TPS and hot structure approaches. The primary portion of the paper will discuss issues and design options for CMC TPS and hot structure components, including leading edges, acreage TPS, and control surfaces. The current state-of-the-art will be briefly discussed for some of the components. The two primary technical challenges impacting the use of CMC TPS and hot structures for hypersonic vehicles are environmental durability and fabrication, and will be discussed briefly.

  14. Affordable Flight Demonstration of the GTX Air-Breathing SSTO Vehicle Concept

    NASA Technical Reports Server (NTRS)

    Krivanek, Thomas M.; Roche, Joseph M.; Riehl, John P.; Kosareo, Daniel N.

    2002-01-01

    The rocket based combined cycle (RBCC) powered single-stage-to-orbit (SSTO) reusable launch vehicle has the potential to significantly reduce the total cost per pound for orbital payload missions. To validate overall system performance, a flight demonstration must be performed. This paper presents an overview of the first phase of a flight demonstration program for the GTX SSTO vehicle concept. Phase 1 will validate the propulsion performance of the vehicle configuration over the supersonic and hypersonic airbreathing portions of the trajectory. The focus and goal of Phase 1 is to demonstrate the integration and performance of the propulsion system flowpath with the vehicle aerodynamics over the air-breathing trajectory. This demonstrator vehicle will have dual mode ramjet/scramjets, which include the inlet, combustor, and nozzle with geometrically scaled aerodynamic surface outer mold lines (OML) defining the forebody, boundary layer diverter, wings, and tail. The primary objective of this study is to demonstrate propulsion system performance and operability including the ram to scram transition, as well as to validate vehicle aerodynamics and propulsion airframe integration. To minimize overall risk and development cost the effort will incorporate proven materials, use existing turbomachinery in the propellant delivery systems, launch from an existing unmanned remote launch facility, and use basic vehicle recovery techniques to minimize control and landing requirements. A second phase would demonstrate propulsion performance across all critical portions of a space launch trajectory (lift off through transition to all-rocket) integrated with flight-like vehicle systems.

  15. Hypersonic propulsion flight tests as essential to air-breathing aerospace plane development

    NASA Technical Reports Server (NTRS)

    Mehta, U.

    1995-01-01

    Hypersonic air-breathing propulsion utilizing scramjets can fundamentally change transatmospheric acclerators for transportation from low Earth orbits (LEOs). The value and limitations of ground tests, of flight tests, and of computations are presented, and scramjet development requirements are discussed. Near-full-scale hypersonic propulsion flight tests are essential for developing a prototype hypersonic propulsion system and for developing computation-design technology that can be used in designing that system. In order to determine how these objectives should be achieved, some lessons learned from past programs are presented. A conceptual two-stage-to-orbit (TSTO) prototype/experimental aerospace plane is recommended as a means of providing access-to-space and for conducting flight tests. A road map for achieving these objectives is also presented.

  16. Air-breathing aerospace plane development essential: Hypersonic propulsion flight tests

    NASA Technical Reports Server (NTRS)

    Mehta, Unmeel B.

    1994-01-01

    Hypersonic air-breathing propulsion utilizing scramjets can fundamentally change transatmospheric accelerators for low earth-to-orbit and return transportation. The value and limitations of ground tests, of flight tests, and of computations are presented, and scramjet development requirements are discussed. It is proposed that near full-scale hypersonic propulsion flight tests are essential for developing a prototype hypersonic propulsion system and for developing computational-design technology so that it can be used for designing this system. In order to determine how these objectives should be achieved, some lessons learned from past programs are presented. A conceptual two-stage-to-orbit (TSTO) prototype/experimental aerospace plane is recommended as a means of providing access-to-space and for conducting flight tests. A road map for achieving these objectives is also presented.

  17. Airframe-integrated propulsion system for hypersonic cruise vehicles

    NASA Technical Reports Server (NTRS)

    Jones, R. A.; Huber, P. W.

    1978-01-01

    Research on a new, hydrogen burning, airbreathing engine concept which offers good potential for efficient hypersonic cruise vehicles is considered. Features of the engine which lead to good performance include; extensive engine-airframe integration, fixed geometry, low cooling, and the control of heat release in the supersonic combustor by mixed-modes of fuel injection from the combustor entrance. The engine concept is described along with results from inlet tests, direct-connect combustor tests, and tests of two subscale boiler-plate research engines presently underway at conditions which simulate flight at Mach 4 and 7.

  18. On Stability and Control of Hypersonic Vehicles

    DTIC Science & Technology

    2009-11-01

    stability and control issues relating to hypersonic vehicles that are deemed to be significantly different from those of the conventional (subsonic or...supersonic) air vehicles. In particular we have addressed issues that are relevant to stability and control of hypersonic vehicles. This report should

  19. Soviet applications for hypersonic vehicles

    NASA Astrophysics Data System (ADS)

    Johnson, Reuben F.

    Consistent with Soviet historical practice in the fields of space vehicle design and aircraft procurement, a hypersonic-cruise aerospacecraft will be used only for specialized missions. Most Soviet surveillance of seas and landmasses will be conducted via satellite, while military missions will be carried out by conventional aircraft and ASAT missions will be performed by unmanned systems. An aerospace vehicle will fill the gaps that have ineluctably emerged between these established systems' capabilities due to their comparatively modest technological sophistication. It can be confidently expected that the aerospaceplane's design will involve the synergistic integration of existing technologies.

  20. Calibration of the Langley 8-Foot High Temperature Tunnel for Hypersonic Airbreathing Propulsion Testing

    NASA Technical Reports Server (NTRS)

    Huebner, Lawrence D.; Rock, Kenneth E.; Voland, Randall T.; Wieting, Allan R.

    1996-01-01

    The NASA Langley 8-Foot High Temperature Tunnel has recently been modified to produce a unique testing capability for hypersonic airbreathing propulsion systems. Prior to these modifications, the facility was used primarily for aerothermal loads and structural verification testing at true flight total enthalpy conditions for Mach numbers between 6 and 7. One of the recent modifications was an oxygen replenishment system which allows operating airbreathing propulsion systems to be tested at true flight total enthalpies. Following the modifications to the facility, calibration runs were performed at total enthalpies corresponding to flight Mach numbers of 6.3 and 6.8 to establish the flow characteristics of the facility with its new capabilities. The results of this calibration, as well as modifications to tunnel combustor hardware prior to calibration to improve tunnel flow quality, are described in this paper.

  1. Implementation of magnetohydrodynamic energy bypass process for hypersonic vehicles

    NASA Astrophysics Data System (ADS)

    Lee, Ying Ming; Czysz, Paul A.; Bruno, Claudio

    2004-08-01

    The global political structure has changed dramatically since the breakup of the former Soviet Union, and world changes have caused the United States to reprioritize its national hypersonic needs. The US Government has looked at the needs of the future, and the hypersonic aerospace plane is one of the systems included in alternative force structures. One hypersonic aerospace plane concept would involve magnetohydrodynamic (MHD) technology (i.e., the AJAX hypersonic flight vehicle concept) originally proposed by Russian scientist Vladimir Fraishtadt. This paper reports on the current progress and findings of an air-breathing horizontal takeoff and landing design concept using an MHD energy bypass injector ramjet engine being studied at MSE Technology Applications, Inc., HyperTech Concepts, and several universities for the National Aeronautics and Space Administration Langley Research Center under a Phase II Small Business Innovation Research project. The areas that are addressed in this paper include: (1) ionization required to achieve the required energy bypass, (2) utilization of a nonequilibrium model to calculate nonequilibrium engine ionization conditions, (3) hydrocarbon fuel reforming, and (4) vehicle performance and sizing. A quasi-onedimensional electromagnetic code combined with a new scramjet model, as well as other tools, were used to examine total system performance.

  2. Hypersonic airbreathing missile concepts under study at NASA Langley Research Center

    NASA Technical Reports Server (NTRS)

    Hunt, J. L.; Johnston, P. J.; Cubbage, J. M.; Dillon, J. L.; Richie, C. B.; Marcum, D. C., Jr.; Carlson, C. H.

    1982-01-01

    The design and performance of several tactical and strategic hypersonic airbreathing missile concepts under study at the NASA Langley Research Center are discussed from an evolutionary perspective. A mid- and chin inlet missile design, constrained to the Navy's vertical box launcher, was investigated; a performance comparison is presented that is favorable to the mid-inlet approach. Parasol wing, confined flow field, and spatula-like cruise missile configurations were examined with strategic applications in mind. The preliminary results are encouraging with respect to aerodynamic and volumetric efficiency and choice of engine integration schemes.

  3. Switching control of a hypersonic vehicle based on guardian maps

    NASA Astrophysics Data System (ADS)

    Xiao, Dibo; Liu, Mengying; Liu, Yanbin; Lu, Yuping

    2016-05-01

    For the air-breathing hypersonic vehicles (AHSV) with a broad flight envelope, this paper proposes a new controller design method to ensure robust stability in the presence of external disturbances. The proposed method is based on the guardian maps theory and H∞ Linear Parameter Varying (LPV) technique. In this paper, an LPV model of the AHSV is firstly established using the Jacobian linearization method. By incorporating the guardian maps theory and H∞ technique, a set of controllers is designed such that the closed loop poles lie in the desired area. The major merit of the proposed method is that all controller parameters over the overall flight envelope can be automatically determined in the iterative courses that are started from an arbitrary initial point within the flight envelope. Finally, the proposed approach is demonstrated by an illustrative example. The results show that the designed controller is operated to stabilize the air-breathing hypersonic vehicle over a wide flight envelope, and to exhibit satisfactory tracking performance as well as strong robustness.

  4. Structural Sizing of a 25,000-lb Payload, Air-Breathing Launch Vehicle For Single-Stage-To-Orbit

    NASA Technical Reports Server (NTRS)

    Roche, Joseph M.; Kosareo, Daniel N.; Palac, Don (Technical Monitor)

    2000-01-01

    In support of NASA's Air-Breathing Launch Vehicle (ABLV) study, a 25,000-lb payload version of the GTX (formerly Trailblazer) reference vehicle concept was developed. The GTX is a vertical lift-off, reusable, single-stage-to-orbit launch vehicle concept that uses hypersonic air-breathing propulsion in a rocket-based combined-cycle (RBCC) propulsion system to reduce the required propellant fraction. To achieve this goal the vehicle and propulsion system must be well integrated both aerodynamically and structurally to reduce weight. This study demonstrates the volumetric and structural efficiency of a vertical takeoff, horizontal landing, hypersonic vehicle with a circular cross section. A departure from the lifting body concepts, this design philosophy is even extended to the engines, which have semicircular nacelles symmetrically mounted on the vehicle. Material candidates with a potential for lightweight and simplicity have been selected from a set of near term technologies (5 to 10 years). To achieve the mission trajectory, preliminary weight estimates show the vehicle's gross lift-off weight is 1.26 x 10(exp 6) lb. The structural configuration of the GTX vehicle and its propulsion system are described. The vehicle design benefits are presented, and key technical issues are highlighted.

  5. Structural Sizing of a 25,000-lb Payload, Air-breathing Launch Vehicle for Single-stage-to-orbit

    NASA Technical Reports Server (NTRS)

    Roche, Joseph M.; Kosareo, Daniel N.

    2001-01-01

    In support of NASA's Air-Breathing Launch Vehicle (ABLV) study, a 25,000-lb payload version of the GTX (formerly Trailblazer) reference vehicle concept was developed. The GTX is a vertical lift-off, reusable, single-stage-to-orbit launch vehicle concept that uses hypersonic air-breathing propulsion in a rocket-based combined-cycle (RBCC) propulsion system to reduce the required propellant fraction. To achieve this goal the vehicle and propulsion system must be well integrated both aerodynamically and structurally to reduce weight. This study demonstrates the volumetric and structural efficiency of a vertical takeoff, horizontal landing, hypersonic vehicle with a circular cross section. A departure from the lifting body concepts, this design philosophy is even extended to the engines, which have semicircular nacelles symmetrically mounted on the vehicle. Material candidates with a potential for lightweight and simplicity have been selected from a set of near term technologies (five to ten years). To achieve the mission trajectory, preliminary weight estimates show the vehicle's gross lift-off weight is 1.26 x 10(exp 6) lb. The structural configuration of the GTX vehicle and its propulsion system are described. The vehicle design benefits are presented, and key technical issues are highlighted.

  6. Air-breathing aerospace plane development essential: Hypersonic propulsion flight tests

    NASA Technical Reports Server (NTRS)

    Mehta, Unmeel B.

    1995-01-01

    Hypersonic airbreathing propulsion utilizing scramjets can change transatmospheric accelerators for low earth-to-orbit and return transportation. The value and limitation of ground tests, of flight tests, and of computations are presented, and scramjet development requirements are discussed. It is proposed that near full-scale hypersonic propulsion flight tests are essential for developing computational design technology so that it can be used for designing this system. In order to determine how these objectives should be achieved, some lessons learned from past programs are presented. A conceptual two-stage-to-orbit (TSTO) prototype/experimental aerospace plane is recommended as a means of providing access-to-space and for conducting flight tests.

  7. 2-D Air-Breathing Lightcraft Engine Experiments in Hypersonic Conditions

    NASA Astrophysics Data System (ADS)

    Salvador, Israel I.; Myrabo, Leik N.; Minucci, Marco A. S.; de Oliveira, Antonio C.; Toro, Paulo G. P.; Chanes, José B.; Rego, Israel S.

    2011-11-01

    Experiments were performed with a 2-D, repetitively-pulsed (RP) laser Lightcraft model in hypersonic flow conditions. The main objective was the feasibility analysis for impulse generation with repetitively-pulsed air-breathing laser Lightcraft engines at hypersonic speeds. The future application of interest for this basic research endeavor is the laser launch of pico-, nano-, and micro-satellites (i.e., 0.1-100 kg payloads) into Low-Earth-Orbit, at low-cost and on-demand. The laser propulsion experiments employed a Hypersonic Shock Tunnel integrated with twin gigawatt pulsed Lumonics 620-TEA CO2 lasers (˜ 1 μs pulses), to produce the required test conditions. This hypersonic campaign was carried out at nominal Mach numbers ranging from 6 to 10. Time-dependent surface pressure distributions were recorded together with Schlieren movies of the flow field structure resulting from laser energy deposition. Results indicated laser-induced pressure increases of 0.7-0.9 bar with laser pulse energies of ˜ 170 J, on off-shroud induced breakdown condition, and Mach number of 7.

  8. Preliminary scramjet design for hypersonic airbreathing missile application

    NASA Technical Reports Server (NTRS)

    Carlson, C. H.

    1983-01-01

    A conceptual design study of a scramjet engine was conducted for a hypersonic surface to air missile (HYSAM). The definition of the engine was based upon the requirements of accelerating the HYSAM from Mach 4 at 20,000 feet to Mach 6 at 100,000 feet and the cruise conditions at Mach 6. The resulting external and internal environmental conditions were used by various engineering disciplines performing design, stress and heat transfer analysis. A detailed structural analysis was conducted along with an indepth thermal analysis. Structurally all the components within the system exhibit positive margins of safety. A feasible concept was defined which uses state-of-the-art materials and existing TMC technology. The engine basically consists of a three dimensional carbon/carbon combustor/nozzle secured to an FS-85 columbium inlet. The carbon/carbon liner is sheathed with carbon felt insulation to thermally protect the FS-85 structure and skin. The thermal analysis of the engine indicates that a thermally viable configuration exists.

  9. Affordable Flight Demonstration of the GTX Air-Breathing SSTO Vehicle Concept

    NASA Technical Reports Server (NTRS)

    Krivanek, Thomas M.; Roche, Joseph M.; Riehl, John P.; Kosareo, Daniel N.

    2003-01-01

    The rocket based combined cycle (RBCC) powered single-stage-to-orbit (SSTO) reusable launch vehicle has the potential to significantly reduce the total cost per pound for orbital payload missions. To validate overall system performance, a flight demonstration must be performed. This paper presents an overview of the first phase of a flight demonstration program for the GTX SSTO vehicle concept. Phase 1 will validate the propulsion performance of the vehicle configuration over the supersonic and hypersonic air- breathing portions of the trajectory. The focus and goal of Phase 1 is to demonstrate the integration and performance of the propulsion system flowpath with the vehicle aerodynamics over the air-breathing trajectory. This demonstrator vehicle will have dual mode ramjetkcramjets, which include the inlet, combustor, and nozzle with geometrically scaled aerodynamic surface outer mold lines (OML) defining the forebody, boundary layer diverter, wings, and tail. The primary objective of this study is to demon- strate propulsion system performance and operability including the ram to scram transition, as well as to validate vehicle aerodynamics and propulsion airframe integration. To minimize overall risk and develop ment cost the effort will incorporate proven materials, use existing turbomachinery in the propellant delivery systems, launch from an existing unmanned remote launch facility, and use basic vehicle recovery techniques to minimize control and landing requirements. A second phase would demonstrate propulsion performance across all critical portions of a space launch trajectory (lift off through transition to all-rocket) integrated with flight-like vehicle systems.

  10. Preliminary aerothermodynamic design method for hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Harloff, G. J.; Petrie, S. L.

    1987-01-01

    Preliminary design methods are presented for vehicle aerothermodynamics. Predictions are made for Shuttle orbiter, a Mach 6 transport vehicle and a high-speed missile configuration. Rapid and accurate methods are discussed for obtaining aerodynamic coefficients and heat transfer rates for laminar and turbulent flows for vehicles at high angles of attack and hypersonic Mach numbers.

  11. The Role of Guidance, Navigation, and Control in Hypersonic Vehicle Multidisciplinary Design and Optimization

    NASA Technical Reports Server (NTRS)

    Ouzts, Peter J.; Soloway, Donald I.; Moerder, Daniel D.; Wolpert, David H.; Benavides, Jose Victor

    2009-01-01

    Airbreathing hypersonic systems offer distinct performance advantages over rocket-based systems for space access vehicles. However, these performance advantages are dependent upon advances in current state-of-the-art technologies in many areas such as ram/scramjet propulsion integration, high temperature materials, aero-elastic structures, thermal protection systems, transition to hypersonics and hypersonic control elements within the framework of complex physics and new design methods. The complex interactions between elements of an airbreathing hypersonic vehicle represent a new paradigm in vehicle design to achieve the optimal performance necessary to meet space access mission objectives. In the past, guidance, navigation, and control (GNC) analysis often follows completion of the vehicle conceptual design process. Individual component groups design subsystems which are then integrated into a vehicle configuration. GNC is presented the task of developing control approaches to meet vehicle performance objectives given that configuration. This approach may be sufficient for vehicles where significant performance margins exist. However, for higher performance vehicles engaging the GNC discipline too late in the design cycle has been costly. For example, the X-29 experimental flight vehicle was built as a technology demonstrator. One of the many technologies to be demonstrated was the use of light-weight material composites for structural components. The use of light-weight materials increased the flexibility of the X- 29 beyond that of conventional metal alloy constructed aircraft. This effect was not considered when the vehicle control system was designed and built. The impact of this is that the control system did not have enough control authority to compensate for the effects of the first fundamental structural mode of the vehicle. As a result, the resulting pitch rate response of the vehicle was below specification and no post-design changes could recover the

  12. Sharp Refractory Composite Leading Edges on Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Walker, Sandra P.; Sullivan, Brian J.

    2003-01-01

    On-going research of advanced sharp refractory composite leading edges for use on hypersonic air-breathing vehicles is presented in this paper. Intense magnitudes of heating and of heating gradients on the leading edge lead to thermal stresses that challenge the survivability of current material systems. A fundamental understanding of the problem is needed to further design development. Methodology for furthering the technology along with the use of advanced fiber architectures to improve the thermal-structural response is explored in the current work. Thermal and structural finite element analyses are conducted for several advanced fiber architectures of interest. A tailored thermal shock parameter for sharp orthotropic leading edges is identified for evaluating composite material systems. The use of the tailored thermal shock parameter has the potential to eliminate the need for detailed thermal-structural finite element analyses for initial screening of material systems being considered for a leading edge component.

  13. Hypersonic Airbreathing Propulsion: An Aerodynamics, Aerothermodynamics, and Acoustics Competency White Paper

    NASA Technical Reports Server (NTRS)

    Drummond, J. Philip; Cockrell, Charles E., Jr.; Pellett, Gerald L.; Diskin, Glenn S.; Auslender, Aaron H.; Exton, Reginald J.; Guy, R. Wayne; Hoppe, John C.; Puster, Richard L.; Rogers, R. Clayton

    2002-01-01

    This White Paper examines the current state of Hypersonic Airbreathing Propulsion at the NASA Langley Research Center and the factors influencing this area of work and its personnel. Using this knowledge, the paper explores beyond the present day and suggests future directions and strategies for the field. Broad views are first taken regarding potential missions and applications of hypersonic propulsion. Then, candidate propulsion systems that may be applicable to these missions are suggested and discussed. Design tools and experimental techniques for developing these propulsion systems are then described, and approaches for applying them in the design process are considered. In each case, current strategies are reviewed and future approaches that may improve the techniques are considered. Finally, the paper concentrates on the needs to be addressed in each of these areas to take advantage of the opportunities that lay ahead for both the NASA Langley Research Center and the Aerodynamic Aerothermodynamic, and Aeroacoustics Competency. Recommendations are then provided so that the goals set forth in the paper may be achieved.

  14. Control Design for a Non-Minimum Phase Hypersonic Vehicle Model

    NASA Astrophysics Data System (ADS)

    McKenna, Thomas

    Air-breathing hypersonic vehicles are emerging as a method for cost-efficient access to space. Great strides have recently been made in the field of hypersonic vehicles, however the unique dynamics of the vehicles present challenges for control design. In this thesis, a nonlinear controller for a hypersonic vehicle model is designed using the Indirect Manifold Construction approach. The high fidelity hypersonic vehicle model considered in this thesis includes many of the challenging effects of hypersonic flight. The main challenge to control design is the vehicle's unstable internal dynamics. This non-minimum phase behavior prevents the use of many standard forms of nonlinear control techniques. The nonlinear controller developed in this thesis following the Indirect Manifold Construction approach uses a hierarchical control design to force outputs to commanded values while ensuring the internal dynamics of the system remain stable. The nonlinear controller is shown to be effective in simulation. The closed loop system is also shown to be stable through a Lyapunov based stability analysis.

  15. Robust intelligent flight control for hypersonic vehicles. Ph.D. Thesis - Massachusetts Inst. of Technology

    NASA Technical Reports Server (NTRS)

    Chamitoff, Gregory Errol

    1992-01-01

    Intelligent optimization methods are applied to the problem of real-time flight control for a class of airbreathing hypersonic vehicles (AHSV). The extreme flight conditions that will be encountered by single-stage-to-orbit vehicles, such as the National Aerospace Plane, present a tremendous challenge to the entire spectrum of aerospace technologies. Flight control for these vehicles is particularly difficult due to the combination of nonlinear dynamics, complex constraints, and parametric uncertainty. An approach that utilizes all available a priori and in-flight information to perform robust, real time, short-term trajectory planning is presented.

  16. Static and Hypersonic Experimental Analysis of Impulse Generation in Air-Breathing Laser-Thermal Propulsion

    NASA Astrophysics Data System (ADS)

    Salvador, Israel Irone

    (98 to 161 mm in diameter), probably due to the more efficient delivery of laser-induced blast wave energy across the 2D model's larger impulse surface area. Next, the hypersonic campaign was carried out, subjecting the 2D model to nominal Mach numbers ranging from 6 to 10. Again, time-dependent surface pressure distributions were recorded together with Schlieren movies of the flow field structure resulting from laser energy deposition. These visualizations of inlet and absorption chamber flowfields, enabled the qualitative analysis of important phenomena impacting laser-propelled hypersonic airbreathing flight. The laser-induced breakdown took an elongated vertically-oriented geometry, occurring off-surface and across the inlet's mid-channel---quite different from the static case in which the energy was deposited very near the shroud under-surface. The shroud under-surface pressure data indicated laser-induced increases of 0.7-0.9 bar with laser pulse energies of ˜170 J, off-shroud induced breakdown condition, and Mach number of 7. The results of this research corroborate the feasibility of laser powered, airbreathing flight with infinite specific impulse (Isp=infinity): i.e., without the need for propellant injection at the laser focus. Additionally, it is shown that further reductions in inlet air working fluid velocity---with attendant increases in static pressure and density---is necessary to generate higher absorption chamber pressure and engine impulse. Finally, building on lessons learned from the present work, the future research plan is laid out for: a) the present 2D model with full inlet forebody, exploring higher laser pulse energies and multi-pulse phenomena; b) a smaller, redesigned 2D model; c) a 254 mm diameter axisymmetric Lightcraft model; and, d) a laser-electromagnetic accelerator model, designed around a 2-Tesla pulsed electromagnet contracted under the present program.

  17. Evaluation of an Ejector Ramjet Based Propulsion System for Air-Breathing Hypersonic Flight

    NASA Technical Reports Server (NTRS)

    Thomas, Scott R.; Perkins, H. Douglas; Trefny, Charles J.

    1997-01-01

    A Rocket Based Combined Cycle (RBCC) engine system is designed to combine the high thrust to weight ratio of a rocket along with the high specific impulse of a ramjet in a single, integrated propulsion system. This integrated, combined cycle propulsion system is designed to provide higher vehicle performance than that achievable with a separate rocket and ramjet. The RBCC engine system studied in the current program is the Aerojet strutjet engine concept, which is being developed jointly by a government-industry team as part of the Air Force HyTech program pre-PRDA activity. The strutjet is an ejector-ramjet engine in which small rocket chambers are embedded into the trailing edges of the inlet compression struts. The engine operates as an ejector-ramjet from take-off to slightly above Mach 3. Above Mach 3 the engine operates as a ramjet and transitions to a scramjet at high Mach numbers. For space launch applications the rockets would be re-ignited at a Mach number or altitude beyond which air-breathing propulsion alone becomes impractical. The focus of the present study is to develop and demonstrate a strutjet flowpath using hydrocarbon fuel at up to Mach 7 conditions. Freejet tests of a candidate flowpath for this RBCC engine were conducted at the NASA Lewis Research Center's Hypersonic Tunnel Facility between July and September 1996. This paper describes the engine flowpath and installation, outlines the primary objectives of the program, and describes the overall results of this activity. Through this program 15 full duration tests, including 13 fueled tests were made. The first major achievement was the further demonstration of the HTF capability. The facility operated at conditions up to 1950 K and 7.34 MPa, simulating approximately Mach 6.6 flight. The initial tests were unfueled and focused on verifying both facility and engine starting. During these runs additional aerodynamic appliances were incorporated onto the facility diffuser to enhance starting

  18. Discrete Roughness Transition for Hypersonic Flight Vehicles

    NASA Technical Reports Server (NTRS)

    Berry, Scott A.; Horvath, Thomas J.

    2007-01-01

    The importance of discrete roughness and the correlations developed to predict the onset of boundary layer transition on hypersonic flight vehicles are discussed. The paper is organized by hypersonic vehicle applications characterized in a general sense by the boundary layer: slender with hypersonic conditions at the edge of the boundary layer, moderately blunt with supersonic, and blunt with subsonic. This paper is intended to be a review of recent discrete roughness transition work completed at NASA Langley Research Center in support of agency flight test programs. First, a review is provided of discrete roughness wind tunnel data and the resulting correlations that were developed. Then, results obtained from flight vehicles, in particular the recently flown Hyper-X and Shuttle missions, are discussed and compared to the ground-based correlations.

  19. Conceptual study of space plane powered by hypersonic airbreathing propulsion system

    NASA Astrophysics Data System (ADS)

    Maita, Masataka; Ohkami, Yoshiaki; Yamanaka, Tatsuo; Mori, Takashige

    1990-10-01

    The paper describes the investigations of aerospace plane concept, conducted by the National Aerospace Laboratory (NAL) of Japan, with particular attention given to a concept which integrates a scram/liquid air cycle engine (LACE) hypersonic propulsion system fueling with slush hydrogen. The key requirements in achieving the space plane using scram/LACE propulsion system are described along with the mission requirements and the vehicle characteristics. Typical outputs of SSTO analysis are presented.

  20. A hypersonic vehicle approach to planetary exploration

    NASA Technical Reports Server (NTRS)

    Murbach, Marcus S.

    1993-01-01

    An enhanced Mars network class mission using a lifting hypersonic entry vehicle is proposed. The basic vehicle, derived from a mature hypersonic flight system called SWERVE, offers several advantages over more conventional low L/D or ballistic entry systems. The proposed vehicle has greatly improved lateral and cross range capability (e.g., it is capable of reaching the polar regions during less than optimal mission opportunities), is not limited to surface target areas of low elevation, and is less susceptible to problems caused by Martian dust storms. Further, the integrated vehicle has attractive deployment features and allows for a much improved evolutionary path to larger vehicles with greater science capability. Analysis of the vehicle is aided by the development of a Mars Hypersonic Flight Simulator from which flight trajectories are obtained. Atmospheric entry performance of the baseline vehicle is improved by a deceleration skirt and transpiration cooling system which significantly reduce TPS (Thermal Protection System) and flight battery mass. The use of the vehicle is also attractive in that the maturity of the flight systems make it cost-competitive with the development of a conventional low L/D entry system. Finally, the potential application of similar vehicles to other planetary missions is discussed.

  1. Effect of aeroelastic-propulsive interactions on flight dynamics of a hypersonic vehicle

    NASA Technical Reports Server (NTRS)

    Raney, David L.; Mcminn, John D.; Pototzky, Anthony S.; Wooley, Christine L.

    1993-01-01

    The desire to achieve orbit-on-demand access to space with rapid turn-around capability and aircraft-like processing operations has given rise to numerous hypersonic aerospace plane design concepts which would take off horizontally from a conventional runway and employ air-breathing scramjet propulsion systems for acceleration to orbital speeds. Most of these air-breathing hypersonic vehicle concepts incorporate an elongated fuselage forebody to act as the aerodynamic compression surface for a scramjet combustor module. This type of airframe-integrated scramjet propulsion system tends to be highly sensitive to inlet conditions and angle-of-attack perturbations. Furthermore, the basic configuration of the fuselage, with its elongated and tapered forebody, produces relatively low frequency elastic modes which will cause perturbations in the combustor inlet conditions due to the oscillation of the forebody compression surface. The flexibility of the forebody compression surface, together with sensitivity of scramjet propulsion systems to inlet conditions, creates the potential for an unprecedented form of aeroelastic-propulsive interaction in which deflections of the vehicle fuselage give rise to propulsion transients, producing force and moment variations that may adversely impact the longitudinal flight dynamics and/or excite the elastic modes. These propulsive force and moment variations may have an appreciable impact on the performance, guidance, and control of a hypersonic aerospace plane. The objectives of this research are to quantify the magnitudes of propulsive force and moment perturbations resulting from elastic deformation of a representative hypersonic vehicle, and to assess the potential impact of these perturbations on the vehicle's longitudinal flight dynamics.

  2. Hypersonic drone vehicle design: A multidisciplinary experience

    NASA Technical Reports Server (NTRS)

    1988-01-01

    UCLA's Advanced Aeronautic Design group focussed their efforts on design problems of an unmanned hypersonic vehicle. It is felt that a scaled hypersonic drone is necesary to bridge the gap between present theory on hypersonics and the future reality of the National Aerospace Plane (NASP) for two reasons: (1) to fulfill a need for experimental data in the hypersonic regime, and (2) to provide a testbed for the scramjet engine which is to be the primary mode of propulsion for the NASP. The group concentrated on three areas of great concern to NASP design: propulsion, thermal management, and flight systems. Problem solving in these areas was directed toward design of the drone with the idea that the same design techniques could be applied to the NASP. A 70 deg swept double-delta wing configuration, developed in the 70's at the NASA Langley, was chosen as the aerodynamic and geometric model for the drone. This vehicle would be air launched from a B-1 at Mach 0.8 and 48,000 feet, rocket boosted by two internal engines to Mach 10 and 100,000 feet, and allowed to cruise under power of the scramjet engine until burnout. It would then return to base for an unpowered landing. Preliminary energy calculations based on flight requirements give the drone a gross launch weight of 134,000 pounds and an overall length of 85 feet.

  3. Aeroservoelastic stabilization techniques for hypersonic flight vehicles

    NASA Technical Reports Server (NTRS)

    Cheng, Peter Y.; Chan, Samuel Y.; Myers, Thomas T.; Klyde, David H.; Mcrue, Duane T.

    1991-01-01

    The potential of Hybrid Phase Stabilization (HPS), particularly for highly unstable aircraft, using a hypersonic flight vehicle (HSV) as a relevant example, is discussed. The development of HPS is presented and the result is compared with that generated using a conventional gain stabilization technique. Since HPS was not addressed in the MIL-spec requirements, a preliminary residual response metric was developed to provide guidance in assessing HPS.

  4. System controls challenges of hypersonic combined-cycle engine powered vehicles

    NASA Technical Reports Server (NTRS)

    Morrison, Russell H.; Ianculescu, George D.

    1992-01-01

    Hypersonic aircraft with air-breathing engines have been described as the most complex and challenging air/space vehicle designs ever attempted. This is particularly true for aircraft designed to accelerate to orbital velocities. The propulsion system for the National Aerospace Plane will be an active factor in maintaining the aircraft on course. Typically addressed are the difficulties with the aerodynamic vehicle design and development, materials limitations and propulsion performance. The propulsion control system requires equal materials limitations and propulsion performance. The propulsion control system requires equal concern. Far more important than merely a subset of propulsion performance, the propulsion control system resides at the crossroads of trajectory optimization, engine static performance, and vehicle-engine configuration optimization. To date, solutions at these crossroads are multidisciplinary and generally lag behind the broader performance issues. Just how daunting these demands will be is suggested. A somewhat simplified treatment of the behavioral characteristics of hypersonic aircraft and the issues associated with their air-breathing propulsion control system design are presented.

  5. Materials and structures for hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Tenney, D. R.; Lisagor, W. B.; Dixon, S. C.

    1988-01-01

    Hypersonic vehicles are envisioned to require, in addition to carbon-carbon and ceramic-matrix composites for leading edges heated to above 2000 F, such 600-1800 F operating temperature materials as advanced Ti alloys, nickel aluminides, and metal-matrix composites; these possess the necessary low density and high strength and stiffness. The primary design drivers are maximum vehicle heating rate, total heat load, flight envelope, propulsion system type, mission life requirements, and liquid hydrogen containment system. Attention is presently given to aspects of these materials and structures requiring more intensive development.

  6. Materials and structures for hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Tenney, Darrel R.; Lisagor, W. Barry; Dixon, Sidney C.

    1988-01-01

    Hypersonic vehicles are envisioned to require, in addition to carbon-carbon and ceramic-matrix composities for leading edges heated to above 2000 F, such 600 to 1800 F operating temperature materials as advanced Ti alloys, nickel aluminides, and metal-matrix composited; These possess the necessary low density and high strength and stiffness. The primary design drivers are maximum vehicle heating rate, total heat load, flight envelope, propulsion system type, mission life requirements and liquid hydrogen containment systems. Attention is presently given to aspects of these materials and structures requiring more intensive development.

  7. Hypersonic Vehicle Propulsion System Control Model Development Roadmap and Activities

    NASA Technical Reports Server (NTRS)

    Stueber, Thomas J.; Le, Dzu K.; Vrnak, Daniel R.

    2009-01-01

    The NASA Fundamental Aeronautics Program Hypersonic project is directed towards fundamental research for two classes of hypersonic vehicles: highly reliable reusable launch systems (HRRLS) and high-mass Mars entry systems (HMMES). The objective of the hypersonic guidance, navigation, and control (GN&C) discipline team is to develop advanced guidance and control algorithms to enable efficient and effective operation of these challenging vehicles. The ongoing work at the NASA Glenn Research Center supports the hypersonic GN&C effort in developing tools to aid the design of advanced control algorithms that specifically address the propulsion system of the HRRLSclass vehicles. These tools are being developed in conjunction with complementary research and development activities in hypersonic propulsion at Glenn and elsewhere. This report is focused on obtaining control-relevant dynamic models of an HRRLS-type hypersonic vehicle propulsion system.

  8. Hypersonic Vehicle Propulsion System Simplified Model Development

    NASA Technical Reports Server (NTRS)

    Stueber, Thomas J.; Raitano, Paul; Le, Dzu K.; Ouzts, Peter

    2007-01-01

    This document addresses the modeling task plan for the hypersonic GN&C GRC team members. The overall propulsion system modeling task plan is a multi-step process and the task plan identified in this document addresses the first steps (short term modeling goals). The procedures and tools produced from this effort will be useful for creating simplified dynamic models applicable to a hypersonic vehicle propulsion system. The document continues with the GRC short term modeling goal. Next, a general description of the desired simplified model is presented along with simulations that are available to varying degrees. The simulations may be available in electronic form (FORTRAN, CFD, MatLab,...) or in paper form in published documents. Finally, roadmaps outlining possible avenues towards realizing simplified model are presented.

  9. Evaluation of an Ejector Ramjet Based Propulsion System for Air-Breathing Hypersonic Flight

    NASA Technical Reports Server (NTRS)

    Thomas, Scott R.; Perkins, H. Douglas; Trefny, Charles J.

    1997-01-01

    A Rocket Based Combined Cycle (RBCC) engine system is designed to combine the high thrust to weight ratio of a rocket along with the high specific impulse of a ramjet in a single, integrated propulsion system. This integrated, combined cycle propulsion system is designed to provide higher vehicle performance than that achievable with a separate rocket and ramjet. The RBCC engine system studied in the current program is the Aerojet strutjet engine concept, which is being developed jointly by a government-industry team as part of the Air Force HyTech program pre-PRDA activity. The strutjet is an ejector-ramjet engine in which small rocket chambers are embedded into the trailing edges of the inlet compression struts. The engine operates as an ejector-ramjet from takeoff to slightly above Mach 3. Above Mach 3 the engine operates as a ramjet and transitions to a scramjet at high Mach numbers. For space launch applications the rockets would be re-ignited at a Mach number or altitude beyond which air-breathing propulsion alone becomes impractical. The focus of the present study is to develop and demonstrate a strutjet flowpath using hydrocarbon fuel at up to Mach 7 conditions.

  10. Current Hypersonic and Space Vehicle Flight Test and Instrumentation

    DTIC Science & Technology

    2015-06-22

    412TW-PA-15264 CURRENT HYPERSONIC AND SPACE VEHICLE FLIGHT TEST AND INSTRUMENTATION John J. Spravka* and Timothy R. Jorris† AIR FORCE TEST...DATES COVERED (From - To) 22 – 26 July 2015 4. TITLE AND SUBTITLE Current Hypersonic and Space Vehicle Flight Test and Instrumentation...utility can be leveraged by a wide range of flight test programs. 15. SUBJECT TERMS Hypersonic, flight test, instrumentation, space access, space

  11. Configuration development study of the OSU 1 hypersonic research vehicle

    NASA Technical Reports Server (NTRS)

    Stein, Matthew D.; Frankhauser, Chris; Zee, Warner; Kosanchick, Melvin, III; Nelson, Nick; Hunt, William

    1993-01-01

    In an effort to insure the future development of hypersonic cruise aircraft, the possible vehicle configurations were examined to develop a single-stage-to-orbit hypersonic research vehicle (HRV). Based on the needs of hypersonic research and development, the mission goals and requirements are determined. A body type is chosen. Three modes of propulsion and two liquid rocket fuels are compared, followed by the optimization of the body configuration through aerodynamic, weight, and trajectory studies. A cost analysis is included.

  12. Hyper-X and Pegasus Launch Vehicle: A Three-Foot Model of the Hypersonic Experimental Research Vehic

    NASA Technical Reports Server (NTRS)

    1997-01-01

    The configuration of the X-43A Hypersonic Experimental Research Vehicle, or Hyper-X, attached to a Pegasus launch vehicle is displayed in this three-foot-long model at NASA's Dryden Flight Research Center, Edwards, California. Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43

  13. Trajectory Optimization and Conceptual Study of Small Test Vehicles for Hypersonic Engine Using High-Altitude Balloon

    NASA Astrophysics Data System (ADS)

    Tsuchiya, Takeshi; Takenaka, Youichi; Taguchi, Hideyuki; Sawai, Shujiro

    Japan Aerospace Exploration Agency, JAXA announced a long-term vision recently. In the vision, JAXA aims to develop hypersonic aircrafts. A pre-cooled turbojet engine has great potential as one of newly developed hypersonic air-breathing engines. We also expect the engine to be installed in space transportation vehicles in future. For combustion test in real flight condition of the engines, JAXA has an experimental plan with a small test vehicle falling from a high-altitude balloon. This paper applies numerical analysis and optimization techniques to conceptual designs of the test vehicle in order to obtain the best configuration and trajectory that can achieve the flight test. The results show helpful knowledge when we design prototype vehicles.

  14. Trajectory Optimization and Conceptual Study of Small Test Vehicles for a Hypersonic Engine Using a High-Altitude Balloon

    NASA Astrophysics Data System (ADS)

    Tsuchiya, Takeshi; Takenaka, Youichi; Taguchi, Hideyuki; Sawai, Shujiro

    The Japan Aerospace Exploration Agency, JAXA, announced a long-term vision recently. In the vision, JAXA aims to develop hypersonic aircrafts. A pre-cooled turbojet engine has great potential as one of newly developed hypersonic airbreathing engines. We also expect the engine to be installed in space transportation vehicles in the future. For combustion test in the real flight conditions of the engines, JAXA has an experimental plan where a small test vehicle is released from a high-altitude balloon. This paper applies numerical analysis and optimization techniques to conceptual designs of the test vehicle in order to obtain the best configuration and trajectory for the flight test. The results show helpful knowledge for designing prototype vehicles.

  15. A Comparison of Active and Passive Methods for Control of Hypersonic Boundary Layers on Airbreathing Configurations

    NASA Technical Reports Server (NTRS)

    Berry, Scott A.; Nowak, Robert J.

    2003-01-01

    Active and passive methods for control of hypersonic boundary layers have been experimentally examined in NASA Langley Research Center wind tunnels on a Hyper-X model. Several configurations for forcing transition using passive discrete roughness elements and active mass addition, or blowing, methods were compared in two hypersonic facilities, the 20-Inch Mach 6 Air and the 31-Inch Mach 10 Air tunnels. Heat transfer distributions, obtained via phosphor thermography, shock system details, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. The comparisons between the active and passive methods for boundary layer control were conducted at test conditions that nearly match the nominal Mach 7 flight trajectory of an angle-of-attack of 2-deg and length Reynolds number of 5.6 million. For the passive roughness examination, the primary parametric variation was a range of trip heights within the calculated boundary layer thickness for several trip concepts. The prior passive roughness study resulted in a swept ramp configuration being selected for the Mach 7 flight vehicle that was scaled to be roughly 0.6 of the calculated boundary layer thickness. For the active jet blowing study, the blowing manifold pressure was systematically varied for each configuration, while monitoring the mass flow, to determine the jet penetration height with schlieren and transition movement with the phosphor system for comparison to the passive results. All the blowing concepts tested were adequate for providing transition onset near the trip location with manifold stagnation pressures on the order of 40 times the model static pressure or higher.

  16. Aerodynamic shape optimization of arbitrary hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Dulikravich, George S.; Sheffer, Scott G.

    1991-01-01

    A new method was developed to optimize, in terms of aerodynamic wave drag minimization, arbitrary (nonaxisymmetric) hypersonic vehicles in modified Newtonian flow, while maintaining the initial volume and length of the vehicle. This new method uses either a surface fitted Fourier series to represent the vehicle's geometry or an independent point motion algorithm. In either case, the coefficients of the Fourier series or the spatial locations of the points defining each cross section were varied and a numerical optimization algorithm based on a quasi-Newton gradient search concept was used to determine the new optimal configuration. Results indicate a significant decrease in aerodynamic wave drag for simple and complex geometries at relatively low CPU costs. In the case of a cone, the results agreed well with known analytical optimum ogive shapes. The procedure is capable of accepting more complex flow field analysis codes.

  17. Materials Development for Hypersonic Flight Vehicles

    NASA Technical Reports Server (NTRS)

    Glass, David E.; Dirling, Ray; Croop, Harold; Fry, Timothy J.; Frank, Geoffrey J.

    2006-01-01

    The DARPA/Air Force Falcon program is planning to flight test several hypersonic technology vehicles (HTV) in the next several years. A Materials Integrated Product Team (MIPT) was formed to lead the development of key thermal protection system (TPS) and hot structures technologies. The technologies being addressed by the MIPT are in the following areas: 1) less than 3000 F leading edges, 2) greater than 3000 F refractory composite materials, 3) high temperature multi-layer insulation, 4) acreage TPS, and 5) high temperature seals. Technologies being developed in each of these areas are discussed in this paper.

  18. Trajectory Approaches for Launching Hypersonic Flight Tests (Preprint)

    DTIC Science & Technology

    2014-08-01

    mostly based on boost- glide type vehicles and hypersonic airbreathing vehicles technology needs. The trajectories presented in this paper will be...AFRL-RQ-WP-TP-2014-0184 TRAJECTORY APPROACHES FOR LAUNCHING HYPERSONIC FLIGHT TESTS (PREPRINT) Barry M. Hellman Vehicle Technology...SUBTITLE TRAJECTORY APPROACHES FOR LAUNCHING HYPERSONIC FLIGHT TESTS (PREPRINT) 5a. CONTRACT NUMBER In-house 5b. GRANT NUMBER 5c. PROGRAM

  19. A conceptual design of an unmanned test vehicle using an airbreathing propulsion system

    NASA Technical Reports Server (NTRS)

    1992-01-01

    According to Aviation Week and Space Technology (Nov. 16, 1992), without a redefined approach to the problem of achieving single stage-to-orbit flight, the X-30 program is virtually assured of cancellation. One of the significant design goals of the X-30 program is to achieve single stage to low-earth orbit using airbreathing propulsion systems. In an attempt to avoid cancellation, the NASP Program has decided to design a test vehicle to achieve these goals. This report recommends a conceptual design of an unmanned test vehicle using an airbreathing propulsion system.

  20. Hypersonic Laminar-Turbulent Transition on Slender Cones at Zero Angle of Attack: Measurements in Support of Mechanism-Based Models for Scaling Ground-Test Data to Flight

    DTIC Science & Technology

    2011-03-04

    Public Release, Distribution Unlimited 13. SUPPLEMENTARY NOTES 14. ABSTRACT Maintaining low laminar heating may be critical to hypersonic gliding ...critical to hypersonic gliding reentry vehicles such as the several which have been proposed for the Prompt Global Strike (PGS) mission. Tran- sition...transition is critical to gliding hypersonic reentry vehicles and hyper- sonic airbreathing cruise vehicles, such as those presently being developed under

  1. Conceptual Study on Hypersonic Turbojet Experimental Vehicle (HYTEX)

    NASA Astrophysics Data System (ADS)

    Taguchi, Hideyuki; Murakami, Akira; Sato, Tetsuya; Tsuchiya, Takeshi

    Pre-cooled turbojet engines have been investigated aiming at realization of reusable space transportation systems and hypersonic airplanes. Evaluation methods of these engine performances have been established based on ground tests. There are some plans on the demonstration of hypersonic propulsion systems. JAXA focused on hypersonic propulsion systems as a key technology of hypersonic transport airplane. Demonstrations of Mach 5 class hypersonic technologies are stated as a development target at 2025 in the long term vision. In this study, systems analyses of hypersonic turbojet experiment (HYTEX) with Mach 5 flight capability is performed. Aerodynamic coefficients are obtained by CFD analyses and wind tunnel tests. Small Pre-cooled turbojet is fabricated and tested using liquid hydrogen as fuel. As a result, characteristics of the baseline vehicle shape is clarified, . and effects of pre-cooling are confirmed at the firing test.

  2. Seal Technology for Hypersonic Vehicle and Propulsion: An Overview

    NASA Technical Reports Server (NTRS)

    Steinetz, Bruce M.

    2008-01-01

    Hypersonic vehicles and propulsion systems pose an extraordinary challenge for structures and materials. Airframes and engines require lightweight, high-temperature materials and structural configurations that can withstand the extreme environment of hypersonic flight. Some of the challenges posed include very high temperatures, heating of the whole vehicle, steady-state and transient localized heating from shock waves, high aerodynamic loads, high fluctuating pressure loads, potential for severe flutter, vibration, and acoustic loads and erosion. Correspondingly high temperature seals are required to meet these aggressive requirements. This presentation reviews relevant seal technology for both heritage (e.g. Space Shuttle, X-15, and X-38) vehicles and presents several seal case studies aimed at providing lessons learned for future hypersonic vehicle seal development. This presentation also reviews seal technology developed for the National Aerospace Plane propulsion systems and presents several seal case studies aimed at providing lessons learned for future hypersonic propulsion seal development.

  3. Analysis of Plasma Communication Schemes for Hypersonic Vehicles: Final Report

    DTIC Science & Technology

    2009-02-01

    the ReComm scheme for communications through the plasma sheath surrounding a hypersonic vehicle during re-entry. We demonstrate that the time...physical processes of the ReComm scheme for communications through the plasma sheath surrounding a hypersonic vehicle during re-entry. The ReComm scheme...relation is derived to estimate the plasma heating in the sheath due to plasma waves excited by the antenna. Contents I. Introduction 4 II. Electron

  4. An Overview of the NASA FAP Hypersonics Project Airbreathing Propulsion Research

    NASA Technical Reports Server (NTRS)

    Auslender, A. H.; Suder, Kenneth L.; Thomas, Scott R.

    2009-01-01

    The propulsion research portfolio of the National Aeronautics and Space Administration Fundamental Aeronautics Program Hypersonics Project encompasses a significant number of technical tasks that are aligned to achieve mastery and intellectual stewardship of the core competencies in the hypersonic-flight regime. An overall coordinated programmatic and technical effort has been structured to advance the state-of-the-art, via both experimental and analytical efforts. A subset of the entire hypersonics propulsion research portfolio is presented in this overview paper. To this end, two programmatic research disciplines are discussed; namely, (1) the Propulsion Discipline, including three associated research elements: the X-51A partnership, the HIFiRE-2 partnership, and the Durable Combustor Rig, and (2) the Turbine-Based Combine Cycle Discipline, including three associated research elements: the Combined Cycle Engine Large Scale Inlet Mode Transition Experiment, the small-scale Inlet Mode Transition Experiment, and the High-Mach Fan Rig.

  5. A review of design issues specific to hypersonic flight vehicles

    NASA Astrophysics Data System (ADS)

    Sziroczak, D.; Smith, H.

    2016-07-01

    This paper provides an overview of the current technical issues and challenges associated with the design of hypersonic vehicles. Two distinct classes of vehicles are reviewed; Hypersonic Transports and Space Launchers, their common features and differences are examined. After a brief historical overview, the paper takes a multi-disciplinary approach to these vehicles, discusses various design aspects, and technical challenges. Operational issues are explored, including mission profiles, current and predicted markets, in addition to environmental effects and human factors. Technological issues are also reviewed, focusing on the three major challenge areas associated with these vehicles: aerothermodynamics, propulsion, and structures. In addition, matters of reliability and maintainability are also presented. The paper also reviews the certification and flight testing of these vehicles from a global perspective. Finally the current stakeholders in the field of hypersonic flight are presented, summarizing the active programs and promising concepts.

  6. Hypersonic Materials and Structures

    NASA Technical Reports Server (NTRS)

    Glass, David E.

    2016-01-01

    Thermal protection systems (TPS) and hot structures are required for a range of hypersonic vehicles ranging from ballistic reentry to hypersonic cruise vehicles, both within Earth's atmosphere and non-Earth atmospheres. The focus of this presentation is on air breathing hypersonic vehicles in the Earth's atmosphere. This includes single-stage to orbit (SSTO), two-stage to orbit (TSTO) accelerators, access to space vehicles, and hypersonic cruise vehicles. This paper will start out with a brief discussion of aerodynamic heating and thermal management techniques to address the high heating, followed by an overview of TPS for rocket-launched and air-breathing vehicles. The argument is presented that as we move from rocket-based vehicles to air-breathing vehicles, we need to move away from the insulated airplane approach used on the Space Shuttle Orbiter to a wide range of TPS and hot structure approaches. The primary portion of the paper will discuss issues and design options for CMC TPS and hot structure components, including leading edges, acreage TPS, and control surfaces. The current state-of-the-art will be briefly discussed for some of the components.

  7. An Air-Breathing Launch Vehicle Concept for Single-Stage-to-Orbit

    NASA Technical Reports Server (NTRS)

    Trefny, Charles J.

    1999-01-01

    The "Trailblazer" is a 300-lb payload, single-stage-to-orbit launch vehicle concept that uses air-breathing propulsion to reduce the required propellant fraction. The integration of air-breathing propulsion is done considering performance, structural and volumetric efficiency, complexity, and design risk. The resulting configuration is intended to be viable using near-term materials and structures. The aeropropulsion performance goal for the Trailblazer launch vehicle is an equivalent effective specific impulse (I*) of 500 sec. Preliminary analysis shows that this requires flight in the atmosphere to about Mach 10, and that the gross lift-off weight is 130,000 lb. The Trailblazer configuration and proposed propulsion system operating modes are described. Preliminary performance results are presented, and key technical issues are highlighted. An overview of the proposed program plan is given.

  8. Computational Aerothermodynamic Design Issues for Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Gnoffo, Peter A.; Weilmuenster, K. James; Hamilton, H. Harris, II; Olynick, David R.; Venkatapathy, Ethiraj

    1997-01-01

    A brief review of the evolutionary progress in computational aerothermodynamics is presented. The current status of computational aerothermodynamics is then discussed, with emphasis on its capabilities and limitations for contributions to the design process of hypersonic vehicles. Some topics to be highlighted include: (1) aerodynamic coefficient predictions with emphasis on high temperature gas effects; (2) surface heating and temperature predictions for thermal protection system (TPS) design in a high temperature, thermochemical nonequilibrium environment; (3) methods for extracting and extending computational fluid dynamic (CFD) solutions for efficient utilization by all members of a multidisciplinary design team; (4) physical models; (5) validation process and error estimation; and (6) gridding and solution generation strategies. Recent experiences in the design of X-33 will be featured. Computational aerothermodynamic contributions to Mars Pathfinder, METEOR, and Stardust (Comet Sample return) will also provide context for this discussion. Some of the barriers that currently limit computational aerothermodynamics to a predominantly reactive mode in the design process will also be discussed, with the goal of providing focus for future research.

  9. Computational Aerothermodynamic Design Issues for Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Olynick, David R.; Venkatapathy, Ethiraj

    2004-01-01

    A brief review of the evolutionary progress in computational aerothermodynamics is presented. The current status of computational aerothermodynamics is then discussed, with emphasis on its capabilities and limitations for contributions to the design process of hypersonic vehicles. Some topics to be highlighted include: (1) aerodynamic coefficient predictions with emphasis on high temperature gas effects; (2) surface heating and temperature predictions for thermal protection system (TPS) design in a high temperature, thermochemical nonequilibrium environment; (3) methods for extracting and extending computational fluid dynamic (CFD) solutions for efficient utilization by all members of a multidisciplinary design team; (4) physical models; (5) validation process and error estimation; and (6) gridding and solution generation strategies. Recent experiences in the design of X-33 will be featured. Computational aerothermodynamic contributions to Mars Pathfinder, METEOR, and Stardust (Comet Sample return) will also provide context for this discussion. Some of the barriers that currently limit computational aerothermodynamics to a predominantly reactive mode in the design process will also be discussed, with the goal of providing focus for future research.

  10. Computational Aerothermodynamic Design Issues for Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Gnoffo, Peter A.; Weilmuenster, K. James; Hamilton, H. Harris, II; Olynick, David R.; Venkatapathy, Ethiraj

    2005-01-01

    A brief review of the evolutionary progress in computational aerothermodynamics is presented. The current status of computational aerothermodynamics is then discussed, with emphasis on its capabilities and limitations for contributions to the design process of hypersonic vehicles. Some topics to be highlighted include: (1) aerodynamic coefficient predictions with emphasis on high temperature gas effects; (2) surface heating and temperature predictions for thermal protection system (TPS) design in a high temperature, thermochemical nonequilibrium environment; (3) methods for extracting and extending computational fluid dynamic (CFD) solutions for efficient utilization by all members of a multidisciplinary design team; (4) physical models; (5) validation process and error estimation; and (6) gridding and solution generation strategies. Recent experiences in the design of X-33 will be featured. Computational aerothermodynamic contributions to Mars Path finder, METEOR, and Stardust (Comet Sample return) will also provide context for this discussion. Some of the barriers that currently limit computational aerothermodynamics to a predominantly reactive mode in the design process will also be discussed, with the goal of providing focus for future research.

  11. Hypersonic Vehicle Trajectory Optimization and Control

    NASA Technical Reports Server (NTRS)

    Balakrishnan, S. N.; Shen, J.; Grohs, J. R.

    1997-01-01

    Two classes of neural networks have been developed for the study of hypersonic vehicle trajectory optimization and control. The first one is called an 'adaptive critic'. The uniqueness and main features of this approach are that: (1) they need no external training; (2) they allow variability of initial conditions; and (3) they can serve as feedback control. This is used to solve a 'free final time' two-point boundary value problem that maximizes the mass at the rocket burn-out while satisfying the pre-specified burn-out conditions in velocity, flightpath angle, and altitude. The second neural network is a recurrent network. An interesting feature of this network formulation is that when its inputs are the coefficients of the dynamics and control matrices, the network outputs are the Kalman sequences (with a quadratic cost function); the same network is also used for identifying the coefficients of the dynamics and control matrices. Consequently, we can use it to control a system whose parameters are uncertain. Numerical results are presented which illustrate the potential of these methods.

  12. Effects of hypersonic vehicle's optical dome on infrared imaging

    NASA Astrophysics Data System (ADS)

    Zhang, Zhenjun; Cao, Zhiguo; Wang, Wenwu

    2011-09-01

    When an optically guided hypersonic vehicle flies in the atmosphere, the scene is viewed through an optical dome. Because of hypersonic friction with the atmosphere, the optical dome is inevitably covered by a serious shock wave, which threatens to alter the dome's physical parameters and further induce wavefront distortion and degradation of images. By studying the physical phenomena occurring within the optical dome in such an adverse environment, this paper identifies the relationship between the variation of the dome's optical characteristics and the infrared image degradation. The research indicates that the image quality degrades sharply as the vehicle's Mach number increases. Simulations also show that while the thermo-optic effect, elastic-optic effect, thermal deformation, and variation of transmittance have little effect on the optical system, the thermal radiation severely degrades images when vehicles fly at hypersonic speeds. Photo-Optical Instrumentation Engineers

  13. Analysis of the Effects of Vitiates on Surface Heat Flux in Ground Tests of Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Cuda, Vincent; Gaffney, Richard L

    2008-01-01

    To achieve the high enthalpy conditions associated with hypersonic flight, many ground test facilities burn fuel in the air upstream of the test chamber. Unfortunately, the products of combustion contaminate the test gas and alter gas properties and the heat fluxes associated with aerodynamic heating. The difference in the heating rates between clean air and a vitiated test medium needs to be understood so that the thermal management system for hypersonic vehicles can be properly designed. This is particularly important for advanced hypersonic vehicle concepts powered by air-breathing propulsion systems that couple cooling requirements, fuel flow rates, and combustor performance by flowing fuel through sub-surface cooling passages to cool engine components and preheat the fuel prior to combustion. An analytical investigation was performed comparing clean air to a gas vitiated with methane/oxygen combustion products to determine if variations in gas properties contributed to changes in predicted heat flux. This investigation started with simple relationships, evolved into writing an engineering-level code, and ended with running a series of CFD cases. It was noted that it is not possible to simultaneously match all of the gas properties between clean and vitiated test gases. A study was then conducted selecting various combinations of freestream properties for a vitiated test gas that matched clean air values to determine which combination of parameters affected the computed heat transfer the least. The best combination of properties to match was the free-stream total sensible enthalpy, dynamic pressure, and either the velocity or Mach number. This combination yielded only a 2% difference in heating. Other combinations showed departures of up to 10% in the heat flux estimate.

  14. Hyper-X and Pegasus Launch Vehicle: A Three-Foot Model of the Hypersonic Experimental Research Vehic

    NASA Technical Reports Server (NTRS)

    1997-01-01

    The configuration of the X-43A Hypersonic Experimental Research Vehicle, or Hyper-X, attached to a Pegasus launch vehicle is displayed in this side view of a three-foot-long model of the vehicle/booster combination at NASA's Dryden Flight Research Center, Edwards, California. Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry

  15. Performance Validation Approach for the GTX Air-Breathing Launch Vehicle

    NASA Technical Reports Server (NTRS)

    Trefny, Charles J.; Roche, Joseph M.

    2002-01-01

    The primary objective of the GTX effort is to determine whether or not air-breathing propulsion can enable a launch vehicle to achieve orbit in a single stage. Structural weight, vehicle aerodynamics, and propulsion performance must be accurately known over the entire flight trajectory in order to make a credible assessment. Structural, aerodynamic, and propulsion parameters are strongly interdependent, which necessitates a system approach to design, evaluation, and optimization of a single-stage-to-orbit concept. The GTX reference vehicle serves this purpose, by allowing design, development, and validation of components and subsystems in a system context. The reference vehicle configuration (including propulsion) was carefully chosen so as to provide high potential for structural and volumetric efficiency, and to allow the high specific impulse of air-breathing propulsion cycles to be exploited. Minor evolution of the configuration has occurred as analytical and experimental results have become available. With this development process comes increasing validation of the weight and performance levels used in system performance determination. This paper presents an overview of the GTX reference vehicle and the approach to its performance validation. Subscale test rigs and numerical studies used to develop and validate component performance levels and unit structural weights are outlined. The sensitivity of the equivalent, effective specific impulse to key propulsion component efficiencies is presented. The role of flight demonstration in development and validation is discussed.

  16. Hypersonic vehicle control law development using H infinity and mu-synthesis

    NASA Technical Reports Server (NTRS)

    Gregory, Irene M.; Chowdhry, Rajiv S.; Mcminn, John D.; Shaughnessy, John D.

    1992-01-01

    Applicability and effectiveness of robust control techniques to a single-stage-to-orbit (SSTO) airbreathing hypersonic vehicle on an ascent accelerating path and their effectiveness are explored in this paper. An SSTO control system design problem, requiring high accuracy tracking of velocity and altitude commands while limiting angle of attack oscillations, minimizing control power usage and stabilizing the vehicle all in the presence of atmospheric turbulence and uncertainty in the system, was formulated to compare results of the control designs using H infinity and mu-synthesis procedures. The math model, an integrated flight/propulsion dynamic model of a conical accelerator class vehicle, was linearized as the vehicle accelerated through Mach 8. Controller analysis was conducted using the singular value technique and the mu-analysis approach. Analysis results were obtained in both the frequency and the time domains. The results clearly demonstrate the inherent advantages of the structured singular value framework for this class of problems. Since payload performance margins are so critical for the SSTO mission, it is crucial that adequate stability margins be provided without sacrificing any payload mass.

  17. Hypersonic vehicle control law development using H infinity and mu-synthesis

    NASA Technical Reports Server (NTRS)

    Gregory, Irene M.; Mcminn, John D.; Shaughnessy, John D.; Chowdhry, Rajiv S.

    1992-01-01

    Applicability and effectivness of robust control techniques to a single-stage-to-orbit (SSTO) airbreathing hypersonic vehicle on an ascent accelerating path and their effectivness are explored in this paper. An SSTO control system design problem, requiring high accuracy tracking of velocity and altitude commands while limiting an angle of attack oscillations, minimizing control power usage and stabilizing the vehicle all in the presence of atmospheric turbulence and uncertainty in the system, was formulated to compare results of the control designs using H infinity and mu-synthesis procedures. The math model, an integrated flight/propulsion dynamic model of a conical accelerator class vehicle, was linearized as the vehicle accelerated through Mach 8. Controller analysis was conducted using the singular value technique and the mu-analysis approach. Analysis results were obtained in both the frequency and the time domains. The results clearly demonstrate the inherent advantages of the structured singular value framework for this class of problems. Since payload performance margins are so critical for the SSTO mission, it is crucial that adequate stability margins be provided without sacrificing any payload mass.

  18. Hyper-X and Pegasus Launch Vehicle: A Three-Foot Model of the Hypersonic Experimental Research Vehic

    NASA Technical Reports Server (NTRS)

    1997-01-01

    A close-up view of the X-43A Hypersonic Experimental Research Vehicle, or Hyper-X, portion of a three-foot-long model of the vehicle/booster combination at NASA's Dryden Flight Research Center, Edwards, California. Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is

  19. Air-Breathing Rocket Engines

    NASA Technical Reports Server (NTRS)

    1998-01-01

    This photograph depicts an air-breathing rocket engine prototype in the test bay at the General Applied Science Lab facility in Ronkonkoma, New York. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's Advanced Space Transportation Program at Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

  20. Thermal-Mechanical Testing of Hypersonic Vehicle Structures

    NASA Technical Reports Server (NTRS)

    Hudson, Larry; Stephens, Craig

    2007-01-01

    A viewgraph presentation describing thermal-mechanical tests on the structures of hypersonic vehicles is shown. The topics include: 1) U.S. Laboratories for Hot Structures Testing; 2) NASA Dryden Flight Loads Laboratory; 3) Hot Structures Test Programs; 4) Typical Sequence for Hot Structures Testing; 5) Current Hot Structures Testing; and 6) Concluding Remarks.

  1. Structural dynamic and aeroelastic considerations for hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Cazier, F. W., Jr.; Doggett, Robert V., Jr.; Ricketts, Rodney H.

    1991-01-01

    The specific geometrical, structural, and operational environment characteristics of hypersonic vehicles are discussed with particular reference to aerospace plane type configurations. A discussion of the structural dynamic and aeroelastic phenomena that must be addressed for this class of vehicles is presented. These phenomena are in the aeroservothermoelasticity technical area. Some illustrative examples of recent experimental and analytical work are given. Some examples of current research are pointed out.

  2. Vehicle integration effects on hypersonic waveriders. M.S. Thesis - George Washington Univ.

    NASA Technical Reports Server (NTRS)

    Cockrell, Charles Edward, Jr.

    1994-01-01

    The integration of a class of hypersonic high-lift configurations known as waveriders into hypersonic cruise vehicles was evaluated. Waveriders offer advantages in aerodynamic performance and propulsion/airframe integration (PAI) characteristics over conventional hypersonic shapes. A wind-tunnel model was developed which integrates realistic vehicle components with two waverider shapes, referred to as the 'straight-wing' and 'cranked-wing' shapes. Both shapes were conical-flow-derived waveriders at a design Mach number of 4.0. The cranked-wing shape was designed to provide advantages in subsonic performance and directional stability over conventional waveriders. Experimental data and limited computational fluid dynamics (CFD) predictions were obtained over a Mach number range of 2.3 to 4.63 at a Reynolds number of 2.0x10(exp 6) per foot. The CFD predictions and flow visualization data confirmed the shock attachment characteristics of the baseline waverider shapes and illustrated the waverider flow-field properties. Both CFD predictions and experimental data showed that no significant performance degradations occur at off-design Mach numbers for the waverider shapes and the integrated configurations. The experimental data showed that the effects of adding a realistic canopy were minimal. The effects of adding engine components were to increase the drag and thus degrade the aerodynamic performance of the configuration. A significant degradation in aerodynamic performance was observed when 0 degree control surfaces were added to close the blunt base of the waverider to a sharp trailing edge. A comparison of the fully-integrated waverider models to the baseline shapes showed that the performance was significantly degraded when all of the components were added to the waveriders. The fully-integrated configurations studied here do not offer significant performance advantages over conventional hypersonic vehicles, but still offer advantages in air-breathing propulsion

  3. Computation of Hypersonic Flow about Maneuvering Vehicles with Changing Shapes

    SciTech Connect

    Ferencz, R M; Felker, F F; Castillo, V M

    2004-02-23

    Vehicles moving at hypersonic speeds have great importance to the National Security. Ballistic missile re-entry vehicles (RV's) travel at hypersonic speeds, as do missile defense intercept vehicles. Despite the importance of the problem, no computational analysis method is available to predict the aerodynamic environment of maneuvering hypersonic vehicles, and no analysis is available to predict the transient effects of their shape changes. The present state-of-the-art for hypersonic flow calculations typically still considers steady flow about fixed shapes. Additionally, with present computational methods, it is not possible to compute the entire transient structural and thermal loads for a re-entry vehicle. The objective of this research is to provide the required theoretical development and a computational analysis tool for calculating the hypersonic flow about maneuvering, deforming RV's. This key enabling technology will allow the development of a complete multi-mechanics simulation of the entire RV flight sequence, including important transient effects such as complex flight dynamics. This will allow the computation of the as-delivered state of the payload in both normal and unusual operational environments. This new analysis capability could also provide the ability to predict the nonlinear, transient behavior of endo-atmospheric missile interceptor vehicles to the input of advanced control systems. Due to the computational intensity of fluid dynamics for hypersonics, the usual approach for calculating the flow about a vehicle that is changing shape is to complete a series of steady calculations, each with a fixed shape. However, this quasi-steady approach is not adequate to resolve the frequencies characteristic of a vehicle's structural dynamics. Our approach is to include the effects of the unsteady body shape changes in the finite-volume method by allowing for arbitrary translation and deformation of the control volumes. Furthermore, because the Eulerian

  4. CFD Study of Turbo-Ramjet Interactions in Hypersonic Airbreathing Propulsion System

    NASA Technical Reports Server (NTRS)

    Chang, Ing; Hunter, Louis G.

    1996-01-01

    Advanced airbreathing propulsion systems used in Mach 4-6 mission scenarios, usually involve turbo-ramjet configurations. As the engines transition from turbojet to ramjet, there is an operational envelope where both engines operate simultaneously. In the first phase of our study, an over/under nozzle configuration was analyzed. The two plumes from the turbojet and ramjet interact at the end of a common 2-D cowl, where they both reach an approximate Mach 3.0 condition and then jointly expand to Mach 3.6 at the common nozzle exit plane. For the problem analyzed, the turbojet engine operates at a higher nozzle pressure ratio than the ramjet, causes the turbojet plume overpowers the ramjet plume, deflecting it approximately 12 degrees downward and in turn the turbojet plume is deflected 6 degrees upward. In the process, shocks were formed at the deflections and a shear layer formed at the confluence of the two jets. This particular case was experimentally tested and the data were used to compare with a computational fluid dynamics (CFD) study using the PARC2D code. The CFD results were in good agreement with both static pressure distributions on the cowl separator and on nozzle walls. The thrust coefficients were also in reasonable agreement. In addition, inviscid relationships were developed around the confluence point, where the two exhaust jets meet, and these results compared favorably with the CFD results. In the second phase of our study, a 3-D CFD solution was generated to compare with the 2-D solution. The major difference between the 2-D and 3-D solutions was the interaction of the shock waves, generated by the plume interactions, on the sidewall. When a shock wave interacts with a sidewall and sidewall boundary layer, it is called a glancing shock sidewall interaction. These interactions entrain boundary layer flow down the shockline into a vortical flow pattern. The 3-D plots show the streamlines being entrained down the shockline. The pressure of the flow

  5. Ascent performance of an air-breathing horizontal-takeoff launch vehicle

    NASA Technical Reports Server (NTRS)

    Powell, Richard W.; Shaughnessy, John D.; Cruz, Christopher I.; Naftel, J. C.

    1991-01-01

    Simulations are conducted to investigate a proposed NASA launch vehicle that is fully reusable, takes off horizontally, and uses airbreathing propulsion in a single stage. The propulsion model is based on a cycle analysis method, and the vehicle is assumed to be a rigid structure with distributed fuel, operating under a range of atmospheric conditions. The program to optimize simulated trajectories (POST) is modified to include a predictor-corrector guidance capability and then used to generate the trajectories. Significant errors are encountered during the unpowered coast phase due to uncertainty in the atmospheric density profile. The amount of ascent propellant needed is shown to be directly related to the thrust-vector angle and the location of the center of gravity of the vehicle because of the importance of aim-drag losses to total ideal velocity.

  6. Analysis of possible improvement of acceleration of a high-velocity air-breathing flying vehicle

    NASA Astrophysics Data System (ADS)

    Goonko, Yu. P.; Mazhul, I. I.

    2008-09-01

    Results of parametric calculations of the total aeropropulsive characteristics and characteristics of acceleration of a small-scale high-velocity flying vehicle with an air-breathing engine are presented. Integral parameters of acceleration from the flight Mach number M∞ = 4 to M∞ = 7 are determined, namely, the time required fuel stock, and range. A schematic configuration of the vehicle is considered, which allows studying the basic parameters, such as the forebody shape, the angles of surfaces of compression of the stream captured by the inlet, angles of external aerodynamic surfaces of the airframe, relative planform area of the wing panels, and relative area of the nozzle cross section. A comparative estimate of the effect of these parameters shows that it is possible to improve the characteristics of acceleration of vehicles of the type considered.

  7. Numerical methods for aerothermodynamic design of hypersonic space transport vehicles

    NASA Astrophysics Data System (ADS)

    Wanie, K. M.; Brenneis, A.; Eberle, A.; Heiss, S.

    1993-04-01

    The requirement of the design process of hypersonic vehicles to predict flow past entire configurations with wings, fins, flaps, and propulsion system represents one of the major challenges for aerothermodynamics. In this context computational fluid dynamics has come up as a powerful tool to support the experimental work. A couple of numerical methods developed at MBB designed to fulfill the needs of the design process are described. The governing equations and fundamental details of the solution methods are shortly reviewed. Results are given for both geometrically simple test cases and realistic hypersonic configurations. Since there is still a considerable lack of experience for hypersonic flow calculations an extensive testing and verification is essential. This verification is done by comparison of results with experimental data and other numerical methods. The results presented prove that the methods used are robust, flexible, and accurate enough to fulfill the strong needs of the design process.

  8. Nonlinear Adaptive Control and Guidance for Unstart Recovery for a Generic Hypersonic Vehicle

    NASA Astrophysics Data System (ADS)

    Gunbatar, Yakup

    This work presents the development of an integrated flight controller for a generic model of a hypersonic air-breathing vehicle. The flight control architecture comprises a guidance and trajectory planning module and a nonlinear inner-loop adaptive controller. The emphasis of the controller design is on achieving stable tracking of suitable reference trajectories in the presence of a specific engine fault (inlet unstart), in which sudden and drastic changes in the vehicle aerodynamics and engine performance occur. First, the equations of motion of the vehicle for a rigid body model, taking the rotation of the Earth into account, is provided. Aerodynamic forces and moments and engine data are provided in lookup-table format. This comprehensive model is used for simulations and verification of the control strategies. Then, a simplified control-oriented model is developed for the purpose of control design and stability analysis. The design of the guidance and nonlinear adaptive control algorithms is first carried out on a longitudinal version of the vehicle dynamics. The design is verified in a simulation study aiming at testing the robustness of the inner-loop controller under significant model uncertainty and engine failures. At the same time, the guidance system provides reference trajectories to maximize the vehicle's endurance, which is cast as an optimal control problem. The design is then extended to tackle the significantly more challenging case of the 6-degree-of-freedom (6-DOF) vehicle dynamics. For the full 6-DOF case, the adaptive nonlinear flight controller is tested on more challenging maneuvers, where values of the flight path and bank angles exceed the nominal range defined for the vehicle. Simulation studies show stable operation of the closed-loop system in nominal operating conditions, unstart conditions, and during transition from sustained scramjet propulsion to engine failure mode.

  9. Novel inlet-airframe integration methodology for hypersonic waverider vehicles

    NASA Astrophysics Data System (ADS)

    Ding, Feng; Liu, Jun; Shen, Chi-bing; Huang, Wei

    2015-06-01

    With the aim of integrating a ramjet or scramjet with an airframe, a novel inlet-airframe integration methodology for the hypersonic waverider vehicle is proposed. For this newly proposed design concept and for the specified flight conditions, not only the forebody of the vehicle but also its engine cowl and wings can ride on the bow shock wave, causing the bow shock wave to remain attached to the leading edge for the entire length of the vehicle. Thus, this integrated vehicle can take full advantage of the waverider's high lift-to-drag ratio characteristics and the ideal pre-compression surface for the engine. In this work, a novel inlet-airframe integrated axisymmetric basic flow model that accounts for both external and internal flows is first designed using the method of characteristics and the streamline tracing technique. Subsequently, the design of the inlet-airframe integrated waverider vehicle is generated from the integrated axisymmetric basic flow model using the streamline tracing technique. Then, the design methodologies of both the integrated axisymmetric basic flow model and the integrated waverider vehicle are verified by a computational numerical method. Finally, the viscous effects and performance of both the integrated axisymmetric basic flow model and the integrated waverider vehicle are analysed under the design condition using the numerical computation. The obtained results show that the proposed approach is effective in designing the integrated hypersonic waverider vehicle.

  10. Design of a Flush Airdata System (FADS) for the Hypersonic Air Launched Option (HALO) Vehicle

    NASA Technical Reports Server (NTRS)

    Whitmore, Stephen A.; Moes, Timothy R.; Deets, Dwain A. (Technical Monitor)

    1994-01-01

    This paper presents a design study for a pressure based Flush airdata system (FADS) on the Hypersonic Air Launched Option (HALO) Vehicle. The analysis will demonstrate the feasibility of using a pressure based airdata system for the HALO and provide measurement uncertainty estimates along a candidate trajectory. The HALO is a conceived as a man-rated vehicle to be air launched from an SR-71 platform and is proposed as a testbed for an airbreathing hydrogen scramjet. A feasibility study has been performed and indicates that the proposed trajectory is possible with minimal modifications to the existing SR71 vehicle. The mission consists of launching the HALO off the top of an SR-71 at Mach 3 and 80,000 ft. A rocket motor is then used to accelerate the vehicle to the test condition. After the scramjet test is completed the vehicle will glide to a lakebed runway landing. This option provides reusability of the vehicle and scramjet engine. The HALO design will also allow for various scramjet engine and flowpath designs to be flight tested. For the HALO flights, measurements of freestream airdata are considered to be a mission critical to perform gain scheduling and trajectory optimization. One approach taken to obtaining airdata involves measurement of certain parameters such as external atmospheric winds, temperature, etc to estimate the airdata quantities. This study takes an alternate approach. Here the feasibility of obtaining airdata using a pressure-based flush airdata system (FADS) methods is assessed. The analysis, although it is performed using the HALO configuration and trajectory, is generally applicable to other hypersonic vehicles. The method to be presented offers the distinct advantage of inferring total pressure, Mach number, and flow incidence angles, without stagnating the freestream flow. This approach allows for airdata measurements to be made using blunt surfaces and significantly diminishes the heating load at the sensor. In the FADS concept a

  11. Multi-Objective Trajectory Optimization of a Hypersonic Reconnaissance Vehicle with Temperature Constraints

    NASA Astrophysics Data System (ADS)

    Masternak, Tadeusz J.

    This research determines temperature-constrained optimal trajectories for a scramjet-based hypersonic reconnaissance vehicle by developing an optimal control formulation and solving it using a variable order Gauss-Radau quadrature collocation method with a Non-Linear Programming (NLP) solver. The vehicle is assumed to be an air-breathing reconnaissance aircraft that has specified takeoff/landing locations, airborne refueling constraints, specified no-fly zones, and specified targets for sensor data collections. A three degree of freedom scramjet aircraft model is adapted from previous work and includes flight dynamics, aerodynamics, and thermal constraints. Vehicle control is accomplished by controlling angle of attack, roll angle, and propellant mass flow rate. This model is incorporated into an optimal control formulation that includes constraints on both the vehicle and mission parameters, such as avoidance of no-fly zones and coverage of high-value targets. To solve the optimal control formulation, a MATLAB-based package called General Pseudospectral Optimal Control Software (GPOPS-II) is used, which transcribes continuous time optimal control problems into an NLP problem. In addition, since a mission profile can have varying vehicle dynamics and en-route imposed constraints, the optimal control problem formulation can be broken up into several "phases" with differing dynamics and/or varying initial/final constraints. Optimal trajectories are developed using several different performance costs in the optimal control formulation: minimum time, minimum time with control penalties, and maximum range. The resulting analysis demonstrates that optimal trajectories that meet specified mission parameters and constraints can be quickly determined and used for larger-scale operational and campaign planning and execution.

  12. Nonlinear control of linear parameter varying systems with applications to hypersonic vehicles

    NASA Astrophysics Data System (ADS)

    Wilcox, Zachary Donald

    The focus of this dissertation is to design a controller for linear parameter varying (LPV) systems, apply it specifically to air-breathing hypersonic vehicles, and examine the interplay between control performance and the structural dynamics design. Specifically a Lyapunov-based continuous robust controller is developed that yields exponential tracking of a reference model, despite the presence of bounded, nonvanishing disturbances. The hypersonic vehicle has time varying parameters, specifically temperature profiles, and its dynamics can be reduced to an LPV system with additive disturbances. Since the HSV can be modeled as an LPV system the proposed control design is directly applicable. The control performance is directly examined through simulations. A wide variety of applications exist that can be effectively modeled as LPV systems. In particular, flight systems have historically been modeled as LPV systems and associated control tools have been applied such as gain-scheduling, linear matrix inequalities (LMIs), linear fractional transformations (LFT), and mu-types. However, as the type of flight environments and trajectories become more demanding, the traditional LPV controllers may no longer be sufficient. In particular, hypersonic flight vehicles (HSVs) present an inherently difficult problem because of the nonlinear aerothermoelastic coupling effects in the dynamics. HSV flight conditions produce temperature variations that can alter both the structural dynamics and flight dynamics. Starting with the full nonlinear dynamics, the aerothermoelastic effects are modeled by a temperature dependent, parameter varying state-space representation with added disturbances. The model includes an uncertain parameter varying state matrix, an uncertain parameter varying non-square (column deficient) input matrix, and an additive bounded disturbance. In this dissertation, a robust dynamic controller is formulated for a uncertain and disturbed LPV system. The developed

  13. Modification of NASA Langley 8 Foot High Temperature Tunnel to provide a unique national research facility for hypersonic air-breathing propulsion systems

    NASA Technical Reports Server (NTRS)

    Kelly, H. N.; Wieting, A. R.

    1984-01-01

    A planned modification of the NASA Langley 8-Foot High Temperature Tunnel to make it a unique national research facility for hypersonic air-breathing propulsion systems is described, and some of the ongoing supporting research for that modification is discussed. The modification involves: (1) the addition of an oxygen-enrichment system which will allow the methane-air combustion-heated test stream to simulate air for propulsion testing; and (2) supplemental nozzles to expand the test simulation capability from the current nominal Mach number to 7.0 include Mach numbers 3.0, 4.5, and 5.0. Detailed design of the modifications is currently underway and the modified facility is scheduled to be available for tests of large scale propulsion systems by mid 1988.

  14. Modification of NASA Langley 8 foot high temperature tunnel to provide a unique national research facility for hypersonic air-breathing propulsion systems

    NASA Technical Reports Server (NTRS)

    Kelly, H. N.; Wieting, A. R.

    1984-01-01

    A planned modification of the NASA Langley 8-Foot High Temperature Tunnel to make it a unique national research facility for hypersonic air-breathing propulsion systems is described, and some of the ongoing supporting research for that modification is discussed. The modification involves: (1) the addition of an oxygen-enrichment system which will allow the methane-air combustion-heated test stream to simulate air for propulsion testing; and (2) supplemental nozzles to expand the test simulation capability from the current nominal Mach number to 7.0 include Mach numbers 3.0, 4.5, and 5.0. Detailed design of the modifications is currently underway and the modified facility is scheduled to be available for tests of large scale propulsion systems by mid 1988.

  15. Rapid near-optimal trajectory generation and guidance law development for single-stage-to-orbit airbreathing vehicles

    NASA Technical Reports Server (NTRS)

    Calise, A. J.; Flandro, G. A.; Corban, J. E.

    1990-01-01

    General problems associated with on-board trajectory optimization, propulsion system cycle selection, and with the synthesis of guidance laws were addressed for an ascent to low-earth-orbit of an air-breathing single-stage-to-orbit vehicle. The NASA Generic Hypersonic Aerodynamic Model Example and the Langley Accelerator aerodynamic sets were acquired and implemented. Work related to the development of purely analytic aerodynamic models was also performed at a low level. A generic model of a multi-mode propulsion system was developed that includes turbojet, ramjet, scramjet, and rocket engine cycles. Provisions were made in the dynamic model for a component of thrust normal to the flight path. Computational results, which characterize the nonlinear sensitivity of scramjet performance to changes in vehicle angle of attack, were obtained and incorporated into the engine model. Additional trajectory constraints were introduced: maximum dynamic pressure; maximum aerodynamic heating rate per unit area; angle of attack and lift limits; and limits on acceleration both along and normal to the flight path. The remainder of the effort focused on required modifications to a previously derived algorithm when the model complexity cited above was added. In particular, analytic switching conditions were derived which, under appropriate assumptions, govern optimal transition from one propulsion mode to another for two cases: the case in which engine cycle operations can overlap, and the case in which engine cycle operations are mutually exclusive. The resulting guidance algorithm was implemented in software and exercised extensively. It was found that the approximations associated with the assumed time scale separation employed in this work are reasonable except over the Mach range from roughly 5 to 8. This phenomenon is due to the very large thrust capability of scramjets in this Mach regime when sized to meet the requirement for ascent to orbit. By accounting for flight path

  16. A comparison of hypersonic vehicle flight and prediction results

    NASA Technical Reports Server (NTRS)

    Iliff, Kenneth W.; Shafer, Mary F.

    1995-01-01

    Aerodynamic and aerothermodynamic comparisons between flight and ground test for four hypersonic vehicles are discussed. The four vehicles are the X-15, the Reentry F, the Sandia Energetic Reentry Vehicle Experiment (SWERVE), and the Space Shuttle. The comparisons are taken from papers published by researchers active in the various programs. Aerodynamic comparisons include reaction control jet interaction on the Space Shuttle. Various forms of heating including catalytic, boundary layer, shock interaction and interference, and vortex impingement are compared. Predictions were significantly exceeded for the heating caused by vortex impingement (on the Space Shuttle OMS pods) and for heating caused by shock interaction and interference on the X-15 and the Space Shuttle. Predictions of boundary-layer state were in error on the X-15, the SWERVE, and the Space Shuttle vehicles.

  17. Performance Characterization of Polyimide-Carbon Fiber Composites for Future Hypersonic Vehicles

    DTIC Science & Technology

    2010-08-01

    Fred Arnold, Greg Schoeppner Jr., AFRL, Long Beach , CA, May 2006. • “Polyimide Composite Durability for Hypersonic Vehicles”, Mike Wright, U.S. Navy...Characterization of Polyimide-Carbon Fiber Composites for Future Hypersonic Vehicles 5a. CONTRACT NUMBER C07-00449 5b. GRANT NUMBER FA9550-07-1-0285...and Navy have initiated research and development progress to develop hypersonic vehicles in the Mach 5-15 range. Two broad categories are being

  18. Use of Smart Structures for Control and Performance Improvement of Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    August, James A.; Joshi, Shiv

    1996-01-01

    The objective of this presentation was to point out the fact that there are many promising applications for smart structures technology on hypersonic vehicles. This is not inherently obvious due to the real and perceived operating environments of hypersonic vehicles. The idea behind this project was to talk to hypersonic vehicle designers and academics to find out what sort of problems could be solved with smart structures. Two main conclusions can be drawn: One is that the actual environment inside a hypersonic vehicle is not always as severe as it appears. The second is that the hypersonic community needs a different type of research done on a faster timetable in order to use smart structures technology. Vehicle design cycle times are such that a technology must be proven before the vehicle is designed.

  19. Hypersonic Global Strike Feasibility and Options

    DTIC Science & Technology

    2012-02-15

    materials for one-time use vehicles,123 and breakthroughs in in Low Energy Nuclear Reaction ( LENR ) research could revolutionize high-speed materials for air...breathing vehicles in the next 15 to 20 years.124 More research on possible LENR advances will accelerate hypersonic technology. Defense Industry...Revolution is not materials, is Energetics. LENR is being worked, some 1,000 to 1,000,000 times Chemical Energy Density, Fuel fraction/ weight goes away

  20. Trajectory optimization and guidance for a hypersonic vehicle

    NASA Technical Reports Server (NTRS)

    Lu, Ping

    1991-01-01

    The optimal ascent problem for a hypersonic vehicle is formulated as an inverse dynamic problem. This formulation is essential in solving the trajectory optimization problem via the nonlinear programming approach. Both minimum-fuel and minimax type of performance indices are considered. The results reveal important features of the optimal trajectory and controls, and they are subsequently used to construct a nonlinear feedback midcourse control law. This control law not only greatly simplifies the difficult constrained optimization problem and yields improved solutions, but is also suitable for onboard implementation. Finally, off-nominal trajectory guidance is addressed using combination of feedback compensation and onboard generation of control through the inverse dynamics approach.

  1. Air-breathing Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    1999-01-01

    This Quick Time movie depicts the Rocketdyne static test of an air-breathing rocket. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's advanced Transportation Program at the Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

  2. Employment of hypersonic glide vehicles: Proposed criteria for use

    SciTech Connect

    Olguin, Abel

    2014-07-01

    Hypersonic Glide Vehicles (HGVs) are a type of reentry vehicle that couples the high speed of ballistic missiles with the maneuverability of aircraft. The HGV has been in development since the 1970s, and its technology falls under the category of Conventional Prompt Global Strike (CPGS) weapons. As noted by James M. Acton, a senior associate in the Nuclear Policy Program at the Carnegie Endowment, CPGS is a “missile in search of a mission.” With the introduction of any significant new military capability, a doctrine for use—including specifics regarding how, when and where it would be used, as well as tactics, training and procedures—must be clearly defined and understood by policy makers, military commanders, and planners. In this paper, benefits and limitations of the HGV are presented. Proposed criteria and four scenarios illustrate a possible method for assessing when to use an HGV.

  3. Adaptive integral dynamic surface control of a hypersonic flight vehicle

    NASA Astrophysics Data System (ADS)

    Aslam Butt, Waseem; Yan, Lin; Amezquita S., Kendrick

    2015-07-01

    In this article, non-linear adaptive dynamic surface air speed and flight path angle control designs are presented for the longitudinal dynamics of a flexible hypersonic flight vehicle. The tracking performance of the control design is enhanced by introducing a novel integral term that caters to avoiding a large initial control signal. To ensure feasibility, the design scheme incorporates magnitude and rate constraints on the actuator commands. The uncertain non-linear functions are approximated by an efficient use of the neural networks to reduce the computational load. A detailed stability analysis shows that all closed-loop signals are uniformly ultimately bounded and the ? tracking performance is guaranteed. The robustness of the design scheme is verified through numerical simulations of the flexible flight vehicle model.

  4. Ablative thermal management structural material on the hypersonic vehicles

    SciTech Connect

    Shortland, H.; Tsai, C.

    1995-09-01

    A hypersonic vehicle is designed to fly at high Mach number in the earth`s atmosphere that will result in higher aerodynamic heating loads on specific areas of the vehicle. A thermal protection system is required for these areas that may exceed the operating temperature limit of structural materials. This paper delineates the application of ablative material as the passive type of thermal protection system for the nose or wing leading edges. A simplified quasi-steady-state one-dimensional computer model was developed to evaluate the performance and thermal design of a leading edge. The detailed description of the governing mathematical equations and results are presented. This model provides a quantitative information to support the design estimate, performance optimization, and assess preliminary feasibility of using ablation as a design approach.

  5. Design Evolution and Performance Characterization of the GTX Air-Breathing Launch Vehicle Inlet

    NASA Technical Reports Server (NTRS)

    DeBonis, J. R.; Steffen, C. J., Jr.; Rice, T.; Trefny, C. J.

    2002-01-01

    The design and analysis of a second version of the inlet for the GTX rocket-based combine-cycle launch vehicle is discussed. The previous design did not achieve its predicted performance levels due to excessive turning of low-momentum comer flows and local over-contraction due to asymmetric end-walls. This design attempts to remove these problems by reducing the spike half-angle to 10- from 12-degrees and by implementing true plane of symmetry end-walls. Axisymmetric Reynolds-Averaged Navier-Stokes simulations using both perfect gas and real gas, finite rate chemistry, assumptions were performed to aid in the design process and to create a comprehensive database of inlet performance. The inlet design, which operates over the entire air-breathing Mach number range from 0 to 12, and the performance database are presented. The performance database, for use in cycle analysis, includes predictions of mass capture, pressure recovery, throat Mach number, drag force, and heat load, for the entire Mach range. Results of the computations are compared with experimental data to validate the performance database.

  6. Real time guidance and propulsion control for single-stage-to-orbit airbreathing vehicles

    NASA Astrophysics Data System (ADS)

    Corban, John Eric

    1989-03-01

    Problems associated with on-board trajectory optimization and with the synthesis of guidance laws are addressed for ascent to low-Earth-orbit of an air-breathing, single-stage-to-orbit vehicle. A multi-mode propulsion system is assumed which incorporates turbojet, ramjet, SCRAMJET, and rocket engines. An energy state approximation is applied to a four-state dynamic model for flight of a point mass over a spherical non-rotating Earth. An algorithm for generating fuel-optimal climb profiles is derived via singular perturbation theory. This algorithm results from application of the minimum principle to a low order dynamic model that includes general functional dependence on angle of attack and a component of thrust normal to the flight path. Maximum dynamic pressure and maximum aerodynamic heating rate constraints are considered in addition to acceleration limits. Switching conditions are derived which, under appropriate assumptions, govern optimal transition from one propulsion mode to another. The use of bank angle to modulate the magnitude of the vertical component of lift is shown to improve the index of performance slightly. A nonlinear transformation technique is employed to derive a feedback controller for tracking the computed trajectory. Numerical results illustrate the nature of the resulting fuel-optimal climb paths and the performance of the feedback control law.

  7. Two-Dimensional Shape Optimization of Hypersonic Vehicles Considering Transonic Aerodynamic Performance

    NASA Astrophysics Data System (ADS)

    Ueno, Atsushi; Suzuki, Kojiro

    For the success of hypersonic vehicles, their shape must be optimized to achieve a high lift-to-drag ratio as well as a low aerodynamic heating rate in the hypersonic regime. In addition, the transonic lift-to-drag ratio must also be optimized to realize quick acceleration to the hypersonic cruise speed. The three-dimensional lift-to-drag ratio can be improved even by the two-dimensional section shape (i.e., airfoil) optimization in the region where the sweep back angle is small. Here, prior to three-dimensional shape optimization, a study is done to optimize airfoils of hypersonic vehicles based on these three parameters. At optimization, the hypersonic lift-to-drag ratio is maximized while the transonic lift-to-drag ratio and the aerodynamic heating rate are constrained. The optimum lift coefficient for hypersonic cruise at the maximum lift-to-drag ratio is investigated. The relation between the leading edge radius, which determines the aerodynamic heating rate, and the hypersonic lift-to-drag ratio is also investigated. Results show that to improve the hypersonic lift-to-drag ratio, the airfoil thickness around the leading edge should be small as long as an appropriate compromise with the transonic lift-to-drag ratio is achieved. Results also show that the optimum lift coefficient for hypersonic cruise is much lower than that for typical supersonic vehicles. Small cruise lift coefficient suggests that the wing loading of a hypersonic vehicle should be small. The leading edge radius should be determined by a compromise between the hypersonic lift-to-drag ratio and leading edge heating. Airfoil optimization can provide an appropriate initial guess of the three-dimensional optimum shape. By using an appropriate initial guess, the computation time of the three-dimensional shape optimization is expected to be reduced.

  8. PAYCOS: A new multidisciplinary analysis program for hypersonic vehicle design

    NASA Technical Reports Server (NTRS)

    Stubbe, J. R.

    1990-01-01

    The Payload Conceptual Sizing Code (PAYCOS), a new multidisciplinary computer program for use in the conceptual development phase of hypersonic lifting vehicles (HV's), is described. The program allows engineers to rapidly determine the feasibility of an HV concept and then improve upon the concept by means of optimization theory. The code contains analysis modules for aerodynamics, thermodynamics, mass properties, flight stability, controls, loads, structures, and packaging. Motivation for the code lies with the increased complexity of HV's over their body-of-revolution ballistic predecessors. With these new shapes, the need to rapidly screen out poor concepts and actively develop new and better concepts is an even more crucial part of the early design process. Preliminary results are given which demonstrate the optimization capabilities of the code.

  9. A ceramic matrix composite thermal protection system for hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Riccitiello, Salvatore R.; Love, Wendell L.; Pitts, William C.

    1993-01-01

    The next generation of hypersonic vehicles (NASP, SSTO) that require reusable thermal protection systems will experience acreage surface temperatures in excess of 1100 C. More important, they will experience a more severe physical environment than the Space Shuttle due to non-pristine launching and landing conditions. As a result, maintenance, inspection, and replacement factors must be more thoroughly incorporated into the design of the TPS. To meet these requirements, an advanced thermal protection system was conceived, designated 'TOPHAT'. This system consists of a toughened outer ceramic matrix composite (CMC) attached to a rigid reusable surface insulator (RSI) which is directly bonded to the surface. The objective of this effort was to evaluate this concept in an aeroconvective environment, to determine the effect of impacts to the CMC material, and to compare the results with existing thermal protection systems.

  10. Basic materials and structures aspects for hypersonic transport vehicles (HTV)

    NASA Astrophysics Data System (ADS)

    Steinheil, E.; Uhse, W.

    A Mach 5 transport design is used to illustrate structural concepts and criteria for materials selections and also key technologies that must be followed in the areas of computational methods, materials and construction methods. Aside from the primary criteria of low weight, low costs, and conceivable risks, a number of additional requirements must be met, including stiffness and strength, corrosion resistance, durability, and a construction adequate for inspection, maintenance and repair. Current aircraft construction requirements are significantly extended for hypersonic vehicles. Additional consideration is given to long-duration temperature resistance of the airframe structure, the integration of large-volume cryogenic fuel tanks, computational tools, structural design, polymer matrix composites, and advanced manufacturing technologies.

  11. Dynamic interactions between hypersonic vehicle aerodynamics and propulsion system performance

    NASA Technical Reports Server (NTRS)

    Flandro, G. A.; Roach, R. L.; Buschek, H.

    1992-01-01

    Described here is the development of a flexible simulation model for scramjet hypersonic propulsion systems. The primary goal is determination of sensitivity of the thrust vector and other system parameters to angle of attack changes of the vehicle. Such information is crucial in design and analysis of control system performance for hypersonic vehicles. The code is also intended to be a key element in carrying out dynamic interaction studies involving the influence of vehicle vibrations on propulsion system/control system coupling and flight stability. Simple models are employed to represent the various processes comprising the propulsion system. A method of characteristics (MOC) approach is used to solve the forebody and external nozzle flow fields. This results in a very fast computational algorithm capable of carrying out the vast number of simulation computations needed in guidance, stability, and control studies. The three-dimensional fore- and aft body (nozzle) geometry is characterized by the centerline profiles as represented by a series of coordinate points and body cross-section curvature. The engine module geometry is represented by an adjustable vertical grid to accommodate variations of the field parameters throughout the inlet and combustor. The scramjet inlet is modeled as a two-dimensional supersonic flow containing adjustable sidewall wedges and multiple fuel injection struts. The inlet geometry including the sidewall wedge angles, the number of injection struts, their sweepback relative to the vehicle reference line, and strut cross-section are user selectable. Combustion is currently represented by a Rayleigh line calculation including corrections for variable gas properties; improved models are being developed for this important element of the propulsion flow field. The program generates (1) variation of thrust magnitude and direction with angle of attack, (2) pitching moment and line of action of the thrust vector, (3) pressure and temperature

  12. Multidisciplinary aeroelastic analysis of a generic hypersonic vehicle

    NASA Technical Reports Server (NTRS)

    Gupta, K. K.; Petersen, K. L.

    1993-01-01

    This paper presents details of a flutter and stability analysis of aerospace structures such as hypersonic vehicles. Both structural and aerodynamic domains are discretized by the common finite element technique. A vibration analysis is first performed by the STARS code employing a block Lanczos solution scheme. This is followed by the generation of a linear aerodynamic grid for subsequent linear flutter analysis within subsonic and supersonic regimes of the flight envelope; the doublet lattice and constant pressure techniques are employed to generate the unsteady aerodynamic forces. Flutter analysis is then performed for several representative flight points. The nonlinear flutter solution is effected by first implementing a CFD solution of the entire vehicle. Thus, a 3-D unstructured grid for the entire flow domain is generated by a moving front technique. A finite element Euler solution is then implemented employing a quasi-implicit as well as an explicit solution scheme. A novel multidisciplinary analysis is next effected that employs modal and aerodynamic data to yield aerodynamic damping characteristics. Such analyses are performed for a number of flight points to yield a large set of pertinent data that define flight flutter characteristics of the vehicle. This paper outlines the finite-element-based integrated analysis procedures in detail, which is followed by the results of numerical analyses of flight flutter simulation.

  13. Multidisciplinary aeroelastic analysis of a generic hypersonic vehicle

    NASA Astrophysics Data System (ADS)

    Gupta, K. K.; Petersen, K. L.

    1993-10-01

    This paper presents details of a flutter and stability analysis of aerospace structures such as hypersonic vehicles. Both structural and aerodynamic domains are discretized by the common finite element technique. A vibration analysis is first performed by the STARS code employing a block Lanczos solution scheme. This is followed by the generation of a linear aerodynamic grid for subsequent linear flutter analysis within subsonic and supersonic regimes of the flight envelope; the doublet lattice and constant pressure techniques are employed to generate the unsteady aerodynamic forces. Flutter analysis is then performed for several representative flight points. The nonlinear flutter solution is effected by first implementing a CFD solution of the entire vehicle. Thus, a 3-D unstructured grid for the entire flow domain is generated by a moving front technique. A finite element Euler solution is then implemented employing a quasi-implicit as well as an explicit solution scheme. A novel multidisciplinary analysis is next effected that employs modal and aerodynamic data to yield aerodynamic damping characteristics. Such analyses are performed for a number of flight points to yield a large set of pertinent data that define flight flutter characteristics of the vehicle. This paper outlines the finite-element-based integrated analysis procedures in detail, which is followed by the results of numerical analyses of flight flutter simulation.

  14. A Collaborative Analysis Tool for Integrating Hypersonic Aerodynamics, Thermal Protection Systems, and RBCC Engine Performance for Single Stage to Orbit Vehicles

    NASA Technical Reports Server (NTRS)

    Stanley, Thomas Troy; Alexander, Reginald

    1999-01-01

    Presented is a computer-based tool that connects several disciplines that are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system, as is the case of SSTO vehicles with air breathing propulsion, which is currently being studied by NASA. The deficiencies in the scramjet powered concept led to a revival of interest in Rocket-Based Combined-Cycle (RBCC) propulsion systems. An RBCC propulsion system integrates airbreathing and rocket propulsion into a single engine assembly enclosed within a cowl or duct. A typical RBCC propulsion system operates as a ducted rocket up to approximately Mach 3. At this point the transitions to a ramjet mode for supersonic-to-hypersonic acceleration. Around Mach 8 the engine transitions to a scram4jet mode. During the ramjet and scramjet modes, the integral rockets operate as fuel injectors. Around Mach 10-12 (the actual value depends on vehicle and mission requirements), the inlet is physically closed and the engine transitions to an integral rocket mode for orbit insertion. A common feature of RBCC propelled vehicles is the high degree of integration between the propulsion system and airframe. At high speeds the vehicle forebody is fundamentally part of the engine inlet, providing a compression surface for air flowing into the engine. The compressed air is mixed with fuel and burned. The combusted mixture must be expanded to an area larger than the incoming stream to provide thrust. Since a conventional nozzle would be too large, the entire lower after body of the vehicle is used as an expansion surface. Because of the high external temperatures seen during atmospheric flight, the design of an airbreathing SSTO vehicle requires delicate tradeoffs between engine design, vehicle shape, and thermal protection system (TPS) sizing in order to produce an optimum system in terms of weight (and cost) and maximum performance.

  15. Rekindled vision of hypersonic travel

    NASA Technical Reports Server (NTRS)

    Colladay, Raymond S.

    1987-01-01

    NASA has joined with the DOD to conduct the National Aerospace Plane (NASP) program, whose experimental test vehicle will be designated the X-30. NASP will study the X-30's takeoff from a runway under its own power, acceleration to high Mach number on the basis of airbreathing propulsion, emergence into LEO, reentry into the earth atmosphere, and descent to a powered horizontal landing. NASP will thereby generate technology base data for three distinct types of aircraft: upper-atmosphere hypersonic-cruise aircraft, LEO space transports, and military transatmospheric vehicles. The current concept-validation phase of NASP focuses on airbreathing propulsion, lightweight/high-strength heat-resistant materials, and computational fluid dynamics.

  16. HASA: Hypersonic Aerospace Sizing Analysis for the Preliminary Design of Aerospace Vehicles

    NASA Technical Reports Server (NTRS)

    Harloff, Gary J.; Berkowitz, Brian M.

    1988-01-01

    A review of the hypersonic literature indicated that a general weight and sizing analysis was not available for hypersonic orbital, transport, and fighter vehicles. The objective here is to develop such a method for the preliminary design of aerospace vehicles. This report describes the developed methodology and provides examples to illustrate the model, entitled the Hypersonic Aerospace Sizing Analysis (HASA). It can be used to predict the size and weight of hypersonic single-stage and two-stage-to-orbit vehicles and transports, and is also relevant for supersonic transports. HASA is a sizing analysis that determines vehicle length and volume, consistent with body, fuel, structural, and payload weights. The vehicle component weights are obtained from statistical equations for the body, wing, tail, thermal protection system, landing gear, thrust structure, engine, fuel tank, hydraulic system, avionics, electral system, equipment payload, and propellant. Sample size and weight predictions are given for the Space Shuttle orbiter and other proposed vehicles, including four hypersonic transports, a Mach 6 fighter, a supersonic transport (SST), a single-stage-to-orbit (SSTO) vehicle, a two-stage Space Shuttle with a booster and an orbiter, and two methane-fueled vehicles.

  17. Three-dimensional robust diving guidance for hypersonic vehicle

    NASA Astrophysics Data System (ADS)

    Zhu, Jianwen; Liu, Luhua; Tang, Guojian; Bao, Weimin

    2016-01-01

    A novel three-dimensional robust guidance law based on H∞ filter and H∞ control is proposed to meet the constraints of the impact accuracy and the flight direction under process disturbances for the dive phase of hypersonic vehicle. Complete three-dimensional coupling relative motion equations are established and decoupled into linear ones by feedback linearization to simplify the design process of the further guidance law. Based on the linearized equations, H∞ filter is introduced to eliminate the measurement noises of line-of-sight angles and estimate the angular rates. Furthermore, H∞ robust control is well employed to design guidance law, and the filtered information is used to generate guidance commands to meet the guidance goal accurately and robustly. The simulation results of CAV-H indicate that the proposed three-dimensional equations can describe the coupling character more clearly than the traditional decoupling guidance, and the proposed guidance strategy can guide the vehicle to satisfy different multiple constraints with high accuracy and robustness.

  18. Trajectory optimization for a hypersonic vehicle with constraint

    NASA Astrophysics Data System (ADS)

    Morimoto, Hitoshi

    A new approach was developed to solve fuel-optimal problems for a hypersonic vehicle over an entire flight trajectory. A realistic vehicle model was proposed. Although shooting methods are accurate if they converge, initial guesses for costate variables which make the calculation to converge are difficult to obtain. The proposed approach allowed to divide an entire flight trajectory into two segments by selecting a connecting point in the middle of the trajectory. From the connecting point, a maximum-glide problem was solved first. It was proved that when a certain boundary conditions are satisfied, the maximum-glide trajectory constitutes a fuel-optimal trajectory over an entire flight profile. Determination of the maximum-glide trajectory provides with some of the initial values of the costate variables which are needed to start integration backwards from the connecting point. Since shooting methods are very sensitive to the initial guesses of the costates, obtaining accurate initial values of some of the costates is a large advantage. By combining the maximum-glide trajectory and the trajectory obtained by backward integration from the connecting point, an entire flight trajectory was obtained. This trajectory, in general, does not satisfy all the boundary conditions imposed. However, by adjusting the connecting points, the resulting entire flight trajectory approached the final solution. Periodic optimal cruise control problems were also solved with inequality constraints.

  19. Advanced aeroservoelastic stabilization techniques for hypersonic flight vehicles

    NASA Technical Reports Server (NTRS)

    Chan, Samuel Y.; Cheng, Peter Y.; Myers, Thomas T.; Klyde, David H.; Magdaleno, Raymond E.; Mcruer, Duane T.

    1992-01-01

    Advanced high performance vehicles, including Single-Stage-To-Orbit (SSTO) hypersonic flight vehicles, that are statically unstable, require higher bandwidth flight control systems to compensate for the instability resulting in interactions between the flight control system, the engine/propulsion dynamics, and the low frequency structural modes. Military specifications, such as MIL-F-9490D and MIL-F-87242, tend to limit treatment of structural modes to conventional gain stabilization techniques. The conventional gain stabilization techniques, however, introduce low frequency effective time delays which can be troublesome from a flying qualities standpoint. These time delays can be alleviated by appropriate blending of gain and phase stabilization techniques (referred to as Hybrid Phase Stabilization or HPS) for the low frequency structural modes. The potential of using HPS for compensating structural mode interaction was previously explored. It was shown that effective time delay was significantly reduced with the use of HPS; however, the HPS design was seen to have greater residual response than a conventional gain stablized design. Additional work performed to advance and refine the HPS design procedure, to further develop residual response metrics as a basis for alternative structural stability specifications, and to develop strategies for validating HPS design and specification concepts in manned simulation is presented. Stabilization design sensitivity to structural uncertainties and aircraft-centered requirements are also assessed.

  20. Optimal diving maneuver strategy considering guidance accuracy for hypersonic vehicle

    NASA Astrophysics Data System (ADS)

    Zhu, Jianwen; Liu, Luhua; Tang, Guojian; Bao, Weimin

    2014-11-01

    An optimal maneuver strategy considering terminal guidance accuracy for hypersonic vehicle in dive phase is investigated in this paper. First, it derives the complete three-dimensional nonlinear coupled motion equation without any approximations based on diving relative motion relationship directly, and converts it into linear decoupled state space equation with the same relative degree by feedback linearization. Second, the diving guidance law is designed based on the decoupled equation to meet the terminal impact point and falling angle constraints. In order to further improve the interception capability, it constructs maneuver control model through adding maneuver control item to the guidance law. Then, an integrated performance index consisting of maximum line-of-sight angle rate and minimum energy consumption is designed, and optimal control is employed to obtain optimal maneuver strategy when the encounter time is determined and undetermined, respectively. Furthermore, the performance index and suboptimal strategy are reconstructed to deal with the control capability constraint and the serous influence on terminal guidance accuracy caused by maneuvering flight. Finally, the approach is tested using the Common Aero Vehicle-H model. Simulation results demonstrate that the proposed strategy can achieve high precision guidance and effective maneuver at the same time, and the indices are also optimized.

  1. A Collaborative Analysis Tool for Integrated Hypersonic Aerodynamics, Thermal Protection Systems, and RBCC Engine Performance for Single Stage to Orbit Vehicles

    NASA Technical Reports Server (NTRS)

    Stanley, Thomas Troy; Alexander, Reginald; Landrum, Brian

    2000-01-01

    Presented is a computer-based tool that connects several disciplines that are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system, as is the case of SSTO vehicles with air breathing propulsion, which is currently being studied by NASA. An RBCC propulsion system integrates airbreathing and rocket propulsion into a single engine assembly enclosed within a cowl or duct. A typical RBCC propulsion system operates as a ducted rocket up to approximately Mach 3. Then there is a transition to a ramjet mode for supersonic-to-hypersonic acceleration. Around Mach 8 the engine transitions to a scramjet mode. During the ramjet and scramjet modes, the integral rockets operate as fuel injectors. Around Mach 10-12 (the actual value depends on vehicle and mission requirements), the inlet is physically closed and the engine transitions to an integral rocket mode for orbit insertion. A common feature of RBCC propelled vehicles is the high degree of integration between the propulsion system and airframe. At high speeds the vehicle forebody is fundamentally part of the engine inlet, providing a compression surface for air flowing into the engine. The compressed air is mixed with fuel and burned. The combusted mixture must be expanded to an area larger than the incoming stream to provide thrust. Since a conventional nozzle would be too large, the entire lower after body of the vehicle is used as an expansion surface. Because of the high external temperatures seen during atmospheric flight, the design of an airbreathing SSTO vehicle requires delicate tradeoffs between engine design, vehicle shape, and thermal protection system (TPS) sizing in order to produce an optimum system in terms of weight (and cost) and maximum performance. To adequately determine the performance of the engine/vehicle, the Hypersonic Flight Inlet Model (HYFIM) module was designed to interface with the RBCC

  2. Numerical Research of Airframe/Engine Integrative Hypersonic Vehicle

    DTIC Science & Technology

    2007-11-02

    paper, an engineering method and a finite volume method based on the center of grid are developed for preliminary research of interested integrative...development of hypersonic technology, advanced experimental, analytical and computational methods are being exploited in the design of hypersonic...configurations to obtain excellent aerodynamic characteristics[5]. Due to the limitation of test capabilities to model all the impossible flight conditions

  3. Method and system for control of upstream flowfields of vehicle in supersonic or hypersonic atmospheric flight

    NASA Technical Reports Server (NTRS)

    Daso, Endwell O. (Inventor); Pritchett, II, Victor E. (Inventor); Wang, Ten-See (Inventor); Farr, Rebecca Ann (Inventor)

    2012-01-01

    The upstream flowfield of a vehicle traveling in supersonic or hypersonic atmospheric flight is actively controlled using attribute(s) experienced by the vehicle. Sensed attribute(s) include pressure along the vehicle's outer mold line, temperature along the vehicle's outer mold line, heat flux along the vehicle's outer mold line, and/or local acceleration response of the vehicle. A non-heated, non-plasma-producing gas is injected into an upstream flowfield of the vehicle from at least one surface location along the vehicle's outer mold line. The pressure of the gas so-injected is adjusted based on the attribute(s) so-sensed.

  4. Hypersonic vehicle simulation model: Winged-cone configuration

    NASA Technical Reports Server (NTRS)

    Shaughnessy, John D.; Pinckney, S. Zane; Mcminn, John D.; Cruz, Christopher I.; Kelley, Marie-Louise

    1990-01-01

    Aerodynamic, propulsion, and mass models for a generic, horizontal-takeoff, single-stage-to-orbit (SSTO) configuration are presented which are suitable for use in point mass as well as batch and real-time six degree-of-freedom simulations. The simulations can be used to investigate ascent performance issues and to allow research, refinement, and evaluation of integrated guidance/flight/propulsion/thermal control systems, design concepts, and methodologies for SSTO missions. Aerodynamic force and moment coefficients are given as functions of angle of attack, Mach number, and control surface deflections. The model data were estimated by using a subsonic/supersonic panel code and a hypersonic local surface inclination code. Thrust coefficient and engine specific impulse were estimated using a two-dimensional forebody, inlet, nozzle code and a one-dimensional combustor code and are given as functions of Mach number, dynamic pressure, and fuel equivalence ratio. Rigid-body mass moments of inertia and center of gravity location are functions of vehicle weight which is in turn a function of fuel flow.

  5. Aerothermodynamics of expert ballistic vehicle at hypersonic speeds

    NASA Astrophysics Data System (ADS)

    Kharitonov, A. M.; Adamov, N. P.; Chirkashenko, V. F.; Mazhul, I. I.; Shpak, S. I.; Shiplyuk, A. N.; Vasenyov, L. G.; Zvegintsev, V. I.; Muylaert, J. M.

    2012-01-01

    The European EXPErimental Re-entry Test bed (EXPERT) vehicle is intended for studying various basic phenomena, such as the boundary-layer transition on blunted bodies, real gas effects during shock wave/boundary layer interaction, and effect of surface catalycity. Another task is to develop methods for recalculating the results of windtunnel experiments to flight conditions. The EXPERT program implies large-scale preflight research, in particular, various calculations with the use of advanced numerical methods, experimental studies of the models in various wind tunnels, and comparative analysis of data obtained for possible extrapolation of data to in-flight conditions. The experimental studies are performed in various aerodynamic centers of Europe and Russia under contracts with ESA-ESTEC. In particular, extensive experiments are performed at the Von Karman Institute for Fluid Dynamics (VKI, Belgium) and also at the DLR aerospace center in Germany. At ITAM SB RAS, the experimental studies of the EXPERT model characteristic were performed under ISTC Projects 2109, 3151, and 3550, in the T-313 supersonic wind tunnel and AT-303 hypersonic wind tunnel.

  6. Integrated guidance and control with L2 disturbance attenuation for hypersonic vehicles

    NASA Astrophysics Data System (ADS)

    Zhao, Tun; Wang, Peng; Liu, Luhua; Wu, Jie

    2016-06-01

    A robust integrated guidance and control (IGC) approach with L2 gain performance is addressed for a hypersonic vehicle that operates in the dive phase and attacks a fixed target with a terminal angular constraint. A full-state hypersonic vehicle model that adopts the bank-to-turn technique is developed by combining relative motion equations, expressed in the line-of-sight coordinate system, between the vehicle and the target with rotational motion equations. For the proposed model in a strict feedback system, a novel IGC law with L2 gain performance is developed based on the backstepping design procedure by recursively constructing Lyapunov functions of the model subsystems. Numerical simulations conducted for a six degrees of freedom model of the general hypersonic vehicle show that the proposed IGC law is robust against existing uncertainties and satisfies performance requirements.

  7. Tracking Control for an Overactuated Hypersonic Air-Breathing Vehicle with Steady State Constraints (PREPRINT)

    DTIC Science & Technology

    2005-12-01

    choice of a steady state control is completely independent from the choice of a stabilizing control law. This separation is key for the methods we will...develop for steady state optimization in later sections. Combining the steady state with the stabilizing control , we can express the control law as u...for stabilizing control and optimization methods for steady state control, both unconstrained and constrained, we were able to produce promising results

  8. Nonlinear Robust/Adaptive Controller Design for an Air-Breathing Hypersonic Vehicle Model (Preprint)

    DTIC Science & Technology

    2006-12-01

    controller for the altitude dynamics is developed using the flight-path angle as a virtual reference input. Finally, the fast angular dynamics with...states [α,Q] ′ is used as a servo-controller (commanded by the pitch moment) to let γ(t) track the required virtual reference command. Since the COM is

  9. Interactions between Flight Dynamics and Propulsion Systems of Air-Breathing Hypersonic Vehicles

    DTIC Science & Technology

    2013-01-01

    Details of a Mach 14 Waverider Wind Tunnel Test,” 18th AIAA Aerospace Ground Testing Conference, 1994, AIAA Paper 1994-2476. 7 [51] Goldman, R. L...May 1995, Technical Paper 3502. 12, 35 [110] Fernandez, R., Trefny, C. J., Thomas, S. R., and Bulman, M. J., “Parametric Data From a Wind Tunnel Test...Springer, 2009, pp. 69–81. 15 [141] Argodale, J., Bird , K., Breen, T., Christensen, D., and Dellinger, N., “2001 Industry Studies: Space,” Tech. rep

  10. A QMU Approach for Characterizing the Operability Limits of Air-Breathing Hypersonic Vehicles

    DTIC Science & Technology

    2011-04-14

    characterized using numerical simulations and an uncertainty quantification methodology. The time- dependent compressible flow equations with heat release...the flow /thermal interactions that occur when the engine unstarts due to thermal choking. quantification of margins and uncertainty (QMU) isused to...compressible flow equations with heat release are solved in a simplified configuration. Verification, calibration and validation are carried out to

  11. Effects of Cavities and Protuberances on Transition over Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Chang, Chau-Lyan; Choudhari, Meelan M.; Li, Fei; Venkatachari, Balaji

    2011-01-01

    Surface protuberances and cavities on a hypersonic vehicle are known to cause several aerodynamic or aerothermodynamic issues. Most important of all, premature transition due to these surface irregularities can lead to a significant rise in surface heating. To help understand laminar-turbulent transition induced by protuberances or cavities on a Crew Exploration Vehicle (CEV) surface, high-fidelity numerical simulations are carried out for both types of trips on a CEV wind tunnel model. Due to the large bluntness, these surface irregularities reside in an accelerating subsonic boundary layer. For the Mach 6 wind tunnel conditions with a roughness Reynolds number Re(sub kk) of 800, it was found that a protuberance with a height to boundary layer thickness ratio of 0.73 leads to strong wake instability and spontaneous vortex shedding, while a cavity with identical geometry only causes a rather weak flow unsteadiness. The same cavity with a larger Reynolds number also leads to similar spontaneous vortex shedding and wake instability. The wake development and the formation of hairpin vortices for both protuberance and cavity were found to be qualitatively similar to that observed for an isolated hemisphere submerged in a subsonic, low speed flat-plate boundary layer. However, the shed vortices and their accompanying instability waves were found to be slightly stabilized downstream by the accelerating boundary layer along the CEV surface. Despite this stabilizing influence, it was found that the wake instability spreads substantially in both wall-normal and azimuthal directions as the flow is evolving towards a transitional state. Similarities and differences between the wake instability behind a protuberance and a cavity are investigated. Computations for the Mach 6 boundary layer over a slender cylindrical roughness element with a height to the boundary layer thickness of about 1.1 also shows spontaneous vortex shedding and strong wake instability. Comparisons of

  12. FLPP IXV Re-Entry Vehicle, Hypersonic Aerodynamics Characterisation

    NASA Astrophysics Data System (ADS)

    Tran, Ph.; Dormieux, M.; Fontaine, J.; Gülhan, A.; Tribot, J.-P.; Binetti, P.; Walloschek, T.

    2009-01-01

    The general objective of the IXV project (Intermediate eXperimental Vehicle), led by NGL Prime in the framework of the ESA FLPP programme (Future Launchers Preparatory Programme), is to improve European capabilities in the strategic field of atmospheric re-entry for space transportation, exploration, and scientific applications. One of the key objectives and challenges of the IXV project is the vehicle re-entry guidance and control demonstration which requires an accurate determination of the aerodynamic characteristics. This paper deals with all the aerodynamic characterization in the hypersonic flow regime. Wind tunnel tests (WTT) and CFD matrices have been defined in order to provide good coverage of the foreseen flight domain, account for uncertainties, and exploit the synergy between experimental and computational activity. WTT have been performed in DLR-H2K (M=6 and 8.7) and ONERA-S4Ma (M=10) facilities, gathering forces and moment data, as well as pressure in key areas. Consistency of the two campaigns results will be addressed. These results have highlighted some flow peculiarities in the deflected flap region. Comparisons with CFD show good agreement with ground experimental results. For flight conditions, real gas and viscous effects play a significant role in the trim conditions that only CFD can currently address; this identification was supported by different partners involved in the project (CFS engineering, DLR, CIRA, and the University of Rome) providing a valuable description of key flow phenomena affecting aerodynamic characteristics. Moreover, at high altitude, limited DSMC computations have been performed for bridging function correction.

  13. Laser power beaming for rocket propulsion and airbreathing propulsion

    NASA Technical Reports Server (NTRS)

    Hertzberg, A.

    1980-01-01

    The developing technology of laser power beaming is introduced, and two systems are used as examples of the capabilities of the laser for beamed energy. In the first system, the potential of the laser to power flight systems ranging from hypersonic air-breathing launch vehicles to commercial jet transports is examined. Attention is given to the possibility of an air-breathing propulsion which offers the promise of a global air transportation network independent of kerosene and powered by solar energy. In addition, consideration is given to a new type of rocket propulsion based on the laser's ability to concentrate coherent laser energy to high power densities. Focused laser beams would heat the propellants directly to produce specific impulses approaching ion and MHD rocket levels, and would do so without the burden of a heavy electrical power supply.

  14. Air-Breathing Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    2000-01-01

    This photograph depicts an air-breathing rocket engine that completed an hour or 3,600 seconds of testing at the General Applied Sciences Laboratory in Ronkonkoma, New York. Referred to as ARGO by its design team, the engine is named after the mythological Greek ship that bore Jason and the Argonauts on their epic voyage of discovery. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's Advanced SpaceTransportation Program at Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

  15. Robust stabilization control based on guardian maps theory for a longitudinal model of hypersonic vehicle.

    PubMed

    Liu, Yanbin; Liu, Mengying; Sun, Peihua

    2014-01-01

    A typical model of hypersonic vehicle has the complicated dynamics such as the unstable states, the nonminimum phases, and the strong coupling input-output relations. As a result, designing a robust stabilization controller is essential to implement the anticipated tasks. This paper presents a robust stabilization controller based on the guardian maps theory for hypersonic vehicle. First, the guardian maps theories are provided to explain the constraint relations between the open subsets of complex plane and the eigenvalues of the state matrix of closed-loop control system. Then, a general control structure in relation to the guardian maps theories is proposed to achieve the respected design demands. Furthermore, the robust stabilization control law depending on the given general control structure is designed for the longitudinal model of hypersonic vehicle. Finally, a simulation example is provided to verify the effectiveness of the proposed methods.

  16. Robust Stabilization Control Based on Guardian Maps Theory for a Longitudinal Model of Hypersonic Vehicle

    PubMed Central

    Liu, Mengying; Sun, Peihua

    2014-01-01

    A typical model of hypersonic vehicle has the complicated dynamics such as the unstable states, the nonminimum phases, and the strong coupling input-output relations. As a result, designing a robust stabilization controller is essential to implement the anticipated tasks. This paper presents a robust stabilization controller based on the guardian maps theory for hypersonic vehicle. First, the guardian maps theories are provided to explain the constraint relations between the open subsets of complex plane and the eigenvalues of the state matrix of closed-loop control system. Then, a general control structure in relation to the guardian maps theories is proposed to achieve the respected design demands. Furthermore, the robust stabilization control law depending on the given general control structure is designed for the longitudinal model of hypersonic vehicle. Finally, a simulation example is provided to verify the effectiveness of the proposed methods. PMID:24795535

  17. A matching approach to communicate through the plasma sheath surrounding a hypersonic vehicle

    SciTech Connect

    Gao, Xiaotian; Jiang, Binhao

    2015-06-21

    In order to overcome the communication blackout problem suffered by hypersonic vehicles, a matching approach has been proposed for the first time in this paper. It utilizes a double-positive (DPS) material layer surrounding a hypersonic vehicle antenna to match with the plasma sheath enclosing the vehicle. Analytical analysis and numerical results indicate a resonance between the matched layer and the plasma sheath will be formed to mitigate the blackout problem in some conditions. The calculated results present a perfect radiated performance of the antenna, when the match is exactly built between these two layers. The effects of the parameters of the plasma sheath have been researched by numerical methods. Based on these results, the proposed approach is easier to realize and more flexible to the varying radiated conditions in hypersonic flight comparing with other methods.

  18. A matching approach to communicate through the plasma sheath surrounding a hypersonic vehicle

    NASA Astrophysics Data System (ADS)

    Gao, Xiaotian; Jiang, Binhao

    2015-06-01

    In order to overcome the communication blackout problem suffered by hypersonic vehicles, a matching approach has been proposed for the first time in this paper. It utilizes a double-positive (DPS) material layer surrounding a hypersonic vehicle antenna to match with the plasma sheath enclosing the vehicle. Analytical analysis and numerical results indicate a resonance between the matched layer and the plasma sheath will be formed to mitigate the blackout problem in some conditions. The calculated results present a perfect radiated performance of the antenna, when the match is exactly built between these two layers. The effects of the parameters of the plasma sheath have been researched by numerical methods. Based on these results, the proposed approach is easier to realize and more flexible to the varying radiated conditions in hypersonic flight comparing with other methods.

  19. NASA Glenn Research Center's Hypersonic Propulsion Program

    NASA Technical Reports Server (NTRS)

    Palac, Donald T.

    1999-01-01

    NASA Glenn Research Center (GRC), as NASA's lead center for aeropropulsion, is responding to the challenge of reducing the cost of space transportation through the integration of air-breathing propulsion into launch vehicles. Air- breathing launch vehicle (ABLV) propulsion requires a marked departure from traditional propulsion applications. and stretches the technology of both rocket and air-breathing propulsion. In addition, the demands of the space launch mission require an unprecedented level of integration of propulsion and vehicle systems. GRC is responding with a program with rocket-based combined cycle (RBCC) propulsion technology as its main focus. RBCC offers the potential for simplicity, robustness, and performance that may enable low-cost single-stage-to-orbit (SSTO) transportation. Other technologies, notably turbine-based combined cycle (TBCC) propulsion, offer benefits such as increased robustness and greater mission flexibility, and are being advanced, at a slower pace, as part of GRC's program in hypersonics.

  20. Shock-tunnel combustor testing for hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Loomis, Mark P.

    1994-01-01

    Proposed configurations for the next generation of transatmospheric vehicles will rely on air breathing propulsion systems during all or part of their mission. At flight Mach numbers greater than about 7 these engines will operate in the supersonic combustion ramjet mode (scramjet). Ground testing of these engine concepts above Mach 8 requires high pressure, high enthalpy facilities such as shock tunnels and expansion tubes. These impulse, or short duration facilities have test times on the order of a millisecond, requiring high speed instrumentation and data systems. One such facility ideally suited for scramjet testing is the NASA-Ames 16-Inch shock tunnel, which over the last two years has completed a series of tests for the NASP (National Aero-Space Plane) program at simulated flight Mach numbers ranging from 12-16. The focus of the experimental programs consisted of a series of classified tests involving a near-full scale hydrogen fueled scramjet combustor model in the semi-free jet method of engine testing whereby the compressed forebody flow ahead of the cowl inlet is reproduced (see appendix A). The AIMHYE-1 (Ames Integrated Modular Hypersonic Engine) test entry for the NASP program was completed in April 1993, while AIMHYE-2 was completed in May 1994. The test entries were regarded as successful, resulting in some of the first data of its kind on the performance of a near full scale scramjet engine at Mach 12-16. The data was distributed to NASP team members for use in design system verification and development. Due to the classified nature of the hardware and data, the data reports resulting from this work are classified and have been published as part of the NASP literature. However, an unclassified AIAA paper resulted from the work and has been included as appendix A. It contains an overview of the test program and a description of some of the important issues.

  1. Elevator Sizing, Placement, and Control-Relevant Tradeoffs for Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Dickeson, Jeffrey J.; Rodriguez, Armando A.; Sridharan, Srikanth; Korad, Akshay

    2010-01-01

    Within this paper, control-relevant vehicle design concepts are examined using a widely used 3 DOF (plus flexibility) nonlinear model for the longitudinal dynamics of a generic carrot-shaped scramjet powered hypersonic vehicle. The impact of elevator size and placement on control-relevant static properties (e.g. level-flight trimmable region, trim controls, Angle of Attack (AOA), thrust margin) and dynamic properties (e.g. instability and right half plane zero associated with flight path angle) are examined. Elevator usage has been examine for a class of typical hypersonic trajectories.

  2. Flight Test Experiment Design for Characterizing Stability and Control of Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Morelli, Eugene A.

    2008-01-01

    A maneuver design method that is particularly well-suited for determining the stability and control characteristics of hypersonic vehicles is described in detail. Analytical properties of the maneuver design are explained. The importance of these analytical properties for maximizing information content in flight data is discussed, along with practical implementation issues. Results from flight tests of the X-43A hypersonic research vehicle (also called Hyper-X) are used to demonstrate the excellent modeling results obtained using this maneuver design approach. A detailed design procedure for generating the maneuvers is given to allow application to other flight test programs.

  3. Novel approach for designing a hypersonic gliding-cruising dual waverider vehicle

    NASA Astrophysics Data System (ADS)

    Liu, Jun; Ding, Feng; Huang, Wei; Jin, Liang

    2014-09-01

    For a hypersonic gliding-cruising vehicle, the gliding Mach number is larger than the cruising Mach number. It may be useful to design the inlet shroud to act as the compression surface of the waverider, to ensure that the vehicle rides on the shock wave, during both the gliding and cruising phases. A new design concept, namely a gliding-cruising dual waverider, is proposed in the current study. During the gliding phase, the hypersonic vehicle rides on the shock wave at the design gliding Mach number, as the inlet shroud is designed to act as waverider's compression surface. During the cruising phase, when the inlet shroud is cast away or jettisoned, the hypersonic vehicle rides on the shock wave at the design cruising Mach number, as the forebody is designed to act as waverider's compression surface. Thus, the design methodology of the dual-cone-derived waverider is described based on the theory of conical flow. Finally, the numerical methods are utilized to verify the new design method of the aerodynamic configuration. This methodology proposed is useful to design a hypersonic vehicle for two regimes of flight.

  4. Heat pipe and surface mass transfer cooling of hypersonic vehicle structures

    NASA Technical Reports Server (NTRS)

    Colwell, Gene T.; Modlin, James M.

    1992-01-01

    The problem of determining the feasibility of cooling hypersonic vehicle leading-edge structures exposed to severe aerodynamic surface heating using heat pipe and mass transfer cooling techniques is addressed. A description is presented of a numerical finite-difference-based hypersonic leading-edge cooling model incorporating poststartup liquid metal heat pipe cooling with surface transpiration and film cooling to predict the transient structural temperature distributions and maximum surface temperatures of hypersonic vehicle leading edge. An application of this model to the transient cooling of a typical aerospace plane wing leading-edge section. The results of this application indicated that liquid metal heat pipe cooling alone is insufficient to maintain surface temperatures below an assumed maximum level of 1800 K for about one-third of a typical aerospace plane ascent trajectory through the earth's atmosphere.

  5. TBCC Discipline Overview. Hypersonics Project

    NASA Technical Reports Server (NTRS)

    Thomas, Scott R.

    2011-01-01

    The "National Aeronautics Research and Development Policy" document, issued by the National Science and Technology Council in December 2006, stated that one (among several) of the guiding objectives of the federal aeronautics research and development endeavors shall be stable and long-term foundational research efforts. Nearly concurrently, the National Academies issued a more technically focused aeronautics blueprint, entitled: the "Decadal Survey of Civil Aeronautics - Foundations for the Future." Taken together these documents outline the principles of an aeronautics maturation plan. Thus, in response to these overarching inputs (and others), the National Aeronautics and Space Administration (NASA) organized the Fundamental Aeronautics Program (FAP), a program within the NASA Aeronautics Research Mission Directorate (ARMD). The FAP initiated foundational research and technology development tasks to enable the capability of future vehicles that operate across a broad range of Mach numbers, inclusive of the subsonic, supersonic, and hypersonic flight regimes. The FAP Hypersonics Project concentrates on two hypersonic missions: (1) Air-breathing Access to Space (AAS) and (2) the (Planetary Atmospheric) Entry, Decent, and Landing (EDL). The AAS mission focuses on Two-Stage-To-Orbit (TSTO) systems using air-breathing combined-cycle-engine propulsion; whereas, the EDL mission focuses on the challenges associated with delivering large payloads to (and from) Mars. So, the FAP Hypersonic Project investments are aligned to achieve mastery and intellectual stewardship of the core competencies in the hypersonic-flight regime, which ultimately will be required for practical systems with highly integrated aerodynamic/vehicle and propulsion/engine technologies. Within the FAP Hypersonics, the technology management is further divided into disciplines including one targeting Turbine-Based Combine-Cycle (TBCC) propulsion. Additionally, to obtain expertise and support from outside

  6. Application of the multigrid solution technique to hypersonic entry vehicles

    NASA Astrophysics Data System (ADS)

    Greene, Francis A.

    1994-09-01

    Multigrid techniques have been incorporated into an existing hypersonic flow analysis code, the Langley aerothermodynamic upwind relaxation algorithm. The multigrid scheme is based on the full approximation storage approach and uses full multigrid to obtain a well-defined fine-mesh starting solution. Predictions were obtained using standard transfer operators, and a V cycle was used to control grid sequencing. Computed hypersonic flow solutions, compared with experimental data for a 15-deg blunted sphere-cone and a blended-wing body, are presented. It is shown that the algorithm predicts heating rates accurately, and computes solutions in one-third the computational time of the nonmultigrid algorithm.

  7. Detailed modeling of electron emission for transpiration cooling of hypersonic vehicles

    NASA Astrophysics Data System (ADS)

    Hanquist, Kyle M.; Hara, Kentaro; Boyd, Iain D.

    2017-02-01

    Electron transpiration cooling (ETC) is a recently proposed approach to manage the high heating loads experienced at the sharp leading edges of hypersonic vehicles. Computational fluid dynamics (CFD) can be used to investigate the feasibility of ETC in a hypersonic environment. A modeling approach is presented for ETC, which includes developing the boundary conditions for electron emission from the surface, accounting for the space-charge limit effects of the near-wall plasma sheath. The space-charge limit models are assessed using 1D direct-kinetic plasma sheath simulations, taking into account the thermionically emitted electrons from the surface. The simulations agree well with the space-charge limit theory proposed by Takamura et al. for emitted electrons with a finite temperature, especially at low values of wall bias, which validates the use of the theoretical model for the hypersonic CFD code. The CFD code with the analytical sheath models is then used for a test case typical of a leading edge radius in a hypersonic flight environment. The CFD results show that ETC can lower the surface temperature of sharp leading edges of hypersonic vehicles, especially at higher velocities, due to the increase in ionized species enabling higher electron heat extraction from the surface. The CFD results also show that space-charge limit effects can limit the ETC reduction of surface temperatures, in comparison to thermionic emission assuming no effects of the electric field within the sheath.

  8. Advances in hypersonic vehicle synthesis with application to studies of advanced thermal protection system

    NASA Technical Reports Server (NTRS)

    Ardema, Mark D.

    1995-01-01

    This report summarizes the work entitled 'Advances in Hypersonic Vehicle Synthesis with Application to Studies of Advanced Thermal Protection Systems.' The effort was in two areas: (1) development of advanced methods of trajectory and propulsion system optimization; and (2) development of advanced methods of structural weight estimation. The majority of the effort was spent in the trajectory area.

  9. Development and validation of purged thermal protection systems for liquid hydrogen fuel tanks of hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Helenbrook, R. D.; Colt, J. Z.

    1977-01-01

    An economical, lightweight, safe, efficient, reliable, and reusable insulation system was developed for hypersonic cruise vehicle hydrogen fuel tanks. Results indicate that, a nitrogen purged, layered insulation system with nonpermeable closed-cell insulation next to the cryogenic tank and a high service temperature fibrous insulation surrounding it, is potentially an attractive solution to the insulation problem. For the postulated hypersonic flight the average unit weight of the purged insulation system (including insulation, condensate and fuel boil off) is 6.31 kg/sq m (1.29 psf). Limited cyclic tests of large specimens of closed cell polymethacrylimide foam indicate it will withstand the expected thermal cycle.

  10. Effects of thermochemistry, nonequilibrium, and surface catalysis on the design of hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Scott, Carl D.

    1989-01-01

    An account is given of the function of physical aspects of a gas on the characteristics of the flow and of the heating associated with hypersonic flight. At the high temperatures encountered, the thermal and chemical characteristics of the air in a hypersonic vehicle's shock layer are altered in ways which depend on the atomic and molecular structure of N and O and their ions; similar effects exist in scramjet propulsion systems. These properties in turn influence the character of shock waves and expansions, and hence the pressure, temperature, and velocity distributions. Transport properties affecting the boundary-layer structure will also affect heat flux and shear stress.

  11. A parametric sensitivity study for single-stage-to-orbit hypersonic vehicles using trajectory optimization

    NASA Technical Reports Server (NTRS)

    Lovell, T. Alan; Schmidt, D. K.

    1994-01-01

    The class of hypersonic vehicle configurations with single stage-to-orbit (SSTO) capability reflect highly integrated airframe and propulsion systems. These designs are also known to exhibit a large degree of interaction between the airframe and engine dynamics. Consequently, even simplified hypersonic models are characterized by tightly coupled nonlinear equations of motion. In addition, hypersonic SSTO vehicles present a major system design challenge; the vehicle's overall mission performance is a function of its subsystem efficiencies including structural, aerodynamic, propulsive, and operational. Further, all subsystem efficiencies are interrelated, hence, independent optimization of the subsystems is not likely to lead to an optimum design. Thus, it is desired to know the effect of various subsystem efficiencies on overall mission performance. For the purposes of this analysis, mission performance will be measured in terms of the payload weight inserted into orbit. In this report, a trajectory optimization problem is formulated for a generic hypersonic lifting body for a specified orbit-injection mission. A solution method is outlined, and results are detailed for the generic vehicle, referred to as the baseline model. After evaluating the performance of the baseline model, a sensitivity study is presented to determine the effect of various subsystem efficiencies on mission performance. This consists of performing a parametric analysis of the basic design parameters, generating a matrix of configurations, and determining the mission performance of each configuration. Also, the performance loss due to constraining the total head load experienced by the vehicle is evaluated. The key results from this analysis include the formulation of the sizing problem for this vehicle class using trajectory optimization, characteristics of the optimal trajectories, and the subsystem design sensitivities.

  12. A Comparative Analysis of Single-Stage-To-Orbit Rocket and Air-Breathing Vehicles

    DTIC Science & Technology

    2006-06-01

    Weight Ratio TBCC ............................................ Turbine-Based Combined-Cycle TPS...spacecraft and possibly allow for SSTO vehicles with sizable payloads. After substantial design and wind tunnel testing , the Hyper-X program peaked with...the successful testing of two unpiloted vehicles. Powered by NASA-developed hydrogen scramjets, the X-43A craft set the world speed record for

  13. An investigation of the fuel-optimal periodic trajectories of a hypersonic vehicle

    NASA Technical Reports Server (NTRS)

    Dewell, Larry D.; Speyer, Jason L.

    1993-01-01

    Periodic trajectories were found to minimize the range-averaged fuel consumption. For a realistic hypersonic aircraft modeled as a point mass over a nonrotating, spherical Earth, the periodic orbit yielded a 15 percent improvement in fuel consumption over static cruise. Moreover, vehicle dynamic loading was contained within a realistic survivability envelope of 8 g's. The resulting periodic orbit is composed of very distinct flight regimes (Keplerian arc, atmospheric glide and powered climb), which may offer mission advantages over the static path.

  14. A Collaborative Analysis Tool for Integrated Hypersonic Aerodynamics, Thermal Protection Systems, and RBCC Engine Performance for Single Stage to Orbit Vehicles

    NASA Technical Reports Server (NTRS)

    Stanley, Thomas Troy; Alexander, Reginald; Landrum, Brian

    2000-01-01

    Presented is a computer-based tool that connects several disciplines that are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system, as is the case of SSTO vehicles with air breathing propulsion, which is currently being studied by NASA. An RBCC propulsion system integrates airbreathing and rocket propulsion into a single engine assembly enclosed within a cowl or duct. A typical RBCC propulsion system operates as a ducted rocket up to approximately Mach 3. Then there is a transition to a ramjet mode for supersonic-to-hypersonic acceleration. Around Mach 8 the engine transitions to a scramjet mode. During the ramjet and scramjet modes, the integral rockets operate as fuel injectors. Around Mach 10-12 (the actual value depends on vehicle and mission requirements), the inlet is physically closed and the engine transitions to an integral rocket mode for orbit insertion. A common feature of RBCC propelled vehicles is the high degree of integration between the propulsion system and airframe. At high speeds the vehicle forebody is fundamentally part of the engine inlet, providing a compression surface for air flowing into the engine. The compressed air is mixed with fuel and burned. The combusted mixture must be expanded to an area larger than the incoming stream to provide thrust. Since a conventional nozzle would be too large, the entire lower after body of the vehicle is used as an expansion surface. Because of the high external temperatures seen during atmospheric flight, the design of an airbreathing SSTO vehicle requires delicate tradeoffs between engine design, vehicle shape, and thermal protection system (TPS) sizing in order to produce an optimum system in terms of weight (and cost) and maximum performance. To adequately determine the performance of the engine/vehicle, the Hypersonic Flight Inlet Model (HYFIM) module was designed to interface with the RBCC

  15. A concept of a hypersonic flight experiment of a winged vehicle

    NASA Astrophysics Data System (ADS)

    Shirouzu, Masao; Watanabe, Shigeya

    A concept of a flight experiment using a winged hypersonic research vehicle is proposed by the National Aerospace Laboratory (NAL) as one of the flight experiment series preceding to the development of HOPE (H-II Orbiting Plane). The present paper describes the purpose of the experiment, the outline of the flight, the configuration and aerodynamic characteristics of the vehicle, and items of experiment and measurement. The present experiment is to acquire experience on the development and the flight of a hypersonic winged vehicle, in contrast to the ballistic flight of the OREX (Orbital Reentry Experiment) and to collect flight data for validation of tests and simulations on the ground. The vehicle of about 1.5 tons will be launched by a two-stage version of the J-I. The vehicle will be separated at an altitude of 70-80 km at a velocity of Mach 18-20, and inserted to the reentry trajectory of HOPE. The vehicle will be decelerated by parachutes and splash into the ocean south of Japan, where it will be recovered.

  16. JANNAF 24th Airbreathing Propulsion Subcommittee and 36th Combustion Subcommittee Joint Meeting. Volume 1

    NASA Technical Reports Server (NTRS)

    Fry, Ronald S. (Editor); Gannaway, Mary T. (Editor)

    1999-01-01

    Volume 1, the first of three volumes is a compilation of 16 unclassified/unlimited-technical papers presented at the Joint Army-Navy-NASA-Air Force (JANNAF) 24th Airbreathing Propulsion Subcommittee and 36th Combustion Subcommittee held jointly with the 181 Propulsion Systems Hazards Subcommittee. The meeting was held on 18-21 October 1999 at NASA Kennedy Space Center and The DoubleTree Oceanfront Hotel, Cocoa Beach, Florida. Topics covered include overviews of RBCC and PDE hypersonic technology, Hyper-X propulsion ground testing, development of JP-8 for hypersonic vehicle applications, numerical simulation of dual-mode SJ combustion, V&V of M&S computer codes, MHD SJ and Rocket Based Combined Cycle (RBCC) launch vehicle concepts, and Pulse Detonation Engine (PDE) propulsion technology development including fundamental investigations, modeling, aerodynamics, operation and performance.

  17. Skip trajectory flight of a ramjet-powered hypersonic vehicle

    NASA Astrophysics Data System (ADS)

    Fomin, V. M.; Aulchenko, S. M.; Zvegintsev, V. I.

    2010-07-01

    Possible skip trajectories of a flying vehicle with a periodically actuated ramjet are numerically simulated. An optimal choice of ramjet actuation areas and duration is demonstrated to ensure the maximum flight range with a given amount of the fuel. The main advantage of skip trajectories is found to be a significant (by an order of magnitude) decrease in thermal loads on the flying vehicle.

  18. Design of Modular, Shape-transitioning Inlets for a Conical Hypersonic Vehicle

    NASA Technical Reports Server (NTRS)

    Gollan, Rowan J.; Smart, Michael K.

    2010-01-01

    For a hypersonic vehicle, propelled by scramjet engines, integration of the engines and airframe is highly desirable. Thus, the forward capture shape of the engine inlet should conform to the vehicle body shape. Furthermore, the use of modular engines places a constraint on the shape of the inlet sidewalls. Finally, one may desire a combustor cross- section shape that is different from that of the inlet. These shape constraints for the inlet can be accommodated by employing a streamline-tracing and lofting technique. This design technique was developed by Smart for inlets with a rectangular-to-elliptical shape transition. In this paper, we generalise that technique to produce inlets that conform to arbitrary shape requirements. As an example, we show the design of a body-integrated hypersonic inlet on a winged-cone vehicle, typical of what might be used in a three-stage orbital launch system. The special challenge of inlet design for this conical vehicle at an angle-of-attack is also discussed. That challenge is that the bow shock sits relatively close to the vehicle body.

  19. Evaluation of thermo-radiation characteristics of IR windows in hypersonic vehicles

    NASA Astrophysics Data System (ADS)

    Wang, Yahui; Wang, Qiang; Zhang, Bochuan; Liang, Bo; Xu, Li; Yun, Hongquan

    2015-04-01

    When hypersonic vehicles have high-hypersonic flights in the atmosphere, high-temperature IR windows become the main factor of complicated aero-thermo-radiation effects, which would reduce the performance of IR detection systems, or even make these systems fail. By analyzing thermal radiation transfer in IR windows, an experimental platform is established to measure thermo-radiation characteristics of IR window materials. And a method is proposed to evaluate thermo-radiation characteristics of IR windows with uneven temperature distribution. Take a MWIR detection system of a hypersonic vehicle as an example, thermo-radiation characteristics of a sapphire IR window is evaluated. The results indicate that, thermo-radiation characteristics of the sapphire IR window material in 3.7μm-4.8μm have an approximate cubic relationship with temperature at 100°C~350°C. With the rise of temperature, the transmittance of the sapphire material decrease, while the window self-radiation increase. As the sapphire IR window is exposed in high-temperature and high-speed airflow, the transmittance drops 4%, still bigger than 95%, self-radiation enhance about 9 times, while temperature of the window rises rapidly. Self-radiation can drive detector into saturation easily, of which the influence on the MWIR detection system is bigger than that of transmittance.

  20. Validation of engineering methods for predicting hypersonic vehicle controls forces and moments

    NASA Technical Reports Server (NTRS)

    Maughmer, M.; Straussfogel, D.; Long, L.; Ozoroski, L.

    1991-01-01

    This work examines the ability of the aerodynamic analysis methods contained in an industry standard conceptual design code, the Aerodynamic Preliminary Analysis System (APAS II), to estimate the forces and moments generated through control surface deflections from low subsonic to high hypersonic speeds. Predicted control forces and moments generated by various control effectors are compared with previously published wind-tunnel and flight-test data for three vehicles: the North American X-15, a hypersonic research airplane concept, and the Space Shuttle Orbiter. Qualitative summaries of the results are given for each force and moment coefficient and each control derivative in the various speed ranges. Results show that all predictions of longitudinal stability and control derivatives are acceptable for use at the conceptual design stage.

  1. JANNAF 25th Airbreathing Propulsion Subcommittee, 37th Combustion Subcommittee and 1st Modeling and Simulation Subcommittee Joint Meeting. Volume 1

    NASA Technical Reports Server (NTRS)

    Fry, Ronald S.; Becker, Dorothy L.

    2000-01-01

    Volume I, the first of three volumes, is a compilation of 24 unclassified/unlimited-distribution technical papers presented at the Joint Army-Navy-NASA-Air Force (JANNAF) 25th Airbreathing Propulsion Subcommittee, 37th Combustion Subcommittee and 1st Modeling and Simulation Subcommittee (MSS) meeting held jointly with the 19th Propulsion Systems Hazards Subcommittee. The meeting was held 13-17 November 2000 at the Naval Postgraduate School and Hyatt Regency Hotel, Monterey, California. Topics covered include: a Keynote Address on Future Combat Systems, a review of the new JANNAF Modeling and Simulation Subcommittee, and technical papers on Hyper-X propulsion development and verification; GTX airbreathing launch vehicles; Hypersonic technology development, including program overviews, fuels for advanced propulsion, ramjet and scramjet research, hypersonic test medium effects; and RBCC engine design and performance, and PDE and UCAV advanced and combined cycle engine technologies.

  2. NASA's Advanced Space Transportation Hypersonic Program

    NASA Technical Reports Server (NTRS)

    Hueter, Uwe; McClinton, Charles; Cook, Stephen (Technical Monitor)

    2002-01-01

    NASA's has established long term goals for access-to-space. NASA's third generation launch systems are to be fully reusable and operational in approximately 25 years. The goals for third generation launch systems are to reduce cost by a factor of 100 and improve safety by a factor of 10,000 over current conditions. The Advanced Space Transportation Program Office (ASTP) at NASA's Marshall Space Flight Center in Huntsville, AL has the agency lead to develop third generation space transportation technologies. The Hypersonics Investment Area, part of ASTP, is developing the third generation launch vehicle technologies in two main areas, propulsion and airframes. The program's major investment is in hypersonic airbreathing propulsion since it offers the greatest potential for meeting the third generation launch vehicles. The program will mature the technologies in three key propulsion areas, scramjets, rocket-based combined cycle and turbine-based combination cycle. Ground and flight propulsion tests are being planned for the propulsion technologies. Airframe technologies will be matured primarily through ground testing. This paper describes NASA's activities in hypersonics. Current programs, accomplishments, future plans and technologies that are being pursued by the Hypersonics Investment Area under the Advanced Space Transportation Program Office will be discussed.

  3. Prediction of forces and moments for hypersonic flight vehicle control effectors

    NASA Technical Reports Server (NTRS)

    Maughmer, Mark D.; Long, Lyle N.; Guilmette, Neal; Pagano, Peter

    1993-01-01

    This research project includes three distinct phases. For completeness, all three phases of the work are briefly described in this report. The goal was to develop methods of predicting flight control forces and moments for hypersonic vehicles which could be used in a preliminary design environment. The first phase included a preliminary assessment of subsonic/supersonic panel methods and hypersonic local flow inclination methods for such predictions. While these findings clearly indicated the usefulness of such methods for conceptual design activities, deficiencies exist in some areas. Thus, a second phase of research was conducted in which a better understanding was sought for the reasons behind the successes and failures of the methods considered, particularly for the cases at hypersonic Mach numbers. This second phase involved using computational fluid dynamics methods to examine the flow fields in detail. Through these detailed predictions, the deficiencies in the simple surface inclination methods were determined. In the third phase of this work, an improvement to the surface inclination methods was developed. This used a novel method for including viscous effects by modifying the geometry to include the viscous/shock layer.

  4. Prediction of forces and moments for hypersonic flight vehicle control effectors

    NASA Technical Reports Server (NTRS)

    Maughmer, Mark D.; Long, Lyle N.; Pagano, Peter J.

    1991-01-01

    Developing methods of predicting flight control forces and moments for hypersonic vehicles, included a preliminary assessment of subsonic/supersonic panel methods and hypersonic local flow inclination methods for such predictions. While these findings clearly indicate the usefulness of such methods for conceptual design activities, deficiencies exist in some areas. Thus, a second phase of research was proposed in which a better understanding is sought for the reasons of the successes and failures of the methods considered, particularly for the cases at hypersonic Mach numbers. To obtain this additional understanding, a more careful study of the results obtained relative to the methods used was undertaken. In addition, where appropriate and necessary, a more complete modeling of the flow was performed using well proven methods of computational fluid dynamics. As a result, assessments will be made which are more quantitative than those of phase 1 regarding the uncertainty involved in the prediction of the aerodynamic derivatives. In addition, with improved understanding, it is anticipated that improvements resulting in better accuracy will be made to the simple force and moment prediction.

  5. Heat pipes for wing leading edges of hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Boman, B. L.; Citrin, K. M.; Garner, E. C.; Stone, J. E.

    1990-01-01

    Wing leading edge heat pipes were conceptually designed for three types of vehicle: an entry research vehicle, aero-space plane, and advanced shuttle. A full scale, internally instrumented sodium/Hastelloy X heat pipe was successfully designed and fabricated for the advanced shuttle application. The 69.4 inch long heat pipe reduces peak leading edge temperatures from 3500 F to 1800 F. It is internally instrumented with thermocouples and pressure transducers to measure sodium vapor qualities. Large thermal gradients and consequently large thermal stresses, which have the potential of limiting heat pipe life, were predicted to occur during startup. A test stand and test plan were developed for subsequent testing of this heat pipe. Heat pipe manufacturing technology was advanced during this program, including the development of an innovative technique for wick installation.

  6. DSMC Grid Methodologies for Computing Low-Density, Hypersonic Flows About Reusable Launch Vehicles

    NASA Technical Reports Server (NTRS)

    Wilmoth, Richard G.; LeBeau, Gerald J.; Carlson, Ann B.

    1996-01-01

    Two different grid methodologies are studied for application to DSMC simulations about reusable launch vehicles. One method uses an unstructured, tetrahedral grid while the other uses a structured, variable-resolution Cartesian grid. The relative merits of each method are discussed in terms of accuracy, computational efficiency, and overall ease of use. Both methods are applied to the computation of a low-density, hypersonic flow about a winged single-stage-to-orbit reusable launch vehicle concept at conditions corresponding to an altitude of 120 km. Both methods are shown to give comparable results for both surface and flowfield quantities as well as for the overall aerodynamic behavior. For the conditions simulated, the flowfield about the vehicle is very rarefied but the DSMC simulations show significant departure from free-molecular predictions for the surface friction and heat transfer as well as certain aerodynamic quantities.

  7. Optimal heading change with minimum energy loss for a hypersonic gliding vehicle

    NASA Astrophysics Data System (ADS)

    Calise, Anthony J.; Bae, Gyoung H.

    A three state model is presented for analyzing the problem of optimal changes in heading with minimum energy loss for a hypersonic gliding vehicle. A further model order reduction to a single state model is examined using singular perturbation theory. The optimal solution for the reduced problem defines an optimal altitude profile dependent on the current energy of the vehicle, and the corresponding optimal lift and bank angle. A separate boundary layer analysis, based on an expansion of the necessary conditions about the reduced solution, is used to account for altitude and flight path angle dynamics and to derive a guidance law in feedback form. The guidance law is evaluated for a hypothetical vehicle.

  8. Optimal heading change with minimum energy loss for a hypersonic gliding vehicle

    NASA Technical Reports Server (NTRS)

    Calise, Anthony J.; Bae, Gyoung H.

    1987-01-01

    A three state model is presented for analyzing the problem of optimal changes in heading with minimum energy loss for a hypersonic gliding vehicle. A further model order reduction to a single state model is examined using singular perturbation theory. The optimal solution for the reduced problem defines an optimal altitude profile dependent on the current energy of the vehicle, and the corresponding optimal lift and bank angle. A separate boundary layer analysis, based on an expansion of the necessary conditions about the reduced solution, is used to account for altitude and flight path angle dynamics and to derive a guidance law in feedback form. The guidance law is evaluated for a hypothetical vehicle.

  9. Impact of Martian atmosphere parameter uncertainties on entry vehicles aerodynamic for hypersonic rarefied conditions

    NASA Astrophysics Data System (ADS)

    Fei, Huang; Xu-hong, Jin; Jun-ming, Lv; Xiao-li, Cheng

    2016-11-01

    An attempt has been made to analyze impact of Martian atmosphere parameter uncertainties on entry vehicle aerodynamics for hypersonic rarefied conditions with a DSMC code. The code has been validated by comparing Viking vehicle flight data with present computational results. Then, by simulating flows around the Mars Science Laboratory, the impact of errors of free stream parameter uncertainties on aerodynamics is investigated. The validation results show that the present numerical approach can show good agreement with the Viking flight data. The physical and chemical properties of CO2 has strong impact on aerodynamics of Mars entry vehicles, so it is necessary to make proper corrections to the data obtained with air model in hypersonic rarefied conditions, which is consistent with the conclusions drawn in continuum regime. Uncertainties of free stream density and velocity weakly influence aerodynamics and pitching moment. However, aerodynamics appears to be little influenced by free stream temperature, the maximum error of what is below 0.5%. Center of pressure position is not sensitive to free stream parameters.

  10. A fuselage/tank structure study for actively cooled hypersonic cruise vehicles: Active cooling system analysis

    NASA Technical Reports Server (NTRS)

    Stone, J. E.

    1975-01-01

    The effects of fuselage cross section and structural arrangement on the performance of actively cooled hypersonic cruise vehicles are investigated. An active cooling system which maintains the aircraft's entire surface area at temperatures below 394 K at Mach 6 is developed along with a hydrogen fuel tankage thermal protection system. Thermodynamic characteristics of the actively cooled thermal protection systems established are summarized. Design heat loads and coolant flowrate requirements are defined for each major structural section and for the total system. Cooling system weights are summarized at the major component level. Conclusions and recommendations are included.

  11. A numerical model for the platelet heat-pipe-cooled leading edge of hypersonic vehicle

    NASA Astrophysics Data System (ADS)

    Liu, Hongpeng; Liu, Weiqiang

    2016-01-01

    A new design, the platelet heat-pipe-cooled leading edge, is discussed for the thermal management to prevent damage to hypersonic vehicle leading edge component. For calculating the steady state behavior of platelet heat-pipe-cooled leading edge, a numerical model based on the principles of evaporation, convection, and condensation of a working fluid is presented. And then its effectiveness is validated by comparing the wall and vapor temperature against experimental data for a conventional heat pipe. Further investigations indicate that alloy IN718, with sodium as the working fluid is a feasible combination for Mach 8 flight with a 15 mm leading edge radius.

  12. A fuselage/tank structure study for actively cooled hypersonic cruise vehicles: Structural analysis

    NASA Technical Reports Server (NTRS)

    Baker, A. H.

    1975-01-01

    The effects of fuselage cross-section (circular and elliptical) and structural arrangement (integral and nonintegral tanks) on the performance of actively cooled hypersonic cruise vehicles was evaluated. It was found that integrally machined stiffening of the tank walls, while providing the most weight-efficient use of materials, results in higher production costs. Fatigue and fracture mechanics appeared to have little effect on the weight of the three study aircraft. The need for thermal strain relief through insulation is discussed. Aircraft size and magnitude of the internal pressure are seen to be significant factors in tank design.

  13. Control-Relevant Modeling, Analysis, and Design for Scramjet-Powered Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Rodriguez, Armando A.; Dickeson, Jeffrey J.; Sridharan, Srikanth; Benavides, Jose; Soloway, Don; Kelkar, Atul; Vogel, Jerald M.

    2009-01-01

    Within this paper, control-relevant vehicle design concepts are examined using a widely used 3 DOF (plus flexibility) nonlinear model for the longitudinal dynamics of a generic carrot-shaped scramjet powered hypersonic vehicle. Trade studies associated with vehicle/engine parameters are examined. The impact of parameters on control-relevant static properties (e.g. level-flight trimmable region, trim controls, AOA, thrust margin) and dynamic properties (e.g. instability and right half plane zero associated with flight path angle) are examined. Specific parameters considered include: inlet height, diffuser area ratio, lower forebody compression ramp inclination angle, engine location, center of gravity, and mass. Vehicle optimizations is also examined. Both static and dynamic considerations are addressed. The gap-metric optimized vehicle is obtained to illustrate how this control-centric concept can be used to "reduce" scheduling requirements for the final control system. A classic inner-outer loop control architecture and methodology is used to shed light on how specific vehicle/engine design parameter selections impact control system design. In short, the work represents an important first step toward revealing fundamental tradeoffs and systematically treating control-relevant vehicle design.

  14. The stability analysis of rolling motion of hypersonic vehicles and its validations

    NASA Astrophysics Data System (ADS)

    Ye, YouDa; Zhao, ZhongLiang; Tian, Hao; Zhang, XianFeng

    2014-12-01

    The stability of the rolling motion of near space hypersonic vehicles with rudder control is studied using method of qualitative analysis of nonlinear differential equations, and the stability criteria of the deflected rolling motions are improved. The outcomes can serve as the basis for further study regarding the influence of pitching and lateral motion on the stability of rolling motion. To validate the theoretical results, numerical simulations were done for the rolling motion of two hypersonic vehicles with typical configurations. Also, wind tunnel experiments for four aircraft models with typical configurations have been done. The results show that: 1) there exist two dynamic patterns of the rolling motion under statically stable condition. The first one is point attractor, for which the motion of aircraft returns to the original state. The second is periodic attractor, for which the aircraft rolls periodically. 2) Under statically unstable condition, there exist three dynamic patterns of rolling motion, namely, the point attractor, periodic attractor around deflected state of rolling motion, and double periodic attractors or chaotic attractors.

  15. Numerical comparison of convective heat transfer augmentation devices used in cooling channels of hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Maldonado, Jaime J.

    1994-01-01

    Hypersonic vehicles are exposed to extreme thermal conditions compared to subsonic aircraft; therefore, some level of thermal management is required to protect the materials used. Normally, hypersonic vehicles experience the highest temperatures in the nozzle throat, and aircraft and propulsion system leading edges. Convective heat transfer augmentation techniques can be used in the thermal management system to increase heat transfer of the cooling channels in those areas. The techniques studied in this report are pin-fin, offset-fin, ribbed and straight roughened channel. A smooth straight channel is used as the baseline for comparing the techniques. SINDA '85, a lumped parameter finite difference thermal analyzer, is used to model the channels. Subroutines are added to model the fluid flow assuming steady one dimensional compressible flow with heat addition and friction. Correlations for convective heat transfer and friction are used in conjunction with the fluid flow analysis mentioned. As expected, the pin-fin arrangement has the highest heat transfer coefficient and the largest pressure drop. All the other devices fall in between the pin-fin and smooth straight channel. The selection of the best heat augmentation method depends on the design requirements. A good approach may be a channel using a combination of the techniques. For instance, several rows of pin-fins may be located at the region of highest heat flux, surrounded by some of the other techniques. Thus, the heat transfer coefficient is maximized at the region of highest heat flux while the pressure drop is not excessive.

  16. Airbreathing Acceleration Toward Earth Orbit

    SciTech Connect

    Whitehead, J C

    2007-05-09

    As flight speed increases, aerodynamic drag rises more sharply than the availability of atmospheric oxygen. The ratio of oxygen mass flux to dynamic pressure cannot be improved by changing altitude. The maximum possible speed for airbreathing propulsion is limited by the ratio of air capture area to vehicle drag area, approximately Mach 6 at equal areas. Simulation of vehicle acceleration shows that the use of atmospheric oxygen offers a significant potential for minimizing onboard consumables at low speeds. These fundamental calculations indicate that a practical airbreathing launch vehicle would accelerate to near steady-state speed while consuming only onboard fuel, then transition to rocket propulsion. It is suggested that an aircraft carrying a rocket-propelled vehicle to approximately Mach 5 could be a realistic technical goal toward improving access to orbit.

  17. Simulation of hypersonic scramjet exhaust. [pressure distribution on afterbody/nozzle sections of vehicle

    NASA Technical Reports Server (NTRS)

    Oman, R. A.; Foreman, K. M.; Leng, J.; Hopkins, H. B.

    1975-01-01

    A plan and some preliminary analysis for the accurate simulation of pressure distributions on the afterbody/nozzle portions of a hypersonic scramjet vehicle are described. The objectives fulfilled were to establish the standards of similitude for a hydrogen/air scramjet exhaust interacting with a vehicle afterbody, determine an experimental technique for validation of the procedures that will be used in conventional wind tunnel facilities, suggest a program of experiments for proof of the concept, and explore any unresolved problems in the proposed simulation procedures. It is shown that true enthalpy, Reynolds number, and nearly exact chemistry can be provided in the exhaust flow for the flight regime from Mach 4 to 10 by a detonation tube simulation. A detailed discussion of the required similarity parameters leads to the conclusion that substitute gases can be used as the simulated exhaust gas in a wind tunnel to achieve the correct interaction forces and moments.

  18. Methods for estimating pressure and thermal loads induced by elevon deflections on hypersonic-vehicle surfaces with turbulent boundary layers

    NASA Technical Reports Server (NTRS)

    Kaufman, L. G., II; Johnson, C. B.

    1981-01-01

    Empirical anaytic methods are presented for calculating thermal and pressure distributions in three-dimensional, shock-wave turbulent-boundary-layer, interaction-flow regions on the surface of controllable hypersonic aircraft and missiles. The methods, based on several experimental investigations, are useful and reliable for estimating both the extent and magnitude of the increased thermal and pressure loads on the vehicle surfaces.

  19. International Space Station as an Observation Platform for Hypersonic Re-Entry of its Visiting Vehicles

    NASA Technical Reports Server (NTRS)

    Bacon, John B.

    2001-01-01

    The International Space Station (ISS) will receive an armada of visiting supply vehicles during its life in orbit. Over 500 tons of material will be destroyed in targeted re-entries of these vehicles. Because all such re-entries lie in the same orbital plane of the station, and because the visiting vehicles typically deorbit within a few hours of departure, the ISS will usually be within sight of the re-entry process, at a range of only 300-600 kilometers. This vantage point offers an unprecedented opportunity for systematically measuring hypersonic destructive processes. This paper examines the integrated operational constraints of the ISS, its supply vehicles, and candidate sensors which can be employed in the scientific observation of the re-entry process. It is asserted the ISS program has the potential to reduce the worldwide risks from future deorbiting spacecraft, through systematic experimental characterization of the factors which affect the rupture, debris survival, and footprint size of its visiting vehicle fleet.

  20. An Airbreathing Launch Vehicle Design with Turbine-Based Low-Speed Propulsion and Dual Mode Scramjet High-Speed Propulsion

    NASA Technical Reports Server (NTRS)

    Moses, P. L.; Bouchard, K. A.; Vause, R. F.; Pinckney, S. Z.; Ferlemann, S. M.; Leonard, C. P.; Taylor, L. W., III; Robinson, J. S.; Martin, J. G.; Petley, D. H.

    1999-01-01

    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 perspective are addressed. This paper primarily focuses on the design details of the preferred configuration and the analyses performed to assess its performance. The integration of both low-speed and high-speed propulsion is covered. Structural and mechanical designs are described along with materials and technologies used. Propellant and systems packaging are shown and the mission-sized vehicle weights are disclosed.

  1. Australian Air Breathing Propulsion Research for Hypersonic, Beamed Energy-Propelled Vehicles

    NASA Astrophysics Data System (ADS)

    Froning, David

    2010-05-01

    A three year laser-propelled vehicle analysis and design investigation has been begun in June, 2009 by Faculty and graduate students at the University of Adelaide under a Grant/Cooperative Agreement Award to the University of Adelaide by the Asian Office of Aerospace Research and Development (AOARD). The major objectives of thsis investigation are: (a) development of hypersonic, air breathing "lightcraft" with innovative air inlets that enable acceptable airflow capture and combustion, and acceptable cowl-lip heating rates during hot, high-speed, high angle-of-attack hypersonic flight; (b) yest of the most promising lightcraft and inlet design in the high power laser beam that is part of the shock tunnel facility at CTO Instituto in Brazil; and (c) plan a series of laser guided and propelled flights that achieve supersonic or higher speed at the Woomera Test Facility (WTF) in South Australia—using the existing WTF launching and tracking facilities and sponsor-provided laser pointing and tracking and illumination systems.

  2. Flow analysis and design optimization methods for nozzle afterbody of a hypersonic vehicle

    NASA Technical Reports Server (NTRS)

    Baysal, Oktay

    1991-01-01

    This report summarizes the methods developed for the aerodynamic analysis and the shape optimization of the nozzle-afterbody section of a hypersonic vehicle. Initially, exhaust gases were assumed to be air. Internal-external flows around a single scramjet module were analyzed by solving the three dimensional Navier-Stokes equations. Then, exhaust gases were simulated by a cold mixture of Freon and Argon. Two different models were used to compute these multispecies flows as they mixed with the hypersonic airflow. Surface and off-surface properties were successfully compared with the experimental data. In the second phase of this project, the Aerodynamic Design Optimization with Sensitivity analysis (ADOS) was developed. Pre and post optimization sensitivity coefficients were derived and used in this quasi-analytical method. These coefficients were also used to predict inexpensively the flow field around a changed shape when the flow field of an unchanged shape was given. Starting with totally arbitrary initial afterbody shapes, independent computations were converged to the same optimum shape, which rendered the maximum axial thrust.

  3. Global Sliding Mode Control for the Bank-to-Turn of Hypersonic Glide Vehicle

    NASA Astrophysics Data System (ADS)

    Zhang, J.; Yu, Y. F.; Yan, P. P.; Fan, Y. H.; Guo, X. W.

    2017-03-01

    The technology of Bank-to-Turn has been recognized as an attractive direction due to their significance for the control of hypersonic glide vehicle. Strong coupling existing among pitch, yaw and roll channel was a great challenge for banking to turn, and thus a novel global sliding mode controller was designed for hypersonic glider in this paper. Considering the coupling among channels as interference, we can use invariance principle of sliding mode motion to realize the decoupling control. The global sliding mode control system could eliminate the stage of reaching, which can lead to the realization of whole systematic process decoupling control. When the global sliding mode factor was designed, a minimum norm pole assignment method of the sliding mode matrix was introduced to improve the robustness of the system. The method of continuity of symbolic function was adopted to overcome the chatter, which furtherly modify the transient performance of the system. The simulation results show that this method has good performance of three channel decoupling control and guidance command tracking. And it can meet the requirements of the dynamic performance of the system.

  4. Flow analysis and design optimization methods for nozzle-afterbody of a hypersonic vehicle

    NASA Technical Reports Server (NTRS)

    Baysal, O.

    1992-01-01

    This report summarizes the methods developed for the aerodynamic analysis and the shape optimization of the nozzle-afterbody section of a hypersonic vehicle. Initially, exhaust gases were assumed to be air. Internal-external flows around a single scramjet module were analyzed by solving the 3D Navier-Stokes equations. Then, exhaust gases were simulated by a cold mixture of Freon and Ar. Two different models were used to compute these multispecies flows as they mixed with the hypersonic airflow. Surface and off-surface properties were successfully compared with the experimental data. The Aerodynamic Design Optimization with Sensitivity analysis was then developed. Pre- and postoptimization sensitivity coefficients were derived and used in this quasi-analytical method. These coefficients were also used to predict inexpensively the flow field around a changed shape when the flow field of an unchanged shape was given. Starting with totally arbitrary initial afterbody shapes, independent computations were converged to the same optimum shape, which rendered the maximum axial thrust.

  5. Conservation equations and physical models for hypersonic air flows over the aeroassist flight experiment vehicle

    NASA Technical Reports Server (NTRS)

    Gnoffo, Peter A.

    1989-01-01

    The code development and application program for the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA), with emphasis directed toward support of the Aeroassist Flight Experiment (AFE) in the near term and Aeroassisted Space Transfer Vehicle (ASTV) design in the long term is reviewed. LAURA is an upwind-biased, point-implicit relaxation algorithm for obtaining the numerical solution to the governing equations for 3-D, viscous, hypersonic flows in chemical and thermal nonequilibrium. The algorithm is derived using a finite volume formulation in which the inviscid components of flux across cell walls are described with Roe's averaging and Harten's entropy fix with second-order corrections based on Yee's Symmetric Total Variation Diminishing scheme. Because of the point-implicit relaxation strategy, the algorithm remains stable at large Courant numbers without the necessity of solving large, block tri-diagonal systems. A single relaxation step depends only on information from nearest neighbors. Predictions for pressure distributions, surface heating, and aerodynamic coefficients compare well with experimental data for Mach 10 flow over an AFE wind tunnel model. Predictions for the hypersonic flow of air in chemical and thermal nonequilibrium over the full scale AFE configuration obtained on a multi-domain grid are discussed.

  6. The stability of rolling motion of hypersonic vehicles with slender configuration under pitching maneuvering

    NASA Astrophysics Data System (ADS)

    Ye, YouDa; Tian, Hao; Zhang, XianFeng

    2015-06-01

    The configurations of near space hypersonic flying vehicles are considerably different from those of conventional aircrafts. Their configurations are relatively slender; hence their moment of inertia around the longitudinal axis is much smaller than those around the other two axes, resulting in strong coupling of rotations around the three axes. Thus, the stability analysis of rolling motion for such flying vehicles is more complicated than those for conventional aircrafts, and there is no available result of stability analysis which can readily be applied to such cases. This paper is mainly concerned with the stated problem. Considering the practical situation, our investigation is targeted a slightly simpler problem, namely the rolling stability of flying vehicle under known pitching motion. The stability criterion of rolling motion is obtained with and without lateral motions. We also conducted numerical simulation for the pitching-rolling coupled motions of flying vehicles by solving Navier-Stokes equations coupled with dynamic equations of flight. The results of simulation agree well with those of theoretical analysis and experiments.

  7. Design Refinement and Modeling Methods for Highly-Integrated Hypersonic Vehicles

    NASA Astrophysics Data System (ADS)

    Torrez, Sean Michael

    A method for early-stage design of high-speed airplanes is presented based on analysis of vehicle performance, including internal flow in the engine and external flows around the body. Several ways of evaluating vehicle performance are shown, including thrust maps, combustor mode stability concerns, combustor optimization and trajectory optimization. The design performance analysis relies on a routine that computes the thrust of a dual-mode scramjet, which is a geometric-compression (ramjet) engine with a combustor that can operate both subsonically and supersonically. This strategy applies to any internal flow which is predominantly one-dimensional in character. A reduced-order model for mixing and combustion has been developed that is based on non-dimensional scaling of turbulent jets in crossflow and tabulated flamelet chemistry, and is used in conjunction with conventional conservation equations for quasi one-dimensional flow to compute flowpath performance. Thrust is computed by stream-tube momentum analysis. Vehicle lift and drag are computed using a supersonic panel method, developed separately. Comparisons to computational fluid dynamics solutions and experimental data were conducted to determine the validity of the combustion modeling approach, and results of these simulations are shown. Computations for both ram-mode and scram-mode operation are compared to experimental results, and predictions are made for flight conditions of a hypersonic vehicle built around the given flowpath. Trajectory performance of the vehicle is estimated using a collocation method to find the required control inputs and fuel consumption. The combustor is optimized for minimum fuel consumption over a short scram trajectory, and the scram-mode trajectory is optimized for minimum fuel consumption over a space-access-type trajectory. A vehicle design and associated optimized trajectory are shown, and general design principles for steady and efficient operation of vehicles of this

  8. Impact of cabin environment on thermal protection system of crew hypersonic vehicle

    NASA Astrophysics Data System (ADS)

    Zhu, Xiao Wei; Zhao, Jing Quan; Zhu, Lei; Yu, Xi Kui

    2016-05-01

    Hypersonic crew vehicles need reliable thermal protection systems (TPS) to ensure their safety. Since there exists relative large temperature difference between cabin airflow and TPS structure, the TPS shield that covers the cabin is always subjected to a non-adiabatic inner boundary condition, which may influence the heat transfer characteristic of the TPS. However, previous literatures always neglected the influence of the inner boundary by assuming that it was perfectly adiabatic. The present work focuses on studying the impact of cabin environment on the thermal performance. A modified TPS model is created with a mixed thermal boundary condition to connect the cabin environment with the TPS. This helps make the simulation closer to the real situation. The results stress that cabin environment greatly influences the temperature profile inside the TPS, which should not be neglected in practice. Moreover, the TPS size can be optimized during the design procedure if taking the effect of cabin environment into account.

  9. Verification of computational aerodynamic predictions for complex hypersonic vehicles using the INCA{trademark} code

    SciTech Connect

    Payne, J.L.; Walker, M.A.

    1995-01-01

    This paper describes a process of combining two state-of-the-art CFD tools, SPRINT and INCA, in a manner which extends the utility of both codes beyond what is possible from either code alone. The speed and efficiency of the PNS code, SPRING, has been combined with the capability of a Navier-Stokes code to model fully elliptic, viscous separated regions on high performance, high speed flight systems. The coupled SPRINT/INCA capability is applicable for design and evaluation of high speed flight vehicles in the supersonic to hypersonic speed regimes. This paper describes the codes involved, the interface process and a few selected test cases which illustrate the SPRINT/INCA coupling process. Results have shown that the combination of SPRINT and INCA produces correct results and can lead to improved computational analyses for complex, three-dimensional problems.

  10. Comparative Analysis of Two-Stage-to-Orbit Rocket and Airbreathing Reusable Launch Vehicles for Military Applications

    DTIC Science & Technology

    2006-03-01

    cycle ( TBCC ) for the both stages. The primary figures of merit are vehicle empty mass and vehicle wetted area. Empty mass is an indicator of...25 Figure 13. Diagram of vehicle using TBCC propulsion .................................................. 26...stage-to-orbit STS Space Transportation System, the space shuttle TAV Transatmospheric Vehicle TBCC Turbine-based combined-cycle Turb

  11. Investigation of piloting aids for manual control of hypersonic maneuvers

    NASA Technical Reports Server (NTRS)

    Raney, David L.; Phillips, Michael R.; Person, Lee H., Jr.

    1995-01-01

    An investigation of piloting aids designed to provide precise maneuver control for an air-breathing hypersonic vehicle is described. Stringent constraints and nonintuitive high-speed flight effects associated with maneuvering in the hypersonic regime raise the question of whether manual control of such a vehicle should even be considered. The objectives of this research were to determine the extent of manual control that is desirable for a vehicle maneuvering in this regime and to identify the form of aids that must be supplied to the pilot to make such control feasible. A piloted real-time motion-based simulation of a hypersonic vehicle concept was used for this study, and the investigation focused on a single representative cruise turn maneuver. Piloting aids, which consisted of an auto throttle, throttle director, autopilot, flight director, and two head-up display configurations, were developed and evaluated. Two longitudinal control response types consisting of a rate-command/attitude-hold system and a load factor-rate/load-factor-hold system were also compared. The complete set of piloting aids, which consisted of the autothrottle, throttle director, and flight director, improved the average Cooper-Harper flying qualities ratings from 8 to 2.6, even though identical inner-loop stability and control augmentation was provided in all cases. The flight director was determined to be the most critical of these aids, and the cruise turn maneuver was unachievable to adequate performance specifications in the absence of this flight director.

  12. Airbreathing/Rocket Single-Stage-to-Orbit Design Matrix

    NASA Technical Reports Server (NTRS)

    Hunt, James L.

    1995-01-01

    A definitive design/performance study was performed on a single-stage-to-orbit (SSTO) airbreathing propelled orbital vehicle with rocket propulsion augmentation in the Access to Space activities during 1993. A credible reference design was established, but by no means an optimum. The results supported the viability of SSTO airbreathing/rocket vehicles for operational scenarios and indicated compelling reasons to continue to explore the design matrix. This paper will (1) summarize the Access to Space design activity from the SSTO airbreathing/rocket perspective, (2) present an airbreathing/rocket SSTO design matrix established for continued optimization of the design space, and (3) focus on the compelling reasons for airbreathing vehicles in Access to Space scenarios.

  13. Nonlinear fractional order proportion-integral-derivative active disturbance rejection control method design for hypersonic vehicle attitude control

    NASA Astrophysics Data System (ADS)

    Song, Jia; Wang, Lun; Cai, Guobiao; Qi, Xiaoqiang

    2015-06-01

    Near space hypersonic vehicle model is nonlinear, multivariable and couples in the reentry process, which are challenging for the controller design. In this paper, a nonlinear fractional order proportion integral derivative (NFOPIλDμ) active disturbance rejection control (ADRC) strategy based on a natural selection particle swarm (NSPSO) algorithm is proposed for the hypersonic vehicle flight control. The NFOPIλDμ ADRC method consists of a tracking-differentiator (TD), an NFOPIλDμ controller and an extended state observer (ESO). The NFOPIλDμ controller designed by combining an FOPIλDμ method and a nonlinear states error feedback control law (NLSEF) is to overcome concussion caused by the NLSEF and conversely compensate the insufficiency for relatively simple and rough signal processing caused by the FOPIλDμ method. The TD is applied to coordinate the contradiction between rapidity and overshoot. By attributing all uncertain factors to unknown disturbances, the ESO can achieve dynamic feedback compensation for these disturbances and thus reduce their effects. Simulation results show that the NFOPIλDμ ADRC method can make the hypersonic vehicle six-degree-of-freedom nonlinear model track desired nominal signals accurately and fast, has good stability, dynamic properties and strong robustness against external environmental disturbances.

  14. Hypersonic aerospace vehicle leading-edge cooling using heat-pipe, transpiration and film-cooling techniques

    SciTech Connect

    Modlin, J.M.

    1991-01-01

    The feasibility of cooling hypersonic-vehicle leading-edge structures exposed to severe aerodynamic surface heat fluxes was studied, using a combination of liquid-metal heat pipes and surface-mass-transfer cooling techniques. A generalized, transient, finite-difference-based hypersonic leading-edge cooling model was developed that incorporated these effects and was demonstrated on an assumed aerospace plane-type wing leading edge section and a SCRAMJET engine inlet leading-edge section. The hypersonic leading-edge cooling model was developed using an existing, experimentally verified heat-pipe model. Then the existing heat-pipe model was modified by adding both transpiration and film-cooling options as new surface boundary conditions. The models used to predict the leading-edge surface heat-transfer reduction effects of the transpiration and film cooling were modifications of more-generalized, empirically based models obtained from the literature. It is concluded that cooling leading-edge structures exposed to severe hypersonic-flight environments using a combination of liquid-metal heat pipe, surface transpiration, and film cooling methods appears feasible.

  15. Hypersonic aeroheating test of space shuttle vehicle configuration 3 (model 22-OTS) in the NASA-Ames 3.5-foot hypersonic wind tunnel (IH20), volume 1

    NASA Technical Reports Server (NTRS)

    Kingsland, R. B.; Lockman, W. K.

    1975-01-01

    The results of hypersonic wind tunnel testing of an 0.0175 scale version of the vehicle 3 space shuttle configuration are presented. Temperature measurements were made on the launch configuration, orbiter plus tank, orbiter alone, tank alone, and solid rocket booster alone to provide heat transfer data. The test was conducted at free-stream Mach numbers of 5.3 and 7.3 and at free-stream Reynolds numbers of 1.5 million, 3.7 million, 5.0 million, and 7.0 million per foot. The model was tested at angles of attack from -5 deg to 20 deg and side slip angles of -5 deg and 0 deg.

  16. Thermal-structural analysis of the platelet heat-pipe-cooled leading edge of hypersonic vehicle

    NASA Astrophysics Data System (ADS)

    Hongpeng, Liu; Weiqiang, Liu

    2016-10-01

    One of the main challenges for the hypersonic vehicle is its thermal protection, more specifically, the cooling of its leading edge. To investigate the feasibility of a platelet heat-pipe-cooled leading edge structure, thermal/stress distributions for steady-state flight conditions are calculated numerically. Studies are carried on for IN718/Na, C-103/Na and T-111/Li compatible material combinations of heat pipe under nominal operations and a central heat pipe failure cases, and the influence of wall thickness on the design robustness is also investigated. And the heat transfer limits (the sonic limit, the capillary limit and the boiling limit) are also computed to check the operation of platelet heat pipes. The results indicate that, with a 15 mm leading edge radius and a wall thickness of 0.5 mm, C-103/Na and T-111/Li combinations of heat pipe is capable of withstanding both nominal and failure conditions for Mach 8 and Mach 10 flight respectively.

  17. Linear parameter varying switching attitude control for a near space hypersonic vehicle with parametric uncertainties

    NASA Astrophysics Data System (ADS)

    Huang, Yiqing; Sun, Changyin; Qian, Chengshan; Wang, Li

    2015-12-01

    This paper deals with the problem of linear parameter varying (LPV) switching attitude control for a near space hypersonic vehicle (NSHV) with parametric uncertainties. First, due to the enormous complexity of the NSHV nonlinear attitude dynamics, a slow-fast loop polytopic LPV attitude model is developed by using Jacobian linearisation and the tensor product model transformation approach. Second, for the purpose of less conservative attitude controller design, the flight envelope is divided into four subregions. For each parameter subregion, slow-loop and fast-loop LPV controllers are designed. By the defined switching character function, these slow-fast loop LPV controllers are then switched in order to guarantee the closed-loop NSHV system to be asymptotically stable and satisfy a specified tracking performance criterion. The condition of LPV switching attitude controller synthesis is given in terms of linear matrix inequalities, which can be readily solved via standard numerical software, and the robust stability analysis of the closed-loop NSHV system is verified based on multiple Lypapunov functions. Finally, numerical simulations have demonstrated the effectiveness of the proposed approach.

  18. Active disturbance rejection based trajectory linearization control for hypersonic reentry vehicle with bounded uncertainties.

    PubMed

    Shao, Xingling; Wang, Honglun

    2015-01-01

    This paper investigates a novel compound control scheme combined with the advantages of trajectory linearization control (TLC) and alternative active disturbance rejection control (ADRC) for hypersonic reentry vehicle (HRV) attitude tracking system with bounded uncertainties. Firstly, in order to overcome actuator saturation problem, nonlinear tracking differentiator (TD) is applied in the attitude loop to achieve fewer control consumption. Then, linear extended state observers (LESO) are constructed to estimate the uncertainties acting on the LTV system in the attitude and angular rate loop. In addition, feedback linearization (FL) based controllers are designed using estimates of uncertainties generated by LESO in each loop, which enable the tracking error for closed-loop system in the presence of large uncertainties to converge to the residual set of the origin asymptotically. Finally, the compound controllers are derived by integrating with the nominal controller for open-loop nonlinear system and FL based controller. Also, comparisons and simulation results are presented to illustrate the effectiveness of the control strategy.

  19. Sliding mode based trajectory linearization control for hypersonic reentry vehicle via extended disturbance observer.

    PubMed

    Xingling, Shao; Honglun, Wang

    2014-11-01

    This paper proposes a novel hybrid control framework by combing observer-based sliding mode control (SMC) with trajectory linearization control (TLC) for hypersonic reentry vehicle (HRV) attitude tracking problem. First, fewer control consumption is achieved using nonlinear tracking differentiator (TD) in the attitude loop. Second, a novel SMC that employs extended disturbance observer (EDO) to counteract the effect of uncertainties using a new sliding surface which includes the estimation error is integrated to address the tracking error stabilization issues in the attitude and angular rate loop, respectively. In addition, new results associated with EDO are examined in terms of dynamic response and noise-tolerant performance, as well as estimation accuracy. The key feature of the proposed compound control approach is that chattering free tracking performance with high accuracy can be ensured for HRV in the presence of multiple uncertainties under control constraints. Based on finite time convergence stability theory, the stability of the resulting closed-loop system is well established. Also, comparisons and extensive simulation results are presented to demonstrate the effectiveness of the control strategy.

  20. An extended supersonic combustion model for the dynamic analysis of hypersonic vehicles. Interim Task Report

    SciTech Connect

    Bossard, J.A.; Peck, R.E.; Schmidt, D.K.

    1993-03-01

    The development of an advanced dynamic model for aeroelastic hypersonic vehicles powered by air breathing engines requires an adequate engine model. This report provides a discussion of some of the more important features of supersonic combustion and their relevance to the analysis and design of supersonic ramjet engines. Of particular interest are those aspects of combustion that impact the control of the process. Furthermore, the report summarizes efforts to enhance the aeropropulsive/aeroelastic dynamic model developed at the Aerospace Research Center of Arizona State University by focusing on combustion and improved modeling of this flow. The expanded supersonic combustor model described here has the capability to model the effects of friction, area change, and mass addition, in addition to the heat addition process. A comparison is made of the results from four cases: (1) heat addition only; (2) heat addition plus friction; (3) heat addition, friction, and area reduction, and (4) heat addition, friction, area reduction, and mass addition. The relative impact of these effects on the Mach number, static temperature, and static pressure distributions within the combustor are then shown. Finally, the effects of frozen versus equilibrium flow conditions within the exhaust plume is discussed.

  1. Thermal Performance of Composite Flexible Blanket Insulations for Hypersonic Aerospace Vehicles

    NASA Technical Reports Server (NTRS)

    Kourtides, Demetrius A.

    1993-01-01

    This paper describes the thermal performance of a Composite Flexible Blanket Insulation (C.F.B.I.) considered for potential use as a thermal protection system or thermal insulation for future hypersonic vehicles such as the National Aerospace Plane (N.A.S.P.). Thermophysical properties for these insulations were also measured including the thermal conductivity at various temperatures and pressures and the emissivity of the fabrics used in the flexible insulations. The thermal response of these materials subjected to aeroconvective heating from a plasma arc is also described. Materials tested included two surface variations of the insulations, and similar insulations coated with a Protective Ceramic Coating (P.C.C.). Surface and backface temperatures were measured in the flexible insulations and on Fibrous Refractory Composite Insulation (F.R.C.I.) used as a calibration model. The uncoated flexible insulations exhibited good thermal performance up to 35 W/sq cm. The use of a P.C.C. to protect these insulations at higher heating rates is described. The results from a computerized thermal analysis model describing thermal response of those materials subjected to the plasma arc conditions are included. Thermal and optical properties were determined including thermal conductivity for the rigid and flexible insulations and emissivity for the insulation fabrics. These properties were utilized to calculate the thermal performance of the rigid and flexible insulations at the maximum heating rate.

  2. Optimum Design of Hypersonic Airbreathing Propulsion

    NASA Astrophysics Data System (ADS)

    Kobayashi, Hiroaki; Sato, Tetsuya; Tanatsugu, Nobuhiro

    The flight of Spaceplane is always under accelarating in the assent way and always under decelarating in the desent way and yet cruising in the return way. Besides, its flight envelope is considerably wider than that of airplane. Thus the integrated design method is required to build the best transportation system optimized taking into account the propulsion system and the airframe under the entire flight conditions. In this paper it is shown an optimization method on TSTO spaceplane system. Genetic algorithm (GA) was applied to optimize design parameters of engine, airframe, and trajectory simultaneously. Several types of engine were quantitatively compared using payload ratio as an evaluating function. It was concluded that precooled turbojets is the most promising engine for TSTO among Turbine Based Combined Cycle (TBCC) engines.

  3. eLaunch Hypersonics: An Advanced Launch System

    NASA Technical Reports Server (NTRS)

    Starr, Stanley

    2010-01-01

    This presentation describes a new space launch system that NASA can and should develop. This approach can significantly reduce ground processing and launch costs, improve reliability, and broaden the scope of what we do in near earth orbit. The concept (not new) is to launch a re-usable air-breathing hypersonic vehicle from a ground based electric track. This vehicle launches a final rocket stage at high altitude/velocity for the final leg to orbit. The proposal here differs from past studies in that we will launch above Mach 1.5 (above transonic pinch point) which further improves the efficiency of air breathing, horizontal take-off launch systems. The approach described here significantly reduces cost per kilogram to orbit, increases safety and reliability of the boost systems, and reduces ground costs due to horizontal-processing. Finally, this approach provides significant technology transfer benefits for our national infrastructure.

  4. Hypersonic missile propulsion system

    SciTech Connect

    Kazmar, R.R.

    1998-11-01

    Pratt and Whitney is developing the technology for hypersonic components and engines. A supersonic combustion ramjet (scramjet) database was developed using hydrogen fueled propulsion systems for space access vehicles and serves as a point of departure for the current development of hydrocarbon scramjets. The Air Force Hypersonic Technology (HyTech) Program has put programs in place to develop the technologies necessary to demonstrate the operability, performance and structural durability of an expendable, liquid hydrocarbon fueled scramjet system that operates from Mach 4 to 8. This program will culminate in a flight type engine test at representative flight conditions. The hypersonic technology base that will be developed and demonstrated under HyTech will establish the foundation to enable hypersonic propulsion systems for a broad range of air vehicle applications from missiles to space access vehicles. A hypersonic missile flight demonstration is planned in the DARPA Affordable Rapid Response Missile Demonstrator (ARRMD) program in 2001.

  5. Evaluation of on-board hydrogen storage methods for hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Akyurtlu, Ates; Akyurtlu, J. F.; Adeyiga, A. A.; Perdue, Samara; Northam, G. B.

    1989-01-01

    Hydrogen is the foremost candidate as a fuel for use in high speed transport. Since any aircraft moving at hypersonic speeds must have a very slender body, means of decreasing the storage volume requirements below that for liquid hydrogen are needed. The total performance of the hypersonic plane needs to be considered for the evaluation of candidate fuel and storage systems. To accomplish this, a simple model for the performance of a hypersonic plane is presented. To allow for the use of different engines and fuels during different phases of flight, the total trajectory is divided into three phases: subsonic-supersonic, hypersonic and rocket propulsion phase. The fuel fraction for the first phase is found be a simple energy balance using an average thrust to drag ratio for this phase. The hypersonic flight phase is investigated in more detail by taking small altitude increments. This approach allowed the use of flight profiles other than the constant dynamic pressure flight. The effect of fuel volume on drag, structural mass and tankage mass was introduced through simplified equations involving the characteristic dimension of the plane. The propellant requirement for the last phase is found by employing the basic rocket equations. The candidate fuel systems such as the cryogenic fuel combinations and solid and liquid endothermic hydrogen generators are first screened thermodynamically with respect to their energy densities and cooling capacities and then evaluated using the above model.

  6. Computational fluid dynamics application to hypersonic flow over a Martian entry vehicle (MEV) - A correlation with experiment

    NASA Technical Reports Server (NTRS)

    Stewart, John E.; Smith, Robert E.; Ashby, George C., Jr.

    1992-01-01

    A class of vehicles for a mission to Mars are analyzed for aerodynamic characteristics using advanced Computational Fluid Dynamics (CFD). The general configuration is a modified cone-conical-frustum geometry where the nose radius has a large influence on the flowfield. Inviscid-compressible flow using the Euler equations and viscous-compressible flow using the thin-layer Navier-Stokes equations is applied to the configuration. The surface modeling, grid generation and application of state-of-the-art CFD software are described. The effects of nose radius, angle of attack, and hypersonic velocity on the flight characteristics of the vehicle are discussed. The numerical simulations demonstrate the merits of the inviscid and viscous software. Results are compared with wind tunnel experiments.

  7. An Overview of the Role of Systems Analysis in NASA's Hypersonics Project

    NASA Technical Reports Server (NTRS)

    Robinson, Jeffrey S.; Martin John G.; Bowles, Jeffrey V> ; Mehta, Unmeel B.; Snyder, CHristopher A.

    2006-01-01

    NASA's Aeronautics Research Mission Directorate recently restructured its Vehicle Systems Program, refocusing it towards understanding the fundamental physics that govern flight in all speed regimes. Now called the Fundamental Aeronautics Program, it is comprised of four new projects, Subsonic Fixed Wing, Subsonic Rotary Wing, Supersonics, and Hypersonics. The Aeronautics Research Mission Directorate has charged the Hypersonics Project with having a basic understanding of all systems that travel at hypersonic speeds within the Earth's and other planets atmospheres. This includes both powered and unpowered systems, such as re-entry vehicles and vehicles powered by rocket or airbreathing propulsion that cruise in and accelerate through the atmosphere. The primary objective of the Hypersonics Project is to develop physics-based predictive tools that enable the design, analysis and optimization of such systems. The Hypersonics Project charges the systems analysis discipline team with providing it the decision-making information it needs to properly guide research and technology development. Credible, rapid, and robust multi-disciplinary system analysis processes and design tools are required in order to generate this information. To this end, the principal challenges for the systems analysis team are the introduction of high fidelity physics into the analysis process and integration into a design environment, quantification of design uncertainty through the use of probabilistic methods, reduction in design cycle time, and the development and implementation of robust processes and tools enabling a wide design space and associated technology assessment capability. This paper will discuss the roles and responsibilities of the systems analysis discipline team within the Hypersonics Project as well as the tools, methods, processes, and approach that the team will undertake in order to perform its project designated functions.

  8. A fuselage/tank structure study for actively cooled hypersonic cruise vehicles: Aircraft design evaluation

    NASA Technical Reports Server (NTRS)

    Nobe, T.

    1975-01-01

    The effects of fuselage cross sections and structural members on the performance of hypersonic cruise aircraft are evaluated. Representative fuselage/tank area structure was analyzed for strength, stability, fatigue and fracture mechanics. Various thermodynamic and structural tradeoffs were conducted to refine the conceptual designs with the primary objective of minimizing weight and maximizing aircraft range.

  9. A US History of Airbreathing/Rocket Combined-Cycle (RBCC) Propulsion for Powering Future Aerospace Transports, with a Look Ahead to the Year 2020

    NASA Technical Reports Server (NTRS)

    Escher, William J. D.

    1999-01-01

    A technohistorical and forward-planning overview of U.S. developments in combined airbreathing/rocket propulsion for advanced aerospace vehicle applications is presented. Such system approaches fall into one of two categories: (1) Combination propulsion systems (separate, non-interacting engines installed), and (2) Combined-Cycle systems. The latter, and main subject, comprises a large family of closely integrated engine types, made up of both airbreathing and rocket derived subsystem hardware. A single vehicle-integrated, multimode engine results, one capable of operating efficiently over a very wide speed and altitude range, atmospherically and in space. While numerous combination propulsion systems have reached operational flight service, combined-cycle propulsion development, initiated ca. 1960, remains at the subscale ground-test engine level of development. However, going beyond combination systems, combined-cycle propulsion potentially offers a compelling set of new and unique capabilities. These capabilities are seen as enabling ones for the evolution of Spaceliner class aerospace transportation systems. The following combined-cycle hypersonic engine developments are reviewed: (1) RENE (rocket engine nozzle ejector), (2) Cryojet and LACE, (3) Ejector Ramjet and its derivatives, (4) the seminal NASA NAS7-377 study, (5) Air Force/Marquardt Hypersonic Ramjet, (6) Air Force/Lockheed-Marquardt Incremental Scramjet flight-test project, (7) NASA/Garrett Hypersonic Research Engine (HRE), (8) National Aero-Space Plane (NASP), (9) all past projects; and such current and planned efforts as (10) the NASA ASTP-ART RBCC project, (11) joint CIAM/NASA DNSCRAM flight test,(12) Hyper-X, (13) Trailblazer,( 14) W-Vehicle and (15) Spaceliner 100. Forward planning programmatic incentives, and the estimated timing for an operational Spaceliner powered by combined-cycle engines are discussed.

  10. Approach to solution of coupled heat transfer problem on the surface of hypersonic vehicle of arbitrary shape

    NASA Astrophysics Data System (ADS)

    Bocharov, A. N.; Bityurin, V. A.; Golovin, N. N.; Evstigneev, N. M.; Petrovskiy, V. P.; Ryabkov, O. I.; Teplyakov, I. O.; Shustov, A. A.; Solomonov, Yu S.; Fortov, V. E.

    2016-11-01

    In this paper, an approach to solve conjugate heat- and mass-transfer problems is considered to be applied to hypersonic vehicle surface of arbitrary shape. The approach under developing should satisfy the following demands. (i) The surface of the body of interest may have arbitrary geometrical shape. (ii) The shape of the body can change during calculation. (iii) The flight characteristics may vary in a wide range, specifically flight altitude, free-stream Mach number, angle-of-attack, etc. (iv) The approach should be realized with using the high-performance-computing (HPC) technologies. The approach is based on coupled solution of 3D unsteady hypersonic flow equations and 3D unsteady heat conductance problem for the thick wall. Iterative process is applied to account for ablation of wall material and, consequently, mass injection from the surface and changes in the surface shape. While iterations, unstructured computational grids both in the flow region and within the wall interior are adapted to the current geometry and flow conditions. The flow computations are done on HPC platform and are most time-consuming part of the whole problem, while heat conductance problem can be solved on many kinds of computers.

  11. Computational Sensitivity Analysis for the Aerodynamic Design of Supersonic and Hypersonic Air Vehicles

    DTIC Science & Technology

    2015-05-18

    adaptation. The additional mesh resolution in the vicinity of the leading edge of the wing is necessary to resolve a shockwave that forms at hypersonic...speeds. Cart3D also uses a number of high-resolution CFD techniques to capture shockwaves accurately, reduce computation time, and minimize high...to here are spurious spatial oscillations that often occur in flow solutions with discontinuities such as shockwaves . 14 aforementioned

  12. Integrated Aero-Servo-Thermo-Propulso-Elasticity (ASTPE) for Hypersonic Scramjet Vehicle Design/Analysis

    DTIC Science & Technology

    2009-12-04

    experience in the last decade also indicates that large discrepancies exist between measured data ( wind - tunnel model) and flight data. This is so because the... wind - tunnel data. The coefficients of out-of-phase forces with varying Mach numbers for a pitching 10° cone and a 20° cone about the apex are...Aerodynamic Method for Design of Analysis Shock Expansion (CD) Hypersonic/ Euler & Viscous Aerodynamic Methods CFD(A) - CFD - Inlet Design - Yes

  13. Global neural dynamic surface tracking control of strict-feedback systems with application to hypersonic flight vehicle.

    PubMed

    Xu, Bin; Yang, Chenguang; Pan, Yongping

    2015-10-01

    This paper studies both indirect and direct global neural control of strict-feedback systems in the presence of unknown dynamics, using the dynamic surface control (DSC) technique in a novel manner. A new switching mechanism is designed to combine an adaptive neural controller in the neural approximation domain, together with the robust controller that pulls the transient states back into the neural approximation domain from the outside. In comparison with the conventional control techniques, which could only achieve semiglobally uniformly ultimately bounded stability, the proposed control scheme guarantees all the signals in the closed-loop system are globally uniformly ultimately bounded, such that the conventional constraints on initial conditions of the neural control system can be relaxed. The simulation studies of hypersonic flight vehicle (HFV) are performed to demonstrate the effectiveness of the proposed global neural DSC design.

  14. Hypersonic Boundary-Layer Transition for X-33 Phase 2 Vehicle

    NASA Technical Reports Server (NTRS)

    Thompson, Richard A.; Hamilton, Harris H., II; Berry, Scott A.; Horvath, Thomas J.; Nowak, Robert J.

    1998-01-01

    A status review of the experimental and computational work performed to support the X-33 program in the area of hypersonic boundary-layer transition is presented. Global transition fronts are visualized using thermographic phosphor measurements. Results are used to derive transition correlations for "smooth body" and discrete roughness data and a computational tool is developed to predict transition onset for X-33 using these results. The X-33 thermal protection system appears to be conservatively designed for transition effects based on these studies. Additional study is needed to address concerns related to surface waviness. A discussion of future test plans is included.

  15. An Evaluation of High Temperature Airframe Seals for Advanced Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    DeMange, Jeffrey J.; Dunlap, Patrick H.; Steinetz, Bruce M.; Drlik, Gary J.

    2007-01-01

    High temperature seals are required for advanced hypersonic airframe applications. In this study, both spring tube thermal barriers and innovative wafer seal systems were evaluated under relevant hypersonic test conditions (temperatures, pressures, etc.) via high temperature compression testing and room temperature flow assessments. Thermal barriers composed of a Rene 41 spring tube filled with Saffil insulation and overbraided with a Nextel 312 sheath showed acceptable performance at 1500 F in both short term and longer term compression testing. Nextel 440 thermal barriers with Rene 41 spring tubes and Saffil insulation demonstrated good compression performance up to 1750 F. A silicon nitride wafer seal/compression spring system displayed excellent load performance at temperatures as high as 2200 F and exhibited room temperature leakage values that were only 1/3 those for the spring tube rope seals. For all seal candidates evaluated, no significant degradation in leakage resistance was noted after high temperature compression testing. In addition to these tests, a superalloy seal suitable for dynamic seal applications was optimized through finite element techniques.

  16. Hypersonic structures: An aerodynamicist's perspective, or one man's dream is another man's nightmare

    NASA Technical Reports Server (NTRS)

    Watts, J. D.; Jackson, L. R.; Hunt, J. L.

    1978-01-01

    The relationship between hypersonic aerodynamic and structural design is reviewed. The evolution of the hypersonic vehicle design is presented. Propulsion systems, structural materials, and fuels are emphasized.

  17. The Role of Formal Experiment Design in Hypersonic Flight System Technology Development

    NASA Technical Reports Server (NTRS)

    McClinton, Charles R.; Ferlemann, Shelly M.; Rock, Ken E.; Ferlemann, Paul G.

    2002-01-01

    Hypersonic airbreathing engine (scramjet) powered vehicles are being considered to replace conventional rocket-powered launch systems. Effective utilization of scramjet engines requires careful integration with the air vehicle. This integration synergistically combines aerodynamic forces with propulsive cycle functions of the engine. Due to the highly integrated nature of the hypersonic vehicle design problem, the large flight envelope, and the large number of design variables, the use of a statistical design approach in design is effective. Modern Design-of-Experiments (MDOE) has been used throughout the Hyper-X program, for both systems analysis and experimental testing. Application of MDOE fall into four categories: (1) experimental testing; (2) studies of unit phenomena; (3) refining engine design; and (4) full vehicle system optimization. The MDOE process also provides analytical models, which are also used to document lessons learned, supplement low-level design tools, and accelerate future studies. This paper will discuss the design considerations for scramjet-powered vehicles, specifics of MDOE utilized for Hyper-X, and present highlights from the use of these MDOE methods within the Hyper-X Program.

  18. STS 2000 - A reference airbreathing SSTO

    NASA Astrophysics Data System (ADS)

    Wagner, Alain; Thevenot, Regis

    1991-12-01

    The purpose of this paper is to present the most recent results (including wind tunnel testing) dealing with a SSTO using airbreathing propulsion up to March 6. The choice of the external shape of the vehicle is explained then the aerodynamic coefficients of the glider are derived. The integration of the propulsion system (a turborocket-ramjet-rocket) to the vehicle is then presented including forebody effects. The launch trajectory and the way to refine it thanks to a simultaneous optimization of the mass budget, is described. The internal accommodation and the structural concepts are briefly presented.

  19. NASA R and T aerospace plane vehicles: Progress and plans

    NASA Technical Reports Server (NTRS)

    Dixon, S. C.

    1985-01-01

    Progress made in key technologies such as materials, structures, aerothermodynamics, hypersonic aerodynamics, and hypersonic airbreathing propulsion are reported. Advances were made in more generic, areas such as active controls, flight computer hardware and software, and interdisciplinary analytical design methodology. These technology advances coupled with the development of and experiences with the Space Shuttle make feasible aerospace plane-type vehicles that meet the more demanding requirements of various DOD missions and/or an all-weather Shuttle II with reduced launch costs. Technology needs and high payoff technologies, and the technology advancements in propulsion, control-configured-vehicles, aerodynamics, aerothermodynamics, aerothermal loads, and materials and structures were studied. The highest payoff technologies of materials and structures including thermal-structural analysis and high temperature test techniques are emphasized. The high priority technology of propulsion, and plans, of what remains to be done rather than firm program commitments, are briefly discussed.

  20. Boundary Layer Transition over Blunt Hypersonic Vehicles Including Effects of Ablation-Induced Out-Gassing

    NASA Technical Reports Server (NTRS)

    Li, Fei; Choudhari, Meelan; Chang, Chau-Lyan; White, Jeffery

    2011-01-01

    Computations are performed to study the boundary layer instability mechanisms pertaining to hypersonic flow over blunt capsules. For capsules with ablative heat shields, transition may be influenced both by out-gassing associated with surface pyrolysis and the resulting modification of surface geometry including the formation of micro-roughness. To isolate the effects of out-gassing, this paper examines the stability of canonical boundary layer flows over a smooth surface in the presence of gas injection into the boundary layer. For a slender cone, the effects of out-gassing on the predominantly second mode instability are found to be stabilizing. In contrast, for a blunt capsule flow dominated by first mode instability, out-gassing is shown to be destabilizing. Analogous destabilizing effects of outgassing are also noted for both stationary and traveling modes of crossflow instability over a blunt sphere-cone configuration at angle of attack.

  1. AIR VEHICLE INTEGRATION AND TECHNOLOGY RESEARCH (AVIATR) Task Order 0015: Predictive Capability for Hypersonic Structural Response and Life Prediction: Phase 1-Identification of Knowledge Gaps, Volume 1: Nonproprietary Version

    DTIC Science & Technology

    2010-09-01

    AFRL-RB-WP-TR-2010-3068,V1 AIR VEHICLE INTEGRATION AND TECHNOLOGY RESEARCH (AVIATR) Task Order 0015: Predictive Capability for Hypersonic ...Order 0015: Predictive Capability for Hypersonic Structural Response and Life Prediction: Phase 1-Identification of Knowledge Gaps, Volume 1...AND ADDRESS(ES) 8. PERFORMING ORGANIZATION The Boeing Company M/C 110-SK56 2600 Westminster Avenue Seal Beach , CA 90740 REPORT NUMBER 9

  2. Iodine Tagging Velocimetry and Mechanism in the Hypersonic Near Wake of a MultiPurpose Crew Vehicle

    NASA Technical Reports Server (NTRS)

    Balla, R. Jeffrey

    2013-01-01

    This study demonstrates a new molecular tagging velocimetry (MTV) method for velocity measurements of high speed flow. It demonstrates offbody Iodine Tagging Velocimetry (ITV) in the hypersonic near wake of a MultiPurpose Crew Vehicle (MPCV) model. Experiments are performed in the NASA-Langley 31-inch Mach 10 air wind tunnel. A 0.5% I2 / N2 mixture is seeded on the leeward backshell of the model using a pressure tap. I2 laser-induced fluorescence is excited along a 5.5 mm line using an ArF excimer laser near 193 nm. Results indicate I2 absorbs at least 2 photons to produce iodine ions and electrons. These recombine as the tagged region is displaced downstream to produce I (2P3/2) whose emission is monitored at 206 nm. Results at P0 = 2.41 MPa (350 psi), T0 = 990K, and 10 micro-sec transit times produce velocities from 630-820 m/sec across the I2 seeded jet at a distance of 38.2 mm (25.5 jet diameters) downstream from the jet orifice. Maximum wake jet velocities near the shear layer are 59% of freestream velocity.

  3. CAVE: A computer code for two-dimensional transient heating analysis of conceptual thermal protection systems for hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Rathjen, K. A.

    1977-01-01

    A digital computer code CAVE (Conduction Analysis Via Eigenvalues), which finds application in the analysis of two dimensional transient heating of hypersonic vehicles is described. The CAVE is written in FORTRAN 4 and is operational on both IBM 360-67 and CDC 6600 computers. The method of solution is a hybrid analytical numerical technique that is inherently stable permitting large time steps even with the best of conductors having the finest of mesh size. The aerodynamic heating boundary conditions are calculated by the code based on the input flight trajectory or can optionally be calculated external to the code and then entered as input data. The code computes the network conduction and convection links, as well as capacitance values, given basic geometrical and mesh sizes, for four generations (leading edges, cooled panels, X-24C structure and slabs). Input and output formats are presented and explained. Sample problems are included. A brief summary of the hybrid analytical-numerical technique, which utilizes eigenvalues (thermal frequencies) and eigenvectors (thermal mode vectors) is given along with aerodynamic heating equations that have been incorporated in the code and flow charts.

  4. Thermostructural Analysis of Unconventional Wing Structures of a Hyper-X Hypersonic Flight Research Vehicle for the Mach 7 Mission

    NASA Technical Reports Server (NTRS)

    Ko, William L.; Gong, Leslie

    2001-01-01

    Heat transfer, thermal stresses, and thermal buckling analyses were performed on the unconventional wing structures of a Hyper-X hypersonic flight research vehicle (designated as X-43) subjected to nominal Mach 7 aerodynamic heating. A wing midspan cross section was selected for the heat transfer and thermal stress analyses. Thermal buckling analysis was performed on three regions of the wing skin (lower or upper); 1) a fore wing panel, 2) an aft wing panel, and 3) a unit panel at the middle of the aft wing panel. A fourth thermal buckling analysis was performed on a midspan wing segment. The unit panel region is identified as the potential thermal buckling initiation zone. Therefore, thermal buckling analysis of the Hyper-X wing panels could be reduced to the thermal buckling analysis of that unit panel. "Buckling temperature magnification factors" were established. Structural temperature-time histories are presented. The results show that the concerns of shear failure at wing and spar welded sites, and of thermal buckling of Hyper-X wing panels, may not arise under Mach 7 conditions.

  5. Simulating flow around scaled model of a hypersonic vehicle in wind tunnel

    NASA Astrophysics Data System (ADS)

    Markova, T. V.; Aksenov, A. A.; Zhluktov, S. V.; Savitsky, D. V.; Gavrilov, A. D.; Son, E. E.; Prokhorov, A. N.

    2016-11-01

    A prospective hypersonic HEXAFLY aircraft is considered in the given paper. In order to obtain the aerodynamic characteristics of a new construction design of the aircraft, experiments with a scaled model have been carried out in a wind tunnel under different conditions. The runs have been performed at different angles of attack with and without hydrogen combustion in the scaled propulsion engine. However, the measured physical quantities do not provide all the information about the flowfield. Numerical simulation can complete the experimental data as well as to reduce the number of wind tunnel experiments. Besides that, reliable CFD software can be used for calculations of the aerodynamic characteristics for any possible design of the full-scale aircraft under different operation conditions. The reliability of the numerical predictions must be confirmed in verification study of the software. The given work is aimed at numerical investigation of the flowfield around and inside the scaled model of the HEXAFLY-CIAM module under wind tunnel conditions. A cold run (without combustion) was selected for this study. The calculations are performed in the FlowVision CFD software. The flow characteristics are compared against the available experimental data. The carried out verification study confirms the capability of the FlowVision CFD software to calculate the flows discussed.

  6. Hypersonic vehicle model and control law development using H(infinity) and micron synthesis

    NASA Astrophysics Data System (ADS)

    Gregory, Irene M.; Chowdhry, Rajiv S.; McMinn, John D.; Shaughnessy, John D.

    1994-10-01

    The control system design for a Single Stage To Orbit (SSTO) air breathing vehicle will be central to a successful mission because a precise ascent trajectory will preserve narrow payload margins. The air breathing propulsion system requires the vehicle to fly roughly halfway around the Earth through atmospheric turbulence. The turbulence, the high sensitivity of the propulsion system to inlet flow conditions, the relatively large uncertainty of the parameters characterizing the vehicle, and continuous acceleration make the problem especially challenging. Adequate stability margins must be provided without sacrificing payload mass since payload margins are critical. Therefore, a multivariable control theory capable of explicitly including both uncertainty and performance is needed. The H(infinity) controller in general provides good robustness but can result in conservative solutions for practical problems involving structured uncertainty. Structured singular value mu framework for analysis and synthesis is potentially much less conservative and hence more appropriate for problems with tight margins. An SSTO control system requires: highly accurate tracking of velocity and altitude commands while limiting angle-of-attack oscillations, minimized control power usage, and a stabilized vehicle when atmospheric turbulence and system uncertainty are present. The controller designs using H(infinity) and mu-synthesis procedures were compared. An integrated flight/propulsion dynamic mathematical model of a conical accelerator vehicle was linearized as the vehicle accelerated through Mach 8. Vehicle acceleration through the selected flight condition gives rise to parametric variation that was modeled as a structured uncertainty. The mu-analysis approach was used in the frequency domain to conduct controller analysis and was confirmed by time history plots. Results demonstrate the inherent advantages of the mu framework for this class of problems.

  7. Hypersonic vehicle model and control law development using H(infinity) and micron synthesis

    NASA Technical Reports Server (NTRS)

    Gregory, Irene M.; Chowdhry, Rajiv S.; Mcminn, John D.; Shaughnessy, John D.

    1994-01-01

    The control system design for a Single Stage To Orbit (SSTO) air breathing vehicle will be central to a successful mission because a precise ascent trajectory will preserve narrow payload margins. The air breathing propulsion system requires the vehicle to fly roughly halfway around the Earth through atmospheric turbulence. The turbulence, the high sensitivity of the propulsion system to inlet flow conditions, the relatively large uncertainty of the parameters characterizing the vehicle, and continuous acceleration make the problem especially challenging. Adequate stability margins must be provided without sacrificing payload mass since payload margins are critical. Therefore, a multivariable control theory capable of explicitly including both uncertainty and performance is needed. The H(infinity) controller in general provides good robustness but can result in conservative solutions for practical problems involving structured uncertainty. Structured singular value mu framework for analysis and synthesis is potentially much less conservative and hence more appropriate for problems with tight margins. An SSTO control system requires: highly accurate tracking of velocity and altitude commands while limiting angle-of-attack oscillations, minimized control power usage, and a stabilized vehicle when atmospheric turbulence and system uncertainty are present. The controller designs using H(infinity) and mu-synthesis procedures were compared. An integrated flight/propulsion dynamic mathematical model of a conical accelerator vehicle was linearized as the vehicle accelerated through Mach 8. Vehicle acceleration through the selected flight condition gives rise to parametric variation that was modeled as a structured uncertainty. The mu-analysis approach was used in the frequency domain to conduct controller analysis and was confirmed by time history plots. Results demonstrate the inherent advantages of the mu framework for this class of problems.

  8. Solution-Space Screening of a Hypersonic Endurance Demonstrator

    NASA Technical Reports Server (NTRS)

    Chudoba, Bernd; Coleman, Gary; Oza, Amit; Gonzalez, Lex; Czysz, Paul

    2012-01-01

    This report documents a parametric sizing study performed to develop a program strategy for research and development and procurement of a feasible next-generation hypersonic air-breathing endurance demonstrator. Overall project focus has been on complementing technical and managerial decision-making during the earliest conceptual design phase towards minimization of operational, technical, and managerial risks.

  9. Wave combustors for trans-atmospheric vehicles

    NASA Technical Reports Server (NTRS)

    Menees, Gene P.; Adelman, Henry G.; Cambier, Jean-Luc; Bowles, Jeffrey V.

    1989-01-01

    The Wave Combustor is an airbreathing hypersonic propulsion system which utilizes shock and detonation waves to enhance fuel-air mixing and combustion in supersonic flow. In this concept, an oblique shock wave in the combustor can act as a flameholder by increasing the pressure and temperature of the air-fuel mixture and thereby decreasing the ignition delay. If the oblique shock is sufficiently strong, then the combustion front and the shock wave can couple into a detonation wave. In this case, combustion occurs almost instantaneously in a thin zone behind the wave front. The result is a shorter, lighter engine compared to the scramjet. This engine, which is called the Oblique Detonation Wave Engine (ODWE), can then be utilized to provide a smaller, lighter vehicle or to provide a higher payload capability for a given vehicle weight. An analysis of the performance of a conceptual trans-atmospheric vehicle powered by an ODWE is given here.

  10. Non-equilibrium stagnation region aerodynamic heating of hypersonic glide vehicles

    NASA Technical Reports Server (NTRS)

    Rosner, D. E.; Cibrian, R.

    1974-01-01

    A simple method of predicting aerodynamic heating and corresponding radiation equilibrium surface temperature-time histories for critical locations on space shuttle orbiter-type vehicles is presented. The method is based on a generalization of correlation equations developed earlier by Rosner for predicting the energy transfer and radiation equilibrium temperatures of surfaces with arbitrary catalytic activity and total hemispheric emittance. Recently obtained experimental data for O and N atom recombination probabilities on candidate material surfaces above 1000 K are used to assess nonequilibrium effects for a range of nose radii and a specific space shuttle re-entry trajectory. It is concluded that low catalytic activity will be especially important in locations of large effective nose radii by both increasing oxidation-resistant coating lifetime and reducing energy transfer into the vehicle.

  11. Aerodynamics of Hypersonic Lifting Vehicles: Conference Proceedings Held at the Fluid Dynamics Panel Symposium in Bristol, United Kingdom on 6-9 April 1987

    DTIC Science & Technology

    1987-11-01

    AI ILE C * 1IYi AGARD-CP-428 IL to AGARD CONFERENCE PROCEEDINGS No.428 Aerodynamics of Hypersonic Lifting Vehicles DTIC D ~flzrnoNUYATM~r4 A 21 T8...hPPMV~ d to, PubWi vellsa"izizziz Sr o DISTIBUTION AND AVAILABILITY ON BACK COVER 88 1 12 003, AGARI)-CP-4 28 NORTH ATLANTIC TREATY ORGANIZATION ADVISORY...diminution des efforts d ~ployds dans le domaine spatial apres Ic programme Apollo, tant aux Etats Unis quc dans la plupart des autres pays occidentaux

  12. New Hypersonic Shock Tunnel at the Laboratory of Aerothermodynamics and Hypersonics Prof. Henry T. Nagamatsu

    NASA Astrophysics Data System (ADS)

    Toro, P. G. P.; Minucci, M. A. S.; Chanes, J. B.; Oliveira, A. C.; Gomes, F. A. A.; Myrabo, L. N.; Nagamatsu, Henry T.

    2008-04-01

    The new 0.60-m. nozzle exit diameter hypersonic shock tunnel was designed to study advanced air-breathing propulsion system such as supersonic combustion and/or laser technologies. In addition, it may be used for hypersonic flow studies and investigations of the electromagnetic (laser) energy addition for flow control. This new hypersonic shock tunnel was designed and installed at the Laboratory for of Aerothermodynamics and Hypersonics Prof. Henry T. Nagamatsu, IEAv-CTA, Brazil. The design of the tunnel enables relatively long test times, 2-10 milliseconds, suitable for the experiments performed at the laboratory. Free stream Mach numbers ranging from 6 to 25 can be produced and stagnation pressures and temperatures up to 360 atm. and up to 9,000 K, respectively, can be generated. Shadowgraph and schlieren optical techniques will be used for flow visualization.

  13. A comparative study of turbulence models in predicting hypersonic inlet flows

    NASA Technical Reports Server (NTRS)

    Kapoor, Kamlesh

    1993-01-01

    A computational study has been conducted to evaluate the performance of various turbulence models. The NASA P8 inlet, which represents cruise condition of a typical hypersonic air-breathing vehicle, was selected as a test case for the study; the PARC2D code, which solves the full two dimensional Reynolds-averaged Navier-Stokes equations, was used. Results are presented for a total of six versions of zero- and two-equation turbulence models. Zero-equation models tested are the Baldwin-Lomax model, the Thomas model, and a combination of the two. Two-equation models tested are low-Reynolds number models (the Chien model and the Speziale model) and a high-Reynolds number model (the Launder and Spalding model).

  14. Hypersonic and Supersonic Static Aerodynamics of Mars Science Laboratory Entry Vehicle

    NASA Technical Reports Server (NTRS)

    Dyakonov, Artem A.; Schoenenberger, Mark; Vannorman, John W.

    2012-01-01

    This paper describes the analysis of continuum static aerodynamics of Mars Science Laboratory (MSL) entry vehicle (EV). The method is derived from earlier work for Mars Exploration Rover (MER) and Mars Path Finder (MPF) and the appropriate additions are made in the areas where physics are different from what the prior entry systems would encounter. These additions include the considerations for the high angle of attack of MSL EV, ablation of the heatshield during entry, turbulent boundary layer, and other aspects relevant to the flight performance of MSL. Details of the work, the supporting data and conclusions of the investigation are presented.

  15. The NASA Glen Research Center's Hypersonic Tunnel Facility. Chapter 16

    NASA Technical Reports Server (NTRS)

    Woike, Mark R.; Willis, Brian P.

    2001-01-01

    The NASA Glenn Research Center's Hypersonic Tunnel Facility (HTF) is a blow-down, freejet wind tunnel that provides true enthalpy flight conditions for Mach numbers of 5, 6, and 7. The Hypersonic Tunnel Facility is unique due to its large scale and use of non-vitiated (clean air) flow. A 3MW graphite core storage heater is used to heat the test medium of gaseous nitrogen to the high stagnation temperatures required to produce true enthalpy conditions. Gaseous oxygen is mixed into the heated test flow to generate the true air simulation. The freejet test section is 1.07m (42 in.) in diameter and 4.3m (14 ft) in length. The facility is well suited for the testing of large scale airbreathing propulsion systems. In this chapter, a brief history and detailed description of the facility are presented along with a discussion of the facility's application towards hypersonic airbreathing propulsion testing.

  16. Configuration trade and code validation study on a conical hypersonic vehicle

    NASA Technical Reports Server (NTRS)

    Covell, Peter F.; Wood, Richard M.; Bauer, Steven X. S.; Walker, Ira J.

    1988-01-01

    A study has been conducted on a generic wing-cone transatmospheric vehicle at Mach numbers form 2.5 to 4.5. The objectives of the study were to experimentally define the aerodynamic characteristics of the vehicle and evaluate several computational aerodynamic prediction methods through comparison with the experimental results. The baseline wing-cone configuration fuselage consisted of a 5 deg half-angle cone forebody, cylindrical midbody, and 9 deg truncated cone afterbody. The 4-percent-thick diamond airfoil wing had an aspect ratio of 1. Several configuration variables were investigated to provide trade information on canard, wing-position and incidence, vertical tail, and nose bluntness effects. Results of the study show that wing-position and wing-incidence effects on the longitudinal aerodynamic characteristics can be significantly influenced by wing-body interference. The use of positive wing incidence to provide favorable forebody orientation for possible inlet performance improvement is accompanied by trim drag and lift-drag ratio penalties. The lateral-directional stability characteristics were strongly influenced by the location of the vertical tails. The higher-order full-potential method provided better estimates of the aerodynamic characteristics than either the linearized supersonic potential method or the tangent-cone/tangent-wedge/shock-expansion on method.

  17. Current Grid Generation Strategies and Future Requirements in Hypersonic Vehicle Design, Analysis and Testing

    NASA Technical Reports Server (NTRS)

    Papadopoulos, Periklis; Venkatapathy, Ethiraj; Prabhu, Dinesh; Loomis, Mark P.; Olynick, Dave; Arnold, James O. (Technical Monitor)

    1998-01-01

    Recent advances in computational power enable computational fluid dynamic modeling of increasingly complex configurations. A review of grid generation methodologies implemented in support of the computational work performed for the X-38 and X-33 are presented. In strategizing topological constructs and blocking structures factors considered are the geometric configuration, optimal grid size, numerical algorithms, accuracy requirements, physics of the problem at hand, computational expense, and the available computer hardware. Also addressed are grid refinement strategies, the effects of wall spacing, and convergence. The significance of grid is demonstrated through a comparison of computational and experimental results of the aeroheating environment experienced by the X-38 vehicle. Special topics on grid generation strategies are also addressed to model control surface deflections, and material mapping.

  18. Statistical Analysis of the Uncertainty in Pre-Flight Aerodynamic Database of a Hypersonic Vehicle

    NASA Astrophysics Data System (ADS)

    Huh, Lynn

    The objective of the present research was to develop a new method to derive the aerodynamic coefficients and the associated uncertainties for flight vehicles via post- flight inertial navigation analysis using data from the inertial measurement unit. Statistical estimates of vehicle state and aerodynamic coefficients are derived using Monte Carlo simulation. Trajectory reconstruction using the inertial navigation system (INS) is a simple and well used method. However, deriving realistic uncertainties in the reconstructed state and any associated parameters is not so straight forward. Extended Kalman filters, batch minimum variance estimation and other approaches have been used. However, these methods generally depend on assumed physical models, assumed statistical distributions (usually Gaussian) or have convergence issues for non-linear problems. The approach here assumes no physical models, is applicable to any statistical distribution, and does not have any convergence issues. The new approach obtains the statistics directly from a sufficient number of Monte Carlo samples using only the generally well known gyro and accelerometer specifications and could be applied to the systems of non-linear form and non-Gaussian distribution. When redundant data are available, the set of Monte Carlo simulations are constrained to satisfy the redundant data within the uncertainties specified for the additional data. The proposed method was applied to validate the uncertainty in the pre-flight aerodynamic database of the X-43A Hyper-X research vehicle. In addition to gyro and acceleration data, the actual flight data include redundant measurements of position and velocity from the global positioning system (GPS). The criteria derived from the blend of the GPS and INS accuracy was used to select valid trajectories for statistical analysis. The aerodynamic coefficients were derived from the selected trajectories by either direct extraction method based on the equations in

  19. The effects of shock wave precursors ahead of hypersonic entry vehicles

    NASA Technical Reports Server (NTRS)

    Stanley, Scott A.; Carlson, Leland A.

    1991-01-01

    A model has been developed to predict the magnitude and characteristics of the shock wave precursor ahead of a hypervelocity vehicle. This model includes both chemical and thermal nonequilibrium, utilizes detailed mass production rates for the photodissociation and photoionization reactions, and accounts for the effects of radiative absorption and emission on the individual internal energy modes of both atomic and diatomic species. Comparison of the present results with shock tube data indicates that the model is reasonably accurate. A series of test cases representing earth aerocapture return from Mars indicate that there is significant production of atoms, ions and electrons ahead of the shock front due to radiative absorption and that the precursor is characterized by an enhanced electron/electronic temperature and molecular ionization. However, the precursor has a negligible effect on the shock layer flow field.

  20. Wireless Subsurface Microsensors for Health Monitoring of Thermal Protection Systems on Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Milos, Frank S.; Watters, David G.; Pallix, Joan B.; Bahr, Alfred J.; Huestis, David L.; Arnold, Jim (Technical Monitor)

    2001-01-01

    Health diagnostics is an area where major improvements have been identified for potential implementation into the design of new reusable launch vehicles in order to reduce life cycle costs, to increase safety margins, and to improve mission reliability. NASA Ames is leading the effort to develop inspection and health management technologies for thermal protection systems. This paper summarizes a joint project between NASA Ames and SRI International to develop 'SensorTags,' radio frequency identification devices coupled with event-recording sensors, that can be embedded in the thermal protection system to monitor temperature or other quantities of interest. Two prototype SensorTag designs containing thermal fuses to indicate a temperature overlimit are presented and discussed.

  1. X-43A Vehicle During Ground Testing

    NASA Technical Reports Server (NTRS)

    1999-01-01

    The X-43A Hypersonic Experimental Vehicle, or 'Hyper-X' is seen here undergoing ground testing at NASA's Dryden Flight Research Center, Edwards, California in December 1999. The X-43A was developed to research a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only

  2. X-43A Vehicle During Ground Testing

    NASA Technical Reports Server (NTRS)

    1999-01-01

    This photo shows a close-up, rear view of the X-43A Hypersonic Experimental Vehicle, or 'Hyper-X' undergoing ground testing at NASA's Dryden Flight Research Center, Edwards, California in December 1999. The X-43A was developed to research a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology

  3. X-43A Vehicle During Ground Testing

    NASA Technical Reports Server (NTRS)

    1999-01-01

    The X-43A Hypersonic Experimental Vehicle, or 'Hyper-X' is seen here undergoing ground testing at NASA's Dryden Flight Research Center, Edwards, California. The X-43A was developed to research a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By

  4. Hyper-X Vehicle Model - Front View

    NASA Technical Reports Server (NTRS)

    1996-01-01

    A front view of an early desk-top model of NASA's X-43A 'Hyper-X,' or Hypersonic Experimental Vehicle, which has been developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic

  5. HYPERDATA - BASIC HYPERSONIC DATA AND EQUATIONS

    NASA Technical Reports Server (NTRS)

    Mackall, D.

    1994-01-01

    In an effort to place payloads into orbit at the lowest possible costs, the use of air-breathing space-planes, which reduces the need to carry the propulsion system oxidizer, has been examined. As this approach would require the space-plane to fly at hypersonic speeds for periods of time much greater than that required by rockets, many factors must be considered when analyzing its benefits. The Basic Hypersonic Data and Equations spreadsheet provides data gained from three analyses of a space-plane's performance. The equations used to perform the analyses are derived from Newton's second law of physics (i.e. force equals mass times acceleration); the derivation is included. The first analysis is a parametric study of some basic factors affecting the ability of a space-plane to reach orbit. This step calculates the fraction of fuel mass to the total mass of the space-plane at takeoff. The user is able to vary the altitude, the heating value of the fuel, the orbital gravity, and orbital velocity. The second analysis calculates the thickness of a spherical fuel tank, while assuming all of the mass of the vehicle went into the tank's shell. This provides a first order analysis of how much material results from a design where the fuel represents a large portion of the total vehicle mass. In this step, the user is allowed to vary the values for gross weight, material density, and fuel density. The third analysis produces a ratio of gallons of fuel per total mass for various aircraft. It shows that the volume of fuel required by the space-plane relative to the total mass is much larger for a liquid hydrogen space-plane than any other vehicle made. This program is a spreadsheet for use on Macintosh series computers running Microsoft Excel 3.0. The standard distribution medium for this package is a 3.5 inch 800K Macintosh format diskette. Documentation is included in the price of the program. Macintosh is a registered trademark of Apple Computer, Inc. Microsoft is a

  6. Combining MHD Airbreathing and Fusion Rocket Propulsion for Earth-to-Orbit Flight

    SciTech Connect

    Froning, H. D. Jr; Yang, Yang; Momota, H.; Burton, E.; Miley, G. H.; Luo, Nie

    2005-02-06

    Previous studies have shown that Single-State-to-Orbit (SSTO) vehicle propellant can be reduced by Magnets-Hydro-Dynamic (MHD) processes that minimize airbreathing propulsion losses and propellant consumption during atmospheric flight. Similarly additional reduction in SSTO propellant is enabled by Inertial Electrostatic Confinement (IEC) fusion, whose more energetic reactions reduce rocket propellant needs. MHD airbreathing propulsion during an SSTO vehicle's initial atmospheric flight phase and IEC fusion propulsion during its final exo-atmospheric flight phase is therefore being explored. Accomplished work is not yet sufficient for claiming such a vehicle's feasibility. But takeoff and propellant mass for an MHD airbreathing and IEC fusion vehicle could be as much as 25 and 40 percent less than one with ordinary airbreathing and IEC fusion; and as much as 50 and 70 percent less than SSTO takeoff and propellant mass with MHD airbreathing and chemical rocket propulsion. Thus this unusual combined cycle engine shows great promise for performance gains beyond contemporary combined-cycle airbreathing engines.

  7. Wireless Subsurface Sensors for Health Monitoring of Thermal Protection Systems on Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Milos, Frank S.; Arnold, Jim (Technical Monitor)

    2001-01-01

    Health diagnostics is an area where major improvements have been identified for potential implementation into the design of new reusable launch vehicles (RLVs) in order to reduce life cycle costs, to increase safety margins, and to improve mission reliability. NASA Ames is leading the effort to develop inspection and health management technologies for thermal protection systems. This paper summarizes a joint project between NASA Ames and industry partners to develop "wireless" devices that can be embedded in the thermal protection system to monitor temperature or other quantities of interest. These devices are sensors integrated with radio-frequency identification (RFID) microchips to enable non-contact communication of sensor data to an external reader that may be a hand-held scanner or a large portal. Both passive and active prototype devices have been developed. The passive device uses a thermal fuse to indicate the occurrence of excessive temperature. This device has a diameter under 0.13 cm. (suitable for placement in gaps between ceramic TPS tiles on an RLV) and can withstand 370 C for 15 minutes. The active device contains a small battery to provide power to a thermocouple for recording a temperature history during flight. The bulk of the device must be placed beneath the TPS for protection from high temperature, but the thermocouple can be placed in a hot location such as near the external surface.

  8. Health Monitoring Technology for Thermal Protection Systems on Reusable Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Milos, Frank S.; Watters, D. G.; Heinemann, J. M.; Karunaratne, K. S.; Arnold, Jim (Technical Monitor)

    2001-01-01

    Integrated subsystem health diagnostics is an area where major improvements have been identified for potential implementation into the design of new reusable launch vehicles (RLVs) in order to reduce life cycle costs, to increase safety margins, and to improve mission reliability. This talk summarizes a joint effort between NASA Ames and industry partners to develop rapid non-contact diagnostic tools for health and performance monitoring of thermal protection systems (TPS) on future RLVs. The specific goals for TPS health monitoring are to increase the speed and reliability of TPS inspections for improved operability at lower cost. The technology being developed includes a 3-D laser scanner for examining the exterior surface of the TPS, and a subsurface microsensor suite for monitoring the health and performance of the TPS. The sensor suite consists of passive overlimit sensors and sensors for continuous parameter monitoring in flight. The sensors are integrated with radio-frequency identification (RFID) microchips to enable wireless communication of-the sensor data to an external reader that may be a hand-held scanner or a large portal. Prototypes of the laser system and both types of subsurface sensors have been developed. The laser scanner was tested on Shuttle Orbiter Columbia and was able to dimension surface chips and holes on a variety of TPS materials. The temperature-overlimit microsensor has a diameter under 0.05 inch (suitable for placement in gaps between ceramic TPS tiles) and can withstand 700 F for 15 minutes.

  9. Experimental Evaluation of the Effect of Angle-of-attack on the External Aerodynamics and Mass Capture of a Symmetric Three-engine Air-breathing Launch Vehicle Configuration at Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Kim, Hyun D.; Frate, Franco C.

    2001-01-01

    A subscale aerodynamic model of the GTX air-breathing launch vehicle was tested at NASA Glenn Research Center's 10- by 10-Foot Supersonic Wind Tunnel from Mach 2.0 to 3.5 at various angles-of-attack. The objective of the test was to investigate the effect of angle-of-attack on inlet mass capture, inlet diverter effectiveness, and the flowfield at the cowl lip plane. The flow-through inlets were tested with and without boundary-layer diverters. Quantitative measurements such as inlet mass flow rates and pitot-pressure distributions in the cowl lip plane are presented. At a 3deg angle-of-attack, the flow rates for the top and side inlets were within 8 percent of the zero angle-of-attack value, and little distortion was evident at the cowl lip plane. Surface oil flow patterns showing the shock/boundary-layer interaction caused by the inlet spikes are shown. In addition to inlet data, vehicle forebody static pressure distributions, boundary-layer profiles, and temperature-sensitive paint images to evaluate the boundary-layer transition are presented. Three-dimensional parabolized Navier-Stokes computational fluid dynamics calculations of the forebody flowfield are presented and show good agreement with the experimental static pressure distributions and boundary-layer profiles. With the boundary-layer diverters installed, no adverse aerodynamic phenomena were found that would prevent the inlets from operating at the required angles-of-attack. We recommend that phase 2 of the test program be initiated, where inlet contraction ratio and diverter geometry variations will be tested.

  10. Thermomechanical response of metal foam sandwich panels for structural thermal protection systems in hypersonic vehicles

    NASA Astrophysics Data System (ADS)

    Rakow, Joseph F.

    Sandwich panels with metal foam cores are proposed for load-bearing structural components in actively cooled thermal protection systems for aerospace vehicles. Prototype acreage metal foam sandwich panels (MFSP's) are constructed and analyzed with the central goal of characterizing the thermomechanical response of the system. MFSP's are subjected to uniform temperature fields and equibiaxial loading in a novel experimental load frame. The load frame exploits the mismatch of coefficients of thermal expansion and allows for thermostructural experimentation without the endemic conflict of thermal and mechanical boundary conditions. Back-to-back strain gages and distributed thermocouples capture the in-plane response of the panels, including buckling and elastic-plastic post-buckling. The out-of-plane response is captured via moire interferometry, which provides a visualization of evolving mode shapes throughout the post-buckling regime. The experimental results agree with an analytical prediction for critical temperatures in sandwich panels based on a Rayleigh-Ritz minimization of the energy functional for a Reissner-Mindlin plate. In addition, a three-dimensional finite element model of the non-linear thermomechanical response of the panel-frame experimental system is developed and the results are shown to agree well with the experimentally identified response of MFSP's. Central to analytical and numerical characterization of MFSP's is an understanding of the response of metal foam under shear loading. The shear response of metal foam is captured experimentally, providing density-dependent relationships for material stiffness, strength, and energy absorption. Speckle photography is employed to identify microstructural size effects in the distribution of strain throughout metal foam under shear loading. In addition, a micromechanical model is established for the density-dependent shear modulus of metal foam, which allows for the coupling of cell-level imperfections

  11. International Conference on Hypersonic Flight in the 21st Century, 1st, University of North Dakota, Grand Forks, Sept. 20-23, 1988, Proceedings

    SciTech Connect

    Higbea, M.E.; Vedda, J.A.

    1988-01-01

    The present conference on the development status of configurational concepts and component technologies for hypersonic-cruise and transatmospheric vehicles discusses topics relating to the U.S. National Aerospace Plane program, ESA-planned aerospace vehicles, Japanese spaceplane concepts, the integration of hypersonic aircraft into existing infrastructures, hypersonic airframe designs, hypersonic avionics and cockpit AI systems, hypersonic-regime CFD techniques, the economics of hypersonic vehicles, and possible legal implications of hypersonic flight. Also discussed are Soviet spaceplane concepts, propulsion systems involving laser power sources and hypervelocity launch technologies, and the management of support systems operations for hypersonic vehicles.

  12. Innovative Airbreathing Propulsion Concepts for High-speed Applications

    NASA Technical Reports Server (NTRS)

    Whitlow, Woodrow, Jr.

    2002-01-01

    Research and Technology Program mature, they are incorporated into ASTP. One example of this is rocket-based combined cycle systems that are being considered as part of ASTP. The NASA Ultra Efficient Engine Technology (UEET) Program has the goal of developing propulsion system component technology that is relevant to a wide range of vehicle missions. In addition to subsonic and supersonic speed regimes, it includes the hypersonic speed regime. More specifically, component technologies for turbine-based combined cycle engines are being developed as part of UEET.

  13. Modeling, Analysis, and Control of a Hypersonic Vehicle with Significant Aero-Thermo-Elastic-Propulsion Interactions: Elastic, Thermal and Mass Uncertainty

    NASA Astrophysics Data System (ADS)

    Khatri, Jaidev

    This thesis examines themodeling, analysis, and control system design issues for scramjet powered hypersonic vehicles. A nonlinear three degrees of freedom longitudinal model which includes aero-propulsion-elasticity effects was used for all analyses. This model is based upon classical compressible flow and Euler-Bernouli structural concepts. Higher fidelity computational fluid dynamics and finite element methods are needed for more precise intermediate and final evaluations. The methods presented within this thesis were shown to be useful for guiding initial control relevant design. The model was used to examine the vehicle's static and dynamic characteristics over the vehicle's trimmable region. The vehicle has significant longitudinal coupling between the fuel equivalency ratio (FER) and the flight path angle (FPA). For control system design, a two-input two-output plant (FER - elevator to speed-FPA) with 11 states (including 3 flexible modes) was used. Velocity, FPA, and pitch were assumed to be available for feedback. Aerodynamic heat modeling and design for the assumed TPS was incorporated to original Bolender's model to study the change in static and dynamic properties. De-centralized control stability, feasibility and limitations issues were dealt with the change in TPS elasticity, mass and physical dimension. The impact of elasticity due to TPS mass, TPS physical dimension as well as prolonged heating was also analyzed to understand performance limitations of de-centralized control designed for nominal model.

  14. Advanced Guidance and Control for Hypersonics and Space Access

    NASA Technical Reports Server (NTRS)

    Hanson, John M.; Hall, Charles E.; Mulqueen, John A.; Jones, Robert E.

    2003-01-01

    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), whether it is fully rocket-powered or has air- breathing components. This becomes clear upon examining the number of expendable launch vehicle failures in the recent past where AG&C technologies could have saved a RLV with the same failure mode, the additional vehicle problems where t h i s technology applies, and the costs and time 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 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 AWC, current technology efforts, and the additional work needed for making this goal a reality. There are a number of approaches to AG&C that have the potential for achieving the desired goals. For some of these methods, we compare the results of tests designed to demonstrate the achievement of the goals. Tests up to now have been focused on rocket-powered vehicles; application to hypersonic air-breathers is planned. We list the test cases used to demonstrate that the desired results are achieved, briefly describe an automated test scoring method, and display results of the tests. Some of the technology components have reached the maturity level where they are ready for application to a new vehicle concept, while others are not far along in development.

  15. Hypersonic research engine/aerothermodynamic integration model, experimental results. Volume 1: Mach 6 component integration

    NASA Technical Reports Server (NTRS)

    Andrews, E. H., Jr.; Mackley, E. A.

    1976-01-01

    The NASA Hypersonic Research Engine (HRE) Project was initiated for the purpose of advancing the technology of airbreathing propulsion for hypersonic flight. A large component (inlet, combustor, and nozzle) and structures development program was encompassed by the project. The tests of a full-scale (18 in. diameter cowl and 87 in. long) HRE concept, designated the Aerothermodynamic Integration Model (AIM), at Mach numbers of 5, 6, and 7. Computer program results for Mach 6 component integration tests are presented.

  16. Prediction of forces and moments for flight vehicle control effectors. Part 1: Validation of methods for predicting hypersonic vehicle controls forces and moments

    NASA Technical Reports Server (NTRS)

    Maughmer, Mark D.; Ozoroski, L.; Ozoroski, T.; Straussfogel, D.

    1990-01-01

    Many types of hypersonic aircraft configurations are currently being studied for feasibility of future development. Since the control of the hypersonic configurations throughout the speed range has a major impact on acceptable designs, it must be considered in the conceptual design stage. The ability of the aerodynamic analysis methods contained in an industry standard conceptual design system, APAS II, to estimate the forces and moments generated through control surface deflections from low subsonic to high hypersonic speeds is considered. Predicted control forces and moments generated by various control effectors are compared with previously published wind tunnel and flight test data for three configurations: the North American X-15, the Space Shuttle Orbiter, and a hypersonic research airplane concept. Qualitative summaries of the results are given for each longitudinal force and moment and each control derivative in the various speed ranges. Results show that all predictions of longitudinal stability and control derivatives are acceptable for use at the conceptual design stage. Results for most lateral/directional control derivatives are acceptable for conceptual design purposes; however, predictions at supersonic Mach numbers for the change in yawing moment due to aileron deflection and the change in rolling moment due to rudder deflection are found to be unacceptable. Including shielding effects in the analysis is shown to have little effect on lift and pitching moment predictions while improving drag predictions.

  17. Hypersonic aeroheating test of space shuttle vehicle: Configuration 3 (model 22 OTS) in the NASA-Ames 3.5-foot hypersonic wind tunnel (IH20), volume 2

    NASA Technical Reports Server (NTRS)

    Kingsland, R. B.; Lockman, W. K.

    1975-01-01

    The model tested was an 0.0175-scale version of the vehicle 3 space shuttle configuration. Temperature measurements were made on the launch configuration, orbiter plus tank, orbiter alone, tank alone, and solid rocket booster (SRB) alone to provide heat transfer data. The test was conducted at free stream Mach numbers of 5.3 and 7.3 and at free stream Reynolds numbers of 1.5, 3.7, 5.0, and 7.0 million per foot. The model was tested at angles of attack from -5 deg to 20 deg and side slip angles of -5 deg and 0 deg.

  18. Hyper-X Vehicle Model - Top Rear View

    NASA Technical Reports Server (NTRS)

    1996-01-01

    This aft-quarter model view of NASA's X-43A 'Hyper-X' or Hypersonic Experimental Vehicle shows its sleek, geometric design. The X-43A was developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen

  19. Wave combustors for trans-atmospheric vehicles

    NASA Technical Reports Server (NTRS)

    Menees, Gene P.; Bowles, Jeffrey V.; Adelman, Henry G.; Cambier, Jean-Luc

    1989-01-01

    A performance analysis is given of a conceptual transatmospheric vehicle (TAV). The TAV is powered by a an oblique detonation wave engine (ODWE). The ODWE is an airbreathing hypersonic propulsion system which utilizes shock and detonation waves to enhance fuel-air mixing and combustion in supersonic flow. In this wave combustor concept, an oblique shock wave in the combustor can act as a flameholder by increasing the pressure and temperature of the air-fuel mixture, thereby decreasing the ignition delay. If the oblique shock is sufficiently strong, then the combustion front and the shock wave can couple into a detonation wave. In this case, combustion occurs almost instantaneously in a thin zone behind the wave front. The result is a shorter lighter engine compared to the scramjet. The ODWE-powered hypersonic vehicle performance is compared to that of a scramjet-powered vehicle. Among the results outlined, it is found that the ODWE trades a better engine performance above Mach 15 for a lower performance below Mach 15. The overall higher performance of the ODWE results in a 51,000-lb weight savings and a higher payload weight fraction of approximately 12 percent.

  20. Dynamic Testing of the NASA Hypersonic Project Combined Cycle Engine Testbed for Mode Transition Experiments

    NASA Technical Reports Server (NTRS)

    2011-01-01

    NASA is interested in developing technology that leads to more routine, safe, and affordable access to space. Access to space using airbreathing propulsion systems has potential to meet these objectives based on Airbreathing Access to Space (AAS) system studies. To this end, the NASA Fundamental Aeronautics Program (FAP) Hypersonic Project is conducting fundamental research on a Turbine Based Combined Cycle (TBCC) propulsion system. The TBCC being studied considers a dual flow-path inlet system. One flow-path includes variable geometry to regulate airflow to a turbine engine cycle. The turbine cycle provides propulsion from take-off to supersonic flight. The second flow-path supports a dual-mode scramjet (DMSJ) cycle which would be initiated at supersonic speed to further accelerate the vehicle to hypersonic speed. For a TBCC propulsion system to accelerate a vehicle from supersonic to hypersonic speed, a critical enabling technology is the ability to safely and effectively transition from the turbine to the DMSJ-referred to as mode transition. To experimentally test methods of mode transition, a Combined Cycle Engine (CCE) Large-scale Inlet testbed was designed with two flow paths-a low speed flow-path sized for a turbine cycle and a high speed flow-path designed for a DMSJ. This testbed system is identified as the CCE Large-Scale Inlet for Mode Transition studies (CCE-LIMX). The test plan for the CCE-LIMX in the NASA Glenn Research Center (GRC) 10- by 10-ft Supersonic Wind Tunnel (10x10 SWT) is segmented into multiple phases. The first phase is a matrix of inlet characterization (IC) tests to evaluate the inlet performance and establish the mode transition schedule. The second phase is a matrix of dynamic system identification (SysID) experiments designed to support closed-loop control development at mode transition schedule operating points for the CCE-LIMX. The third phase includes a direct demonstration of controlled mode transition using a closed loop control

  1. Studies in hypersonic aeroelasticity

    NASA Astrophysics Data System (ADS)

    Nydick, Ira Harvey

    2000-11-01

    This dissertation describes the aeroelastic analysis of a generic hypersonic vehicle, focusing on two specific problems: (1) hypersonic panel flutter, and (2) aeroelastic behavior of a complete unrestrained generic hypersonic vehicle operating at very high Mach numbers. The panels are modeled as shallow shells using Marguerre nonlinear shallow shell theory for orthotropic panels and the aerodynamic loads are obtained from third order piston theory. Two models of curvature, several applied temperature distributions, and the presence of a shock are also included in the model. Results indicate that the flutter speed of the panel is significantly reduced by temperature variations comparable to the buckling temperature and by the presence of a shock. A panel with initial curvature can be more stable than the flat panel but the increase in stability depends in a complex way on the material properties of the panel and the amount of curvature. At values of dynamic pressure above critical, aperiodic motion was observed. The value of dynamic pressure for which this occurs in both heated panels and curved panels is much closer to the critical dynamic pressure than for the flat, unheated panel. A comparison of piston theory aerodynamics and Euler and Navier-Stokes aerodynamics was performed for a two dimensional panel with prescribed motion and the results indicate that while 2nd or higher order piston theory agrees very well with the Euler solution for the frequencies seen in hypersonic panel flutter, it differs substantially from the Navier-Stokes solution. The aeroelastic behavior of the complete vehicle was simulated using the unrestrained equations of motion, utilizing the method of quasi-coordinates. The unrestrained mode shapes of the vehicle were obtained from an equivalent plate analysis using an available code (ELAPS). The effects of flexible trim and rigid body degrees of freedom are carefully incorporated in the mathematical model. This model was applied to a

  2. Viscous-shock-layer analysis of hypersonic flows over long slender vehicles. Ph.D. Thesis, 1988

    NASA Technical Reports Server (NTRS)

    Lee, Kam-Pui; Gupta, Roop N.

    1992-01-01

    An efficient and accurate method for solving the viscous shock layer equations for hypersonic flows over long slender bodies is presented. The two first order equations, continuity and normal momentum, are solved simultaneously as a coupled set. The flow conditions included are from high Reynolds numbers at low altitudes to low Reynolds numbers at high altitudes. For high Reynolds number flows, both chemical nonequilibrium and perfect gas cases are analyzed with surface catalytic effects and different turbulence models, respectively. At low Reynolds number flow conditions, corrected slip models are implemented with perfect gas case. Detailed comparisons are included with other predictions and experimental data.

  3. Structures and materials technology issues for reusable launch vehicles

    NASA Technical Reports Server (NTRS)

    Dixon, S. C.; Tenney, D. R.; Rummler, D. R.; Wieting, A. R.; Bader, R. M.

    1985-01-01

    Projected space missions for both civil and defense needs require significant improvements in structures and materials technology for reusable launch vehicles: reductions in structural weight compared to the Space Shuttle Orbiter of up to 25 percent or more, a possible factor of 5 or more increase in mission life, increases in maximum use temperature of the external surface, reusable containment of cryogenic hydrogen and oxygen, significant reductions in operational costs, and possibly less lead time between technology readiness and initial operational capability. In addition, there is increasing interest in hypersonic airbreathing propulsion for launch and transmospheric vehicles, and such systems require regeneratively cooled structure. The technology issues are addressed, giving brief assessments of the state-of-the-art and proposed activities to meet the technology requirements in a timely manner.

  4. Structures and materials technology issues for reusable launch vehicles

    NASA Technical Reports Server (NTRS)

    Dixon, S. C.; Tenney, D. R.; Rummler, D. R.; Wieting, A. R.; Bader, R. M.

    1985-01-01

    Projected space missions for both civil and defense needs require significant improvements in structures and materials technology for reusable launch vehicles: reductions in structural weight compared to the Space Shuttle Orbiter of up to 25% or more, a possible factor of 5 or more increase in mission life, increases in maximum use temperature of the external surface, reusable containment of cryogenic hydrogen and oxygen, significant reductions in operational costs, and possibly less lead time between technology readiness and initial operational capability. In addition, there is increasing interest in hypersonic airbreathing propulsion for launch and transmospheric vehicles, and such systems require regeneratively cooled structure. The technology issues are addressed, giving brief assessments of the state-of-the-art and proposed activities to meet the technology requirements in a timely manner.

  5. NASA Hypersonic Propulsion: Overview of Progress from 1995 to 2005

    NASA Technical Reports Server (NTRS)

    Cikanek, Harry A., III; Bartolotta, Paul A.; Klem, Mark D.; Rausch, Vince L.

    2007-01-01

    Hypersonic propulsion work supported by the United States National Aeronautics and Space Administration had a primary focus on Space Transportation during the period from 1995 to 2005. The framework for these advances was established by policy and pursued with substantial funding. Many noteworthy advances were made, highlighted by the pinnacle flights of the X-43. This paper reviews and summarizes the programs and accomplishments of this era. The accomplishments are compared to the goals and objectives to lend an overarching perspective to what was achieved. At least dating back to the early days of the Space Shuttle program, NASA has had the objective of reducing the cost of access to space and concurrently improving safety and reliability. National Space Transportation Policy in 1994 coupled with a base of prior programs such as the National Aerospace Plane and the need to look beyond the Space Shuttle program set the stage for NASA to pursue Space Transportation Advances. Programs defined to pursue the advances represented a broad approach addressing classical rocket propulsion as well as airbreathing propulsion in various combinations and forms. The resulting portfolio of activities included systems analysis and design studies, discipline research and technology, component technology development, propulsion system ground test demonstration and flight demonstration. The types of propulsion systems that were pursued by these programs included classical rocket engines, "aerospike" rocket engines, high performance rocket engines, scram jets, rocket based combined cycles, and turbine based combined cycles. Vehicle architectures included single and two stage vehicles. Either single types of propulsion systems or combinations of the basic propulsion types were applied to both single and two stage vehicle design concepts. Some of the propulsion system design concepts were built and tested at full scale, large scale and small scale. Many flight demonstrators were

  6. X-33 Hypersonic Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.

    1999-01-01

    Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will design, build, and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604BOO02G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate the aerodynamic flight database for the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. Al these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.

  7. X-33 Hypersonic Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.

    1999-01-01

    Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database i n the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.

  8. X-33 Hypersonic Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.

    1999-01-01

    Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database in the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.

  9. X-33 Hypersonic Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.

    1999-01-01

    Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database in the hypersonic regime, The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.

  10. Integrated System Test of an Airbreathing Rocket

    NASA Technical Reports Server (NTRS)

    Mack, Gregory; Beaudry, Charles; Ketchum, Andrew; McArthur, J. Craig (Technical Monitor)

    2002-01-01

    This viewgraph presentation provides information on NASA's attempts to develop an air-breathing propulsion in an effort to make future space transportation safer, more reliable and significantly less expensive than today's missions. Spacecraft powered by air-breathing rocket engines would be completely reusable, able to take off and land at airport runways and ready to fly again within days. A radical new engine project is called the Integrated System Tests of an Air-breathing Rocket, or ISTAR.

  11. A methodology for hypersonic transport technology planning

    NASA Technical Reports Server (NTRS)

    Repic, E. M.; Olson, G. A.; Milliken, R. J.

    1973-01-01

    A systematic procedure by which the relative economic value of technology factors affecting design, configuration, and operation of a hypersonic cruise transport can be evaluated is discussed. Use of the methodology results in identification of first-order economic gains potentially achievable by projected advances in each of the definable, hypersonic technologies. Starting with a baseline vehicle, the formulas, procedures and forms which are integral parts of this methodology are developed. A demonstration of the methodology is presented for one specific hypersonic vehicle system.

  12. Hypersonic Flight Test Windows for Technology Development Testing

    DTIC Science & Technology

    2013-11-01

    hypersonic vehicles requires the application of a significant amount of thermal protection or use of a hot structures concept, which can be a major cost...AFRL-RQ-WP-TM-2013-0260 HYPERSONIC FLIGHT TEST WINDOWS FOR TECHNOLOGY DEVELOPMENT TESTING Barry M. Hellman Vehicle Technology Branch...DATES COVERED (From - To) November 2013 Final 01 November 2013 – 25 November 2013 4. TITLE AND SUBTITLE HYPERSONIC FLIGHT TEST WINDOWS FOR

  13. Results of heat transfer tests of an 0.0175-scale space shuttle vehicle model 22 OTS in the NASA-Ames 3.5 foot hypersonic wind tunnel (IH3), volume 1

    NASA Technical Reports Server (NTRS)

    Foster, T. F.; Lockman, W. K.

    1975-01-01

    Heat transfer data for the 0.0175-scale space shuttle vehicle 3 are presented. Interference heating effects were investigated by a model build-up technique of orbiter alone, tank alone, second, and first stage configurations. The test program was conducted in the NASA-Ames 3.5-foot hypersonic wind tunnel at Mach 5.3 for nominal free stream Reynolds number per foot values of 1.5, and 5.0 million.

  14. Visualization of Hypersonic Flat-Plate Boundary Layer in Shock Tunnel

    NASA Astrophysics Data System (ADS)

    Zhang, Qinghu; Yi, Shihe; Zhi, Chen; Zhu, Yangzh; Yu, Wu

    In order to design the future aerospace vehicles, it is essential to experimentally investigate the hypersonic boundary layer [1]. Many aspects of hypersonic turbulent boundary layer and transition process are poorly understood.

  15. X-43A hypersonic research aircraft mated to its modified Pegasus booster rocket.

    NASA Technical Reports Server (NTRS)

    2001-01-01

    The first of three X-43A hypersonic research aircraft was mated to its modified Pegasus booster rocket in late January at NASA's Dryden Flight Research Center, Edwards, Calif. FIRST X-43A MATED TO BOOSTER -- The first of three X-43A hypersonic research aircraft was mated to its modified Pegasus booster rocket in late January at NASA's Dryden Flight Research Center, Edwards, Calif. Mating of the X-43A and its specially-designed adapter to the first stage of the booster rocket marks a major milestone in the Hyper-X hypersonic research program. The 12-foot, unpiloted research vehicle was developed and built by MicroCraft Inc., Tullahoma, Tenn., for NASA. The booster, built by Orbital Sciences Corp., Dulles, Va., will accelerate the X-43A after the X-43A booster 'stack' is air-launched from NASA's venerable NB-52 mothership. The X-43A will separate from the rocket at a predetermined altitude and speed and fly a pre-programmed trajectory, conducting aerodynamic and propulsion experiments until it impacts into the Pacific Ocean. Three research flights are planned, two at Mach 7 and one at Mach 10 (seven and 10 times the speed of sound respectively) with the first tentatively scheduled for early summer of 2001. The X-43A is powered by a revolutionary supersonic-combustion ramjet ('scramjet') engine, and will use the underbody of the aircraft to form critical elements of the engine. The forebody shape helps compress the intake airflow, while the aft section acts as a nozzle to direct thrust. The X-43A flights will be the first actual flight tests of an aircraft powered by an air-breathing scramjet engine.

  16. Multi-Scale Modeling of Hypersonic Gas Flow

    NASA Astrophysics Data System (ADS)

    Boyd, Iain D.

    On March 27, 2004, NASA successfully flew the X-43A hypersonic test flight vehicle at a velocity of 5000 mph to break the aeronautics speed record that had stood for over 35 years. The final flight of the X-43A on November 16, 2004 further increased the speed record to 6,600 mph which is almost ten times the speed of sound. The very high speed attainable by hypersonic airplanes could revolutionize air travel by dramatically reducing inter-continental flight times. For example, a hypersonic flight from New York to Sydney, Australia, a distance of 10,000 miles, would take less than 2 h. Reusable hypersonic vehicles are also being researched to significantly reduce the cost of access to space. Computer modeling of the gas flows around hypersonic vehicles will play a critical part in their development. This article discusses the conditions that can prevail in certain hypersonic gas flows that require a multi-scale modeling approach.

  17. Computational Fluid Dynamics (CFD) Image of Hyper-X Research Vehicle at Mach 7 with Engine Operating

    NASA Technical Reports Server (NTRS)

    1997-01-01

    This computational fluid dynamics (CFD) image shows the Hyper-X vehicle at a Mach 7 test condition with the engine operating. The solution includes both internal (scramjet engine) and external flow fields, including the interaction between the engine exhaust and vehicle aerodynamics. The image illustrates surface heat transfer on the vehicle surface (red is highest heating) and flowfield contours at local Mach number. The last contour illustrates the engine exhaust plume shape. This solution approach is one method of predicting the vehicle performance, and the best method for determination of vehicle structural, pressure and thermal design loads. The Hyper-X program is an ambitious series of experimental flights to expand the boundaries of high-speed aeronautics and develop new technologies for space access. When the first of three aircraft flies, it will be the first time a non-rocket engine has powered a vehicle in flight at hypersonic speeds--speeds above Mach 5, equivalent to about one mile per second or approximately 3,600 miles per hour at sea level. Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly

  18. Minimum-fuel ascent to orbit using air-breathing propulsion

    NASA Technical Reports Server (NTRS)

    Van Buren, Mark A.; Mease, Kenneth D.

    1989-01-01

    Single-stage vehicles using air-breathing propulsion hold promise for more economical delivery of payloads to orbit. The characterization of minimum-fuel trajectories over the range of possible engine and aerodynamic performance of such vehicles provides useful feedback to engine and vehicle designers and paves the way for the development of guidance logic. The minimum-fuel trajectory problem is formulated, propulsion system and aerodynamic models are presented, a numerical solution approach is described, and some preliminary results are discussed.

  19. JANNAF Airbreathing Propulsion Subcommittee and 35th Combustion Subcommittee Meeting. Volume 1

    NASA Technical Reports Server (NTRS)

    Fry, Ronald S. (Editor); Gannaway, Mary T. (Editor); Rognan, Melanie (Editor)

    1998-01-01

    This document, CPIA Publication 682, Volume 1, is a compilation of 5 unclassified/unlimited technical papers (approved for public release) which were presented at the 1 998 meeting of the Joint Army-Navy-NASA-Air Force (JANNAF) Airbreathing Propulsion Subcommittee (APS) and Combustion Subcommittee (CS) held jointly with the Propulsion Systems Hazards Subcommittee (PSHS). The meeting was held on 7-11 December 1998 at Raytheon Systems Company and the Marriott Hotel, Tucson, AZ. Topics covered include HyTech technology development, hydrocarbon fuel development for hypersonic applications, pulse detonation propulsion system development and arc heaters for direct-connect scramjet testing.

  20. Advanced hypersonic aircraft design

    NASA Technical Reports Server (NTRS)

    Utzinger, Rob; Blank, Hans-Joachim; Cox, Craig; Harvey, Greg; Mckee, Mike; Molnar, Dave; Nagy, Greg; Petersen, Steve

    1992-01-01

    The objective of this design project is to develop the hypersonic reconnaissance aircraft to replace the SR-71 and to complement existing intelligence gathering devices. The initial design considerations were to create a manned vehicle which could complete its mission with at least two airborne refuelings. The aircraft must travel between Mach 4 and Mach 7 at an altitude of 80,000 feet for a maximum range of 12,000 nautical miles. The vehicle should have an air breathing propulsion system at cruise. With a crew of two, the aircraft should be able to take off and land on a 10,000 foot runway, and the yearly operational costs were not to exceed $300 million. Finally, the aircraft should exhibit stealth characteristics, including a minimized radar cross-section (RCS) and a reduced sonic boom. The technology used in this vehicle should allow for production between the years 1993 and 1995.

  1. Review of Leading Approaches for Mitigating Hypersonic Vehicle Communications Blackout and a Method of Ceramic Particulate Injection Via Cathode Spot Arcs for Blackout Mitigation

    NASA Technical Reports Server (NTRS)

    Gillman, Eric D.; Foster, John E.; Blankson, Isaiah M.

    2010-01-01

    Vehicles flying at hypersonic velocities within the atmosphere become enveloped in a "plasma sheath" that prevents radio communication, telemetry, and most importantly, GPS signal reception for navigation. This radio "blackout" period has been a problem since the dawn of the manned space program and was an especially significant hindrance during the days of the Apollo missions. An appropriate mitigation method must allow for spacecraft to ground control and ground control to spacecraft communications through the reentry plasma sheath. Many mitigation techniques have been proposed, including but not limited to, aerodynamic shaping, magnetic windows, and liquid injection. The research performed on these mitigation techniques over the years will be reviewed and summarized, along with the advantages and obstacles that each technique will need to overcome to be practically implemented. A unique approach for mitigating the blackout communications problem is presented herein along with research results associated with this method. The novel method involves the injection of ceramic metal-oxide particulate into a simulated reentry plasma to quench the reentry plasma. Injection of the solid ceramic particulates is achieved by entrainment within induced, energetic cathode spot flows.

  2. Hypersonic rarefied wake characterization

    NASA Technical Reports Server (NTRS)

    Brewer, E. B.

    1993-01-01

    Results of a numerical study using the direct simulation Monte Carlo (DSMC) method are presented for hypersonic rarefied flow over an aeroassisted space transfer vehicle (ASTV). The emphasis of the study is the characterization of the near wake region which includes the ASTV payload. The study covered the transitional flow regime from near continuum to free molecular. Calculations show that the character of the near wake is significantly affected by the presence of the payload. Flow separation occurs when an afterbody is present throughout the transitional flow regime. In contrast, when no afterbody is present, no separation is observed until the flow approaches continuum.

  3. X-43A/Hyper-X Vehicle Arrives at NASA Dryden

    NASA Technical Reports Server (NTRS)

    1999-01-01

    A close-up of the X-43A Hypersonic Experimental Vehicle, or 'Hyper-X,' in its protective shipping framework as it arrives at the Dryden Flight Research Center in October 1999. The X-43A was developed to research a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only

  4. Arc-heated gas flow experiments for hypersonic propulsion applications

    NASA Astrophysics Data System (ADS)

    Roseberry, Christopher Matthew

    Although hydrogen is an attractive fuel for a hypersonic air-breathing vehicle in terms of reaction rate, flame temperature, and energy content per unit mass, the substantial tank volume required to store hydrogen imposes a drag penalty to performance that tends to offset these advantages. An alternative approach is to carry a hydrocarbon fuel and convert it on-board into a hydrogen-rich gas mixture to be injected into the engine combustors. To investigate this approach, the UTA Arc-Heated Wind Tunnel facility was modified to run on methane rather than the normally used nitrogen. Previously, this facility was extensively developed for the purpose of eventually performing experiments simulating scramjet engine flow along a single expansion ramp nozzle (SERN) in addition to more generalized applications. This formidable development process, which involved modifications to every existing subsystem along with the incorporation of new subsystems, is described in detail. Fortunately, only a minor plumbing reconfiguration was required to prepare the facility for the fuel reformation research. After a failure of the arc heater power supply, a 5.6 kW plasma-cutting torch was modified in order to continue the arc pyrolysis experiments. The outlet gas flow from the plasma torch was sampled and subsequently analyzed using gas chromatography. The experimental apparatus converted the methane feedstock almost completely into carbon, hydrogen and acetylene. A high yield of hydrogen, consisting of a product mole fraction of roughly 0.7, was consistently obtained. Unfortunately, the energy consumption of the apparatus was too excessive to be feasible for a flight vehicle. However, other researchers have pyrolyzed hydrocarbons using electric arcs with much less power input per unit mass.

  5. Hypersonic propulsion - Breaking the thermal barrier

    NASA Technical Reports Server (NTRS)

    Weidner, J. P.

    1993-01-01

    The challenges of hypersonic propulsion impose unique features on the hypersonic vehicle - from large volume requirements to contain cryogenic fuel to airframe-integrated propulsion required to process sufficient quantities of air. Additional challenges exist in the design of the propulsion module that must be capable of efficiently processing air at very high enthalpies, adding and mixing fuel at supersonic speeds and expanding the exhaust products to generate thrust greater than drag. The paper explores the unique challenges of the integrated hypersonic propulsion system, addresses propulsion cycle selection to cope with the severe thermal environment and reviews the direction of propulsion research at hypervelocity speeds.

  6. Research in robust control for hypersonic aircraft

    NASA Technical Reports Server (NTRS)

    Calise, A. J.

    1993-01-01

    The research during the second reporting period has focused on robust control design for hypersonic vehicles. An already existing design for the Hypersonic Winged-Cone Configuration has been enhanced. Uncertainty models for the effects of propulsion system perturbations due to angle of attack variations, structural vibrations, and uncertainty in control effectiveness were developed. Using H(sub infinity) and mu-synthesis techniques, various control designs were performed in order to investigate the impact of these effects on achievable robust performance.

  7. Local and System Level Considerations for Plasma-Based Techniques in Hypersonic Flight

    NASA Astrophysics Data System (ADS)

    Suchomel, Charles; Gaitonde, Datta

    2007-01-01

    The harsh environment encountered due to hypersonic flight, particularly when air-breathing propulsion devices are utilized, poses daunting challenges to successful maturation of suitable technologies. This has spurred the quest for revolutionary solutions, particularly those exploiting the fact that air under these conditions can become electrically conducting either naturally or through artificial enhancement. Optimized development of such concepts must emphasize not only the detailed physics by which the fluid interacts with the imposed electromagnetic fields, but must also simultaneously identify system level issues integration and efficiencies that provide the greatest leverage. This paper presents some recent advances at both levels. At the system level, an analysis is summarized that incorporates the interdependencies occurring between weight, power and flow field performance improvements. Cruise performance comparisons highlight how one drag reduction device interacts with the vehicle to improve range. Quantified parameter interactions allow specification of system requirements and energy consuming technologies that affect overall flight vehicle performance. Results based on on the fundamental physics are presented by distilling numerous computational studies into a few guiding principles. These highlight the complex non-intuitive relationships between the various fluid and electromagnetic fields, together with thermodynamic considerations. Generally, energy extraction is an efficient process, while the reverse is accompanied by significant dissipative heating and inefficiency. Velocity distortions can be detrimental to plasma operation, but can be exploited to tailor flows through innovative electromagnetic configurations.

  8. Transpiration Cooling Of Hypersonic Blunt Body

    NASA Technical Reports Server (NTRS)

    Henline, William D.

    1991-01-01

    Results on analytical approximation and numerical simulation compared. Report presents theoretical study of degree to which transpiration blocks heating of blunt, axisymmetric body by use of injected air. Transpiration cooling proposed to reduce operating temperatures on nose cones of proposed hypersonic aerospace vehicles. Analyses important in design of thermal protection for such vehicles.

  9. Hypersonic Entry

    NASA Technical Reports Server (NTRS)

    2008-01-01

    [figure removed for brevity, see original site] Click on the image for the animation

    NASA's Phoenix Mars Lander will enter the Martian atmosphere at hypersonic speeds. Friction will heat the forward-facing surface of the heat shield to a peak of about 1,420 degrees Celsius (2,600 degrees Fahrenheit) at an altitude of 41 kilometers (25.5 miles).

    This illustration is part of the animation featured above.

    The Phoenix Mission is led by the University of Arizona, Tucson, on behalf of NASA. Project management of the mission is by NASA's Jet Propulsion Laboratory, Pasadena, Calif. Spacecraft development is by Lockheed Martin Space Systems, Denver.

  10. High-Temperature Properties of Ceramic Fibers and Insulations for Thermal Protection of Atmospheric Entry and Hypersonic Cruise Vehicles

    NASA Technical Reports Server (NTRS)

    Kourtides, Demetrius A.; Pitts, William C.; Araujo, Myrian; Zimmerman, R. S.

    1988-01-01

    Multilayer insulations which will operate in the 500C to 1000C temperature range are being considered for possible applications on aerospace vehicles subject to convective and radiative heating during atmospheric entry. The insulations described in this paper consist of ceramic fabrics, insulations, and metal foils quilted together using ceramic thread. As these types of insulations have highly anisotropic properties, the total heat transfer characteristics of these insulations must be determined. Data are presented on the thermal diffusivity and thermal conductivity of four types of multilayer insulations and are compared to the baseline Advanced Flexible Reusable Surface Insulation

  11. Hyper-X Research Vehicle - Artist Concept in Flight with Scramjet Engine Firing

    NASA Technical Reports Server (NTRS)

    1997-01-01

    This is an artist's depiction of a Hyper-X research vehicle under scramjet power in free-flight following separation from its booster rocket. The X-43A was developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need

  12. A CFD validation roadmap for hypersonic flows

    NASA Technical Reports Server (NTRS)

    Marvin, Joseph G.

    1992-01-01

    A roadmap for computational fluid dynamics (CFD) code validation is developed. The elements of the roadmap are consistent with air-breathing vehicle design requirements and related to the important flow path components: forebody, inlet, combustor, and nozzle. Building block and benchmark validation experiments are identified along with their test conditions and measurements. Based on an evaluation criteria, recommendations for an initial CFD validation data base are given and gaps identified where future experiments would provide the needed validation data.

  13. High-temperature properties of ceramic fibers and insulations for thermal protection of atmospheric entry and hypersonic cruise vehicles

    NASA Technical Reports Server (NTRS)

    Kourtides, Demetrius A.; Pitts, William C.; Araujo, Myrian; Zimmerman, R. S.

    1988-01-01

    Multilayer insulations (MIs) which will operate in the 500 to 1000 C temperature range are being considered for possible applications on aerospace vehicles subject to convective and radiative heating during atmospheric entry. The insulations described consist of ceramic fibers, insulations, and metal foils quilted together with ceramic thread. As these types of insulations have highly anisotropic properties, the total heat transfer characteristics must be determined. Data are presented on the thermal diffusivity and thermal conductivity of four types of MIs and are compared to the baseline Advanced Flexible Reusable Surface Insulation currently used on the Space Shuttle Orbiter. In addition, the high temperature properties of the fibers used in these MIs are discussed. The fibers investigated included silica and three types of aluminoborosilicate (ABS). Static tension tests were performed at temperatures up to 1200 C and the ultimate strain, tensile strength, and tensile modulus of single fibers were determined.

  14. High temperature properties of ceramic fibers and insulations for thermal protection of atmospheric entry and hypersonic cruise vehicles

    NASA Technical Reports Server (NTRS)

    Kourtides, Demetrius A.; Pitts, William C.; Araujo, Myrian; Zimmerman, R. S.

    1988-01-01

    Multilayer insulations (MIs) which will operate in the 500 to 1000 C temperature range are being considered for possible applications on aerospace vehicles subject to convective and radiative heating during atmospheric entry. The insulations described consist of ceramic fibers, insulations, and metal foils quilted together with ceramic thread. As these types of insulations have highly anisotropic properties, the total heat transfer characteristics must be determined. Data are presented on the thermal diffusivity and thermal conductivity of four types of MIs and are compared to the baseline Advanced Flexible Reusable Surface Insulation currently used on the Space Shuttle Orbiter. In addition, the high temperature properties of the fibers used in these MIs are discussed. The fibers investigated included silica and three types of aluminoborosilicate (ABS). Static tension tests were performed at temperatures up to 1200 C and the ultimate strain, tensile strength, and tensile modulus of single fibers were determined.

  15. A fuselage/tank structure study for actively cooled hypersonic cruise vehicles, summary. [aircraft design of aircraft fuel systems

    NASA Technical Reports Server (NTRS)

    Pirrello, C. J.; Baker, A. H.; Stone, J. E.

    1976-01-01

    A detailed analytical study was made to investigate the effects of fuselage cross section (circular and elliptical) and the structural arrangement (integral and nonintegral tanks) on aircraft performance. The vehicle was a 200 passenger, liquid hydrogen fueled Mach 6 transport designed to meet a range goal of 9.26 Mn (5000 NM). A variety of trade studies were conducted in the area of configuration arrangement, structural design, and active cooling design in order to maximize the performance of each of three point design aircraft: (1) circular wing-body with nonintegral tanks, (2) circular wing-body with integral tanks and (3) elliptical blended wing-body with integral tanks. Aircraft range and weight were used as the basis for comparison. The resulting design and performance characteristics show that the blended body integral tank aircraft weights the least and has the greatest range capability, however, producibility and maintainability factors favor nonintegral tank concepts.

  16. High-temperature properties of ceramic fibers and insulations for thermal protection of atmospheric entry and hypersonic cruise vehicles

    SciTech Connect

    Kourtides, D.A.; Pitts, W.C.; Araujo, M.; Zimmerman, R.S.

    1988-02-01

    Multilayer insulations (MIs) which will operate in the 500 to 1000 C temperature range are being considered for possible applications on aerospace vehicles subject to convective and radiative heating during atmospheric entry. The insulations described consist of ceramic fibers, insulations, and metal foils quilted together with ceramic thread. As these types of insulations have highly anisotropic properties, the total heat transfer characteristics must be determined. Data are presented on the thermal diffusivity and thermal conductivity of four types of MIs and are compared to the baseline Advanced Flexible Reusable Surface Insulation currently used on the Space Shuttle Orbiter. In addition, the high temperature properties of the fibers used in these MIs are discussed. The fibers investigated included silica and three types of aluminoborosilicate (ABS). Static tension tests were performed at temperatures up to 1200 C and the ultimate strain, tensile strength, and tensile modulus of single fibers were determined.

  17. The Syracuse University Center for Training and Research in Hypersonics

    NASA Technical Reports Server (NTRS)

    LaGraff, John; Blankson, Isaiah (Technical Monitor); Robinson, Stephen K. (Technical Monitor); Walsh, Michael J. (Technical Monitor); Anderson, Griffin Y. (Technical Monitor)

    2000-01-01

    In Fall 1993, NASA Headquarters established Centers for Hypersonics at the University of Maryland, the University of Texas-Arlington, and Syracuse University. These centers are dedicated to research and education in hypersonic technologies and have the objective of educating the next generation of engineers in this critical field. At the Syracuse University Center for Hypersonics this goal is being realized by focusing resources to: Provide an environment in which promising undergraduate students can learn the fundamental engineering principles of hypersonics so that they may make a seamless transition to graduate study and research in this field; Provide graduate students with advanced training in hypersonics and an opportunity to interact with leading authorities in the field in both research and instructional capacities; and Perform fundamental research in areas that will impact hypersonic vehicle design and development.

  18. Experimental aerothermodynamic research of hypersonic aircraft

    NASA Technical Reports Server (NTRS)

    Cleary, Joseph W.

    1990-01-01

    Wind tunnel tests were conducted to establish a benchmark experimental data base for a genetic hypersonic aircraft vehicle. Comprehensive measurements were made at Mach 7 to give flow visualization, surface pressure, surface convective heat transfer, and flow field Pitot pressure for a delta platform all-body vehicle. The tests were conducted in the NASA/Ames 3.5-Foot Hypersonic Wind Tunnel at Reynolds numbers sufficient to give turbulent flow. Comparisons are made of the experimental results with computational solutions of the flow by an upwind parabolized Navier-Stokes code developed at Ames. Good agreement of experiment with solutions by the code is demonstrated.

  19. Airbreathing Propulsion System Analysis Using Multithreaded Parallel Processing

    NASA Technical Reports Server (NTRS)

    Schunk, Richard Gregory; Chung, T. J.; Rodriguez, Pete (Technical Monitor)

    2000-01-01

    In this paper, parallel processing is used to analyze the mixing, and combustion behavior of hypersonic flow. Preliminary work for a sonic transverse hydrogen jet injected from a slot into a Mach 4 airstream in a two-dimensional duct combustor has been completed [Moon and Chung, 1996]. Our aim is to extend this work to three-dimensional domain using multithreaded domain decomposition parallel processing based on the flowfield-dependent variation theory. Numerical simulations of chemically reacting flows are difficult because of the strong interactions between the turbulent hydrodynamic and chemical processes. The algorithm must provide an accurate representation of the flowfield, since unphysical flowfield calculations will lead to the faulty loss or creation of species mass fraction, or even premature ignition, which in turn alters the flowfield information. Another difficulty arises from the disparity in time scales between the flowfield and chemical reactions, which may require the use of finite rate chemistry. The situations are more complex when there is a disparity in length scales involved in turbulence. In order to cope with these complicated physical phenomena, it is our plan to utilize the flowfield-dependent variation theory mentioned above, facilitated by large eddy simulation. Undoubtedly, the proposed computation requires the most sophisticated computational strategies. The multithreaded domain decomposition parallel processing will be necessary in order to reduce both computational time and storage. Without special treatments involved in computer engineering, our attempt to analyze the airbreathing combustion appears to be difficult, if not impossible.

  20. A hypersonic shock tunnel test of the Aeroassist Flight Experiment (AFE) vehicle at high altitude test conditions

    NASA Technical Reports Server (NTRS)

    Wittliff, C. E.; O'Connor, A. C.; Sweet, S. J.

    1992-01-01

    The Aeroassist Flight Experiment (AFE) vehicle is a 4.26 m-diameter raked and blunted elliptic cone that is intended to obtain aerodynamic and aerothermodynamic data on blunt bodies at velocities near 10 km/s and altitudes above 75 km. Previous wind tunnel tests of this configuration have been at Mach numbers of 10 or lower. The present test program consisted of five test conditions at Mach numbers between 15 and 20 and simulated altitudes from 75 km to 97 km. Static longitudinal aerodynamic coefficients, the lift-to-drag ratio, and the trim angle-of-attack were obtained for angles-of-attack between 7 deg and 27 deg. The effects of angle-of-attack and shock Reynolds number, Re2, have been studied. These results indicate the AFE configuration is longitudinally stable. The variation of normal force slope, CNalpha and lift-to-drag slope, (L/D)alpha as well as the trim angle-of-attack with Re2 are documented. The trim angle-of-attack increases with decreasing values of shock Reynolds number (increasing altitude) while the lift-to-drag ratio at the trimmed conditions is relatively constant.

  1. Impact of Ion Acoustic Wave Instabilities in the Flow Field of a Hypersonic Vehicle on EM Signals

    NASA Astrophysics Data System (ADS)

    Mudaliar, Saba; Sotnikov, Vladimir

    2016-10-01

    Flow associated with a high speed air vehicle (HSAV) can get partially ionized. In the absence of external magnetic field the flow field turbulence is due to ion acoustic wave (IAW) instabilities. Our interest is in studying the impact of this turbulence on the radiation characteristics of EM signals from the HSAV. We decompose the radiated signal into coherent and diffuse parts. We find that the coherent part has the same spectrum as that of the source signal, but it is distorted because of dispersive coherent attenuation. The diffuse part is expressed as a convolution (in wavenumber and frequency) of the source signal with the spectrum of electron density fluctuations. This is a constrained convolution in the sense that the spectrum has to satisfy the IAW dispersion relation. A quantity that characterizes the flow is the mean free path (MFP). When the MFP is large compared to the thickness of the flow the coherent part is significant. If the MFP is larger than the thickness of the flow the diffuse part is the dominant part of the received signal. In the special case when the source signal frequency is close the electron plasma frequency, there can exist in the flow region Langmuir modes in addition to the EM modes. The radiation characteristics of EM source signals from the HSAV in this case are quite different.

  2. NASA's Hypersonic Investment Area

    NASA Technical Reports Server (NTRS)

    Hueter, Uwe; Hutt, John; McClinton, Charles

    2002-01-01

    NASA has established long term goals for access to space. The third generation launch systems are to be fully reusable and operational around 2025. The goal for third-generation launch systems represents significant reduction in cost and improved safety over the current first generation system. The Advanced Space Transportation Office (ASTP) at NASA s Marshall Space Flight Center (MSFC) has the agency lead to develop space transportation technologies. Within ASTP, under the Hypersonic Investment Area (HIA), third generation technologies are being pursued in the areas of propulsion, airframe, integrated vehicle health management (IVHM), avionics, power, operations and system analysis. These technologies are being matured through research and both ground and flight-testing. This paper provides an overview of the HIA program plans and recent accomplishments.

  3. Emerging hypersonic propulsion technology

    NASA Technical Reports Server (NTRS)

    Curran, E. T.; Beach, H. L., Jr.

    1988-01-01

    Currently there is a renewal of interest in the utilization of air breathing engines for hypersonic flight. The use of such engines in accelerative missions is discussed, and the nature of the trade-off between engine thrust-to-weight ratio and specific impulse is highlighted. It is also pointed out that the use of a cryogenic fuel such as liquid hydrogen offers the opportunity to develop both precooled derivatives of turboaccelerator engines and new cryogenic engine cycles, where the heat exchange process plays a significant role in the engine concept. The continuing challenges of developing high speed supersonic combustion ramjet engines are discussed. The paper concludes with a brief review of the difficult discipline of vehicle integration, and the challenges of both ground and flight testing.

  4. Integrated Propulsion/Vehicle System Structurally Optimized

    NASA Technical Reports Server (NTRS)

    Hunter, James E.; McCurdy, David R.

    2003-01-01

    Ongoing research and testing are essential in the development of air-breathing hypersonic propulsion technology, and this year some positive advancement was made at the NASA Glenn Research Center. Recent work performed for GTX, a rocket-based combined-cycle, single-stage-to-orbit concept, included structural assessments of both the engine and flight vehicle. In the development of air-breathing engine technology, it is impractical to design and optimize components apart from the fully integrated system because tradeoffs must be made between performance and structural capability. Efforts were made to control the flight trajectory, for example, to minimize the aerodynamic heating effects. Structural optimization was applied to evaluate concept feasibility and was instrumental in the determination of the gross liftoff weight of the integrated system. Achieving low Earth orbit with even a small payload requires an aggressive approach to weight minimization through the use of lightweight, oxidation-resistant composite materials. Assessing the integrated system involved investigating the flight trajectory to determine where the critical load events occur in flight and then generating the corresponding environment at each of these events. Structural evaluation requires the mapping of the critical flight loads to finite element models, including the combined effects of aerodynamic, inertial, combustion, and other loads. NASA s APAS code was used to generate aerodynamic pressure and temperature profiles at each critical event. The radiation equilibrium surface temperatures from APAS were used to predict temperatures through the thickness. Heat transfer solutions using NASA's MINIVER code and the SINDA code (Cullimore & Ring Technologies, Littleton, CO) were calculated at selective points external to the integrated vehicle system and then extrapolated over the entire exposed surface. FORTRAN codes were written to expedite the finite element mapping of the aerodynamic heating

  5. Design and Fabrication of the ISTAR Direct-Connect Combustor Experiment at the NASA Hypersonic Tunnel Facility

    NASA Technical Reports Server (NTRS)

    Lee, Jin-Ho; Krivanek, Thomas M.

    2005-01-01

    The Integrated Systems Test of an Airbreathing Rocket (ISTAR) project was a flight demonstration project initiated to advance the state of the art in Rocket Based Combined Cycle (RBCC) propulsion development. The primary objective of the ISTAR project was to develop a reusable air breathing vehicle and enabling technologies. This concept incorporated a RBCC propulsion system to enable the vehicle to be air dropped at Mach 0.7 and accelerated up to Mach 7 flight culminating in a demonstration of hydrocarbon scramjet operation. A series of component experiments was planned to reduce the level of risk and to advance the technology base. This paper summarizes the status of a full scale direct connect combustor experiment with heated endothermic hydrocarbon fuels. This is the first use of the NASA GRC Hypersonic Tunnel facility to support a direct-connect test. The technical and mechanical challenges involved with adapting this facility, previously used only in the free-jet configuration, for use in direct connect mode will be also described.

  6. Multidisciplinary Modeling Software for Analysis, Design, and Optimization of HRRLS Vehicles

    NASA Technical Reports Server (NTRS)

    Spradley, Lawrence W.; Lohner, Rainald; Hunt, James L.

    2011-01-01

    The concept for Highly Reliable Reusable Launch Systems (HRRLS) under the NASA Hypersonics project is a two-stage-to-orbit, horizontal-take-off / horizontal-landing, (HTHL) architecture with an air-breathing first stage. The first stage vehicle is a slender body with an air-breathing propulsion system that is highly integrated with the airframe. The light weight slender body will deflect significantly during flight. This global deflection affects the flow over the vehicle and into the engine and thus the loads and moments on the vehicle. High-fidelity multi-disciplinary analyses that accounts for these fluid-structures-thermal interactions are required to accurately predict the vehicle loads and resultant response. These predictions of vehicle response to multi physics loads, calculated with fluid-structural-thermal interaction, are required in order to optimize the vehicle design over its full operating range. This contract with ResearchSouth addresses one of the primary objectives of the Vehicle Technology Integration (VTI) discipline: the development of high-fidelity multi-disciplinary analysis and optimization methods and tools for HRRLS vehicles. The primary goal of this effort is the development of an integrated software system that can be used for full-vehicle optimization. This goal was accomplished by: 1) integrating the master code, FEMAP, into the multidiscipline software network to direct the coupling to assure accurate fluid-structure-thermal interaction solutions; 2) loosely-coupling the Euler flow solver FEFLO to the available and proven aeroelasticity and large deformation (FEAP) code; 3) providing a coupled Euler-boundary layer capability for rapid viscous flow simulation; 4) developing and implementing improved Euler/RANS algorithms into the FEFLO CFD code to provide accurate shock capturing, skin friction, and heat-transfer predictions for HRRLS vehicles in hypersonic flow, 5) performing a Reynolds-averaged Navier-Stokes computation on an HRRLS

  7. The optimum hypersonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Trimmer, L. L.; Cary, A., Jr.; Voisinet, R. L. P.

    1986-01-01

    The capabilities of existing hypersonic wind tunnels in the U.S. are assessed to form a basis for recommendations for a new, costly facility which would provide data for modeling the hypervelocity aerodynamics envisioned for the new generation of aerospace vehicles now undergoing early studies. Attention is given to the regimes, both entry and aerodynamic, which the new vehicles will encounter, and the shortcomings of data generated for the Orbiter before flight are discussed. The features of foreign-gas, impulse, aeroballistic range, arc-heated and combustion-heated facilities are examined, noting that in any hypersonic wind tunnel the flow must be preheated to prevent liquefaction upon expansion in the test channel. The limitations of the existing facilities and the identification of the regimes which must be studied lead to a description of the characteristics of an optimum hypersonic wind tunnel, including the operations and productivity, the instrumentation, the nozzle design and the flow quality. Three different design approaches are described, each costing at least $100 million to achieve workability.

  8. Facility requirements for hypersonic propulsion system testing

    NASA Astrophysics Data System (ADS)

    Dunn, M. G.; Lordi, J. A.; Wittliff, C. E.; Holden, M. S.

    Facility requirements and capabilities for hypersonic propulsion system testing are reviewed with emphasis on short-duration test facilities. Past and current hypersonic facility studies are reviewed, and some of the many problems currently associated with wing-body hypersonic aircraft and several currently operational ground-based facilities or facilities in the development stage are described. Limitations on the short-duration shock tunnel are examined, including problem areas where this device can make significant contributions to the type of unified computational, ground-test, and flight-experiment program that will be necessary to resolve complex issues associated with the development of either a SSTO vehicle or an air-breathing/rocket-assist-to-orbit vehicle.

  9. Hyper-X Research Vehicle - Artist Concept Mounted on Pegasus Rocket Attached to B-52 Launch Aircraft

    NASA Technical Reports Server (NTRS)

    1997-01-01

    This artist's concept depicts the Hyper-X research vehicle riding on a booster rocket prior to being launched by the Dryden Flight Research Center's B-52 at about 40,000 feet. The X-43A was developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry

  10. Hypersonic flight performance improvements by overfueled ramjet combustion

    NASA Astrophysics Data System (ADS)

    Sachs, G.; Bayer, R.; Lederer, R.; Schaber, R.

    1991-12-01

    The performance characteristics of hypersonic airbreathing engines are examined with emphasis on the effect of overfueled combustion on thrust and specific fuel-consumption, as well as on the combustion temperature, real gas effects, and pollution due to exhaust gas. It is shown that overfueled ramjet combustion can provide a means for improving flight performance at hypersonic speed and, consequently, reduce the mission fuel burn and the propulsion system weight. It is also shown that, in the separation flight maneuver, the separation condition for the upper stage can be improved by overfueled ramjet combustion of the first stage, making it possible to increase the payload which the upper stage can deliver into orbit. The flight mechanics modeling considerations are presented.

  11. Study of Forebody Injection and Mixing with Application to Hypervelocity Airbreathing Propulsion

    NASA Technical Reports Server (NTRS)

    Axdahl, Erik; Kumar, Ajay; Wilhite, Alan

    2012-01-01

    The use of premixed, shock-induced combustion in the context of a hypervelocity, airbreathing vehicle requires effective injection and mixing of hydrogen fuel and air on the vehicle forebody. Three dimensional computational simulations of fuel injection and mixing from flush-wall and modified ramp and strut injectors are reported in this study. A well-established code, VULCAN, is used to conduct nonreacting, viscous, turbulent simulations on a flat plate at conditions relevant to a Mach 12 flight vehicle forebody. In comparing results of various fuel injection strategies, it is found that strut injection provides the greatest balance of performance between mixing efficiency and stream thrust potential.

  12. Investigation of antimatter air-breathing propulsion for single-stage-to-orbit ships

    NASA Astrophysics Data System (ADS)

    Froning, H. D.

    Because the mutual annihilation of matter and antimatter releases all the energy that is stored within the physical structure of material mass, it provides the most powerful reaction that is possible for propulsive thrust. This paper considers the use of such annihilation energy for single-stage-to-orbit vehicles that would be powered by rocket and air-breathing propulsion and would reach and return from orbit with a single propulsive stage.

  13. Hypersonic Force Application and Launch Technology Demonstration

    DTIC Science & Technology

    2004-09-14

    region Regardless of anti-access threats In a single or multi-theater environment Distribution: Gov & Gov Contractors, ITAR Restricted 4 SMALL SATELLITE...demand Distribution: Gov & Gov Contractors, ITAR Restricted Objective: CAV Technology Demonstration Flight Test Description of CAV: Lifting aeroshell...Common Aero Vehicle (CAV) Distribution: Gov & Gov Contractors, ITAR Restricted 7 HYPERSONIC TECHNOLOGY EVOLUTION Building Block Tech Development and

  14. NASA hypersonic flight demonstrators—overview, status, and future plans

    NASA Astrophysics Data System (ADS)

    L. Moses, Paul; L. Rausch, Vincent; T. Nguyen, Luat; R. Hill, Jeryl

    2004-08-01

    NASA's Next Generation Launch Technology (NGLT) program is developing and maturing advanced propulsion and vehicle systems technologies and flight vehicle concepts to enable future development of safer and more economical launch systems. Within NGLT, NASA is developing advanced air breathing propulsion systems and demonstrating these systems in hypersonic flight vehicles. The flight demonstrations are necessary to fully validate these technologies for application to future space launch vehicles and other flight systems. NASA's Hyper-X Program (X-43A) began the effort to flight demonstrate hypersonic air breathing propulsion systems to provide technologies that will enable development of safer and more economic space access vehicles in the future. Following X-43A, NASA, in collaboration with the United States (US) Department of Defense (DoD), is developing additional, progressively more complex hypersonic X-vehicles that will demonstrate new air breathing propulsion systems, propulsion-airframe integration, and other vehicle systems technologies required for high speed flight up to Mach 15. These technologies will contribute to safer, more reliable and more economic future launch systems and hypersonic aircraft/missiles. This paper describes NASA's current hypersonic flight demonstration projects, status of the efforts, and plans for future vehicles.

  15. Analysis of Hypersonic Vehicle Wakes

    DTIC Science & Technology

    2015-09-17

    ong x 2 ! 2 i f viscous, a i f inviscid ivisc irest = 1 ! 9 fo r f resh start, 1 fo r restar t ncyc 1999e ! gl obal steps j update = 29 1 steps...turbul ent prandt l number schmidt _ turb = 6. 9 ! turbul ent schmidt number xt r = a. ! t ransition l ocat ion fo r a l gebraic model al ong x...number schmidt_turb = 8.9 ! t urbu l ent schmidt number xtr = e. ! transition location for al gebraic model al ong x ivisc 2 ! 2 if viscous, e if

  16. Uncertainty Assessment of Hypersonic Aerothermodynamics Prediction Capability

    NASA Technical Reports Server (NTRS)

    Bose, Deepak; Brown, James L.; Prabhu, Dinesh K.; Gnoffo, Peter; Johnston, Christopher O.; Hollis, Brian

    2011-01-01

    The present paper provides the background of a focused effort to assess uncertainties in predictions of heat flux and pressure in hypersonic flight (airbreathing or atmospheric entry) using state-of-the-art aerothermodynamics codes. The assessment is performed for four mission relevant problems: (1) shock turbulent boundary layer interaction on a compression corner, (2) shock turbulent boundary layer interaction due a impinging shock, (3) high-mass Mars entry and aerocapture, and (4) high speed return to Earth. A validation based uncertainty assessment approach with reliance on subject matter expertise is used. A code verification exercise with code-to-code comparisons and comparisons against well established correlations is also included in this effort. A thorough review of the literature in search of validation experiments is performed, which identified a scarcity of ground based validation experiments at hypersonic conditions. In particular, a shortage of useable experimental data at flight like enthalpies and Reynolds numbers is found. The uncertainty was quantified using metrics that measured discrepancy between model predictions and experimental data. The discrepancy data is statistically analyzed and investigated for physics based trends in order to define a meaningful quantified uncertainty. The detailed uncertainty assessment of each mission relevant problem is found in the four companion papers.

  17. Aerodynamic prediction techniques for hypersonic configuration design

    NASA Technical Reports Server (NTRS)

    1981-01-01

    An investigation of approximate theoretical techniques for predicting aerodynamic characteristics and surface pressures for relatively slender vehicles at moderate hypersonic speeds was performed. Emphasis was placed on approaches that would be responsive to preliminary configuration design level of effort. Potential theory was examined in detail to meet this objective. Numerical pilot codes were developed for relatively simple three dimensional geometries to evaluate the capability of the approximate equations of motion considered. Results from the computations indicate good agreement with higher order solutions and experimental results for a variety of wing, body, and wing-body shapes for values of the hypersonic similarity parameter M delta approaching one.

  18. Body weight of hypersonic aircraft, part 1

    NASA Technical Reports Server (NTRS)

    Ardema, Mark D.

    1988-01-01

    The load bearing body weight of wing-body and all-body hypersonic aircraft is estimated for a wide variety of structural materials and geometries. Variations of weight with key design and configuration parameters are presented and discussed. Both hot and cool structure approaches are considered in isotropic, organic composite, and metal matrix composite materials; structural shells are sandwich or skin-stringer. Conformal and pillow-tank designs are investigated for the all-body shape. The results identify the most promising hypersonic aircraft body structure design approaches and their weight trends. Geometric definition of vehicle shapes and structural analysis methods are presented in appendices.

  19. Vibrational relaxation in hypersonic flow fields

    NASA Technical Reports Server (NTRS)

    Meador, Willard E.; Miner, Gilda A.; Heinbockel, John H.

    1993-01-01

    Mathematical formulations of vibrational relaxation are derived from first principles for application to fluid dynamic computations of hypersonic flow fields. Relaxation within and immediately behind shock waves is shown to be substantially faster than that described in current numerical codes. The result should be a significant reduction in nonequilibrium radiation overshoot in shock layers and in radiative heating of hypersonic vehicles; these results are precisely the trends needed to bring theoretical predictions more in line with flight data. Errors in existing formulations are identified and qualitative comparisons are made.

  20. Hypersonic Aerodynamics Fellowships

    DTIC Science & Technology

    1991-02-11

    J.D. and Capriotti , D., "Viscous Optimized Hypersonic Waveriders," AIAA Paper No. 87-0272. Appendix B: Corpening, G. and Anderson,J., "Numerical...Hypersonic Waveriders K. G. Bowcutt, J. D. Anderson and D. Capriotti , Univ. of Maryland, College Park, MD AIAA 25th Aerospace Sciences Meeting January...OPTIMIZED HYPERSONIC WAVERIDERS by Kevin G. Rowcutt,* John D. Anderson, Jr.," and Diego Capriotti "** Department of Aerospace Engineering University of

  1. The X-43A hypersonic research aircraft and its modified Pegasus booster rocket mounted to NASA's NB

    NASA Technical Reports Server (NTRS)

    2001-01-01

    The first of three X-43A hypersonic research aircraft and its modified Pegasus booster rocket recently underwent combined systems testing while mounted to NASA's NB-52B carrier aircraft at the Dryden Flight Research Center, Edwards, California. The combined systems test was one of the last major milestones in the Hyper-X research program before the first X-43A flight. One of the major goals of the Hyper-X program is flight validation of airframe-integrated, air-breathing propulsion system, which so far have only been tested in ground facilities, such as wind tunnels. The X-43A flights will be the first actual flight tests of an aircraft powered by a revolutionary supersonic-combustion ramjet ('scramjet') engine capable of operating at hypersonic speeds above Mach 5 (five times the speed of sound). The X-43A design uses the underbody of the aircraft to form critical elements of the engine. The forebody shape helps compress the intake airflow, while the aft section acts as a nozzle to direct thrust. The 12-foot, unpiloted research vehicle was developed and built by MicroCraft Inc., Tullahoma, Tenn., under NASA contract. The booster, built by Orbital Sciences Corp., Dulles, Va., will accelerate the X-43A after the X-43A/booster 'stack' is air-launched from NASA's venerable NB-52 mothership. The X-43A will separate from the rocket at a predetermined altitude and speed and fly a pre-programmed trajectory, conducting aerodynamic and propulsion experiments until it descends into the Pacific Ocean. Three research flights are planned, two at Mach 7 and one at Mach 10.

  2. The QED engine spectrum - Fusion-electric propulsion for air-breathing to interstellar flight

    NASA Technical Reports Server (NTRS)

    Bussard, Robert W.; Jameson, Lorin W.

    1993-01-01

    A new inertial-electrostatic-fusion direct electric power source can be used to drive a relativistic e-beam to heat propellant. The resulting system is shown to yield specific impulse and thrust/mass ratio 2-3 orders of magnitude larger than from other advanced propulsion concepts. This QED system can be applied to aerospace vehicles from air-breathing to near-interstellar flight. Examples are given for Earth/Mars flight missions, that show transit times of 40 d with 20 percent payload in single-stage vehicles.

  3. Assessment of nonequilibrium radiation computation methods for hypersonic flows

    NASA Technical Reports Server (NTRS)

    Sharma, Surendra

    1993-01-01

    The present understanding of shock-layer radiation in the low density regime, as appropriate to hypersonic vehicles, is surveyed. Based on the relative importance of electron excitation and radiation transport, the hypersonic flows are divided into three groups: weakly ionized, moderately ionized, and highly ionized flows. In the light of this division, the existing laboratory and flight data are scrutinized. Finally, an assessment of the nonequilibrium radiation computation methods for the three regimes in hypersonic flows is presented. The assessment is conducted by comparing experimental data against the values predicted by the physical model.

  4. Supersonic Air-Breathing Stage For Commercial Launch Rocket

    NASA Technical Reports Server (NTRS)

    Martin, James A.

    1993-01-01

    Concept proposed to expand use of air-breathing, reusable stages to put more payload into orbit at less cost. Stage with supersonic air-breathing engines added to carry expendable stages from subsonic airplane to supersonic velocity. Carry payload to orbit. Expendable stages and payload placed in front of supersonic air-breathing stage. After releasing expendable stages, remotely piloted supersonic air-breathing stage returns to takeoff site and land for reuse. New concept extends use of low-cost reusable hardware and increases payload delivered from B-52.

  5. Swimming in air-breathing fishes.

    PubMed

    Lefevre, S; Domenici, P; McKenzie, D J

    2014-03-01

    Fishes with bimodal respiration differ in the extent of their reliance on air breathing to support aerobic metabolism, which is reflected in their lifestyles and ecologies. Many freshwater species undertake seasonal and reproductive migrations that presumably involve sustained aerobic exercise. In the six species studied to date, aerobic exercise in swim flumes stimulated air-breathing behaviour, and there is evidence that surfacing frequency and oxygen uptake from air show an exponential increase with increasing swimming speed. In some species, this was associated with an increase in the proportion of aerobic metabolism met by aerial respiration, while in others the proportion remained relatively constant. The ecological significance of anaerobic swimming activities, such as sprinting and fast-start manoeuvres during predator-prey interactions, has been little studied in air-breathing fishes. Some species practise air breathing during recovery itself, while others prefer to increase aquatic respiration, possibly to promote branchial ion exchange to restore acid-base balance, and to remain quiescent and avoid being visible to predators. Overall, the diversity of air-breathing fishes is reflected in their swimming physiology as well, and further research is needed to increase the understanding of the differences and the mechanisms through which air breathing is controlled and used during exercise.

  6. Aerodynamic heating in hypersonic flows

    NASA Technical Reports Server (NTRS)

    Reddy, C. Subba

    1993-01-01

    Aerodynamic heating in hypersonic space vehicles is an important factor to be considered in their design. Therefore the designers of such vehicles need reliable heat transfer data in this respect for a successful design. Such data is usually produced by testing the models of hypersonic surfaces in wind tunnels. Most of the hypersonic test facilities at present are conventional blow-down tunnels whose run times are of the order of several seconds. The surface temperatures on such models are obtained using standard techniques such as thin-film resistance gages, thin-skin transient calorimeter gages and coaxial thermocouple or video acquisition systems such as phosphor thermography and infrared thermography. The data are usually reduced assuming that the model behaves like a semi-infinite solid (SIS) with constant properties and that heat transfer is by one-dimensional conduction only. This simplifying assumption may be valid in cases where models are thick, run-times short, and thermal diffusivities small. In many instances, however, when these conditions are not met, the assumption may lead to significant errors in the heat transfer results. The purpose of the present paper is to investigate this aspect. Specifically, the objectives are as follows: (1) to determine the limiting conditions under which a model can be considered a semi-infinite body; (2) to estimate the extent of errors involved in the reduction of the data if the models violate the assumption; and (3) to come up with correlation factors which when multiplied by the results obtained under the SIS assumption will provide the results under the actual conditions.

  7. Diagnostics for Hypersonic Engine Control

    DTIC Science & Technology

    2015-02-01

    hypersonic engine control are changes in isolator margin, inlet mass capture, and performance (thrust, combustion efficiency, etc.). 15. SUBJECT TERMS...Most interesting for hypersonic engine control are changes in isolator margin, inlet mass capture, and performance (thrust, combustion efficiency, etc...AFRL-RQ-WP-TR-2015-0037 DIAGNOSTICS FOR HYPERSONIC ENGINE CONTROL Michael S. Brown and Jeffrey M. Donbar Hypersonic Sciences Branch

  8. Supersonic/hypersonic aerodynamic methods for aircraft design and analysis

    NASA Technical Reports Server (NTRS)

    Torres, Abel O.

    1992-01-01

    A methodology employed in engineering codes to predict aerodynamic characteristics over arbitrary supersonic/hypersonic configurations is considered. Engineering codes use a combination of simplified methods, based on geometrical impact angle and freestream conditions, to compute pressure distribution over the vehicle's surface in an efficient and timely manner. These approximate methods are valid for both hypersonic (Mach greater than 4) and lower speeds (Mach down to 2). It is concluded that the proposed methodology enables the user to obtain reasonable estimates of vehicle performance and engineering methods are valuable in the design process of these type of vehicles.

  9. Hypersonic drone design: A multidisciplinary experience

    NASA Technical Reports Server (NTRS)

    1988-01-01

    Efforts were focused on design problems of an unmanned hypersonic vehicle. It is felt that a scaled hypersonic drone is necessary to bridge the gap between present theory on hypersonics and the future reality of the National Aerospace Plane (NASP) for two reasons: to fulfill a need for experimental data in the hypersonic regime, and to provide a testbed for the scramjet engine which is to be the primary mode of propulsion for the NASP. Three areas of great concern to NASP design were examined: propulsion, thermal management, and flight systems. Problem solving in these areas was directed towards design of the drone with the idea that the same design techniques could be applied to the NASP. A seventy degree swept double delta wing configuration, developed in the 70's at NASA Langley, was chosen as the aerodynamic and geometric model for the drone. This vehicle would be air-launched from a B-1 at Mach 0.8 and 48,000 feet, rocket boosted by two internal engines to Mach 10 and 100,000 feet, and allowed to cruise under power of the scramjet engine until burnout. It would then return to base for an unpowered landing. Preliminary energy calculations based upon the flight requirements give the drone a gross launch weight of 134,000 lb. and an overall length of 85 feet.

  10. A predication model for combustion modes of the scramjet-powered aerospace vehicle based on the nonlinear features of the isolator flow field

    NASA Astrophysics Data System (ADS)

    Yang, Qingchun; Wang, Hongxin; Chetehouna, Khaled; Gascoin, Nicolas

    2017-01-01

    The supersonic combustion ramjet (scramjet) engine remains the most promising airbreathing engine cycle for hypersonic flight, particularly the high-performance dual-mode scramjet in the range of flight Mach number from 4 to 7, because it can operates under different combustion modes. Isolator is a very key component of the dual-mode scramjet engine. In this paper, nonlinear characteristics of combustion mode transition is theoretically analyzed. The discontinuous sudden changes of static pressure and Mach number are obtained as the mode transition occurs, which emphasizing the importance of predication and control of combustion modes. In this paper, a predication model of different combustion modes is developed based on these these nonlinear features in the isolator flow field. it can provide a valuable reference for control system design of the scramjet-powered aerospace vehicle.

  11. Airbreathing combined cycle engine systems

    NASA Technical Reports Server (NTRS)

    Rohde, John

    1992-01-01

    The Air Force and NASA share a common interest in developing advanced propulsion systems for commercial and military aerospace vehicles which require efficient acceleration and cruise operation in the Mach 4 to 6 flight regime. The principle engine of interest is the turboramjet; however, other combined cycles such as the turboscramjet, air turborocket, supercharged ejector ramjet, ejector ramjet, and air liquefaction based propulsion are also of interest. Over the past months careful planning and program implementation have resulted in a number of development efforts that will lead to a broad technology base for those combined cycle propulsion systems. Individual development programs are underway in thermal management, controls materials, endothermic hydrocarbon fuels, air intake systems, nozzle exhaust systems, gas turbines and ramjet ramburners.

  12. Experiments on hypersonic ramjet propulsion cycles using a ram accelerator

    NASA Technical Reports Server (NTRS)

    Chew, G.; Knowlen, C.; Burnham, E. A.; Hertzberg, A.; Bruckner, A. P.

    1991-01-01

    Work on hypersonic propulsion research using a ram accelerator is presented. Several different ram accelerator propulsive cycles have been experimentally demonstrated over the Mach number range of 3 to 8.5. The subsonic, thermally choked combustion mode has accelerated projectiles to near the Chapman-Jouguet (C-J) detonation velocity within many different propellant mixtures. In the transdetonative velocity regime (85 to 115 percent of C-J speed), projectiles have established a propulsive cycle which allows them to transition smoothly from subdetonative to superdetonative velocities. Luminosity data indicate that the combustion process moves forward onto the projectile body as it approaches the C-J speed. In the superdetonative velocity range, the projectiles accelerate while always traveling faster than the C-J velocity. Ram accelerator projectiles operating continuously through these velocity regimes generate distinctive hypersonic phenomena which can be studied very effectively in the laboratory. These results would be very useful for validating sophisticated CFD computer codes and in collecting engineering data for potential airbreathing hypersonic propulsive systems.

  13. Hypersonic Flow Computations on Unstructured Meshes

    NASA Technical Reports Server (NTRS)

    Bibb, K. L.; Riley, C. J.; Peraire, J.

    1997-01-01

    A method for computing inviscid hypersonic flow over complex configurations using unstructured meshes is presented. The unstructured grid solver uses an edge{based finite{volume formulation. Fluxes are computed using a flux vector splitting scheme that is capable of representing constant enthalpy solutions. Second{order accuracy in smooth flow regions is obtained by linearly reconstructing the solution, and stability near discontinuities is maintained by locally forcing the scheme to reduce to first-order accuracy. The implementation of the algorithm to parallel computers is described. Computations using the proposed method are presented for a sphere-cone configuration at Mach numbers of 5.25 and 10.6, and a complex hypersonic re-entry vehicle at Mach numbers of 4.5 and 9.8. Results are compared to experimental data and computations made with established structured grid methods. The use of the solver as a screening tool for rapid aerodynamic assessment of proposed vehicles is described.

  14. A guidance concept for hypersonic aerospacecrafts

    NASA Astrophysics Data System (ADS)

    Ishimoto, Shinji

    In this paper a guidance concept for hypersonic re-entry flights is presented. The method uses a closed-form guidance technique based on a drag acceleration reference profile. A guidance law for range control is developed. It employs a physical relation between vehicle energy and range instead of a prediction-correction technique used for Shuttle entry guidance. Simulation results show that the algorithm provides good performance.

  15. NASA's hypersonic fluid and thermal physics program (Aerothermodynamics)

    NASA Technical Reports Server (NTRS)

    Graves, R. A.; Hunt, J. L.

    1985-01-01

    This survey paper gives an overview of NASA's hypersonic fluid and thermal physics program (recently renamed aerothermodynamics). The purpose is to present the elements of, example results from, and rationale and projection for this program. The program is based on improving the fundamental understanding of aerodynamic and aerothermodynamic flow phenomena over hypersonic vehicles in the continuum, transitional, and rarefied flow regimes. Vehicle design capabilities, computational fluid dynamics, computational chemistry, turbulence modeling, aerothermal loads, orbiter flight data analysis, orbiter experiments, laser photodiagnostics, and facilities are discussed.

  16. Optimization of Air-Breathing Engine Concept

    NASA Technical Reports Server (NTRS)

    Patnaik, Surya N.; Lavelle, Thomas M.; Hopkins, Dale A.

    1996-01-01

    The design optimization of air-breathing propulsion engine concepts has been accomplished by soft-coupling the NASA Engine Performance Program (NEPP) analyzer with the NASA Lewis multidisciplinary optimization tool COMETBOARDS. Engine problems, with their associated design variables and constraints, were cast as nonlinear optimization problems with thrust as the merit function. Because of the large number of mission points in the flight envelope, the diversity of constraint types, and the overall distortion of the design space; the most reliable optimization algorithm available in COMETBOARDS, when used by itself, could not produce satisfactory, feasible, optimum solutions. However, COMETBOARDS' unique features-which include a cascade strategy, variable and constraint formulations, and scaling devised especially for difficult multidisciplinary applications-successfully optimized the performance of subsonic and supersonic engine concepts. Even when started from different design points, the combined COMETBOARDS and NEPP results converged to the same global optimum solution. This reliable and robust design tool eliminates manual intervention in the design of air-breathing propulsion engines and eases the cycle analysis procedures. It is also much easier to use than other codes, which is an added benefit. This paper describes COMETBOARDS and its cascade strategy and illustrates the capabilities of the combined design tool through the optimization of a high-bypass- turbofan wave-rotor-topped subsonic engine and a mixed-flow-turbofan supersonic engine.

  17. A computational fluid dynamics simulation of the hypersonic flight of the Pegasus(TM) vehicle using an artificial viscosity model and a nonlinear filtering method. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Mendoza, John Cadiz

    1995-01-01

    The computational fluid dynamics code, PARC3D, is tested to see if its use of non-physical artificial dissipation affects the accuracy of its results. This is accomplished by simulating a shock-laminar boundary layer interaction and several hypersonic flight conditions of the Pegasus(TM) launch vehicle using full artificial dissipation, low artificial dissipation, and the Engquist filter. Before the filter is applied to the PARC3D code, it is validated in one-dimensional and two-dimensional form in a MacCormack scheme against the Riemann and convergent duct problem. For this explicit scheme, the filter shows great improvements in accuracy and computational time as opposed to the nonfiltered solutions. However, for the implicit PARC3D code it is found that the best estimate of the Pegasus experimental heat fluxes and surface pressures is the simulation utilizing low artificial dissipation and no filter. The filter does improve accuracy over the artificially dissipative case but at a computational expense greater than that achieved by the low artificial dissipation case which has no computational time penalty and shows better results. For the shock-boundary layer simulation, the filter does well in terms of accuracy for a strong impingement shock but not as well for weaker shock strengths. Furthermore, for the latter problem the filter reduces the required computational time to convergence by 18.7 percent.

  18. Results of tests to determine the aerodynamic characteristics of two potential aeromaneuvering orbit-to-orbit shuttle (AMOOS) vehicle configurations in the NASA-Ames 3.5 foot hypersonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Ketter, F. C., Jr.

    1974-01-01

    An aerodynamic wind tunnel investigation was conducted in the NASA-Ames Research Center (ARC) 3.5-foot hypersonic facility to provide data for use in obtaining experimental force and static stability characteristics of two potential aeromaneuvering orbit-to-orbit shuttle (AMOOS) vehicle configurations. The experimental data were compared with the aerodynamic characteristics estimated using Newtonian theory, thus establishing the usefulness of these predictions. The candidate AMOOS configurations selected for the wind tunnel tests were the AMOOS 5B and HB configurations. Two flap configurations were tested for each candidate - a forward or compression surface flap and an aft or expansion flap. Photographs and sketches of the two configurations with different control surfaces are shown. It was determined that Newtonian theory generally predicted the aerodynamics of the 5B configuration with acceptable accuracy for all expansion flap deflections and for compression flap deflections less than or equal to 10 degrees. Flow separation upstream of large compression flap deflections was detected from the experimental data.

  19. The 2D and 3D hypersonic flows with unstructured meshes

    NASA Technical Reports Server (NTRS)

    Thareja, Rajiv

    1993-01-01

    Viewgraphs on 2D and 3D hypersonic flows with unstructured meshes are presented. Topics covered include: mesh generation, mesh refinement, shock-shock interaction, velocity contours, mesh movement, vehicle bottom surface, and adapted meshes.

  20. Recombination Catalysts for Hypersonic Fuels

    NASA Technical Reports Server (NTRS)

    Chinitz, W.

    1998-01-01

    The goal of commercially-viable access to space will require technologies that reduce propulsion system weight and complexity, while extracting maximum energy from the products of combustion. This work is directed toward developing effective nozzle recombination catalysts for the supersonic and hypersonic aeropropulsion engines used to provide such access to space. Effective nozzle recombination will significantly reduce rk=le length (hence, propulsion system weight) and reduce fuel requirements, further decreasing the vehicle's gross lift-off weight. Two such catalysts have been identified in this work, barium and antimony compounds, by developing chemical kinetic reaction mechanisms for these materials and determining the engine performance enhancement for a typical flight trajectory. Significant performance improvements are indicated, using only 2% (mole or mass) of these compounds in the combustor product gas.

  1. Analysis of various descent trajectories for a hypersonic-cruise, cold-wall research airplane

    NASA Technical Reports Server (NTRS)

    Lawing, P. L.

    1975-01-01

    The probable descent operating conditions for a hypersonic air-breathing research airplane were examined. Descents selected were cruise angle of attack, high dynamic pressure, high lift coefficient, turns, and descents with drag brakes. The descents were parametrically exercised and compared from the standpoint of cold-wall (367 K) aircraft heat load. The descent parameters compared were total heat load, peak heating rate, time to landing, time to end of heat pulse, and range. Trends in total heat load as a function of cruise Mach number, cruise dynamic pressure, angle-of-attack limitation, pull-up g-load, heading angle, and drag-brake size are presented.

  2. Development and validation of CNS (compressible Navier-Stokes) for hypersonic external flows

    NASA Technical Reports Server (NTRS)

    Flores, Jolen; Chow, Chuen-Yen; Ryan, James S.

    1989-01-01

    CNS, a new computational fluid dynamics procedure, has been developed to aid in hypersonic vehicle design. The code can be used to model the entire external flow around hypersonic vehicle shapes, from the captured shock at the nose to the beginning of the wake. Unlike space-marching codes, the technique allows axially separated flow regions to be modeled. Validation trials using sphere-cone data reveal good solution accuracy for the surface pressure and flowfield temperature.

  3. Trim drag reduction concepts for horizontal takeoff single-stage-to-Orbit vehicles

    NASA Technical Reports Server (NTRS)

    Shaughnessy, John D.; Gregory, Irene M.

    1991-01-01

    The results of a study to investigate concepts for minimizing trim drag of horizontal takeoff single-stage-to-orbit (SSTO) vehicles are presented. A generic hypersonic airbreathing conical configuration was used as the subject aircraft. The investigation indicates that extreme forward migration of the aerodynamic center as the vehicle accelerates to orbital velocities causes severe aerodynamic instability and trim moments that must be counteracted. Adequate stability can be provided by active control of elevons and rudder, but use of elevons to produce trim moments results in excessive trim drag and fuel consumption. To alleviate this problem, two solution concepts are examined. Active control of the center of gravity (COG) location to track the aerodynamic center decreases trim moment requirements, reduces elevon deflections, and leads to significant fuel savings. Active control of the direction of the thrust vector produces required trim moments, reduces elevon deflections, and also results in significant fuel savings. It is concluded that the combination of active flight control to provide stabilization, (COG) position control to minimize trim moment requirements, and thrust vectoring to generate required trim moments has the potential to significantly reduce fuel consumption during ascent to orbit of horizontal takeoff SSTO vehicles.

  4. An engineering code to analyze hypersonic thermal management systems

    NASA Technical Reports Server (NTRS)

    Vangriethuysen, Valerie J.; Wallace, Clark E.

    1993-01-01

    Thermal loads on current and future aircraft are increasing and as a result are stressing the energy collection, control, and dissipation capabilities of current thermal management systems and technology. The thermal loads for hypersonic vehicles will be no exception. In fact, with their projected high heat loads and fluxes, hypersonic vehicles are a prime example of systems that will require thermal management systems (TMS) that have been optimized and integrated with the entire vehicle to the maximum extent possible during the initial design stages. This will not only be to meet operational requirements, but also to fulfill weight and performance constraints in order for the vehicle to takeoff and complete its mission successfully. To meet this challenge, the TMS can no longer be two or more entirely independent systems, nor can thermal management be an after thought in the design process, the typical pervasive approach in the past. Instead, a TMS that was integrated throughout the entire vehicle and subsequently optimized will be required. To accomplish this, a method that iteratively optimizes the TMS throughout the vehicle will not only be highly desirable, but advantageous in order to reduce the manhours normally required to conduct the necessary tradeoff studies and comparisons. A thermal management engineering computer code that is under development and being managed at Wright Laboratory, Wright-Patterson AFB, is discussed. The primary goal of the code is to aid in the development of a hypersonic vehicle TMS that has been optimized and integrated on a total vehicle basis.

  5. Aerodynamic analysis of hypersonic waverider aircraft

    NASA Technical Reports Server (NTRS)

    Sandlin, Doral R.; Pessin, David N.

    1993-01-01

    The purpose of this study is to validate two existing codes used by the Systems Analysis Branch at NASA ARC, and to modify the codes so they can be used to generate and analyze waverider aircraft at on-design and off-design conditions. To generate waverider configurations and perform the on-design analysis, the appropriately named Waverider code is used. The Waverider code is based on the Taylor-Maccoll equations. Validation is accomplished via a comparison with previously published results. The Waverider code is modified to incorporate a fairing to close off the base area of the waverider configuration. This creates a more realistic waverider. The Hypersonic Aircraft Vehicle Optimization Code (HAVOC) is used to perform the off-design analysis of waverider configurations generated by the Waverider code. Various approximate analysis methods are used by HAVOC to predict the aerodynamic characteristics, which are validated via a comparison with experimental results from a hypersonic test model.

  6. Combined LAURA-UPS hypersonic solution procedure

    NASA Technical Reports Server (NTRS)

    Wood, William A.; Thompson, Richard A.

    1993-01-01

    A combined solution procedure for hypersonic flowfields around blunted slender bodies was implemented using a thin-layer Navier-Stokes code (LAURA) in the nose region and a parabolized Navier-Stokes code (UPS) on the after body region. Perfect gas, equilibrium air, and non-equilibrium air solutions to sharp cones and a sharp wedge were obtained using UPS alone as a preliminary step. Surface heating rates are presented for two slender bodies with blunted noses, having used LAURA to provide a starting solution to UPS downstream of the sonic line. These are an 8 deg sphere-cone in Mach 5, perfect gas, laminar flow at 0 and 4 deg angles of attack and the Reentry F body at Mach 20, 80,000 ft equilibrium gas conditions for 0 and 0.14 deg angles of attack. The results indicate that this procedure is a timely and accurate method for obtaining aerothermodynamic predictions on slender hypersonic vehicles.

  7. CFD on hypersonic flow geometries with aeroheating

    NASA Astrophysics Data System (ADS)

    Sohail, Muhammad Amjad; Chao, Yan; Hui, Zhang Hui; Ullah, Rizwan

    2012-11-01

    The hypersonic flowfield around a blunted cone and cone-flare exhibits some of the major features of the flows around space vehicles, e.g. a detached bow shock in the stagnation region and the oblique shock wave/boundary layer interaction at the cone-flare junction. The shock wave/boundary layer interaction can produce a region of separated flow. This phenomenon may occur, for example, at the upstream-facing corner formed by a deflected control surface on a hypersonic entry vehicle, where the length of separation has implications for control effectiveness. Computational fluid-dynamics results are presented to show the flowfield around a blunted cone and cone-flare configurations in hypersonic flow with separation. This problem is of particular interest since it features most of the aspects of the hypersonic flow around planetary entry vehicles. The region between the cone and the flare is particularly critical with respect to the evaluation of the surface pressure and heat flux with aeroheating. Indeed, flow separation is induced by the shock wave boundary layer interaction, with subsequent flow reattachment, that can dramatically enhance the surface heat transfer. The exact determination of the extension of the recirculation zone is a particularly delicate task for numerical codes. Laminar flow and turbulent computations have been carried out using a full Navier-Stokes solver, with freestream conditions provided by the experimental data obtained at Mach 6, 8, and 16.34 wind tunnel. The numerical results are compared with the measured pressure and surface heat flux distributions in the wind tunnel and a good agreement is found, especially on the length of the recirculation region and location of shock waves. The critical physics of entropy layer, boundary layers, boundary layers and shock wave interaction and flow behind shock are properly captured and elaborated.. Hypersonic flows are characterized by high Mach number and high total enthalpy. An elevated

  8. Design Processes and Criteria for the X-51A Flight Vehicle Airframe

    DTIC Science & Technology

    2007-11-01

    office [1]. The fundamental objective of the X-51A flight test program is to flight demonstrate the USAF Hypersonic Technology (HyTech) scramjet engine...external vehicle configuration utilizes a waverider-type aerodynamic mold line to maximize hypersonic lift to drag (L/D) ratios. The hypersonic cruise...capabilities in both 1) time critical, long range standoff response, 2) hypersonic reach/reconnaissance vehicles and 3) reusable on-demand access to

  9. Air-breathing adaptation in a marine Devonian lungfish

    PubMed Central

    Clement, Alice M.; Long, John A.

    2010-01-01

    Recent discoveries of tetrapod trackways in 395 Myr old tidal zone deposits of Poland (Niedźwiedzki et al. 2010 Nature 463, 43–48 (doi:10.1038/nature.08623)) indicate that vertebrates had already ventured out of the water and might already have developed some air-breathing capacity by the Middle Devonian. Air-breathing in lungfishes is not considered to be a shared specialization with tetrapods, but evolved independently. Air-breathing in lungfishes has been postulated as starting in Middle Devonian times (ca 385 Ma) in freshwater habitats, based on a set of skeletal characters involved in air-breathing in extant lungfishes. New discoveries described herein of the lungfish Rhinodipterus from marine limestones of Australia identifies the node in dipnoan phylogeny where air-breathing begins, and confirms that lungfishes living in marine habitats had also developed specializations to breathe air by the start of the Late Devonian (ca 375 Ma). While invasion of freshwater habitats from the marine realm was previously suggested to be the prime cause of aerial respiration developing in lungfishes, we believe that global decline in oxygen levels during the Middle Devonian combined with higher metabolic costs is a more likely driver of air-breathing ability, which developed in both marine and freshwater lungfishes and tetrapodomorph fishes such as Gogonasus. PMID:20147310

  10. X-43A Project Overview: Adventures in Hypersonics

    NASA Technical Reports Server (NTRS)

    Davis, Mark; Grindle, Laurie

    2007-01-01

    A viewgraph presentation describing the hypersonics program at NASA Dryden Flight Research Center is shown. The topics include: 1) X-43A Program Overview; 2) Vehicle Description; 3) Flight 1, MIB & Return to Flight; 4) Flight 2 and Results; 5) Flight 3 and Results; and 6) Concluding Remarks

  11. Carbon-carbon composites: Emerging materials for hypersonic flight

    NASA Technical Reports Server (NTRS)

    Maahs, Howard G.

    1989-01-01

    An emerging class of high temperature materials called carbon-carbon composites are being developed to help make advanced aerospace flight become a reality. Because of the high temperature strength and low density of carbon-carbon composites, aerospace engineers would like to use these materials in even more advanced applications. One application of considerable interest is as the structure of the aerospace vehicle itself rather than simply as a protective heat shield as on Space Shuttle. But suitable forms of these materials have yet to be developed. If this development can be successfully accomplished, advanced aerospace vehicles such as the National Aero-Space Plane (NASP) and other hypersonic vehicles will be closer to becoming a reality. A brief definition is given of C-C composites. Fabrication problems and oxidation protection concepts are examined. Applications of C-C composites in the Space Shuttle and in advanced hypersonic vehicles as well as other applications are briefly discussed.

  12. Nonequilibrium radiative hypersonic flow simulation

    NASA Astrophysics Data System (ADS)

    Shang, J. S.; Surzhikov, S. T.

    2012-08-01

    Nearly all the required scientific disciplines for computational hypersonic flow simulation have been developed on the framework of gas kinetic theory. However when high-temperature physical phenomena occur beneath the molecular and atomic scales, the knowledge of quantum physics and quantum chemical-physics becomes essential. Therefore the most challenging topics in computational simulation probably can be identified as the chemical-physical models for a high-temperature gaseous medium. The thermal radiation is also associated with quantum transitions of molecular and electronic states. The radiative energy exchange is characterized by the mechanisms of emission, absorption, and scattering. In developing a simulation capability for nonequilibrium radiation, an efficient numerical procedure is equally important both for solving the radiative transfer equation and for generating the required optical data via the ab-initio approach. In computational simulation, the initial values and boundary conditions are paramount for physical fidelity. Precise information at the material interface of ablating environment requires more than just a balance of the fluxes across the interface but must also consider the boundary deformation. The foundation of this theoretic development shall be built on the eigenvalue structure of the governing equations which can be described by Reynolds' transport theorem. Recent innovations for possible aerospace vehicle performance enhancement via an electromagnetic effect appear to be very attractive. The effectiveness of this mechanism is dependent strongly on the degree of ionization of the flow medium, the consecutive interactions of fluid dynamics and electrodynamics, as well as an externally applied magnetic field. Some verified research results in this area will be highlighted. An assessment of all these most recent advancements in nonequilibrium modeling of chemical kinetics, chemical-physics kinetics, ablation, radiative exchange

  13. Transition at hypersonic speeds

    NASA Technical Reports Server (NTRS)

    Morkovin, Mark V.

    1987-01-01

    Certain conjectures on the physics of instabilities in high-speed flows are discussed and the state of knowledge of hypersonic transition summarized. The case is made for an unpressured systematic research program in this area consisting of controlled microscopic experiments, theory, and numerical simulations.

  14. Analytical and computational investigations of a magnetohydrodynamics (MHD) energy-bypass system for supersonic gas turbine engines to enable hypersonic flight

    NASA Astrophysics Data System (ADS)

    Benyo, Theresa Louise

    Historically, the National Aeronautics and Space Administration (NASA) has used rocket-powered vehicles as launch vehicles for access to space. A familiar example is the Space Shuttle launch system. These vehicles carry both fuel and oxidizer onboard. If an external oxidizer (such as the Earth's atmosphere) is utilized, the need to carry an onboard oxidizer is eliminated, and future launch vehicles could carry a larger payload into orbit at a fraction of the total fuel expenditure. For this reason, NASA is currently researching the use of air-breathing engines to power the first stage of two-stage-to-orbit hypersonic launch systems. Removing the need to carry an onboard oxidizer leads also to reductions in total vehicle weight at liftoff. This in turn reduces the total mass of propellant required, and thus decreases the cost of carrying a specific payload into orbit or beyond. However, achieving hypersonic flight with air-breathing jet engines has several technical challenges. These challenges, such as the mode transition from supersonic to hypersonic engine operation, are under study in NASA's Fundamental Aeronautics Program. One propulsion concept that is being explored is a magnetohydrodynamic (MHD) energy- bypass generator coupled with an off-the-shelf turbojet/turbofan. It is anticipated that this engine will be capable of operation from takeoff to Mach 7 in a single flowpath without mode transition. The MHD energy bypass consists of an MHD generator placed directly upstream of the engine, and converts a portion of the enthalpy of the inlet flow through the engine into electrical current. This reduction in flow enthalpy corresponds to a reduced Mach number at the turbojet inlet so that the engine stays within its design constraints. Furthermore, the generated electrical current may then be used to power aircraft systems or an MHD accelerator positioned downstream of the turbojet. The MHD accelerator operates in reverse of the MHD generator, re-accelerating the

  15. Optical skin friction measurement technique in hypersonic wind tunnel

    NASA Astrophysics Data System (ADS)

    Chen, Xing; Yao, Dapeng; Wen, Shuai; Pan, Junjie

    2016-10-01

    Shear-sensitive liquid-crystal coatings (SSLCCs) have an optical characteristic that they are sensitive to the applied shear stress. Based on this, a novel technique is developed to measure the applied shear stress of the model surface regarding both its magnitude and direction in hypersonic flow. The system of optical skin friction measurement are built in China Academy of Aerospace Aerodynamics (CAAA). A series of experiments of hypersonic vehicle is performed in wind tunnel of CAAA. Global skin friction distribution of the model which shows complicated flow structures is discussed, and a brief mechanism analysis and an evaluation on optical measurement technique have been made.

  16. Hypersonic Shock/Boundary-Layer Interaction Database

    NASA Technical Reports Server (NTRS)

    Settles, G. S.; Dodson, L. J.

    1991-01-01

    Turbulence modeling is generally recognized as the major problem obstructing further advances in computational fluid dynamics (CFD). A closed solution of the governing Navier-Stokes equations for turbulent flows of practical consequence is still far beyond grasp. At the same time, the simplified models of turbulence which are used to achieve closure of the Navier-Stokes equations are known to be rigorously incorrect. While these models serve a definite purpose, they are inadequate for the general prediction of hypersonic viscous/inviscid interactions, mixing problems, chemical nonequilibria, and a range of other phenomena which must be predicted in order to design a hypersonic vehicle computationally. Due to the complexity of turbulence, useful new turbulence models are synthesized only when great expertise is brought to bear and considerable intellectual energy is expended. Although this process is fundamentally theoretical, crucial guidance may be gained from carefully-executed basic experiments. Following the birth of a new model, its testing and validation once again demand comparisons with data of unimpeachable quality. This report concerns these issues which arise from the experimental aspects of hypersonic modeling and represents the results of the first phase of an effort to develop compressible turbulence models.

  17. Advanced vehicle concepts systems and design analysis studies

    NASA Technical Reports Server (NTRS)

    Waters, Mark H.; Huynh, Loc C.

    1994-01-01

    The work conducted by the ELORET Institute under this Cooperative Agreement includes the modeling of hypersonic propulsion systems and the evaluation of hypersonic vehicles in general and most recently hypersonic waverider vehicles. This work in hypersonics was applied to the design of a two-stage to orbit launch vehicle which was included in the NASA Access to Space Project. Additional research regarded the Oblique All-Wing (OAW) Project at NASA ARC and included detailed configuration studies of OAW transport aircraft. Finally, work on the modeling of subsonic and supersonic turbofan engines was conducted under this research program.

  18. Hypersonic Research Vehicle (HRV) real-time flight test support feasibility and requirements study. Part 2: Remote computation support for flight systems functions

    NASA Technical Reports Server (NTRS)

    Rediess, Herman A.; Hewett, M. D.

    1991-01-01

    The requirements are assessed for the use of remote computation to support HRV flight testing. First, remote computational requirements were developed to support functions that will eventually be performed onboard operational vehicles of this type. These functions which either cannot be performed onboard in the time frame of initial HRV flight test programs because the technology of airborne computers will not be sufficiently advanced to support the computational loads required, or it is not desirable to perform the functions onboard in the flight test program for other reasons. Second, remote computational support either required or highly desirable to conduct flight testing itself was addressed. The use is proposed of an Automated Flight Management System which is described in conceptual detail. Third, autonomous operations is discussed and finally, unmanned operations.

  19. I(sup STAR), NASA's Next Step in Air-Breathing Propulsion for Space Access

    NASA Technical Reports Server (NTRS)

    Hutt, John J.; McArthur, Craig; Cook, Stephen (Technical Monitor)

    2001-01-01

    The United States' National Aeronautics and Space Administration (NASA) has established a strategic plan for future activities in space. A primary goal of this plan is to make drastic improvements in the cost and safety of earth to low-earth-orbit transportation. One approach to achieving this goal is through the development of highly reusable, highly reliable space transportation systems analogous to the commercial airline system. In the year 2000, NASA selected the Rocket Based Combined Cycle (RBCC) engine as the next logical step towards this goal. NASA will develop a complete flight-weight, pump-fed engine system under the Integrated System Test of an Airbreathing Rocket (I(sup STAR)) Project. The objective of this project is develop a reusable engine capable of self-powering a vehicle through the air-augmented rocket, ramjet and scramjet modes required in all RBCC based operational vehicle concepts. The project is currently approved and funded to develop the engine through ground test demonstration. Plans are in place to proceed with flight demonstration pending funding approval. The project is in formulation phase and the Preliminary Requirements Review has been completed. The engine system and vehicle have been selected at the conceptual level. The I(sup STAR) engine concept is based on an air-breathing flowpath downselected from three configurations evaluated in NASA's Advanced Reusable Technology contract. The selected flowpath features rocket thrust chambers integrated into struts separating modular flowpath ducts, a variable geometry inlet, and a thermally choked throat. The engine will be approximately 220 inches long and 79 inches wide and fueled with a hydrocarbon fuel using liquid oxygen as the primary oxidizer candidate. The primary concept for the pump turbine drive is pressure-fed catalyzed hydrogen peroxide. In order to control costs, the flight demonstration vehicle will be launched from a B-52 aircraft. The vehicle concept is based on the Air

  20. Hypersonic and planetary entry flight mechanics

    NASA Technical Reports Server (NTRS)

    Vinh, N. X.; Busemann, A.; Culp, R. D.

    1980-01-01

    The book treats hypersonic flight trajectories and atmospheric entry flight mechanics in light of their importance for space shuttle entry. Following a review of the structures of planetary atmospheres and aerodynamic forces, equations are derived for flight over a spherical planet, and the performance of long-range hypervelocity vehicles in extra-atmospheric flight is analyzed. Consideration is then given to vehicle trajectories in the powered and atmospheric reentry phases of flight, and several first-order solutions are derived for various planetary entry situations. The second-order theory of Loh for entry trajectories is presented along with the classical theories of Yaroshevskii and Chapman for entry into planetary atmospheres, and the thermal problems encountered in hypersonic flight are analyzed. A unified theory for entry into planetary atmospheres is then introduced which allows the performance of a general type of lifting vehicle to be studied, and applied to the analysis of orbit contraction due to atmospheric drag, flight with lift modulation and lateral maneuvers.

  1. Computational Fluid Dynamics Technology for Hypersonic Applications

    NASA Technical Reports Server (NTRS)

    Gnoffo, Peter A.

    2003-01-01

    Several current challenges in computational fluid dynamics and aerothermodynamics for hypersonic vehicle applications are discussed. Example simulations are presented from code validation and code benchmarking efforts to illustrate capabilities and limitations. Opportunities to advance the state-of-art in algorithms, grid generation and adaptation, and code validation are identified. Highlights of diverse efforts to address these challenges are then discussed. One such effort to re-engineer and synthesize the existing analysis capability in LAURA, VULCAN, and FUN3D will provide context for these discussions. The critical (and evolving) role of agile software engineering practice in the capability enhancement process is also noted.

  2. Magnetohydrodynamics Accelerator Research into Advanced Hypersonics (MARIAH). Part 2

    NASA Technical Reports Server (NTRS)

    Baughman, Jack A.; Micheletti, David A.; Nelson, Gordon L.; Simmons, Gloyd A.

    1997-01-01

    This report documents the activities, results, conclusions and recommendations of the Magnetohydrodynamics Accelerator Research Into Advanced Hypersonics (MARIAH) Project in which the use of magnetohydrodynamics (MHD) technology is investigated for its applicability to augment hypersonic wind tunnels. The long range objective of this investigation is to advance the development of ground test facilities to support the development of hypervelocity flight vehicles. The MHD accelerator adds kinetic energy directly to the wind tunnel working fluid, thereby increasing its Mach number to hypervelocity levels. Several techniques for MHD augmentation, as well as other physical characteristics of the process are studied to enhance the overall performance of hypersonic wind tunnel design. Specific recommendations are presented to improve the effectiveness of ground test facilities. The work contained herein builds on nearly four decades of research and experimentation by the aeronautics ground test and evaluation community, both foreign and domestic.

  3. Magnetohydrodynamics Accelerator Research Into Advanced Hypersonics (MARIAH). Part 1

    NASA Technical Reports Server (NTRS)

    Micheletti, David A.; Baughman, Jack A.; Nelson, Gordon L.; Simmons, Gloyd A.

    1997-01-01

    This report documents the activities, results, conclusions and recommendations of the Magnetohydrodynamics Accelerator Research Into Advanced Hypersonics (MARIAH) Project in which the use of magnetohydrodynamics (MHD) technology is investigated for its applicability to augment hypersonic wind tunnels. The long range objective of this investigation is to advance the development of ground test facilities to support the development of hypervelocity flight vehicles. The MHD accelerator adds kinetic energy directly to the wind tunnel working fluid, thereby increasing its Mach number to hypervelocity levels. Several techniques for MHD augmentation, as well as other physical characteristics of the process are studied to enhance the overall performance of hypersonic wind tunnel design. Specific recommendations are presented to improve the effectiveness of ground test facilities. The work contained herein builds on nearly four decades of research and experimentation by the aeronautics ground test and evaluation community, both foreign and domestic.

  4. Applications of underexpanded jets in hypersonic aerothermodynamics research

    NASA Astrophysics Data System (ADS)

    Riabov, Vladimir V.

    2012-11-01

    A method of underexpanded hypersonic viscous jets has been developed to acquire experimental aerodynamic data for simple-shape bodies (plates, wedges, disks, and others) in the transitional regime between free-molecular and continuum flow regimes. The kinetic, viscous, and nonequilibrium processes in the jets of He, Ar, N2, and CO2 under various experimental conditions have been analyzed by asymptotic methods and numerical techniques. Fundamental laws for the aerodynamic characteristics and similarity parameters are revealed. In the case of hypersonic stabilization, the Reynolds number and temperature factor are the main similarity parameters. This research has discovered those conditions, which allow the significant influence of other parameters (specific heat ratio, viscosity parameter, Mach number). The acquired data could be used effectively for research and prediction of aerodynamic characteristics of vehicles during hypersonic flights under the rarefied atmospheric conditions of Earth, Mars, Venus, and other planets.

  5. Wind Tunnel Test of Mach 5 Class Hypersonic Airplane

    NASA Astrophysics Data System (ADS)

    Nakatani, Hiroki; Taguchi, Hideyuki; Fujita, Kazuhisa; Shindo, Shigemi; Honami, Shinji

    JAXA is currently performing studies on a Hypersonic Turbojet Experimental Vehicle, which involve a hypersonic flight test of a Small Pre-cooled Turbojet Engine. The aerodynamic performance of this airplane was examined at the JAXA hypersonic, supersonic, and transonic wind tunnel facilities. The 6-degrees-of-freedom forces and pressure distribution around the model were measured and evaluated. This airplane satisfies the lift-to-drag ratio requirement for a flight test at Mach 5. In addition, the results indicate that this airplane has longitudinal and directional static stability if the moment reference point is x/l smaller than 0.35. A separation occurs at the external expanding nozzle. Therefore, a redesign is necessary to solve these problems.

  6. Three-Dimensional Aeroelastic and Aerothermoelastic Behavior in Hypersonic Flow

    NASA Technical Reports Server (NTRS)

    McNamara, Jack J.; Friedmann, Peretz P.; Powell, Kenneth G.; Thuruthimattam, Biju J.; Bartels, Robert E.

    2005-01-01

    The aeroelastic and aerothermoelastic behavior of three-dimensional configurations in hypersonic flow regime are studied. The aeroelastic behavior of a low aspect ratio wing, representative of a fin or control surface on a generic hypersonic vehicle, is examined using third order piston theory, Euler and Navier-Stokes aerodynamics. The sensitivity of the aeroelastic behavior generated using Euler and Navier-Stokes aerodynamics to parameters governing temporal accuracy is also examined. Also, a refined aerothermoelastic model, which incorporates the heat transfer between the fluid and structure using CFD generated aerodynamic heating, is used to examine the aerothermoelastic behavior of the low aspect ratio wing in the hypersonic regime. Finally, the hypersonic aeroelastic behavior of a generic hypersonic vehicle with a lifting-body type fuselage and canted fins is studied using piston theory and Euler aerodynamics for the range of 2.5 less than or equal to M less than or equal to 28, at altitudes ranging from 10,000 feet to 80,000 feet. This analysis includes a study on optimal mesh selection for use with Euler aerodynamics. In addition to the aeroelastic and aerothermoelastic results presented, three time domain flutter identification techniques are compared, namely the moving block approach, the least squares curve fitting method, and a system identification technique using an Auto-Regressive model of the aeroelastic system. In general, the three methods agree well. The system identification technique, however, provided quick damping and frequency estimations with minimal response record length, and therefore o ers significant reductions in computational cost. In the present case, the computational cost was reduced by 75%. The aeroelastic and aerothermoelastic results presented illustrate the applicability of the CFL3D code for the hypersonic flight regime.

  7. Hypersonic airframe structures: Technology needs and flight test requirements

    NASA Technical Reports Server (NTRS)

    Stone, J. E.; Koch, L. C.

    1979-01-01

    Hypersonic vehicles, that may be produced by the year 2000, were identified. Candidate thermal/structural concepts that merit consideration for these vehicles were described. The current status of analytical methods, materials, manufacturing techniques, and conceptual developments pertaining to these concepts were reviewed. Guidelines establishing meaningful technology goals were defined and twenty-eight specific technology needs were identified. The extent to which these technology needs can be satisfied, using existing capabilities and facilities without the benefit of a hypersonic research aircraft, was assessed. The role that a research aircraft can fill in advancing this technology was discussed and a flight test program was outlined. Research aircraft thermal/structural design philosophy was also discussed. Programs, integrating technology advancements with the projected vehicle needs, were presented. Program options were provided to reflect various scheduling and cost possibilities.

  8. Innovative Airbreathing Propulsion Concepts for Access to Space

    NASA Technical Reports Server (NTRS)

    Whitlow, Jr., Woodrow; Blech, Richard A.; Blankson, Isaiah M.

    2001-01-01

    This paper will present technologies and concepts for novel aeropropulsion systems. These technologies will enhance the safety of operations, reduce life cycle costs, and contribute to reduced costs of air travel and access to space. One of the goals of the NASA program is to reduce the carbon-dioxide emissions of aircraft engines. Engine concepts that use highly efficient fuel cell/electric drive technologies in hydrogen-fueled engines will be presented in the proposed paper. Carbon-dioxide emissions will be eliminated by replacing hydrocarbon fuel with hydrogen, and reduce NOx emissions through better combustion process control. A revolutionary exoskeletal engine concept, in which the engine drum is rotated, will be shown. This concept has the potential to allow a propulsion system that can be used for subsonic through hypersonic flight. Dual fan concepts that have ultra-high bypass ratios, low noise, and low drag will be presented. Flow-controlled turbofans and control-configured turbofans also will be discussed. To increase efficiency, a system of microengines distributed along lifting surfaces and on the fuselage is being investigated. This concept will be presented in the paper. Small propulsion systems for affordable, safe personal transportation vehicles will be discussed. These low-oil/oilless systems use technologies that enable significant cost and weight reductions. Pulse detonation engine-based hybrid-cycle and combined-cycle propulsion systems for aviation and space access will be presented.

  9. The potential for a new era of supersonic and hypersonic aviation

    NASA Technical Reports Server (NTRS)

    Harris, Roy V.

    1990-01-01

    A new era of supersonic and hypersonic aviation is envisioned. The potential for supersonic and hypersonic flight vehicles in this new era is analyzed. Technology challenges that must be met in order to bring in this new era of flight are discussed. The current technical status and future potential are cited in the areas of aerodynamics, propulsion, and structural materials. A next major step in the development of high-speed air transportation is suggested.

  10. Joint USAF/NASA hypersonic research aircraft study

    NASA Technical Reports Server (NTRS)

    Kirkham, F. S.; Jones, R. A.; Buck, M. L.; Zima, W. P.

    1975-01-01

    A joint USAF/NASA study has developed a conceptual design for a new high-speed research airplane (X-24C) and identified candidate flight research experiments in the Mach 3 to 6 speed range. Four major categories of high priority research experiments are described as well as the X-24C design concept. The vehicle, a rocket-boosted, delta planform aircraft, is air launched from a B-52 and is capable of forty seconds of rocket cruise at Mach 6 with a research scramjet. Research provisions include a dedicated 10-foot long research experiments section, removable fins and strakes, and provisions for testing integrated airbreathing propulsion systems.

  11. Aerodynamic Characteristics of Two Waverider-Derived Hypersonic Cruise Configurations

    NASA Technical Reports Server (NTRS)

    Cockrell, Charles E., Jr.; Huebner, Lawrence D.; Finley, Dennis B.

    1996-01-01

    An evaluation was made on the effects of integrating the required aircraft components with hypersonic high-lift configurations known as waveriders to create hypersonic cruise vehicles. Previous studies suggest that waveriders offer advantages in aerodynamic performance and propulsion/airframe integration (PAI) characteristics over conventional non-waverider hypersonic shapes. A wind-tunnel model was developed that integrates vehicle components, including canopies, engine components, and control surfaces, with two pure waverider shapes, both conical-flow-derived waveriders for a design Mach number of 4.0. Experimental data and limited computational fluid dynamics (CFD) solutions were obtained over a Mach number range of 1.6 to 4.63. The experimental data show the component build-up effects and the aerodynamic characteristics of the fully integrated configurations, including control surface effectiveness. The aerodynamic performance of the fully integrated configurations is not comparable to that of the pure waverider shapes, but is comparable to previously tested hypersonic models. Both configurations exhibit good lateral-directional stability characteristics.

  12. Structural weight analysis of hypersonic aircraft

    NASA Technical Reports Server (NTRS)

    Ardema, M. D.

    1972-01-01

    The weights of major structural components of hypersonic, liquid hydrogen fueled aircraft are estimated and discussed. The major components are the body structure, body thermal protection system tankage and wing structure. The method of estimating body structure weight is presented in detail while the weights of the other components are estimated by methods given in referenced papers. Two nominal vehicle concepts are considered. The advanced concept employs a wing-body configuration and hot structure with a nonintegral tank, while the potential concept employs an all body configuration and cold, integral pillow tankage structure. Characteristics of these two concepts are discussed and parametric data relating their weight fractions to variations in vehicle shape and size design criteria and mission requirements, and structural arrangement are presented. Although the potential concept is shown to have a weight advantage over the advanced, it involves more design uncertainties since it is farther removed in design from existing aircraft.

  13. Feasibility study of air-breathing turbo-engines for horizontal take-off and landing space plane

    SciTech Connect

    Minoda, M.; Sakata, K.; Tamaki, T.; Saitoh, T.; Yasuda, A.

    1989-01-01

    Various concepts of air-breathing engines (ABEs) that could be used for a horizontal take-off and landing SSTO vehicle are investigated. The performances (with respect to thrust and the specific fuel consumption) of turboengines based on various technologies, including a turbojet with and without afterburner (TJ), turboramjet, and air-turbo-ram jet engines are compared. The mission capabilities of these ABEs for SSTO and TSTO vehicles is examined in terms of the ratio of the effective remaining weight (i.e., the weight on the orbit) to the take-off gross weight, using two-dimensional flight analysis. It was found that the dry TJ with the turbine inlet temperature 2000 C is one of the most promising candidates for the propulsion system of the SSTO vehicle, because of its small weight and high specific impulse. 6 refs.

  14. Application of CFD to a generic hypersonic flight research study

    NASA Technical Reports Server (NTRS)

    Green, Michael J.; Lawrence, Scott L.; Dilley, Arthur D.; Hawkins, Richard W.; Walker, Mary M.; Oberkampf, William L.

    1993-01-01

    Computational analyses have been performed for the initial assessment of flight research vehicle concepts that satisfy requirements for potential hypersonic experiments. Results were obtained from independent analyses at NASA Ames, NASA Langley, and Sandia National Labs, using sophisticated time-dependent Navier-Stokes and parabolized Navier-Stokes methods. Careful study of a common problem consisting of hypersonic flow past a slightly blunted conical forebody was undertaken to estimate the level of uncertainty in the computed results, and to assess the capabilities of current computational methods for predicting boundary-layer transition onset. Results of this study in terms of surface pressure and heat transfer comparisons, as well as comparisons of boundary-layer edge quantities and flow-field profiles are presented here. Sensitivities to grid and gas model are discussed. Finally, representative results are presented relating to the use of Computational Fluid Dynamics in the vehicle design and the integration/support of potential experiments.

  15. Research on hypersonic aircraft using pre-cooled turbojet engines

    NASA Astrophysics Data System (ADS)

    Taguchi, Hideyuki; Kobayashi, Hiroaki; Kojima, Takayuki; Ueno, Atsushi; Imamura, Shunsuke; Hongoh, Motoyuki; Harada, Kenya

    2012-04-01

    Systems analysis of a Mach 5 class hypersonic aircraft is performed. The aircraft can fly across the Pacific Ocean in 2 h. A multidisciplinary optimization program for aerodynamics, structure, propulsion, and trajectory is used in the analysis. The result of each element model is improved using higher accuracy analysis tools. The aerodynamic performance of the hypersonic aircraft is examined through hypersonic wind tunnel tests. A thermal management system based on the data of the wind tunnel tests is proposed. A pre-cooled turbojet engine is adopted as the propulsion system for the hypersonic aircraft. The engine can be operated continuously from take-off to Mach 5. This engine uses a pre-cooling cycle using cryogenic liquid hydrogen. The high temperature inlet air of hypersonic flight would be cooled by the same liquid hydrogen used as fuel. The engine is tested under sea level static conditions. The engine is installed on a flight test vehicle. Both liquid hydrogen fuel and gaseous hydrogen fuel are supplied to the engine from a tank and cylinders installed within the vehicle. The designed operation of major components of the engine is confirmed. A large amount of liquid hydrogen is supplied to the pre-cooler in order to make its performance sufficient for Mach 5 flight. Thus, fuel rich combustion is adopted at the afterburner. The experiments are carried out under the conditions that the engine is mounted upon an experimental airframe with both set up either horizontally or vertically. As a result, the operating procedure of the pre-cooled turbojet engine is demonstrated.

  16. Hypersonic Wind Tunnel Test of a Flare-type Membrane Aeroshell for Atmospheric Entry Capsules

    NASA Astrophysics Data System (ADS)

    Yamada, Kazuhiko; Koyama, Masashi; Kimura, Yusuke; Suzuki, Kojiro; Abe, Takashi; Koichi Hayashi, A.

    A flexible aeroshell for atmospheric entry vehicles has attracted attention as an innovative space transportation system. In this study, hypersonic wind tunnel tests were carried out to investigate the behavior, aerodynamic characteristics and aerodynamic heating environment in hypersonic flow for a previously developed capsule-type vehicle with a flare-type membrane aeroshell made of ZYLON textile sustained by a rigid torus frame. Two different models with different flare angles (45º and 60º) were tested to experimentally clarify the effect of flare angle. Results indicate that flare angle of aeroshell has significant and complicate effect on flow field and aerodynamic heating in hypersonic flow at Mach 9.45 and the flare angle is very important parameter for vehicle design with the flare-type membrane aeroshell.

  17. Permanent-Change Thermal Paints for Hypersonic Flight-Test

    DTIC Science & Technology

    2010-09-24

    initial test case for post recovery handling and analysis of the painted surfaces. • Investigate the bonding of the paints to a range of materials and...heating during flight and can quickly reach elevated temperature. All engineering materials degrade in mechanical performance at elevated temperature...and this can be severe for the candidate materials and conditions for the airframes of hypersonic flight-test vehicles. To maintain structural mass

  18. Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Technology Development Overview

    NASA Technical Reports Server (NTRS)

    Hughes, Stephen J.; Cheatwood, F. McNeil; Calomino, Anthony M.; Wright, Henry S.; Wusk, Mary E.; Hughes, Monica F.

    2013-01-01

    The successful flight of the Inflatable Reentry Vehicle Experiment (IRVE)-3 has further demonstrated the potential value of Hypersonic Inflatable Aerodynamic Decelerator (HIAD) technology. This technology development effort is funded by NASA's Space Technology Mission Directorate (STMD) Game Changing Development Program (GCDP). This paper provides an overview of a multi-year HIAD technology development effort, detailing the projects completed to date and the additional testing planned for the future.

  19. Hypersonic aircraft design

    NASA Technical Reports Server (NTRS)

    Alkamhawi, Hani; Greiner, Tom; Fuerst, Gerry; Luich, Shawn; Stonebraker, Bob; Wray, Todd

    1990-01-01

    A hypersonic aircraft is designed which uses scramjets to accelerate from Mach 6 to Mach 10 and sustain that speed for two minutes. Different propulsion systems were considered and it was decided that the aircraft would use one full scale turbofan-ramjet. Two solid rocket boosters were added to save fuel and help the aircraft pass through the transonic region. After considering aerodynamics, aircraft design, stability and control, cooling systems, mission profile, and landing systems, a conventional aircraft configuration was chosen over that of a waverider. The conventional design was chosen due to its landing characteristics and the relative expense compared to the waverider. Fuel requirements and the integration of the engine systems and their inlets are also taken into consideration in the final design. A hypersonic aircraft was designed which uses scramjets to accelerate from Mach 6 to Mach 10 and sustain that speed for two minutes. Different propulsion systems were considered and a full scale turbofan-ramjet was chosen. Two solid rocket boosters were added to save fuel and help the aircraft pass through the transonic reqion. After the aerodynamics, aircraft design, stability and control, cooling systems, mission profile, landing systems, and their physical interactions were considered, a conventional aircraft configuration was chosen over that of a waverider. The conventional design was chosen due to its landing characteristics and the relative expense compared to the waverider. Fuel requirements and the integration of the engine systems and their inlets were also considered in the designing process.

  20. Hypersonic transport aircraft

    NASA Technical Reports Server (NTRS)

    1987-01-01

    A hypersonic transport aircraft design project was selected as a result of interactions with NASA Lewis Research Center personnel and fits the Presidential concept of the Orient Express. The Graduate Teaching Assistant (GTA) and an undergraduate student worked at the NASA Lewis Research Center during the 1986 summer conducting a literature survey, and relevant literature and useful software were collected. The computer software was implemented in the Computer Aided Design Laboratory of the Mechanical and Aerospace Engineering Department. In addition to the lectures by the three instructors, a series of guest lectures was conducted. The first of these lectures 'Anywhere in the World in Two Hours' was delivered by R. Luidens of NASA Lewis Center. In addition, videotaped copies of relevant seminars obtained from NASA Lewis were also featured. The first assignment was to individually research and develop the mission requirements and to discuss the findings with the class. The class in consultation with the instructors then developed a set of unified mission requirements. Then the class was divided into three design groups (1) Aerodynamics Group, (2) Propulsion Group, and (3) Structures and Thermal Analyses Group. The groups worked on their respective design areas and interacted with each other to finally come up with an integrated conceptual design. The three faculty members and the GTA acted as the resource persons for the three groups and aided in the integration of the individual group designs into the final design of a hypersonic aircraft.

  1. Special Course on Aerothermodynamics of Hypersonic Vehicles

    DTIC Science & Technology

    1989-06-01

    Etude des propriftis thermodynamiques et de transport dans le plasmas thermiques A l’lquilibre et hors d’iquilibre thermodynamiques : Applications aut...Tannehill J.G. Marvin A new PNS Code for Chemical NonequitibriumFlows AIAA Paper 87- 0284 [931 S. Pn.franel Propriftisa thermodynamique . et de transport

  2. Galerkin CFD solvers for use in a multi-disciplinary suite for modeling advanced flight vehicles

    NASA Astrophysics Data System (ADS)

    Moffitt, Nicholas J.

    This work extends existing Galerkin CFD solvers for use in a multi-disciplinary suite. The suite is proposed as a means of modeling advanced flight vehicles, which exhibit strong coupling between aerodynamics, structural dynamics, controls, rigid body motion, propulsion, and heat transfer. Such applications include aeroelastics, aeroacoustics, stability and control, and other highly coupled applications. The suite uses NASA STARS for modeling structural dynamics and heat transfer. Aerodynamics, propulsion, and rigid body dynamics are modeled in one of the five CFD solvers below. Euler2D and Euler3D are Galerkin CFD solvers created at OSU by Cowan (2003). These solvers are capable of modeling compressible inviscid aerodynamics with modal elastics and rigid body motion. This work reorganized these solvers to improve efficiency during editing and at run time. Simple and efficient propulsion models were added, including rocket, turbojet, and scramjet engines. Viscous terms were added to the previous solvers to create NS2D and NS3D. The viscous contributions were demonstrated in the inertial and non-inertial frames. Variable viscosity (Sutherland's equation) and heat transfer boundary conditions were added to both solvers but not verified in this work. Two turbulence models were implemented in NS2D and NS3D: Spalart-Allmarus (SA) model of Deck, et al. (2002) and Menter's SST model (1994). A rotation correction term (Shur, et al., 2000) was added to the production of turbulence. Local time stepping and artificial dissipation were adapted to each model. CFDsol is a Taylor-Galerkin solver with an SA turbulence model. This work improved the time accuracy, far field stability, viscous terms, Sutherland?s equation, and SA model with NS3D as a guideline and added the propulsion models from Euler3D to CFDsol. Simple geometries were demonstrated to utilize current meshing and processing capabilities. Air-breathing hypersonic flight vehicles (AHFVs) represent the ultimate

  3. Technical Seminar: Exploring Hypersonic Flow

    NASA Video Gallery

    NASA Aeronautics is developing a method for 2D and 3D imaging of hypersonic flows, called Nitric Oxide Planar Laser-Induced Fluorescence (NO-PLIF). NO-PLIF has been used to study basic transition f...

  4. Electron-Beam Diagnostic Methods for Hypersonic Flow Diagnostics

    NASA Technical Reports Server (NTRS)

    1994-01-01

    The purpose of this work was the evaluation of the use of electron-bean fluorescence for flow measurements during hypersonic flight. Both analytical and numerical models were developed in this investigation to evaluate quantitatively flow field imaging concepts based upon the electron beam fluorescence technique for use in flight research and wind tunnel applications. Specific models were developed for: (1) fluorescence excitation/emission for nitrogen, (2) rotational fluorescence spectrum for nitrogen, (3) single and multiple scattering of electrons in a variable density medium, (4) spatial and spectral distribution of fluorescence, (5) measurement of rotational temperature and density, (6) optical filter design for fluorescence imaging, and (7) temperature accuracy and signal acquisition time requirements. Application of these models to a typical hypersonic wind tunnel flow is presented. In particular, the capability of simulating the fluorescence resulting from electron impact ionization in a variable density nitrogen or air flow provides the capability to evaluate the design of imaging instruments for flow field mapping. The result of this analysis is a recommendation that quantitative measurements of hypersonic flow fields using electron-bean fluorescence is a tractable method with electron beam energies of 100 keV. With lower electron energies, electron scattering increases with significant beam divergence which makes quantitative imaging difficult. The potential application of the analytical and numerical models developed in this work is in the design of a flow field imaging instrument for use in hypersonic wind tunnels or onboard a flight research vehicle.

  5. On Challenges for Hypersonic Turbulent Simulations

    SciTech Connect

    Yee, H C; Sjogreen, B

    2009-01-14

    This short note discusses some of the challenges for design of suitable spatial numerical schemes for hypersonic turbulent flows, including combustion, and thermal and chemical nonequilibrium flows. Often, hypersonic turbulent flows in re-entry space vehicles and space physics involve mixed steady strong shocks and turbulence with unsteady shocklets. Material mixing in combustion poses additional computational challenges. Proper control of numerical dissipation in numerical methods beyond the standard shock-capturing dissipation at discontinuities is an essential element for accurate and stable simulations of the subject physics. On one hand, the physics of strong steady shocks and unsteady turbulence/shocklet interactions under the nonequilibrium environment is not well understood. On the other hand, standard and newly developed high order accurate (fourth-order or higher) schemes were developed for homogeneous hyperbolic conservation laws and mixed hyperbolic and parabolic partial differential equations (PDEs) (without source terms). The majority of finite rate chemistry and thermal nonequilibrium simulations employ methods for homogeneous time-dependent PDEs with a pointwise evaluation of the source terms. The pointwise evaluation of the source term might not be the best choice for stability, accuracy and minimization of spurious numerics for the overall scheme.

  6. Development of Supersonic Vehicle for Demonstration of a Precooled Turbojet Engine

    NASA Astrophysics Data System (ADS)

    Sawai, Shujiro; Fujita, Kazuhisa; Kobayashi, Hiroaki; Sakai, Shin'ichiro; Bando, Nobutaka; Kadooka, Shouhei; Tsuboi, Nobuyuki; Miyaji, Koji; Uchiyama, Taku; Hashimoto, Tatsuaki

    JAXA is developing Mach 5 hypersonic turbojet engine technology that can be applied in a future hypersonic transport. Now, Jet Engine Technology Research Center of JAXA conducts the experimental study using a 1 / 10 scale-model engine. In parallel to engine development activities, a new supersonic flight-testing vehicle for the hypersonic turbojet engine is under development since 2004. In this paper, the system configuration of the flight-testing vehicle is outlined and development status is reported.

  7. Optimization of Air-Breathing Propulsion Engine Concepts

    NASA Technical Reports Server (NTRS)

    Patnaik, Surya N.; Hopkins, Dale A.

    1997-01-01

    Air-breathing propulsion engines play an important role in the development of both civil and military aircraft Design optimization of such engines can lead to higher power, or more thrust for less fuel consumption. A multimission propulsion engine design can be modeled mathematically as a multivariable global optimization problem, with a sequence of subproblems, which are specific to the mission events defined through Mach number, altitude, and power setting combinations.

  8. Hypersonic reconnaissance aircraft

    NASA Technical Reports Server (NTRS)

    Bulk, Tim; Chiarini, David; Hill, Kevin; Kunszt, Bob; Odgen, Chris; Truong, Bon

    1992-01-01

    A conceptual design of a hypersonic reconnaissance aircraft for the U.S. Navy is discussed. After eighteen weeks of work, a waverider design powered by two augmented turbofans was chosen. The aircraft was designed to be based on an aircraft carrier and to cruise 6,000 nautical miles at Mach 4;80,000 feet and above. As a result the size of the aircraft was only allowed to have a length of eighty feet, fifty-two feet in wingspan, and roughly 2,300 square feet in planform area. Since this is a mainly cruise aircraft, sixty percent of its 100,000 pound take-off weight is JP fuel. At cruise, the highest temperature that it will encounter is roughly 1,100 F, which can be handled through the use of a passive cooling system.

  9. Assessment of flying-quality criteria for air-breathing aerospacecraft

    NASA Technical Reports Server (NTRS)

    Mcruer, Duane T.; Myers, Thomas T.; Hoh, Roger H.; Ashkenas, Irving L.; Johnston, Donald E.

    1992-01-01

    A study of flying quality requirements for air breathing aerospacecraft gives special emphasis to the unusual operational requirements and characteristics of these aircraft, including operation at hypersonic speed. The report considers distinguishing characteristics of these vehicles, including dynamic deficiencies and their implications for control. Particular emphasis is given to the interaction of the airframe and propulsion system, and the requirements for dynamic systems integration. Past operational missions are reviewed to define tasks and maneuvers to be considered for this class of aircraft. Areas of special concern with respect to vehicle dynamics and control are identified. Experience with the space shuttle orbiter is reviewed with respect to flight control system mechanization and flight experience in approach and landing flying qualities for the National Aerospace Plane (NASP).

  10. The Lightcraft project: Flight technology for a hypersonic mass transit system

    NASA Technical Reports Server (NTRS)

    Myrabo, Leik; Bouchard, Kenneth

    1992-01-01

    Rensselaer Polytechnic Institute has been developing transatmospheric 'Lightcraft' technology aimed at creating an efficient, economically affordable, hypersonic mass transportation system. The system utilizes laser-energized airbreathing engines to accelerate minimum-volume passenger capsules. The system gains a high level of reliability by using remote 'centralized' space power sources, e.g., satellite solar power stations. The most critical portion of the Lightcraft's acceleration trajectory involves flight propulsion at hypersonic velocities within the Earth's atmosphere, using a 'Magneto-Hydro-Dynamic (MHD) Fanjet' mode. Of all the propulsion modes proposed for the Lightcraft's combined-cycle engine, the MHD-Fanjet mode has received the least critical inquiry, largely because of its complexity. During the 1991-1992 academic year, Rensselaer's ADP teams produced a detailed conceptual design for the MHD-Fanjet engine, including the specific details of its integration with the other three propulsive modes. To facilitate this process, students built a full-scale mockup of a 1/12th section of this annular engine, complete with a working model of the shroud translation system. The class also made preliminary design calculations for the double-dipole, 'cuspfield' superconducting magnets that provide the external magnetic field needed by the MHD air accelerator, as well as for an onboard microwave power system to enhance the electrical conductivity of the air plasma working fluid. In addition, a large hypersonic model of the MHD accelerator was designed for future tests in RPI's Hypersonic Shock Tunnel in order to validate present analytical performance models. Another group continued design work on a full-sized prototype of a one-person 'Mercury Lightcraft' (a transatmospheric flight simulator), with major emphasis on the detailed design of the major structure, robotic landing gear, and exterior aeroshell.

  11. Optical Window Materials For Hypersonic Flow

    NASA Astrophysics Data System (ADS)

    Au, Robert H.

    1989-09-01

    Optical window materials were investigated for infrared sensor systems used in observing ground targets from a hypersonic-glide vehicle. The equilibrium temperature of the window in the glide region depends on the emissivity and varied between 1,370 and 2,250 K. The high temperatures showed that a protective cover over the window is required during the entire glide region of the trajectory. Ejection of the window cover at 70-kft altitude in the terminal region was assumed, resulting in maximum window temperatures of 565 K and 592 K for magnesium oxide and diamond windows, respectively, both 0.8-in thick. The window temperatures for germanium and sapphire were also calculated. Thermal shock, thermal expansion, the effects of the window radiation on the infrared detectors and methods to reduce the hot window problem were examined.

  12. Review and assessment of turbulence models for hypersonic flows

    NASA Astrophysics Data System (ADS)

    Roy, Christopher J.; Blottner, Frederick G.

    2006-10-01

    Accurate aerodynamic prediction is critical for the design and optimization of hypersonic vehicles. Turbulence modeling remains a major source of uncertainty in the computational prediction of aerodynamic forces and heating for these systems. The first goal of this article is to update the previous comprehensive review of hypersonic shock/turbulent boundary-layer interaction experiments published in 1991 by Settles and Dodson (Hypersonic shock/boundary-layer interaction database. NASA CR 177577, 1991). In their review, Settles and Dodson developed a methodology for assessing experiments appropriate for turbulence model validation and critically surveyed the existing hypersonic experiments. We limit the scope of our current effort by considering only two-dimensional (2D)/axisymmetric flows in the hypersonic flow regime where calorically perfect gas models are appropriate. We extend the prior database of recommended hypersonic experiments (on four 2D and two 3D shock-interaction geometries) by adding three new geometries. The first two geometries, the flat plate/cylinder and the sharp cone, are canonical, zero-pressure gradient flows which are amenable to theory-based correlations, and these correlations are discussed in detail. The third geometry added is the 2D shock impinging on a turbulent flat plate boundary layer. The current 2D hypersonic database for shock-interaction flows thus consists of nine experiments on five different geometries. The second goal of this study is to review and assess the validation usage of various turbulence models on the existing experimental database. Here we limit the scope to one- and two-equation turbulence models where integration to the wall is used (i.e., we omit studies involving wall functions). A methodology for validating turbulence models is given, followed by an extensive evaluation of the turbulence models on the current hypersonic experimental database. A total of 18 one- and two-equation turbulence models are reviewed

  13. Hypersonic research engine project. Phase 2: Some combustor test results of NASA aerothermodynamic integration model

    NASA Technical Reports Server (NTRS)

    Sun, Y. H.; Gaede, A. E.; Sainio, W. C.

    1975-01-01

    Combustor test results of the NASA Aerothermodynamic Integration Model are presented of a ramjet engine developed for operation between Mach 3 and 8. Ground-based and flight experiments which provide the data required to advance the technology of hypersonic air-breathing propulsion systems as well as to evaluate facility and testing techniques are described. The engine was tested with synthetic air at Mach 5, 6, and 7. The hydrogen fuel was heated to 1500 R prior to injection to simulate a regeneratively cooled system. Combustor efficiencies up to 95 percent at Mach 6 were achieved. Combustor process in terms of effectiveness, pressure integral factor, total pressure recovery and Crocco's pressure-area relationship are presented and discussed. Interactions between inlet-combustor, combustor stages, combustor-nozzle, and the effects of altitude, combustor step, and struts are observed and analyzed.

  14. Airbreathing nuclear propulsion: A new look

    NASA Technical Reports Server (NTRS)

    Rom, F. E.

    1971-01-01

    Nuclear-powered air-cushion vehicles using lightweight aircraft-type nuclear powerplants show promise of carrying transoceanic cargo at cost-per-metric-ton-kilometer (cost-per-ton-n mi) rates comparable to railroad rates. These rates are independent of the distance traveled. Cargo rates for nonstop distances of 4000 n mi are expected to be less than one-half those for similar fossil-fueled air-cushion vehicles. For 6000-n mi nonstop distances, the rates are expected to be less than one-sixth as much. There are no fundamental technical reasons why subsonic nuclear aircraft cannot be made to fly successfully if the gross weight is over 1 million lb. Public safety of airborne nuclear powerplants is receiving the greatest attention in low-level experimental and analytical investigations. Idealized model containment vessels which have been impacted on reinforced concrete showed no leaks after impact at velocities to 400 mph. The experiments indicate feasibility of impacting at speeds over 600 mph with no leaks.

  15. Analysis of Pulsed Laser-Generated Impulse in AN Advanced Airbreathing Thruster.

    NASA Astrophysics Data System (ADS)

    Richard, Jacques Constant

    This thesis describes the study of an advanced beam-powered propulsion system, called an External Radiation -Heated (ERH) thruster. The repetitively-pulsed, airbreathing engine develops thrust by expanding high pressure, radiation -heated gas over an annular shroud surface. The blast waves are generated by laser radiation heating of air using Laser Supported Detonation (LSD) waves. The phenomenology of LSD waves will be described in detail, as will the blast waves and resultant impulse they produce. Analytical simulation of the ERH thruster is accomplished with a one-dimensional model of blast waves propagating uniformly and radially outward from a laser -generated "line source" of high pressure, high temperature gas. Cylindrical blast wave scaling relationships developed by Sedov are employed in this model. The possibility of including other physical phenomena (e.g., viscosity, radiation, conduction or real gas effects) in the analysis will be reviewed. The analyses for the ERH thruster model are performed for a sample vehicle point design. This vehicle, known as the "Lightcraft Technology Demonstrator" (LTD), may be constructed within the next five years to illustrate the potential of Earth-to-Orbit laser propulsion. The external flow over the LTD vehicle was analyzed to determine basic drag characteristics, inlet total pressure recovery and captured air mass flow rate--all projected as functions of flight Mach number and altitude. The ERH thruster performance analysis indicates that the optimum LTD inlet air gap is about 3 cm around the 100 cm diameter centerbody, for transonic "refresh" air flow over the impulse surface. In this analysis, the principal indicator used to predict engine performance was the "impulse coupling coefficient (CC)"; i.e., the thrust developed per unit laser power input. Coupling coefficients up to 600-700 Newtons/Megawatt were found to be feasible, which are an order of magnitude larger than those for laser-heated rockets. For maximum

  16. Turbulence Models for Accurate Aerothermal Prediction in Hypersonic Flows

    NASA Astrophysics Data System (ADS)

    Zhang, Xiang-Hong; Wu, Yi-Zao; Wang, Jiang-Feng

    Accurate description of the aerodynamic and aerothermal environment is crucial to the integrated design and optimization for high performance hypersonic vehicles. In the simulation of aerothermal environment, the effect of viscosity is crucial. The turbulence modeling remains a major source of uncertainty in the computational prediction of aerodynamic forces and heating. In this paper, three turbulent models were studied: the one-equation eddy viscosity transport model of Spalart-Allmaras, the Wilcox k-ω model and the Menter SST model. For the k-ω model and SST model, the compressibility correction, press dilatation and low Reynolds number correction were considered. The influence of these corrections for flow properties were discussed by comparing with the results without corrections. In this paper the emphasis is on the assessment and evaluation of the turbulence models in prediction of heat transfer as applied to a range of hypersonic flows with comparison to experimental data. This will enable establishing factor of safety for the design of thermal protection systems of hypersonic vehicle.

  17. Flap effectiveness appraisal for winged re-entry vehicles

    NASA Astrophysics Data System (ADS)

    de Rosa, Donato; Pezzella, Giuseppe; Donelli, Raffaele S.; Viviani, Antonio

    2016-05-01

    The interactions between shock waves and boundary layer are commonplace in hypersonic aerodynamics. They represent a very challenging design issue for hypersonic vehicle. A typical example of shock wave boundary layer interaction is the flowfield past aerodynamic surfaces during control. As a consequence, such flow interaction phenomena influence both vehicle aerodynamics and aerothermodynamics. In this framework, the present research effort describes the numerical activity performed to simulate the flowfield past a deflected flap in hypersonic flowfield conditions for a winged re-entry vehicle.

  18. Overview of X-38 Hypersonic Aerothermodynamic Wind Tunnel Data and Comparison with Numerical Results

    NASA Technical Reports Server (NTRS)

    Campbell, C.; Caram, J.; Berry, S.; Horvath, T.; Merski, N.; Loomis, M.; Venkatapathy, E.

    2004-01-01

    A NASA team of engineers has been organized to design a crew return vehicle for returning International Space Station crew members from orbit. The hypersonic aerothermodynamic characteristics of the X-23/X-24A derived X-38 crew return vehicle are being evaluated in various wind tunnels in support of this effort. Aerothermodynamic data from two NASA hypersonic tunnels at Mach 6 and Mach 10 has been obtained with cast ceramic models and a thermographic phosphorus digital imaging system. General windward surface heating features are described based on experimental surface heating images and surface oil flow patterns for the nominal hypersonic aerodynamic orientation. Body flap reattachment heating levels are examined. Computational Fluid Dynamics tools have been applied at the appropriate wind tunnel conditions to make comparisons with this data.

  19. Analysis of Laser-Generated Impulse In An Airbreathing Pulsed Detonation Engine: Part 1

    NASA Astrophysics Data System (ADS)

    Richard, Jacques C.; Myrabo, Leik N.

    2005-04-01

    An investigation is performed on an airbreathing laser propulsion (LP) system designed to propel a 1.4 m diameter, 120-kg (dry mass) vehicle called the Lightcraft Technology Demonstrator (LTD) into low Earth orbit, along with its opto-electronics payload. The LTD concept led directly to the model ♯200 lightcraft — recently demonstrated in laboratory and flight experiments at White Sands Missile Range, NM at the High Energy Laser Systems Test Facility (HELSTF), using the 10-kW PLVTS CO2 laser. The pulsed detonation wave engine (PDE) employs repetitively ignited, laser-supported detonation (LSD) waves to develop thrust by expanding high pressure blast waves over an annular, interior shroud surface. Numerical simulation of thruster impulse is accomplished with a 1-D cylindrical model of blast waves propagating radially outward from a laser-generated `line-source' of high temperature, high pressure air. External airflow over the LTD structure is also analyzed to predict basic engine/vehicle drag characteristics, including inlet total pressure recovery, and captured air mass flow rate — all projected vs. flight Mach number and altitude.

  20. Actively cooled plate fin sandwich structural panels for hypersonic aircraft

    NASA Technical Reports Server (NTRS)

    Smith, L. M.; Beuyukian, C. S.

    1979-01-01

    An unshielded actively cooled structural panel was designed for application to a hypersonic aircraft. The design was an all aluminum stringer-stiffened platefin sandwich structure which used a 60/40 mixture of ethylene glycol/water as the coolant. Eight small test specimens of the basic platefin sandwich concept and three fatigue specimens from critical areas of the panel design was fabricated and tested (at room temperature). A test panel representative of all features of the panel design was fabricated and tested to determine the combined thermal/mechanical performance and structural integrity of the system. The overall findings are that; (1) the stringer-stiffened platefin sandwich actively cooling concept results in a low mass design that is an excellent contender for application to a hypersonic vehicle, and (2) the fabrication processes are state of the art but new or modified facilities are required to support full scale panel fabrication.

  1. Filtered Rayleigh scattering measurements in supersonic/hypersonic facilities

    NASA Technical Reports Server (NTRS)

    Miles, Richard B.; Forkey, Joseph N.; Lempert, Walter R.

    1992-01-01

    Preliminary measurements are presented of flow field properties in Mach 3 and Mach 5 flows using filtered Rayleigh scattering. Filter properties have been characterized by high resolution spectroscopy in order to optimize the selection of laser frequency and filter operating conditions, as well as for the development of an accurate filter modeling program. An optimized filter is used the background suppression feature of this technique to image the boundary layer structure in a Mach 3 high Reynolds number facility and the shock structure in a Mach 5 overexpanded jet. This had been achieved using a visible laser source. By frequency scanning the laser, time-averaged velocity measurements in the Mach 3 and Mach 5 flows are made. Data acquisition at 10 torr and below indicates that this approach can be extrapolated for use in hypersonic flow facilities and is applicable as an in-flight optical air data device for hypersonic vehicles.

  2. Formulation of aerodynamic prediction techniques for hypersonic configuration design

    NASA Technical Reports Server (NTRS)

    1979-01-01

    An investigation of approximate theoretical techniques for predicting aerodynamic characteristics and surface pressures for relatively slender vehicles at moderate hypersonic speeds was performed. Emphasis was placed on approaches that would be responsive to preliminary configuration design level of effort. Supersonic second order potential theory was examined in detail to meet this objective. Shock layer integral techniques were considered as an alternative means of predicting gross aerodynamic characteristics. Several numerical pilot codes were developed for simple three dimensional geometries to evaluate the capability of the approximate equations of motion considered. Results from the second order computations indicated good agreement with higher order solutions and experimental results for a variety of wing like shapes and values of the hypersonic similarity parameter M delta approaching one.

  3. Modeling, Measurements, and Fundamental Database Development for Nonequilibrium Hypersonic Aerothermodynamics

    NASA Technical Reports Server (NTRS)

    Bose, Deepak

    2012-01-01

    The design of entry vehicles requires predictions of aerothermal environment during the hypersonic phase of their flight trajectories. These predictions are made using computational fluid dynamics (CFD) codes that often rely on physics and chemistry models of nonequilibrium processes. The primary processes of interest are gas phase chemistry, internal energy relaxation, electronic excitation, nonequilibrium emission and absorption of radiation, and gas-surface interaction leading to surface recession and catalytic recombination. NASAs Hypersonics Project is advancing the state-of-the-art in modeling of nonequilibrium phenomena by making detailed spectroscopic measurements in shock tube and arcjets, using ab-initio quantum mechanical techniques develop fundamental chemistry and spectroscopic databases, making fundamental measurements of finite-rate gas surface interactions, implementing of detailed mechanisms in the state-of-the-art CFD codes, The development of new models is based on validation with relevant experiments. We will present the latest developments and a roadmap for the technical areas mentioned above

  4. An assessment of laser velocimetry in hypersonic flow

    NASA Technical Reports Server (NTRS)

    1992-01-01

    Although extensive progress has been made in computational fluid mechanics, reliable flight vehicle designs and modifications still cannot be made without recourse to extensive wind tunnel testing. Future progress in the computation of hypersonic flow fields is restricted by the need for a reliable mean flow and turbulence modeling data base which could be used to aid in the development of improved empirical models for use in numerical codes. Currently, there are few compressible flow measurements which could be used for this purpose. In this report, the results of experiments designed to assess the potential for laser velocimeter measurements of mean flow and turbulent fluctuations in hypersonic flow fields are presented. Details of a new laser velocimeter system which was designed and built for this test program are described.

  5. Nonlinear potential analysis techniques for supersonic-hypersonic configuration design

    NASA Technical Reports Server (NTRS)

    Clever, W. C.; Shankar, V.

    1983-01-01

    Approximate nonlinear inviscid theoretical techniques for predicting aerodynamic characteristics and surface pressures for relatively slender vehicles at moderate hypersonic speeds were developed. Emphasis was placed on approaches that would be responsive to preliminary configuration design level of effort. Second order small disturbance and full potential theory was utilized to meet this objective. Numerical pilot codes were developed for relatively general three dimensional geometries to evaluate the capability of the approximate equations of motion considered. Results from the computations indicate good agreement with higher order solutions and experimental results for a variety of wing, body and wing-body shapes for values of the hypersonic similarity parameter M delta approaching one. Case computational times of a minute were achieved for practical aircraft arrangements.

  6. Heating and flow-field studies on a straight-wing hypersonic reentry vehicle at angles of attack from 20 to 80 deg with simulation of real-gas trends

    NASA Technical Reports Server (NTRS)

    Hunt, J. L.

    1973-01-01

    Data are presented from a series of phase-change heat transfer and flow visualization tests at Mach 7.4, 8, and 10.3 in air, Mach 19.5 in nitrogen, Mach 20.3 in helium, and Mach 6 in tetrafluoromethane (CF4) on the windward surface of a straight wing hypersonic reentry configuration for angles of attack from 20 deg to 80 deg. The results indicate that: (1) for hypersonic stream Mach numbers, the flow field over the straight-wing configuration is essentially independent of Mach number, (2) transition Reynolds number decreases with increasing angle of attack, (3) at some critical angle of attack, the wing-shock standoff distance is greatly increased and the stagnation line moves downstream from the wing leading edge, (4) value of the critical angle of attack is very sensitive to the flow shock density ratio or effective gamma, and (5) at angles of attack above the critical value for all gases, the nondimensional level of heat transfer to the wing is higher for the higher shock density ratio flows.

  7. A comparison of hypersonic flight and prediction results

    NASA Technical Reports Server (NTRS)

    Iliff, Kenneth W.; Shafer, Mary F.

    1993-01-01

    Aerodynamic and aerothermodynamic comparisons between flight and ground test for four hypersonic vehicles are discussed. The four vehicles are the X-15, the Reentry F, the Sandia Energetic Reentry Vehicle Experiment (SWERVE), and the Space Shuttle. The comparisons are taken from papers published by researchers active in the various programs. Aerodynamic comparisons include reaction control jet interaction on the Space Shuttle. Various forms of heating including catalytic, boundary layer, shock interaction and interference, and vortex impingement are compared. Predictions were significantly exceeded for the heating caused by vortex impingement (on the Space Shuttle OMS pods) and for heating caused by shock interaction and interference on the X-15 and the Space Shuttle. Predictions of boundary-layer state were in error on the X-15, the SWERVE, and the Space Shuttle vehicles.

  8. Design and off-design performance analysis of a maximum compression/minimum drag hypersonic forebody

    NASA Technical Reports Server (NTRS)

    Hawkins, Richard W.; Richardson, Pamela F.

    1991-01-01

    The performance of a funnel-shaped hypersonic vehicle forebody is shown at its design Mach number of 15 and at two lower off-design Mach numbers. Inlet flow quality is presented by contour plots of flow angularity and pressure. Performance is indicated by mass capture ratio, total pressure recovery, kinetic energy efficiency and drag.

  9. Multigrid for hypersonic inviscid flows

    NASA Technical Reports Server (NTRS)

    Decker, Naomi H.; Turkel, Eli

    1990-01-01

    The use of multigrid methods to solve the Euler equations for hypersonic flow is discussed. The steady state equations are considered with a Runge-Kutta smoother based on the time accurate equations together with local time stepping and residual smoothing. The effect of the Runge-Kutta coefficients on the convergence rate was examined considering both damping characteristics and convection properties. The importance of boundary conditions on the convergence rate for hypersonic flow is discussed. Also of importance are the switch between the second and fourth difference viscosity. Solutions are given for flow around the bump in a channel and flow around a biconic section.

  10. Sustained periodic terrestrial locomotion in air-breathing fishes.

    PubMed

    Pace, C M; Gibb, A C

    2014-03-01

    While emergent behaviours have long been reported for air-breathing osteichthyians, only recently have researchers undertaken quantitative analyses of terrestrial locomotion. This review summarizes studies of sustained periodic terrestrial movements by air-breathing fishes and quantifies the contributions of the paired appendages and the axial body to forward propulsion. Elongate fishes with axial-based locomotion, e.g. the ropefish Erpetoichthys calabaricus, generate an anterior-to-posterior wave of undulation that travels down the axial musculoskeletal system and pushes the body against the substratum at multiple points. In contrast, appendage-based locomotors, e.g. the barred mudskipper Periophthalmus argentilineatus, produce no axial bending during sustained locomotion, but instead use repeated protraction-retraction cycles of the pectoral fins to elevate the centre of mass and propel the entire body anteriorly. Fishes that use an axial-appendage-based mechanism, e.g. walking catfishes Clarias spp., produce side-to-side, whole-body bending in co-ordination with protraction-retraction cycles of the pectoral fins. Once the body is maximally bent to one side, the tail is pressed against the substratum and drawn back through the mid-sagittal plane, which elevates the centre of mass and rotates it about a fulcrum formed by the pectoral fin and the ground. Although appendage-based terrestrial locomotion appears to be rare in osteichthyians, many different species appear to have converged upon functionally similar axial-based and axial-appendage-based movements. Based on common forms observed across divergent taxa, it appears that dorsoventral compression of the body, elongation of the axial skeleton or the presence of robust pectoral fins can facilitate effective terrestrial movement by air-breathing fishes.

  11. Fluid dynamic problems associated with air-breathing propulsive systems

    NASA Technical Reports Server (NTRS)

    Chow, W. L.

    1979-01-01

    A brief account of research activities on problems related to air-breathing propulsion is made in this final report for the step funded research grant NASA NGL 14-005-140. Problems include the aircraft ejector-nozzle propulsive system, nonconstant pressure jet mixing process, recompression and reattachment of turbulent free shear layer, supersonic turbulent base pressure, low speed separated flows, transonic boattail flow with and without small angle of attack, transonic base pressures, Mach reflection of shocks, and numerical solution of potential equation through hodograph transformation.

  12. Aerodynamic and aerothermodynamic trade-off analysis of a small hypersonic flying test bed

    NASA Astrophysics Data System (ADS)

    Pezzella, Giuseppe

    2011-08-01

    This paper deals with the aerodynamic and aerothermodynamic trade-off analysis aiming to design a small hypersonic flying test bed with a relatively simple vehicle architecture. Such vehicle will have to be launched with a sounding rocket and shall re-enter the Earth atmosphere allowing to perform several experiments on critical re-entry technologies such as boundary-layer transition and shock-shock interaction phenomena. The flight shall be conducted at hypersonic Mach number, in the range 6-8 at moderate angles of attack. In the paper some design analyses are shown as, for example, the longitudinal and lateral-directional stability analysis. A preliminary optimization of the configuration has been also done to improve the aerodynamic performance and stability of the vehicle. Several design results, based both on engineering approach and computational fluid dynamics, are reported and discussed in the paper. The aerodynamic model of vehicle is also provided.

  13. Results of investigations on a 0.004-scale 140C modified configuration space shuttle vehicle orbiter model (74-0) in the NASA/Langley Research Center hypersonic helium tunnel

    NASA Technical Reports Server (NTRS)

    Hawthorne, P. J.

    1975-01-01

    Data obtained during a wind tunnel test of a 0.004-scale 140C modified configuration SSV orbiter are documented. The test was conducted during August 1974 with 80 occupancy hours charged, and all runs were conducted at a nominal Mach number of 20 and at Reynolds numbers of 0.7, 1.0, 1.8, and 1,100,000 based on body length. The complete -140C modified model was tested with various elevon settings at angles of attack from 10 to 50 degrees at zero yaw and from angles of sideslip of -10 to +10 at 35 deg angle of attack. The purpose of this test was to obtain high hypersonic longitudinal and lateral-directional stability and control characteristics of the updated SSV configuration.

  14. Hypersonic Boundary/Shear Layer Transition for Blunt to Slender Configurations - A NASA Langley Experimental Perspective

    DTIC Science & Technology

    2004-10-01

    scramjet propulsion system performance. The development of numerical tools for the reliable and rapid prediction of boundary layer transition on complex...vehicles, and planetary aerocapture /entry vehicles). Characterization of surface heating on complex shapes and deflected control surfaces and fluid... simulation (6-18), unit Reynolds number (0.01-8 million/ft), and normal shock density ratio (6 and 12). This range of hypersonic simulation

  15. Advanced controls for airbreathing engines, volume 3: Allison gas turbine

    NASA Technical Reports Server (NTRS)

    Bough, R. M.

    1993-01-01

    The application of advanced control concepts to airbreathing engines may yield significant improvements in aircraft/engine performance and operability. Screening studies of advanced control concepts for airbreathing engines were conducted by three major domestic aircraft engine manufacturers to determine the potential impact of concepts on turbine engine performance and operability. The purpose of the studies was to identify concepts which offered high potential yet may incur high research and development risk. A target suite of proposed advanced control concepts was formulated and evaluated in a two-phase study to quantify each concept's impact on desired engine characteristics. To aid in the evaluation specific aircraft/engine combinations were considered: a Military High Performance Fighter mission, a High Speed Civil Transport mission, and a Civil Tiltrotor mission. Each of the advanced control concepts considered in the study are defined and described. The concept potential impact on engine performance was determined. Relevant figures of merit on which to evaluate the concepts are determined. Finally, the concepts are ranked with respect to the target aircraft/engine missions. A final report describing the screening studies was prepared by each engine manufacturer. Volume 3 of these reports describes the studies performed by the Allison Gas Turbine Division.

  16. Simultaneous Global Pressure and Temperature Measurement Technique for Hypersonic Wind Tunnels

    NASA Technical Reports Server (NTRS)

    Buck, Gregory M.

    2000-01-01

    High-temperature luminescent coatings are being developed and applied for simultaneous pressure and temperature mapping in conventional-type hypersonic wind tunnels, providing global pressure as well as Global aeroheating measurements. Together, with advanced model fabrication and analysis methods, these techniques will provide a more rapid and complete experimental aerodynamic and aerothermodynamic database for future aerospace vehicles. The current status in development of simultaneous pressure- and temperature-sensitive coatings and measurement techniques for hypersonic wind tunnels at Langley Research Center is described. and initial results from a feasibility study in the Langley 31-Inch Mach 10 Tunnel are presented.

  17. Energy management and vehicle synthesis

    NASA Technical Reports Server (NTRS)

    Czysz, P.; Murthy, S. N. B.

    1995-01-01

    The major drivers in the development of launch vehicles for the twenty-first century are reduction in cost of vehicles and operations, continuous reusability, mission abort capability with vehicle recovery, and readiness. One approach to the design of such vehicles is to emphasize energy management and propulsion as being the principal means of improvements given the available industrial capability and the required freedom in selecting configuration concept geometries. A methodology has been developed for the rational synthesis of vehicles based on the setting up and utilization of available data and projections, and a reference vehicle. The application of the methodology is illustrated for a single stage to orbit (SSTO) with various limits for the use of airbreathing propulsion.

  18. Status of Turbulence Modeling for Hypersonic Propulsion Flowpaths

    NASA Technical Reports Server (NTRS)

    Georgiadis, Nicholas J.; Yoder, Dennis A.; Vyas, Manan A.; Engblom, William A.

    2012-01-01

    This report provides an assessment of current turbulent flow calculation methods for hypersonic propulsion flowpaths, particularly the scramjet engine. Emphasis is placed on Reynolds-averaged Navier-Stokes (RANS) methods, but some discussion of newer meth- ods such as Large Eddy Simulation (LES) is also provided. The report is organized by considering technical issues throughout the scramjet-powered vehicle flowpath including laminar-to-turbulent boundary layer transition, shock wave / turbulent boundary layer interactions, scalar transport modeling (specifically the significance of turbulent Prandtl and Schmidt numbers) and compressible mixing. Unit problems are primarily used to conduct the assessment. In the combustor, results from calculations of a direct connect supersonic combustion experiment are also used to address the effects of turbulence model selection and in particular settings for the turbulent Prandtl and Schmidt numbers. It is concluded that RANS turbulence modeling shortfalls are still a major limitation to the accuracy of hypersonic propulsion simulations, whether considering individual components or an overall system. Newer methods such as LES-based techniques may be promising, but are not yet at a maturity to be used routinely by the hypersonic propulsion community. The need for fundamental experiments to provide data for turbulence model development and validation is discussed.

  19. Instrumentation Development for Large Scale Hypersonic Inflatable Aerodynamic Decelerator Characterization

    NASA Technical Reports Server (NTRS)

    Swanson, Gregory T.; Cassell, Alan M.

    2011-01-01

    Hypersonic Inflatable Aerodynamic Decelerator (HIAD) technology is currently being considered for multiple atmospheric entry applications as the limitations of traditional entry vehicles have been reached. The Inflatable Re-entry Vehicle Experiment (IRVE) has successfully demonstrated this technology as a viable candidate with a 3.0 m diameter vehicle sub-orbital flight. To further this technology, large scale HIADs (6.0 8.5 m) must be developed and tested. To characterize the performance of large scale HIAD technology new instrumentation concepts must be developed to accommodate the flexible nature inflatable aeroshell. Many of the concepts that are under consideration for the HIAD FY12 subsonic wind tunnel test series are discussed below.

  20. Aerodynamic and Aerothermodynamic Layout of the Hypersonic Flight Experiment Shefex

    NASA Astrophysics Data System (ADS)

    Eggers, Th.

    2005-02-01

    The purpose of the SHarp Edge Flight EXperiment SHEFEX is the investigation of possible new shapes for future launcher or reentry vehicles [1]. The main focus is the improvement of common space vehicle shapes by application of facetted surfaces and sharp edges. The experiment will enable the time accurate investigation of the flow effects and their structural answer during the hypersonic flight from 90 km down to an altitude of 20 km. The project, being performed under responsibility of the German Aerospace Center (DLR) is scheduled to fly on top of a two-stage solid propellant sounding rocket for the first half of 2005. The paper contains a survey of the aerodynamic and aerothermodynamic layout of the experimental vehicle. The results are inputs for the definition of the structural layout, the TPS and the flight instrumentation as well as for the preparation of the flight test performed by the Mobile Rocket Base of DLR.

  1. Study of hypersonic propulsion/airframe integration technology

    NASA Technical Reports Server (NTRS)

    Hartill, W. R.; Goebel, T. P.; Vancamp, V. V.

    1978-01-01

    An assessment is done of current and potential ground facilities, and analysis and flight test techniques for establishing a hypersonic propulsion/airframe integration technology base. A mach 6 cruise prototype aircraft incorporating integrated Scramjet engines was considered the baseline configuration, and the assessment focused on the aerodynamic and configuration aspects of the integration technology. The study describes the key technology milestones that must be met to permit a decision on development of a prototype vehicle, and defines risk levels for these milestones. Capabilities and limitations of analysis techniques, current and potential ground test facilities, and flight test techniques are described in terms of the milestones and risk levels.

  2. Nonlinear potential analysis techniques for supersonic-hypersonic aerodynamic design

    NASA Technical Reports Server (NTRS)

    Shankar, V.; Clever, W. C.

    1984-01-01

    Approximate nonlinear inviscid theoretical techniques for predicting aerodynamic characteristics and surface pressures for relatively slender vehicles at supersonic and moderate hypersonic speeds were developed. Emphasis was placed on approaches that would be responsive to conceptual configuration design level of effort. Second order small disturbance and full potential theory was utilized to meet this objective. Numerical codes were developed for relatively general three dimensional geometries to evaluate the capability of the approximate equations of motion considered. Results from the computations indicate good agreement with experimental results for a variety of wing, body, and wing-body shapes.

  3. Hypersonic Flight: Time To Go Operational

    DTIC Science & Technology

    2013-02-14

    AIR WAR COLLEGE AIR UNIVERSITY HYPERSONIC FLIGHT: TIME TO GO OPERATIONAL by Robert A. Dietrick, Lt Col, USAF A Research Report...threats are increasing and could jeopardize the ability of the US Air Force to effectively conduct global strike by 2032. Scramjet powered hypersonic ...flight could be a key capability by reducing time to strike and increasing survivability. Historically, the key challenges preventing hypersonic

  4. Development of a Multi-Disciplinary Aerothermostructural Model Applicable to Hypersonic Flight

    NASA Technical Reports Server (NTRS)

    Kostyk, Chris; Risch, Tim

    2013-01-01

    The harsh and complex hypersonic flight environment has driven design and analysis improvements for many years. One of the defining characteristics of hypersonic flight is the coupled, multi-disciplinary nature of the dominant physics. In an effect to examine some of the multi-disciplinary problems associated with hypersonic flight engineers at the NASA Dryden Flight Research Center developed a non-linear 6 degrees-of-freedom, full vehicle simulation that includes the necessary model capabilities: aerothermal heating, ablation, and thermal stress solutions. Development of the tool and results for some investigations will be presented. Requirements and improvements for future work will also be reviewed. The results of the work emphasize the need for a coupled, multi-disciplinary analysis to provide accurate

  5. HIAD-2 (Hypersonic Inflatable Aerodynamic Decelerator)

    NASA Video Gallery

    The Hypersonic Inflatable Aerodynamic Decelerator (HIAD) project is a disruptive technology that will accommodate the atmospheric entry of heavy payloads to planetary bodies such as Mars. HIAD over...

  6. Mathematical Models Of Turbulence In Hypersonic Flow

    NASA Technical Reports Server (NTRS)

    Marvin, J. G.; Coakley, T. J.

    1991-01-01

    Report discusses mathematical models of turbulence used in numerical simulations of complicated viscous, hypersonic flows. Includes survey of essential features of models and their statuses in applications.

  7. Earth-to-orbit reusable launch vehicles: A comparative assessment

    NASA Technical Reports Server (NTRS)

    Chase, R. L.

    1978-01-01

    A representative set of space systems, functions, and missions for NASA and DoD from which launch vehicle requirements and characteristics was established as well as a set of air-breathing launch vehicles based on graduated technology capabilities corresponding to increasingly higher staging Mach numbers. The utility of the air-breathing launch vehicle candidates based on lift-off weight, performance, technology needs, and risk was assessed and costs were compared to alternative concepts. The results indicate that a fully reusable launch vehicle, whether two stage or one stage, could potentially reduce the cost per flight 60-80% compared to that for a partially reusable vehicle but would require advances in thermal protection system technology. A two-stage-to-orbit, parallel-lift vehicle with an air-breathing booster would cost approximately the same as a single-stage-to-orbit vehicle, but the former would have greater flexibility and a significantly reduced developmental risk. A twin-booster, subsonic-staged, parallel-lift vehicle represents the lowest system cost and developmental risk. However, if a large supersonic turbojet engine in the 350,000-N thrust class were available, supersonic staging would be preferred, and the investment in development would be returned in reduced program cost.

  8. Aero-thermal analysis of lifting body configurations in hypersonic flow

    NASA Astrophysics Data System (ADS)

    Kumar, Sachin; Mahulikar, Shripad P.

    2016-09-01

    The aero-thermal analysis of a hypersonic vehicle is of fundamental interest for designing its thermal protection system. The aero-thermal environment predictions over several critical regions of the hypothesized lifting body vehicle, including the stagnation region of the nose-cap, cylindrically swept leading edges, fuselage-upper, and fuselage-lower surfaces, are discussed. The drag (Λ=70°) and temperature (Λ=80°) minimized sweepback angles are considered in the configuration design of the two hypothesized lifting body shape hypersonic vehicles. The main aim of the present study is to analyze and compare the aero-thermal characteristics of these two lifting body configurations at same heat capacity. Accordingly, a Computational Fluid Dynamics simulation has been carried out at Mach number (M∞=7), H=35 km altitude with zero Angle of Attack. Finally, the material selection for thermal protection system based on these predictions and current methodology is described.

  9. Wall catalytic recombination and boundary conditions in nonequilibrium hypersonic flows - With applications

    NASA Technical Reports Server (NTRS)

    Scott, Carl D.

    1992-01-01

    The meaning of catalysis and its relation to aerodynamic heating in nonequilibrium hypersonic flows are discussed. The species equations are described and boundary conditions for them are derived for a multicomponent gas and for a binary gas. Slip effects are included for application of continuum methods to low-density flows. Measurement techniques for determining catalytic wall recombination rates are discussed. Among them are experiments carried out in arc jets as well as flow reactors. Diagnostic methods for determining the atom or molecule concentrations in the flow are included. Results are given for a number of materials of interest to the aerospace community, including glassy coatings such as the RCG coating of the Space Shuttle and for high temperature refractory metals such as coated niobium. Methods of calculating the heat flux to space vehicles in nonequilibrium flows are described. These methods are applied to the Space Shuttle, the planned Aeroassist Flight Experiment, and a hypersonic slender vehicle such as a transatmospheric vehicle.

  10. Experimental investigation of a unique airbreathing pulsed laser propulsion concept

    NASA Technical Reports Server (NTRS)

    Myrabo, L. N.; Nagamatsu, H. T.; Manka, C.; Lyons, P. W.; Jones, R. A.

    1991-01-01

    Investigations were conducted into unique methods of converting pulsed laser energy into propulsive thrust across a flat impulse surface under atmospheric conditions. The propulsion experiments were performed with a 1-micron neodymium-glass laser at the Space Plasma Branch of the Naval Research Laboratory. Laser-induced impulse was measured dynamically by ballistic pendulums and statically using piezoelectric pressure transducers on a stationary impulse surface. The principal goal was to explore methods for increasing the impulse coupling performance of airbreathing laser-propulsion engines. A magnetohydrodynamic thrust augmentation effect was discovered when a tesla-level magnetic field was applied perpendicular to the impulse surface. The impulse coupling coefficient performance doubled and continued to improve with increasing laser-pulse energies. The resultant performance of 180 to 200 N-s/MJ was found to be comparable to that of the earliest afterburning turbojets.

  11. Theme and variations: amphibious air-breathing intertidal fishes.

    PubMed

    Martin, K L

    2014-03-01

    Over 70 species of intertidal fishes from 12 families breathe air while emerging from water. Amphibious intertidal fishes generally have no specialized air-breathing organ but rely on vascularized mucosae and cutaneous surfaces in air to exchange both oxygen and carbon dioxide. They differ from air-breathing freshwater fishes in morphology, physiology, ecology and behaviour. Air breathing and terrestrial activity are present to varying degrees in intertidal fish species, correlated with the tidal height of their habitat. The gradient of amphibious lifestyle includes passive remainers that stay in the intertidal zone as tides ebb, active emergers that deliberately leave water in response to poor aquatic conditions and highly mobile amphibious skipper fishes that may spend more time out of water than in it. Normal terrestrial activity is usually aerobic and metabolic rates in air and water are similar. Anaerobic metabolism may be employed during forced exercise or when exposed to aquatic hypoxia. Adaptations for amphibious life include reductions in gill surface area, increased reliance on the skin for respiration and ion exchange, high affinity of haemoglobin for oxygen and adjustments to ventilation and metabolism while in air. Intertidal fishes remain close to water and do not travel far terrestrially, and are unlikely to migrate or colonize new habitats at present, although in the past this may have happened. Many fish species spawn in the intertidal zone, including some that do not breathe air, as eggs and embryos that develop in the intertidal zone benefit from tidal air emergence. With air breathing, amphibious intertidal fishes survive in a variable habitat with minimal adjustments to existing structures. Closely related species in different microhabitats provide unique opportunities for comparative studies.

  12. Air-breathing fishes in aquaculture. What can we learn from physiology?

    PubMed

    Lefevre, S; Wang, T; Jensen, A; Cong, N V; Huong, D T T; Phuong, N T; Bayley, M

    2014-03-01

    During the past decade, the culture of air-breathing fish species has increased dramatically and is now a significant global source of protein for human consumption. This development has generated a need for specific information on how to maximize growth and minimize the environmental effect of culture systems. Here, the existing data on metabolism in air-breathing fishes are reviewed, with the aim of shedding new light on the oxygen requirements of air-breathing fishes in aquaculture, reaching the conclusion that aquatic oxygenation is much more important than previously assumed. In addition, the possible effects on growth of the recurrent exposure to deep hypoxia and associated elevated concentrations of carbon dioxide, ammonia and nitrite, that occurs in the culture ponds used for air-breathing fishes, are discussed. Where data on air-breathing fishes are simply lacking, data for a few water-breathing species will be reviewed, to put the physiological effects into a growth perspective. It is argued that an understanding of air-breathing fishes' respiratory physiology, including metabolic rate, partitioning of oxygen uptake from air and water in facultative air breathers, the critical oxygen tension, can provide important input for the optimization of culture practices. Given the growing importance of air breathers in aquaculture production, there is an urgent need for further data on these issues.

  13. Surface Heat Flux and Pressure Distribution on a Hypersonic Blunt Body With DEAS

    NASA Astrophysics Data System (ADS)

    Salvador, I. I.; Minucci, M. A. S.; Toro, P. G. P.; Oliveira, A. C.; Channes, J. B.

    2008-04-01

    With the currently growing interest for advanced technologies to enable hypersonic flight comes the Direct Energy Air Spike concept, where pulsed beamed laser energy is focused upstream of a blunt flight vehicle to disrupt the flow structure creating a virtual, slender body geometry. This allies in the vehicle both advantages of a blunt body (lower thermal stresses) to that of a slender geometry (lower wave drag). The research conducted at the Henry T. Nagamatsu Laboratory for Aerodynamics and Hypersonics focused on the measurement of the surface pressure and heat transfer rates on a blunt model. The hypersonic flight conditions were simulated at the HTN Laboratory's 0.3 m T2 Hypersonic Shock Tunnel. During the tests, the laser energy was focused upstream the model by an infrared telescope to create the DEAS effect, which was supplied by a TEA CO2 laser. Piezoelectric pressure transducers were used for the pressure measurements and fast response coaxial thermocouples were used for the measurement of surface temperature, which was later used for the estimation of the wall heat transfer using the inverse heat conduction theory.

  14. Performance analysis of rocket-ramjet propelled SSTO-vehicles

    NASA Astrophysics Data System (ADS)

    Schoettle, U. M.

    1985-10-01

    Winged single-stage-to-orbit (SSTO) vehicles designed for vertical or horizontal takeoff and horizontal landing with both rocket and rocket-ramjet propulsion concepts are analyzed and their performance and costs are compared. For this purpose, LOX/LH2 rocket baseline vehicles with payload and mission requirements similar to the Space Shuttle system are modified to accommodate hydrogen-fueled ramjet engines. The use of the airbreathing engines results in a substantial decrease in propellant consumption but is heavily penalized by the added weights of the ramjet engine and structural reinforcements to resist higher aero-thermodynamic loads. The results suggest that for vehicles of the same gross lift-off weight the total system costs of the airbreathing vehicles are 19 percent higher compared to rocket systems; however, due to increased payload capabilities, the specific transportation costs are lower.

  15. Non-Equilibrium Effects on Hypersonic Turbulent Boundary Layers

    NASA Astrophysics Data System (ADS)

    Kim, Pilbum

    Understanding non-equilibrium effects of hypersonic turbulent boundary layers is essential in order to build cost efficient and reliable hypersonic vehicles. It is well known that non-equilibrium effects on the boundary layers are notable, but our understanding of the effects are limited. The overall goal of this study is to improve the understanding of non-equilibrium effects on hypersonic turbulent boundary layers. A new code has been developed for direct numerical simulations of spatially developing hypersonic turbulent boundary layers over a flat plate with finite-rate reactions. A fifth-order hybrid weighted essentially non-oscillatory scheme with a low dissipation finite-difference scheme is utilized in order to capture stiff gradients while resolving small motions in turbulent boundary layers. The code has been validated by qualitative and quantitative comparisons of two different simulations of a non-equilibrium flow and a spatially developing turbulent boundary layer. With the validated code, direct numerical simulations of four different hypersonic turbulent boundary layers, perfect gas and non-equilibrium flows of pure oxygen and nitrogen, have been performed. In order to rule out uncertainties in comparisons, the same inlet conditions are imposed for each species, and then mean and turbulence statistics as well as near-wall turbulence structures are compared at a downstream location. Based on those comparisons, it is shown that there is no direct energy exchanges between internal and turbulent kinetic energies due to thermal and chemical non-equilibrium processes in the flow field. Instead, these non-equilibria affect turbulent boundary layers by changing the temperature without changing the main characteristics of near-wall turbulence structures. This change in the temperature induces the changes in the density and viscosity and the mean flow fields are then adjusted to satisfy the conservation laws. The perturbation fields are modified according to

  16. Hypersonic Composites Resist Extreme Heat and Stress

    NASA Technical Reports Server (NTRS)

    2007-01-01

    Through research contracts with NASA, Materials and Electrochemical Research Corporation (MER), of Tucson, Arizona, contributed a number of technologies to record-breaking hypersonic flights. Through this research, MER developed a coating that successfully passed testing to simulate Mach 10 conditions, as well as provide several additional carbon-carbon (C-C) composite components for the flights. MER created all of the leading edges for the X-43A test vehicles at Dryden-considered the most critical parts of this experimental craft. In addition to being very heat resistant, the coating had to be very lightweight and thin, as the aircraft was designed to very precise specifications and could not afford to have a bulky coating. MER patented its carbon-carbon (C-C) composite process and then formed a spinoff company, Frontier Materials Corporation (FMC), also based in Tucson. FMC is using the patent in conjunction with low-cost PAN (polyacrylonitrile)-based fibers to introduce these materials to the commercial markets. The C-C composites are very lightweight and exceptionally strong and stiff, even at very high temperatures. The composites have been used in industrial heating applications, the automotive and aerospace industries, as well as in glass manufacturing and on semiconductors. Applications also include transfer components for glass manufacturing and structural members for carrier support in semiconductor processing.

  17. Scaled Rocket Testing in Hypersonic Flow

    NASA Technical Reports Server (NTRS)

    Dufrene, Aaron; MacLean, Matthew; Carr, Zakary; Parker, Ron; Holden, Michael; Mehta, Manish

    2015-01-01

    NASA's Space Launch System (SLS) uses four clustered liquid rocket engines along with two solid rocket boosters. The interaction between all six rocket exhaust plumes will produce a complex and severe thermal environment in the base of the vehicle. This work focuses on a recent 2% scale, hot-fire SLS base heating test. These base heating tests are short-duration tests executed with chamber pressures near the full-scale values with gaseous hydrogen/oxygen engines and RSRMV analogous solid propellant motors. The LENS II shock tunnel/Ludwieg tube tunnel was used at or near flight duplicated conditions up to Mach 5. Model development was strongly based on the Space Shuttle base heating tests with several improvements including doubling of the maximum chamber pressures and duplication of freestream conditions. Detailed base heating results are outside of the scope of the current work, rather test methodology and techniques are presented along with broader applicability toward scaled rocket testing in supersonic and hypersonic flow.

  18. Results of heat transfer tests of an 0.0175-scale space shuttle vehicle model 22 OTS in the NASA-Ames 3.5-foot hypersonic wind tunnel (IH3), volume 3

    NASA Technical Reports Server (NTRS)

    Foster, T. F.; Lockman, W. K.

    1975-01-01

    Heat-transfer data for the 0.0175-scale space shuttle vehicle 3 are presented, and interference heating effects were investigated by a model build-up technique of the orbiter. The test program was conducted at Mach 5.3 for nominal free-stream Reynolds number per foot values of 1,500,000 and 5,000,000.

  19. HIFiRE-5 Flight Vehicle Design

    NASA Technical Reports Server (NTRS)

    Kimmel, Roger L.; Adamczak, David; Berger, Karen; Choudhari, Meelan

    2010-01-01

    The Hypersonic International Flight Research Experimentation (HIFiRE) program is a hypersonic flight test program executed by the Air Force Research Laboratories (AFRL) and Australian Defence Science and Technology Organization (DSTO). HIFiRE flight 5 is devoted to measuring transition on a three-dimensional body. This paper summarizes payload configuration, trajectory, vehicle stability limits and roughness tolerances. Results show that the proposed configuration is suitable for testing transition on a three-dimensional body. Transition is predicted to occur within the test window, and a design has been developed that will allow the vehicle to be manufactured within prescribed roughness tolerances

  20. Unstructured Mesh Methods for the Simulation of Hypersonic Flows

    NASA Technical Reports Server (NTRS)

    Peraire, Jaime; Bibb, K. L. (Technical Monitor)

    2001-01-01

    This report describes the research work undertaken at the Massachusetts Institute of Technology. The aim of this research is to identify effective algorithms and methodologies for the efficient and routine solution of hypersonic viscous flows about re-entry vehicles. For over ten years we have received support from NASA to develop unstructured mesh methods for Computational Fluid Dynamics. As a result of this effort a methodology based on the use, of unstructured adapted meshes of tetrahedra and finite volume flow solvers has been developed. A number of gridding algorithms flow solvers, and adaptive strategies have been proposed. The most successful algorithms developed from the basis of the unstructured mesh system FELISA. The FELISA system has been extensively for the analysis of transonic and hypersonic flows about complete vehicle configurations. The system is highly automatic and allows for the routine aerodynamic analysis of complex configurations starting from CAD data. The code has been parallelized and utilizes efficient solution algorithms. For hypersonic flows, a version of the, code which incorporates real gas effects, has been produced. One of the latest developments before the start of this grant was to extend the system to include viscous effects. This required the development of viscous generators, capable of generating the anisotropic grids required to represent boundary layers, and viscous flow solvers. In figures I and 2, we show some sample hypersonic viscous computations using the developed viscous generators and solvers. Although these initial results were encouraging, it became apparent that in order to develop a fully functional capability for viscous flows, several advances in gridding, solution accuracy, robustness and efficiency were required. As part of this research we have developed: 1) automatic meshing techniques and the corresponding computer codes have been delivered to NASA and implemented into the GridEx system, 2) a finite

  1. 2-D Airbreathing Lightcraft Engine Experiments in Quiescent Conditions

    NASA Astrophysics Data System (ADS)

    Salvador, Israel I.; Myrabo, Leik N.; Minucci, Marco A. S.; de Oliveira, Antonio C.; Toro, Paulo G. P.; Chanes, José B.; Rego, Israel S.

    2011-11-01

    Ground-breaking laser propulsion (LP) experiments were performed under quiescent conditions with a 25 cm wide, two-dimensional Lightcraft model using a Lumonics TEA-622 CO2 laser emitting ˜ 1 μs pulses. In preparation for subsequent hypersonic experiments, this static test campaign was conducted at ambient pressures of 0.06, 0.15, 0.30 and 1 bar with laser pulse energies of 150 to 230 J. Time-variant pressure distributions, generated over engine "absorption chamber" walls, were integrated to obtain total impulse and momentum coupling coefficients (Cm) representative of a single propulsion cycle. Schlieren visualization of laser-induced air breakdown and expanding blast waves was also accomplished. Surprisingly, the Cm results of 600-3000 Ns/MJ were 2.5x to 5x greater than previous results from smaller Lightcraft models; this suggests that higher static Cm performance can likely be achieved in larger scale LP engines. This research collaboration, forged between the USAF and Brazilian Air Force, was carried out at the Henry T. Nagamatsu Laboratory of Aerothermodynamics and Hypersonics in Brazil.

  2. Transpiration cooling in hypersonic flight

    NASA Technical Reports Server (NTRS)

    Tavella, Domingo; Roberts, Leonard

    1989-01-01

    A preliminary numerical study of transpiration cooling applied to a hypersonic configuration is presented. Air transpiration is applied to the NASA all-body configuration flying at an altitude of 30500 m with a Mach number of 10.3. It was found that the amount of heat disposal by convection is determined primarily by the local geometry of the aircraft for moderate rates of transpiration. This property implies that different areas of the aircraft where transpiration occurs interact weakly with each other. A methodology for quick assessments of the transpiration requirements for a given flight configuration is presented.

  3. Hermes vehicle

    NASA Astrophysics Data System (ADS)

    Cretenet, J.-C.

    1984-10-01

    Projected mission profiles and computational models of the Hermes winged manned reentry vehicle are discussed. Launched into a low orbit with a crew of at most four by the Ariane 5 vehicle, Hermes will serve transportation, crew relief, safety, and freight carriage functions. It will have a nominal 300-500 km circular or 600-900 km heliosynchronous orbit, weigh from 13,500-16,700 kg, and return in hypersonic glide to a wheeled landing. The vehicle dimensions will be 15-18 m length, 6 m height at the tail fin, and a 10 m span. The L/D ratio will be 1.5-1.6, thereby furnishing a cross range of 2500 km. Hermes will have a Shuttle-type fuselage, carbon-carbon nose, and insulation on the extrados designed to keep the structure at 175 C or lower. The avionics will have 3-4 levels of redundance, each mission phase-dependent. Power will be supplied by three fuel cells and a bank of four Ag-Zn batteries. Hardware development is scheduled to begin in 1988.

  4. Hypersonic ground test capabilities for T and E testing above mach 8 ''a case where S and T meets T and E''

    SciTech Connect

    Constantino, M; Miles, R; Brown, G; Laster, M; Nelson, G

    1999-10-05

    Simulation of hypersonic flight in ground test and evaluation (T and E) facilities is a challenging and formidable task, especially to fully duplicate the flight environment above approximately Mach 8 for most all hypersonic flight systems that have been developed, conceived, or envisioned. Basically, and for many years, the enabling technology to build such a ground test wind tunnel facility has been severely limited in the area of high-temperature, high-strength materials and thermal protection approaches. To circumvent the problems, various approaches have been used, including partial simulation and use of similarity laws and reduced test time. These approaches often are not satisfactory, i.e. operability and durability testing for air-breathing propulsion development and thermal protection development of many flight systems. Thus, there is a strong need for science and technology (S and T) community involvement in technology development to address these problems. This paper discusses a specific case where this need exists and where significant S and T involvement has made and continues to make significant contributions. The case discussed will be an Air Force research program currently underway to develop enabling technologies for a Mach 8-15 hypersonic true temperature wind tunnel with relatively long run time. The research is based on a concept proposed by princeton University using radiant or beamed energy into the supersonic nozzle flow.

  5. Study of heat sink thermal protection systems for hypersonic research aircraft

    NASA Technical Reports Server (NTRS)

    Vahl, W. A.; Edwards, C. L. W.

    1978-01-01

    The feasibility of using a single metallic heat sink thermal protection system (TPS) over a projected flight test program for a hypersonic research vehicle was studied using transient thermal analyses and mission performance calculations. Four materials, aluminum, titanium, Lockalloy, and beryllium, as well as several combinations, were evaluated. Influence of trajectory parameters were considered on TPS and mission performance for both the clean vehicle configuration as well as with an experimental scramjet mounted. From this study it was concluded that a metallic heat sink TPS can be effectively employed for a hypersonic research airplane flight envelope which includes dash missions in excess of Mach 8 and 60 seconds of cruise at Mach numbers greater than 6. For best heat sink TPS match over the flight envelope, Lockalloy and titanium appear to be the most promising candidates

  6. The NASA-sponsored Maryland center for hypersonic education and research

    NASA Technical Reports Server (NTRS)

    Lewis, Mark J.; Gupta, Ashwani K.

    1995-01-01

    The Office of Aeronautics of the National Aeronautics and Space Administration has established a program to support university programs in the field of hypersonic flight. Beginning in the fall of 1993, three universities, including the University of Maryland at College Park, were selected to participate in this activity. The program at the University of Maryland includes faculty in the Department of Aerospace Engineering and Department of Mechanical Engineering, and provides a multidisciplinary environment for graduate and undergraduate students to study and conduct research in the field of hypersonic flight. Ongoing projects cover the range of applications from cruisers through transatmospheric and reentry vehicles. Research activities, focused on propulsion, fluid dynamics, inverse design, and vehicle optimization and integration, are conducted in conjuntion with industrial partners and government laboratories.

  7. Overview of X-38 Hypersonic Wind Tunnel Data and Comparison with Numerical Results

    NASA Technical Reports Server (NTRS)

    Campbell, Charles H.; Caram, Jose; Berry, Scott; DiFulvio, Michael; Horvath, Tom

    1997-01-01

    A NASA team of engineers has been organized to design a crew return vehicle for returning International Space Station crew members from orbit. The hypersonic characteristics of this X-23/X-2&4 derived crew return vehicle (designated X-38) are being evaluated in various wind tunnels in support of this effort. Aerodynamic data has been acquired in three NASA hypersonic facilities at Mach 20, and Mach 6. Computational Fluid Dynamics tools have been applied at the appropriate wind tunnel conditions to make comparisons with portions of this data. Experimental data from the Mach 6 Air and CF4 facilities illustrate a net positive pitching moment increment due to density ratio, as well as increased elevon effectiveness. Chemical nonequilibrium computational fluid dynamics solutions at flight conditions reinforce this conclusion.

  8. Hypersonic transports: Economics and environmental effects

    NASA Technical Reports Server (NTRS)

    Petersen, R. H.; Waters, M. H.

    1972-01-01

    An economic analysis of hypersonic transports is presented to show projected operating costs (direct and indirect) and return on investment. Important assumptions are varied to determine the probable range of values for operating costs and return on investment. The environmental effects of hypersonic transports are discussed and compared to current supersonic transports. Estimates of sideline and fly-over noise are made for a typical hypersonic transport, and the sonic boom problem is analyzed and discussed. Since the exhaust products from liquid hydrogen-fueled engines differ from those of kerosene-fueled aircraft, a qualitative assessment of air pollution effects is made.

  9. Hypersonic transports - Economics and environmental effects.

    NASA Technical Reports Server (NTRS)

    Petersen, R. H.; Waters, M. H.

    1972-01-01

    An economic analysis of hypersonic transports is presented to show projected operating costs (direct and indirect) and return on investment. Important assumptions are varied to determine the probable range of values for operating costs and return on investment. The environmental effects of hypersonic transports are discussed and compared to current supersonic transports. Estimates of sideline and flyover noise are made for a typical hypersonic transport, and the sonic boom problem is analyzed and discussed. Since the exhaust products from liquid hydrogen-fueled engines differ from those of kerosene-fueled aircraft, a qualitative assessment of air pollution effects is made.

  10. Hypersonic transports - Economics and environmental effects.

    NASA Technical Reports Server (NTRS)

    Petersen, R. H.; Waters, M. H.

    1973-01-01

    An economic analysis of hypersonic transports is presented to show projected operating costs (direct and indirect) and return on investment. Important assumptions are varied to determine the probable range of values for operating costs and return on investment. The environmental effects of hypersonic transports are discussed and compared to current supersonic transports. Estimates of sideline and flyover noise are made for a typical hypersonic transport, and the sonic boom problem is analyzed and discussed. Since the exhaust products from liquid hydrogen-fueled engines differ from those of kerosene-fueled aircraft, a qualitative assessment of air pollution effects is made.

  11. An Overview of SBIR Phase 2 Airbreathing Propulsion Technologies

    NASA Technical Reports Server (NTRS)

    Nguyen, Hung D.; Steele, Gynelle C.; Bitler, Dean W.

    2014-01-01

    Technological innovation is the overall focus of NASA's Small Business Innovation Research (SBIR) program. The program invests in the development of innovative concepts and technologies to help NASA's mission directorates address critical research and development needs for agency projects. This report highlights innovative SBIR Phase II projects from 2007-2012 specifically addressing areas in Airbreathing Propulsion which is one of six core competencies at NASA Glenn Research Center. There are twenty technologies featured with emphasis on a wide spectrum of applications such as with a Turbo-Brayton cryocooler for aircraft superconducting systems, braided composite rotorcraft structures, engine air brake, combustion control valve, flexible composite driveshaft, and much more. Each article in this booklet describes an innovation, technical objective, and highlights NASA commercial and industrial applications. This report serves as an opportunity for NASA personnel including engineers, researchers, and program managers to learn of NASA SBIR's capabilities that might be crosscutting into this technology area. As the result, it would cause collaborations and partnerships between the small companies and NASA Programs and Projects resulting in benefit to both SBIR companies and NASA.

  12. Stack air-breathing membraneless glucose microfluidic biofuel cell

    NASA Astrophysics Data System (ADS)

    Galindo-de-la-Rosa, J.; Moreno-Zuria, A.; Vallejo-Becerra, V.; Arjona, N.; Guerra-Balcázar, M.; Ledesma-García, J.; Arriaga, L. G.

    2016-11-01

    A novel stacked microfluidic fuel cell design comprising re-utilization of the anodic and cathodic solutions on the secondary cell is presented. This membraneless microfluidic fuel cell employs porous flow-through electrodes in a “V”-shape cell architecture. Enzymatic bioanodic arrays based on glucose oxidase were prepared by immobilizing the enzyme onto Toray carbon paper electrodes using tetrabutylammonium bromide, Nafion and glutaraldehyde. These electrodes were characterized through the scanning electrochemical microscope technique, evidencing a good electrochemical response due to the electronic transference observed with the presence of glucose over the entire of the electrode. Moreover, the evaluation of this microfluidic fuel cell with an air-breathing system in a double-cell mode showed a performance of 0.8951 mWcm-2 in a series connection (2.2822mAcm-2, 1.3607V), and 0.8427 mWcm-2 in a parallel connection (3.5786mAcm-2, 0.8164V).

  13. Modeling and Evaluating the Environmental Degradation of UHTCs under Hypersonic Flow (Preprint)

    DTIC Science & Technology

    2014-02-01

    34Oxidation behavior of zirconium diboride silicon carbide produced by the spark plasma sintering method," J Amer. Ceram. Soc., 92 [9] 2046-2052...vehicles. It uses a plasma generated gas mixture which contains a high fraction of ionized atoms of oxygen and nitrogen. Recombination of the charged...species is catalyzed by the sample surface to release heat. The plasma is accelerated to high velocities (supersonic but not hypersonic) using an

  14. FGM (Functionally Graded Material) Thermal Barrier Coatings for Hypersonic Structures - Design and Thermal Structural Analysis

    DTIC Science & Technology

    2007-06-29

    models were developed and implemented in the transient dynamics solver of the finite element program. Experimental results from an induction heating ...with little or no spallation. The optimum time, heat flux and applied pressure level for sintering were deduced. TBCs produced from a 40-45% ceramic, 4...tests in 2004. Hypersonic / high supersonic vehicles will experience increased heating , aerodynamic and aeroacoustic loads. Severe aerodynamic heating

  15. Transpiration cooling of hypersonic blunt bodies with finite rate surface reactions

    NASA Technical Reports Server (NTRS)

    Henline, William D.

    1989-01-01

    The convective heat flux blockage to blunt body and hypersonic vehicles by transpiration cooling are presented. The general problem of mass addition to laminar boundary layers is reviewed. Results of similarity analysis of the boundary layer problem are provided for surface heat flux with transpiration cooling. Detailed non-similar results are presented from the numerical program, BLIMPK. Comparisons are made with the similarity theory. The effects of surface catalysis are investigated.

  16. High capacity for extracellular acid-base regulation in the air-breathing fish Pangasianodon hypophthalmus.

    PubMed

    Damsgaard, Christian; Gam, Le Thi Hong; Tuong, Dang Diem; Thinh, Phan Vinh; Huong Thanh, Do Thi; Wang, Tobias; Bayley, Mark

    2015-05-01

    The evolution of accessory air-breathing structures is typically associated with reduction of the gills, although branchial ion transport remains pivotal for acid-base and ion regulation. Therefore, air-breathing fishes are believed to have a low capacity for extracellular pH regulation during a respiratory acidosis. In the present study, we investigated acid-base regulation during hypercapnia in the air-breathing fish Pangasianodon hypophthalmus in normoxic and hypoxic water at 28-30°C. Contrary to previous studies, we show that this air-breathing fish has a pronounced ability to regulate extracellular pH (pHe) during hypercapnia, with complete metabolic compensation of pHe within 72 h of exposure to hypoxic hypercapnia with CO2 levels above 34 mmHg. The high capacity for pHe regulation relies on a pronounced ability to increase levels of HCO3(-) in the plasma. Our study illustrates the diversity in the physiology of air-breathing fishes, such that generalizations across phylogenies may be difficult.

  17. Aerodynamically landing reentry vehicles

    NASA Astrophysics Data System (ADS)

    Widjaja, I.

    This article represents a continuation of a paper in the preceding edition of this journal. The longitudinal stability of the reentry vehicle configuration 24B is discussed, taking into account an evaluation of the possibilities for lateral control, aileron effectiveness, and rudder effectiveness. It is pointed out that regarding the selection of the characteristics of the descent trajectory, there are apparently no constraints related to stability or controllability limits. In the hypersonic range, large reciprocal lift drag ratios can also be obtained without positive flap displacement. Attention is given to angle of sideslip relations, a cylindrical body with flat nose and trim tabs, the flow characteristics in the case of a cylinder with a flat nose, graphical relations describing longitudinal stability and controllability in the hypersonic range, and relations involving lift, drag, and the lift drag ratio.

  18. Results of heat transfer tests of a 0.0175-scale space shuttle vehicle 5 model (60-OTS) in the NASA-Ames Research Center 3.5-foot hypersonic wind tunnel (test IH48)

    NASA Technical Reports Server (NTRS)

    Dye, W. H.; Lockman, W. K.

    1976-01-01

    Heat transfer data are presented for a .0175-scale model of the Rockwell International Space Shuttle Vehicle 5. The primary purpose of these tests was to obtain aerodynamic interference heating data on the external tank in the tank alone, second-, and first-stage configurations. Data were also obtained on the Orbiter and solid rocket boosters. Nominal Mach Nos. of 5.2 and 5.3 at nominal freestream unit Reynolds numbers of 1.5 and 5.0 million per foot, respectively, were investigated. Photographs of the tested configurations and test equipment are shown.

  19. Measurement uncertainty and feasibility study of a flush airdata system for a hypersonic flight experiment

    NASA Technical Reports Server (NTRS)

    Whitmore, Stephen A.; Moes, Timothy R.

    1994-01-01

    Presented is a feasibility and error analysis for a hypersonic flush airdata system on a hypersonic flight experiment (HYFLITE). HYFLITE heating loads make intrusive airdata measurement impractical. Although this analysis is specifically for the HYFLITE vehicle and trajectory, the problems analyzed are generally applicable to hypersonic vehicles. A layout of the flush-port matrix is shown. Surface pressures are related airdata parameters using a simple aerodynamic model. The model is linearized using small perturbations and inverted using nonlinear least-squares. Effects of various error sources on the overall uncertainty are evaluated using an error simulation. Error sources modeled include boundarylayer/viscous interactions, pneumatic lag, thermal transpiration in the sensor pressure tubing, misalignment in the matrix layout, thermal warping of the vehicle nose, sampling resolution, and transducer error. Using simulated pressure data for input to the estimation algorithm, effects caused by various error sources are analyzed by comparing estimator outputs with the original trajectory. To obtain ensemble averages the simulation is run repeatedly and output statistics are compiled. Output errors resulting from the various error sources are presented as a function of Mach number. Final uncertainties with all modeled error sources included are presented as a function of Mach number.

  20. Spaceliner Class Operability Gains Via Combined Airbreathing/ Rocket Propulsion: Summarizing an Operational Assessment of Highly Reusable Space Transports

    NASA Technical Reports Server (NTRS)

    Nix, Michael B.; Escher, William J. d.

    1999-01-01

    In discussing a new NASA initiative in advanced space transportation systems and technologies, the Director of the NASA Marshall Space Flight Center, Arthur G. Stephenson, noted that, "It would use new propulsion technology, air-breathing engine so you don't have to carry liquid oxygen, at least while your flying through the atmosphere. We are calling it Spaceliner 100 because it would be 100 times cheaper, costing $ 100 dollars a pound to orbit." While airbreathing propulsion is directly named, rocket propulsion is also implied by, "... while you are flying through the atmosphere." In-space final acceleration to orbital speed mandates rocket capabilities. Thus, in this informed view, Spaceliner 100 will be predicated on combined airbreathing/rocket propulsion, the technical subject of this paper. Interestingly, NASA's recently concluded Highly Reusable Space Transportation (HRST) study focused on the same affordability goal as that of the Spaceliner 100 initiative and reflected the decisive contribution of combined propulsion as a way of expanding operability and increasing the design robustness of future space transports, toward "aircraft like" capabilities. The HRST study built on the Access to Space Study and the Reusable Launch Vehicle (RLV) development activities to identify and characterize space transportation concepts, infrastructure and technologies that have the greatest potential for reducing delivery cost by another order of magnitude, from $1,000 to $100-$200 per pound for 20,000 lb. - 40.000 lb. payloads to low earth orbit (LEO). The HRST study investigated a number of near-term, far-term, and very far-term launch vehicle concepts including all-rocket single-stage-to-orbit (SSTO) concepts, two-stage-to-orbit (TSTO) concepts, concepts with launch assist, rocket-based combined cycle (RBCC) concepts, advanced expendable vehicles, and more far term ground-based laser powered launchers. The HRST study consisted of preliminary concept studies, assessments