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.

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.

Hypersonic Boost Glider

NASA Image and Video Library

1957-04-15

Hypersonic Boost Glider in 11 Inch Hypersonic Tunnel L57-1681 In 1957 Langley tested its HYWARDS design in the 11 Inch Hypersonic Tunnel. Photograph published in Engineer in Charge: A History of the Langley Aeronautical Laboratory, 1917-1958 by James R. Hansen. Page 369.

Hypersonic research engine project. Phase 2: Aerothermodynamic integration model development, data item no. 55-4-21

NASA Technical Reports Server (NTRS)

Jilly, L. F. (Editor)

1975-01-01

The design and development of the Aerothermodynamic Integration Model (AIM) of the Hypersonic Research Engine (HRE) is described. The feasibility of integrating the various analytical and experimental data available for the design of the hypersonic ramjet engine was verified and the operational characteristic and the overall performance of the selected design was determined. The HRE-AIM was designed for operation at speeds of Mach 3 through Mach 8.

Polymer, metal and ceramic matrix composites for advanced aircraft engine applications

NASA Technical Reports Server (NTRS)

Mcdanels, D. L.; Serafini, T. T.; Dicarlo, J. A.

1985-01-01

Advanced aircraft engine research within NASA Lewis is being focused on propulsion systems for subsonic, supersonic, and hypersonic aircraft. Each of these flight regimes requires different types of engines, but all require advanced materials to meet their goals of performance, thrust-to-weight ratio, and fuel efficiency. The high strength/weight and stiffness/weight properties of resin, metal, and ceramic matrix composites will play an increasingly key role in meeting these performance requirements. At NASA Lewis, research is ongoing to apply graphite/polyimide composites to engine components and to develop polymer matrices with higher operating temperature capabilities. Metal matrix composites, using magnesium, aluminum, titanium, and superalloy matrices, are being developed for application to static and rotating engine components, as well as for space applications, over a broad temperature range. Ceramic matrix composites are also being examined to increase the toughness and reliability of ceramics for application to high-temperature engine structures and components.

Unstart Coupling Mechanism Analysis of Multiple-Modules Hypersonic Inlet

PubMed Central

Wang, Lei; Cao, Shibin

2013-01-01

The combination of multiplemodules in parallel manner is an important way to achieve the much higher thrust of scramjet engine. For the multiple-modules scramjet engine, when inlet unstarted oscillatory flow appears in a single-module engine due to high backpressure, how to interact with each module by massflow spillage, and whether inlet unstart occurs in other modules are important issues. The unstarted flowfield and coupling characteristic for a three-module hypersonic inlet caused by center module II and side module III were, conducted respectively. The results indicate that the other two hypersonic inlets are forced into unstarted flow when unstarted phenomenon appears on a single-module hypersonic inlet due to high backpressure, and the reversed flow in the isolator dominates the formation, expansion, shrinkage, and disappearance of the vortexes, and thus, it is the major factor of unstart coupling of multiple-modules hypersonic inlet. The coupling effect among multiple modules makes hypersonic inlet be more likely unstarted. PMID:24348146

Scramjet analysis, testing

NASA Technical Reports Server (NTRS)

Leingang, J. L.; Stull, F. D.

1992-01-01

A survey of supersonic combustion ramjet (scramjet) engine development in the US covers development of this unique engine cycle from its inception in the early 1960's through the various programs currently being pursued and, in some instances, describing the future direction of the programs. These include developmental efforts supported by the US Navy, NASA, and US Air Force. Results of inlet, combustor, and nozzle component tests, free-jet engine tests, analytical techniques developed to analyze and predict component and engine performance, and flight-weight hardware development are presented. These results show that efficient scramjet propulsion is attainable in a variety of flight configurations with a variety of fuels. Since the scramjet is the most efficient engine cycle for hypersonic flight within the atmosphere, it should be given serious consideration in future propulsion schemes.

NASA's hypersonic propulsion program: History and direction

NASA Technical Reports Server (NTRS)

Wander, Steve

1992-01-01

Research into hypersonic propulsion; i.e., supersonic combustion, was seriously initiated at the Langley Research Center in the 1960's with the Hypersonic Research Engine (HRE) project. This project was designed to demonstrate supersonic combustion within the context of an engine module consisting of an inlet, combustor, and nozzle. In addition, the HRE utilized both subsonic and supersonic combustion (dual-mode) to demonstrate smooth operation over a Mach 4 to 7 speed range. The propulsion program thus concentrated on fundamental supersonic combustion studies and free jet propulsion tests for the three dimensional fixed geometry engine design to demonstrate inlet and combustor integration and installed performance potential. The developmental history of the program is presented. Additionally, the HRE program's effect on the current state of hypersonic propulsion is discussed.

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 (including industry and academia) the hypersonic uses both NASA Research Announcements (NRAs) and a jointly sponsored, Air Force Office of Scientific Research and NASA, National Hypersonic Science Center that are focused on propulsion research. Finally, these two disciplines use selected external partnership agreements with both governmental agencies and industrial entities. The TBCC discipline is comprised of analytic and experimental tasks, and is structured into the following two research topic areas: (1) TBCC Integrated Flowpath Technologies, and (2) TBCC Component Technologies. These tasks will provide experimental data to support design and analysis tool development and validation that will enable advances in TBCC technology.

The NASA hypersonic research engine program

NASA Technical Reports Server (NTRS)

Rubert, Kennedy F.; Lopez, Henry J.

1992-01-01

An overview is provided of the NASA Hypersonic Research Engine Program. The engine concept is described which was evolved, and the accomplishments of the program are summarized. The program was undertaken as an in-depth program of hypersonic airbreathing propulsion research to provide essential inputs to future prototype engine development and decision making. An airbreathing liquid hydrogen fueled research oriented scramjet was to be developed to certain performance goals. The work was many faceted, required aerodynamic design evaluation, structures development, and development of flight systems such as the fuel and control system, but the main objective was the study of the internal aerothermodynamics of the propulsion system.

Scramjet integration on hypersonic research airplane concepts

NASA Technical Reports Server (NTRS)

Weidner, J. P.; Small, W. J.; Penland, J. A.

1976-01-01

Several rocket-boosted research airplane concepts were evaluated with a research scramjet engine to determine their potential to provide research on critical aspects of airframe-integrated hypersonic systems. Extensive calculations to determine the force and moment contributions of the scramjet inlet, combustor, nozzle, and airframe were conducted to evaluate the overall performance of the combined engine/airframe system at hypersonic speeds. Results of both wind-tunnel tests and analysis indicate that it is possible to develop a research airplane configuration that will cruise at hypersonic speed on scramjet power alone, and will also have acceptable low-speed aerodynamic characteristics for landing.

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

NASA Astrophysics Data System (ADS)

Modlin, James Michael

An investigation was conducted to study the feasibility of cooling hypersonic vehicle leading edge structures exposed to severe aerodynamic surface heat fluxes 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. Two applications of the hypersonic leading edge cooling model were examined. An assumed aerospace plane-type wing leading edge section exposed to a severe laminar, hypersonic aerodynamic surface heat flux was studied. A second application of the hypersonic leading edge cooling model was conducted on an assumed one-quarter inch nose diameter SCRAMJET engine inlet leading edge section exposed to both a transient laminar, hypersonic aerodynamic surface heat flux and a type 4 shock interference surface heat flux. The investigation led to the conclusion 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 appeared feasible.

Frequencies of inaudible high-frequency sounds differentially affect brain activity: positive and negative hypersonic effects.

PubMed

Fukushima, Ariko; Yagi, Reiko; Kawai, Norie; Honda, Manabu; Nishina, Emi; Oohashi, Tsutomu

2014-01-01

The hypersonic effect is a phenomenon in which sounds containing significant quantities of non-stationary high-frequency components (HFCs) above the human audible range (max. 20 kHz) activate the midbrain and diencephalon and evoke various physiological, psychological and behavioral responses. Yet important issues remain unverified, especially the relationship existing between the frequency of HFCs and the emergence of the hypersonic effect. In this study, to investigate the relationship between the hypersonic effect and HFC frequencies, we divided an HFC (above 16 kHz) of recorded gamelan music into 12 band components and applied them to subjects along with an audible component (below 16 kHz) to observe changes in the alpha2 frequency component (10-13 Hz) of spontaneous EEGs measured from centro-parieto-occipital regions (Alpha-2 EEG), which we previously reported as an index of the hypersonic effect. Our results showed reciprocal directional changes in Alpha-2 EEGs depending on the frequency of the HFCs presented with audible low-frequency component (LFC). When an HFC above approximately 32 kHz was applied, Alpha-2 EEG increased significantly compared to when only audible sound was applied (positive hypersonic effect), while, when an HFC below approximately 32 kHz was applied, the Alpha-2 EEG decreased (negative hypersonic effect). These findings suggest that the emergence of the hypersonic effect depends on the frequencies of inaudible HFC.

Review of NASA's Hypersonic Research Engine Project

NASA Technical Reports Server (NTRS)

Andrews, Earl H.; Mackley, Ernest A.

1993-01-01

The goals of the NASA Hypersonic Research Engine (HRE) Project, which began in 1964, were to design, develop, and construct a hypersonic research ramjet/scramjet engine for high performance and to flight-test the developed concept over the speed range from Mach 3 to 8. The project was planned to be accomplished in three phases: project definition, research engine development, and flight test using the X-15A-2 research aircraft, which was modified to carry hydrogen fuel for the research engine. The project goal of an engine flight test was eliminated when the X-15 program was canceled in 1968. Ground tests of engine models then became the focus of the project. Two axisymmetric full-scale engine models having 18-inch-diameter cowls were fabricated and tested: a structural model and a combustion/propulsion model. A brief historical review of the project with salient features, typical data results, and lessons learned is presented.

X-34 Main Propulsion System-Selected Subsystem Analyses

NASA Technical Reports Server (NTRS)

Brown, T. M.; McDonald, J. P.; Knight, K. C.; Champion, R. H., Jr.

1998-01-01

The X-34 hypersonic flight vehicle is currently under development by Orbital Sciences Corporation (Orbital). The Main Propulsion System (MPS) has been designed around the liquid propellant Fastrac rocket engine currently under development at NASA Marshall Space Flight Center. This paper presents selected analyses of MPS subsystems and components. Topics include the integration of component and system level modeling of the LOX dump subsystem and a simple terminal bubble velocity analysis conducted to guide propellant feed line design.

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.

Status of turbulence modeling for hypersonic propulsion flowpaths

NASA Astrophysics Data System (ADS)

Georgiadis, Nicholas J.; Yoder, Dennis A.; Vyas, Manan A.; Engblom, William A.

2014-06-01

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 methods 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.

Hypersonic research engine/aerothermodynamic integration model, experimental results. Volume 2: Mach 6 performance

NASA Technical Reports Server (NTRS)

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

1976-01-01

Computer program performance results of a Mach 6 hypersonic research engine during supersonic and subsonic combustion modes were presented. The combustion mode transition was successfully performed, exit surveys made, and effects of altitude, angle of attack, and inlet spike position were determined during these tests.

NASA's Hypersonic Research Engine Project: A review

NASA Technical Reports Server (NTRS)

Andrews, Earl H.; Mackley, Ernest A.

1994-01-01

The goals of the NASA Hypersonic Research Engine (HRE) Project, which began in 1964, were to design, develop, and construct a high-performance hypersonic research ramjet/scramjet engine for flight tests of the developed concept over the speed range of Mach 4 to 8. The project was planned to be accomplished in three phases: project definition, research engine development, and flight test using the X-15A-2 research airplane, which was modified to carry hydrogen fuel for the research engine. The project goal of an engine flight test was eliminated when the X-15 program was canceled in 1968. Ground tests of full-scale engine models then became the focus of the project. Two axisymmetric full-scale engine models, having 18-inch-diameter cowls, were fabricated and tested: a structural model and combustion/propulsion model. A brief historical review of the project, with salient features, typical data results, and lessons learned, is presented. An extensive number of documents were generated during the HRE Project and are listed.

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.

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

NASA Astrophysics Data System (ADS)

Higbea, Mary E.; Vedda, James A.

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.

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

DTIC Science & Technology

2010-08-01

already being processed for leading edge primary structures and engine components for present and future stealth aircraft. In addition to describing our...The form of entry is the last name, first name, middle initial, and additional qualifiers separated by commas, e.g. Smith, Richard, Jr. 7...availability or distribution limitations of the report. If additional limitations/restrictions or special markings are indicated, follow agency

STUDY PROGRAM FOR TURBO-COOLER FOR PRODUCING ENGINE COOLING AIR.

DTIC Science & Technology

VANES , STAGNATION POINT, DECELERATION, ACCELERATION, SUPERSONIC DIFFUSERS, TURBINE BLADES , EVAPOTRANSPIRATION, LIQUID COOLED, HEAT TRANSFER, GAS BEARINGS, SEALS...HYPERSONIC AIRCRAFT , COOLING + VENTILATING EQUIPMENT), (*GAS TURBINES , COOLING + VENTILATING EQUIPMENT), HYPERSONIC FLOW, AIR COOLED, AIRCRAFT ... ENGINES , FEASIBILITY STUDIES, PRESSURE, SUPERSONIC CHARACTERISTICS, DESIGN, HEAT EXCHANGERS, COOLING (U) AXIAL FLOW TURBINES , DUCT INLETS, INLET GUIDE

The Need for Speed. Hypersonic Aircraft and the Transformation of Long Range Airpower

DTIC Science & Technology

2005-06-01

lxxii Heppenheimer , Thomas A . “Origins of Hypersonic Propulsion: A Personal History.” Air Power History 47, no. 3 (Fall 2000), 14-27. Hicks, John W...early Aerospace plane research. 59 Thomas A . Heippenheimer, “Origins of Hypersonic Propulsion: A Personal History,” Air Power History 47, no. 3 (Fall...experiments with hypersonic engines. 65 Ibid. 66 Thomas A . Heippenheimer, 15. 67 Ibid., 17-18. Lockheed and General Applied Science Laboratories

Hypersonic engine component experiments in high heat flux, supersonic flow environment

NASA Technical Reports Server (NTRS)

Gladden, Herbert J.; Melis, Matthew E.

1993-01-01

A major concern in advancing the state-of-the-art technologies for hypersonic vehicles is the development of an aeropropulsion system capable of withstanding the sustained high thermal loads expected during hypersonic flight. Even though progress has been made in the computational understanding of fluid dynamics and the physics/chemistry of high speed flight, there is also a need for experimental facilities capable of providing a high heat flux environment for testing component concepts and verifying/calibrating these analyses. A hydrogen/oxygen rocket engine heat source was developed at the NASA Lewis Research Center as one element in a series of facilities at national laboratories designed to fulfill this need. This 'Hot Gas Facility' is capable of providing heat fluxes up to 450 w/sq cm on flat surfaces and up to 5,000 w/sq cm at the leading edge stagnation point of a strut in a supersonic flow stream. Gas temperatures up to 3050 K can also be attained. Two recent experimental programs conducted in this facility are discussed. The objective of the first experiment is to evaluate the erosion and oxidation characteristics of a coating on a cowl leading edge (or strut leading edge) in a supersonic, high heat flux environment. Macrophotographic data from a coated leading edge model show progressive degradation over several thermal cycles at aerothermal conditions representative of high Mach number flight. The objective of the second experiment is to assess the capability of cooling a porous surface exposed to a high temperature, high velocity flow environment and to provide a heat transfer data base for a design procedure. Experimental results from transpiration cooled surfaces in a supersonic flow environment are presented.

Scramjet nozzle design and analysis as applied to a highly integrated hypersonic research airplane

NASA Technical Reports Server (NTRS)

Small, W. J.; Weidner, J. P.; Johnston, P. J.

1976-01-01

Engine-nozzle airframe integration at hypersonic speeds was conducted by using a high-speed research aircraft concept as a focus. Recently developed techniques for analysis of scramjet-nozzle exhaust flows provide a realistic analysis of complex forces resulting from the engine-nozzle airframe coupling. By properly integrating the engine-nozzle propulsive system with the airframe, efficient, controlled and stable flight results over a wide speed range.

Hypersonic propulsion. [scramjet technology

NASA Technical Reports Server (NTRS)

Anderson, G. Y.

1975-01-01

The NASA research on scramjet technology for speeds above Mach 5 is reported. A brief overview of the NASA Hypersonic Research Engine (HRE) project is presented with emphasis on the most significant contributions on the HRE to scramjet technology. The work on high-performance airframe-integrated scramjet engines is described, and a new versatile research airplane is discussed with emphasis on propulsion.

Compact Analyzer/Controller For Oxygen-Enrichment System

NASA Technical Reports Server (NTRS)

Puster, Richard L.; Singh, Jag J.; Sprinkle, Danny R.

1990-01-01

System controls hypersonic air-breathing engine tests. Compact analyzer/controller developed, built, and tested in small-scale wind tunnel prototype of the 8' HTT (High-Temperature Tunnel). Monitors level of oxygen and controls addition of liquid oxygen to enrich atmosphere for combustion. Ensures meaningful ground tests of hypersonic engines in range of speeds from mach 4 to mach 7.

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 integration. Additionally, areas are identified in this study where improvements could be made to enhance the performance. Both fully-integrated configurations are longitudinally unstable over the Mach number range studied for unpowered conditions. The cranked-wing fully-integrated configuration provided significantly better lateral-directional stability characteristics than the straight-wing configuration.

NASA Hypersonics Overview

NASA Technical Reports Server (NTRS)

Dryer, Jay

2017-01-01

This briefing is an overview of NASA's hypersonic portfolio and core capabilities. The scope of work is fundamental research spanning technology readiness and system complexity levels; critical technologies enabling re-usable hypersonic systems; system-level research, design, analysis, validation; and, engage, invigorate and train the next generation of engineers. This briefing was requested by the Aeronautics Subcommittee of the NASA Advisory Council.

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.

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.

Advanced Concept

NASA Image and Video Library

2002-01-01

An artist's rendering of the air-breathing, hypersonic X-43B, the third and largest of NASA's Hyper-X series flight demonstrators, which could fly later this decade. Revolutionizing the way we gain access to space is NASA's primary goal for the Hypersonic Investment Area, managed for NASA by the Advanced Space Transportation Program at the Marshall Space Flight Center in Huntsville, Alabama. The Hypersonic Investment area, which includes leading-edge partners in industry and academia, will support future generation reusable vehicles and improved access to space. These technology demonstrators, intended for flight testing by decade's end, are expected to yield a new generation of vehicles that routinely fly about 100,000 feet above Earth's surface and reach sustained speeds in excess of Mach 5 (3,750 mph), the point at which "supersonic" flight becomes "hypersonic" flight. The flight demonstrators, the Hyper-X series, will be powered by air-breathing rocket or turbine-based engines, and ram/scramjets. Air-breathing engines, known as combined-cycle systems, achieve their efficiency gains over rocket systems by getting their oxygen for combustion from the atmosphere, as opposed to a rocket that must carry its oxygen. Once a hypersonic vehicle has accelerated to more than twice the speed of sound, the turbine or rockets are turned off, and the engine relies solely on oxygen in the atmosphere to burn fuel. When the vehicle has accelerated to more than 10 to 15 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 series of hypersonic flight demonstrators includes three air-breathing vehicles: the X-43A, X-43B and X-43C.

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.

Survey of supersonic combustion ramjet research at Langley

NASA Technical Reports Server (NTRS)

Northam, G. B.; Anderson, G. Y.

1986-01-01

The Hypersonic Propulsion Branch at NASA Langley Research Center has maintained an active research program in supersonic combustion ramjet (scramjet) and high speed ramjet propulsion since the 1960s. The focus for this research has centered on propulsion for manned reuseable vehicles with cryogenic hydrogen fuel. This paper presents some highlights of this research. The design philosophy of the Langley fixed-geometry airframe-integrated modular scramjet is discussed. The component development and research programs that have supported the successful demonstration of the engine concept using subscale engine module hardware is reviewed and a brief summary of the engine tests presented. An extensive bibliography of research supported by the Langley program is also included.

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.

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.

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.

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.

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.

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.

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.

Controls and guidance research

NASA Technical Reports Server (NTRS)

Homaifar, Abdollah; Dunn, Derome; Song, Yong-Duan; Lai, Steven H.-Y.

1992-01-01

The objectives of the control group are concentrated on research and education. The control problem of the hypersonic space vehicle represents an important and challenging issue in aerospace engineering. The work described in this report is part of our effort in developing advanced control strategies for such a system. In order to achieve the objectives stated in the NASA-CORE proposal, the tasks were divided among the group based upon their educational expertise. Within the educational component we are offering a Linear Systems and Control course for students in electrical and mechanical engineering. Also, we are proposing a new course in Digital Control Systems with a corresponding laboratory.

Hypersonic trajectory control of aerospace plane with integrated SCRAMJET engine

NASA Astrophysics Data System (ADS)

Yonemoto, Koichi

The aerospace plane is an airbreathing 'propulsion configured' vehicle having proper forebody contour for inflow pre-compression to the inlet and afterbody that operates as an external expansion nozzle. Since the whole lower side of the body acts as important compression and expansion elements for the airbreathing engine, the flight attitude influences its performance such as specific impulse and thrust coefficient considerably. The stability of ascent trajectory controlling dynamic pressure or heat-input rate is analyzed considering the performance change due to attitude fluctuation. The performance of scramjet engine, a typical hypersonic airbreathing engine, is estimated by a rapid prediction methodology of the combustor proposed by Ikawa.

Analysis of experimental results of the inlet for the NASA hypersonic research engine aerothermodynamic integration model. [wind tunnel tests of ramjet engine hypersonic inlets

NASA Technical Reports Server (NTRS)

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

1976-01-01

An aerodynamic engine inlet analysis was performed on the experimental results obtained at nominal Mach numbers of 5, 6, and 7 from the NASA Hypersonic Research Engine (HRE) Aerothermodynamic Integration Model (AIM). Incorporation on the AIM of the mixed-compression inlet design represented the final phase of an inlet development program of the HRE Project. The purpose of this analysis was to compare the AIM inlet experimental results with theoretical results. Experimental performance was based on measured surface pressures used in a one-dimensional force-momentum theorem. Results of the analysis indicate that surface static-pressure measurements agree reasonably well with theoretical predictions except in the regions where the theory predicts large pressure discontinuities. Experimental and theoretical results both based on the one-dimensional force-momentum theorem yielded inlet performance parameters as functions of Mach number that exhibited reasonable agreement. Previous predictions of inlet unstart that resulted from pressure disturbances created by fuel injection and combustion appeared to be pessimistic.

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.

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.

Engineers and technicians in the control room at the Dryden Flight Research Center must constantly monitor critical operations and checks during research projects like NASA's hypersonic X-43A

NASA Image and Video Library

2004-01-24

Engineers and technicians in the control room at the Dryden Flight Research Center must constantly monitor critical operations and checks during research projects like NASA's hypersonic X-43A. Visible in the photo, taken two days before the X-43's captive carry flight in January 2004, are [foreground to background]; Tony Kawano (Range Safety Officer), Brad Neal (Mission Controller), and Griffin Corpening (Test Conductor).

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.

Hypersonic Glider Model in Full Scale Tunnel 1957

NASA Image and Video Library

1957-09-07

L57-1439 A model based on Langley s concept of a hypersonic glider was test flown on an umbilical cord inside the Full Scale Tunnel in 1957. Photograph published in Engineer in Charge: A History of the Langley Aeronautical Laboratory, 1917-1958 by James R. Hansen. Page 374.

X-43D Conceptual Design and Feasibility Study

NASA Technical Reports Server (NTRS)

Johnson, Donald B.; Robinson, Jeffrey S.

2005-01-01

NASA s Next Generation Launch Technology (NGLT) Program, in conjunction with the office of the Director of Defense Research and Engineering (DDR&E), developed an integrated hypersonic technology demonstration roadmap. This roadmap is an integral part of the National Aerospace Initiative (NAI), a multi-year, multi-agency cooperative effort to invest in and develop, among other things, hypersonic technologies. This roadmap contains key ground and flight demonstrations required along the path to developing a reusable hypersonic space access system. One of the key flight demonstrations required for systems that will operate in the high Mach number regime is the X-43D. As currently conceived, the X-43D is a Mach 15 flight test vehicle that incorporates a hydrogen-fueled scramjet engine. The purpose of the X-43D is to gather high Mach number flight environment and engine operability information which is difficult, if not impossible, to gather on the ground. During 2003, the NGLT Future Hypersonic Flight Demonstration Office initiated a feasibility study on the X-43D. The objective of the study was to develop a baseline conceptual design, assess its performance, and identify the key technical issues. The study also produced a baseline program plan, schedule, and cost, along with a list of key programmatic risks.

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.

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.

Analysis of a hypersonic waverider research vehicle with a hydrocarbon scramjet engine

NASA Technical Reports Server (NTRS)

Molvik, Gregory A.; Bowles, Jeffrey V.; Huynh, Loc C.

1993-01-01

The results of a feasibility study of a hypersonic waverider research vehicle with a hydrocarbon scramjet engine are presented. The integrated waverider/scramjet geometry is first optimized with a vehicle synthesis code to produce a maximum product of the lift-to-drag ratio and the cycle specific impulse, hence cruise range. Computational fluid dynamics (CFD) is then employed to provide a nose-to-tail analysis of the system at the on-design conditions. Some differences are noted between the results of the two analysis techniques. A comparison of experimental, engineering analysis and CFD results on a waverider forebody are also included for validation.

A survey of heating and turbulent boundary layer characteristics of several hypersonic research aircraft configurations

NASA Technical Reports Server (NTRS)

Lawing, P. L.

1981-01-01

Four of the configurations investigated during a proposed NASA-Langley hypersonic research aircraft program were selected for phase-change-paint heat-transfer testing and forebody boundary layer pitot surveys. In anticipation of future hypersonic aircraft, both published and unpublished data and results are reviewed and presented with the purpose of providing a synoptic heat-transfer data base from the research effort. Engineering heat-transfer predictions are compared with experimental data on both a global and a local basis. The global predictions are shown to be sufficient for purposes of configuration development, and even the local predictions can be adequate when interpreted in light of the proper flow field. In that regard, cross flow in the forebody boundary layers was examined for significant heating and aerodynamic effect on the scramjet engines. A design philosophy which evolved from the research airplane effort is used to design a forebody shape that produces thin, uniform, forebody boundary layers on a hypersonic airbreathing missile. Finally, heating/boundary layer phenomena which are not predictable with state-of-the-art knowledge and techniques are shown and discussed.

Analytical and experimental studies of heat pipe radiation cooling of hypersonic propulsion systems

NASA Technical Reports Server (NTRS)

Martin, R. A.; Merrigan, M. A.; Elder, M. G.; Sena, J. T.; Keddy, E. S.; Silverstein, C. C.

1992-01-01

Analytical and experimental studies were completed to assess the feasibility of using high-temperature heat pipes to cool hypersonic engine components. This new approach involves using heat pipes to transport heat away from the combustor, nozzle, or inlet regions, and to reject it to the environment by thermal radiation from an external heat pipe nacelle. For propulsion systems using heat pipe radiation cooling (HPRC), it is possible to continue to use hydrocarbon fuels into the Mach 4 to Mach 6 speed range, thereby enhancing the economic attractiveness of commercial or military hypersonic flight. In the second-phase feasibility program recently completed, it is found that heat loads produced by considering both convection and radiation heat transfer from the combustion gas can be handled with HPRC design modifications. The application of thermal insulation to ramburner and nozzle walls was also found to reduce the heat load by about one-half and to reduce peak HPRC system temperatures to below 2700 F. In addition, the operation of HPRC at cruise conditions of around Mach 4.5 and at an altitude of 90,000 ft lowers the peak hot-section temperatures to around 2800 F. An HPRC heat pipe was successfully fabricated and tested at Mach 5 conditions of heat flux, heat load, and temperature.

Supersonic combustion engine testbed, heat lightning

NASA Technical Reports Server (NTRS)

Hoying, D.; Kelble, C.; Langenbahn, A.; Stahl, M.; Tincher, M.; Walsh, M.; Wisler, S.

1990-01-01

The design of a supersonic combustion engine testbed (SCET) aircraft is presented. The hypersonic waverider will utilize both supersonic combustion ramjet (SCRAMjet) and turbofan-ramjet engines. The waverider concept, system integration, electrical power, weight analysis, cockpit, landing skids, and configuration modeling are addressed in the configuration considerations. The subsonic, supersonic and hypersonic aerodynamics are presented along with the aerodynamic stability and landing analysis of the aircraft. The propulsion design considerations include: engine selection, turbofan ramjet inlets, SCRAMjet inlets and the SCRAMjet diffuser. The cooling requirements and system are covered along with the topics of materials and the hydrogen fuel tanks and insulation system. A cost analysis is presented and the appendices include: information about the subsonic wind tunnel test, shock expansion calculations, and an aerodynamic heat flux program.

Airframe-integrated propulsion system for hypersonic cruise vehicles

NASA Technical Reports Server (NTRS)

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

1978-01-01

The paper describes a new hydrogen-burning airframe-integrated scramjet concept which offers good potential for efficient hypersonic cruise vehicles. The characteristics of the engine which assure good performance are extensive engine-airframe integration, fixed geometry, low cooling, and control of heat release in the supersonic combustor by mixed modes of fuel injection from the combustor entrance. The present paper describes the concept and presents results from inlet tests, direct-connect combustor tests, and tests of two subscale boiler-plate research engines currently underway under conditions which simulate flight at Mach 4 and 7. It is concluded that this engine concept has the potential for high thrust and efficiency, low drag and weight, low cooling requirement, and application to a wide range of vehicle sizes.

Hypersonic CFD applications for the National Aero-Space Plane

NASA Technical Reports Server (NTRS)

Richardson, Pamela F.; Mcclinton, Charles R.; Bittner, Robert D.; Dilley, A. Douglas; Edwards, Kelvin W.

1989-01-01

Design and analysis of the NASP depends heavily upon developing the critical technology areas that cover the entire engineering design of the vehicle. These areas include materials, structures, propulsion systems, propellants, integration of airframe and propulsion systems, controls, subsystems, and aerodynamics areas. Currently, verification of many of the classical engineering tools relies heavily on computational fluid dynamics. Advances are being made in the development of CFD codes to accomplish nose-to-tail analyses for hypersonic aircraft. Additional details involving the partial development, analysis, verification, and application of the CFL3D code and the SPARK combustor code are discussed. A nonequilibrium version of CFL3D that is presently being developed and tested is also described. Examples are given of portion calculations for research hypersonic aircraft geometries and comparisons with experiment data show good agreement.

An inlet analysis for the NASA hypersonic research engine aerothermodynamic integration model

NASA Technical Reports Server (NTRS)

Andrews, E. H., Jr.; Russell, J. W.; Mackley, E. A.; Simmonds, A. L.

1974-01-01

A theoretical analysis for the inlet of the NASA Hypersonic Research Engine (HRE) Aerothermodynamic Integration Model (AIM) has been undertaken by use of a method-of-characteristics computer program. The purpose of the analysis was to obtain pretest information on the full-scale HRE inlet in support of the experimental AIM program (completed May 1974). Mass-flow-ratio and additive-drag-coefficient schedules were obtained that well defined the range effected in the AIM tests. Mass-weighted average inlet total-pressure recovery, kinetic energy efficiency, and throat Mach numbers were obtained.

Analytical and experimental validation of the Oblique Detonation Wave Engine concept

NASA Technical Reports Server (NTRS)

Adelman, Henry G.; Cambier, Jean-Luc; Menees, Gene P.; Balboni, John A.

1988-01-01

The Oblique Detonation Wave Engine (ODWE) for hypersonic flight has been analytically studied by NASA using the CFD codes which fully couple finite rate chemistry with fluid dynamics. Fuel injector designs investigated included wall and strut injectors, and the in-stream strut injectors were chosen to provide good mixing with minimal stagnation pressure losses. Plans for experimentally validating the ODWE concept in an arc-jet hypersonic wind tunnel are discussed. Measurements of the flow field properties behind the oblique wave will be compared to analytical predictions.

Development of braided rope seals for hypersonic engine applications. Part 2: Flow modeling

NASA Technical Reports Server (NTRS)

Mutharasan, Rajakkannu; Steinetz, Bruce M.; Tao, Xiaoming; Ko, Frank

1991-01-01

Two models based on the Kozeny-Carmen equation were developed to analyze the fluid flow through a new class of braided rope seals under development for advanced hypersonic engines. A hybrid seal geometry consisting of a braided sleeve and a substantial amount of longitudinal fibers with high packing density was selected for development based on its low leakage rates. The models developed allow prediction of the gas leakage rate as a function of fiber diameter, fiber packing density, gas properties, and pressure drop across the seal.

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.

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.

Hypersonic research engine project. Phase 2: Aerothermodynamic Integration Model (AIM) data reduction computer program, data item no. 54.16

NASA Technical Reports Server (NTRS)

Gaede, A. E.; Platte, W. (Editor)

1975-01-01

The data reduction program used to analyze the performance of the Aerothermodynamic Integration Model is described. Routines to acquire, calibrate, and interpolate the test data, to calculate the axial components of the pressure area integrals and the skin function coefficients, and to report the raw data in engineering units are included along with routines to calculate flow conditions in the wind tunnel, inlet, combustor, and nozzle, and the overall engine performance. Various subroutines were modified and used to obtain species concentrations and transport properties in chemical equilibrium at each of the internal and external engine stations. It is recommended that future test plans include the configuration, calibration, and channel assignment data on a magnetic tape generated at the test site immediately before or after a test, and that the data reduction program be designed to operate in a batch environment.

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.

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

Requirements for facilities and measurement techniques to support CFD development for hypersonic aircraft

NASA Technical Reports Server (NTRS)

Sellers, William L., III; Dwoyer, Douglas L.

1992-01-01

The design of a hypersonic aircraft poses unique challenges to the engineering community. Problems with duplicating flight conditions in ground based facilities have made performance predictions risky. Computational fluid dynamics (CFD) has been proposed as an additional means of providing design data. At the present time, CFD codes are being validated based on sparse experimental data and then used to predict performance at flight conditions with generally unknown levels of uncertainty. This paper will discuss the facility and measurement techniques that are required to support CFD development for the design of hypersonic aircraft. Illustrations are given of recent success in combining experimental and direct numerical simulation in CFD model development and validation for hypersonic perfect gas flows.

Development of hypersonic engine seals: Flow effects of preload and engine pressures

NASA Technical Reports Server (NTRS)

Cai, Zhong; Mutharasan, Rajakkannu; Ko, Frank K.; Steinetz, Bruce M.

1993-01-01

A new type of engine seal is being developed to meet the needs of advanced hypersonic engines. A seal braided of emerging high temperature ceramic fibers comprised of a sheath-core construction was selected for study based on its low leakage rates. Flexible, low-leakage, high temperature seals are required to seal the movable engine panels of advanced ramjet-scramjet engines either preventing potentially dangerous leakage into backside engine cavities or limiting the purge coolant flow rates through the seals. To predict the leakage through these flexible, porous seal structures as a function of preload and engine pressures, new analytical flow models are required. An empirical leakage resistance/preload model is proposed to characterize the observed decrease in leakage with increasing preload. Empirically determined compression modulus and preload factor are used to correlate experimental leakage data for a wide range of seal architectures. Good agreement between measured and predicted values are observed over a range of engine pressures and seal preloads.

The Use of Steady and Unsteady Detonation Waves for Propulsion Systems

NASA Technical Reports Server (NTRS)

Adelman, Henry G.; Menees, Gene P.; Cambier, Jean-Luc; Bowles, Jeffrey V.; Cavolowsky, John A. (Technical Monitor)

1995-01-01

Detonation wave enhanced supersonic combustors such as the Oblique Detonation Wave Engine (ODWE) are attractive propulsion concepts for hypersonic flight. These engines utilize detonation waves to enhance fuel-air mixing and combustion. The benefits of wave combustion systems include shorter and lighter engines which require less cooling and generate lower internal drag. These features allow air-breathing operation at higher Mach numbers than the diffusive burning scramjet delaying the need for rocket engine augmentation. A comprehensive vehicle synthesis code has predicted the aerodynamic characteristics and structural size and weight of a typical single-stage-to-orbit vehicle using an ODWE. Other studies have focused on the use of unsteady or pulsed detonation waves. For low speed applications, pulsed detonation engines (PDE) have advantages in low weight and higher efficiency than turbojets. At hypersonic speeds, the pulsed detonations can be used in conjunction with a scramjet type engine to enhance mixing and provide thrust augmentation.

Aerodynamic characteristics at Mach 6 of a hypersonic research airplane concept having a 70 deg swept delta wing

NASA Technical Reports Server (NTRS)

Clark, L. E.; Richie, C. B.

1977-01-01

The hypersonic aerodynamic characteristics of an air-launched, delta-wing research aircraft concept were investigated at Mach 6. The effect of various components such as nose shape, wing camber, wing location, center vertical tail, wing tip fins, forward delta wing, engine nacelle, and speed brakes was also studied. Tests were conducted with a 0.021 scale model at a Reynolds number, based on model length, of 10.5 million and over an angel of attack range from -4 deg to 20 deg. Results show that most configurations with a center vertical tail have static longitudinal stability at trim, static directional stability at angles of attack up to 12 deg, and static lateral stability throughout the angle of attack range. Configurations with wing tip fins generally have static longitudinal stability at trim, have lateral stability at angles of attack above 8 deg, and are directionally unstable over the angle of attack range.

Advanced instrumentation for next-generation aerospace propulsion control systems

NASA Technical Reports Server (NTRS)

Barkhoudarian, S.; Cross, G. S.; Lorenzo, Carl F.

1993-01-01

New control concepts for the next generation of advanced air-breathing and rocket engines and hypersonic combined-cycle propulsion systems are analyzed. The analysis provides a database on the instrumentation technologies for advanced control systems and cross matches the available technologies for each type of engine to the control needs and applications of the other two types of engines. Measurement technologies that are considered to be ready for implementation include optical surface temperature sensors, an isotope wear detector, a brushless torquemeter, a fiberoptic deflectometer, an optical absorption leak detector, the nonintrusive speed sensor, and an ultrasonic triducer. It is concluded that all 30 advanced instrumentation technologies considered can be recommended for further development to meet need of the next generation of jet-, rocket-, and hypersonic-engine control systems.

A Concept for the HIFiRE 8 Flight Test

NASA Astrophysics Data System (ADS)

Alesi, H.; Paull, A.; Smart, M.; Bowcutt, K. G.

2015-09-01

HIFiRE 8 is a hypersonic flight test experiment scheduled for launch in late 2018 from the Woomera Test Center in Australia. This project aims to develop a Flight Test Vehicle that will, for the first time, complete 30 seconds of scramjet powered hypersonic flight at a Mach Number of 7.0. The engine used for this flight will be a rectangular to elliptic shape transition scramjet. It will be fuelled with gaseous hydrogen. The flight test engine configuration will be derived using scientific and engineering evaluation in the UQ shock tunnel T4 and other potential ground-based facilities. This paper presents current plans for the HIFiRE 8 trajectory, mission events, airframe and engine designs and also includes descriptions of critical subsystems and associated modelling, simulation and analysis activities.

Hypersonic Inlet for a Laser Powered Propulsion System

NASA Astrophysics Data System (ADS)

Harrland, Alan; Doolan, Con; Wheatley, Vincent; Froning, Dave

2011-11-01

Propulsion within the lightcraft concept is produced via laser induced detonation of an incoming hypersonic air stream. This process requires suitable engine configurations that offer good performance over all flight speeds and angles of attack to ensure the required thrust is maintained. Stream traced hypersonic inlets have demonstrated the required performance in conventional hydrocarbon fuelled scramjet engines, and has been applied to the laser powered lightcraft vehicle. This paper will outline the current methodology employed in the inlet design, with a particular focus on the performance of the lightcraft inlet at angles of attack. Fully three-dimensional turbulent computational fluid dynamics simulations have been performed on a variety of inlet configurations. The performance of the lightcraft inlets have been evaluated at differing angles of attack. An idealized laser detonation simulation has also been performed to validate that the lightcraft inlet does not unstart during the laser powered propulsion cycle.

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.

Revolutionary Materials for Hypersonic Flight

DTIC Science & Technology

2011-10-01

Transactions A (in preparation ). 16. M.D. Novak, CG. Levi, " Silicide Coatings on Niobium Substrates II: Thermodynamic Analysis and Mechanism of...These conditions impose significant challenges to material selection. While metallic materials are favored for their durability to thermal distortion...those of even the most highly-conducting engineering materials. While metallic systems are desirable for use in hypersonic vehicles for durability and

HyPlane for Space Tourism and Business Transportation

NASA Astrophysics Data System (ADS)

Savino, R.

In the present work a preliminary study on a small hypersonic airplane for a long duration space tourism mission is presented. It is also consistent with a point-to-point medium range (5000-6000 km) hypersonic trip, in the frame of the "urgent business travel" market segment. The main ideas is to transfer technological solutions developed for aeronautical and space atmospheric re-entry systems to the design of such a hypersonic airplane. A winged vehicle characterized by high aerodynamic efficiency and able to manoeuvre along the flight path, in all aerodynamic regimes encountered, is taken into consideration. Rocket-Based Combined Cycle and Turbine-Based Combined Cycle engines are investigated to ensure higher performances in terms of flight duration and range. Different flight-paths are also considered, including sub-orbital parabolic trajectories and steady state hypersonic cruise. The former, in particular, takes advantage of the high aerodynamic efficiency during the unpowered phase, in combination with a periodic engine actuation, to guarantee a long duration oscillating flight path. These trajectories offer Space tourists the opportunity of extended missions, characterized by repeated periods of low-gravity at altitudes high enough to ensure a wide view of the Earth from Space.

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.

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.

X-43A Hypersonic Experimental Vehicle - Artist Concept in Flight

NASA Technical Reports Server (NTRS)

1999-01-01

An artist's conception of the X-43A Hypersonic Experimental Vehicle, or 'Hyper-X' in flight. 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, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

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 the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

Artist Concept of X-43A/Hyper-X Hypersonic Experimental Research Vehicle in Flight

NASA Technical Reports Server (NTRS)

1998-01-01

An artist's conception of the X-43A Hypersonic Experimental Vehicle, or 'Hyper-X' in flight. 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, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

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-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

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 heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

Hot dynamic test rig for measuring hypersonic engine seal flow and durability

NASA Technical Reports Server (NTRS)

Miller, Jeffrey H.; Steinetz, Bruce M.; Sirocky, Paul J.; Kren, Lawrence A.

1994-01-01

A test fixture for measuring the dynamic performance of candidate high-temperature engine seal concepts was developed. The test fixture was developed to evaluate seal concepts under development for advanced hypersonic engines, such as those being considered for the National Aerospace Plane (NASP). The fixture can measure dynamic seal leakage performance from room temperature up to 840 C and air pressure differentials of to 0.7 MPa. Performance of the seals can be measured while sealing against flat or engine-simulated distorted walls. In the fixture, two seals are preloaded against the sides of a 0.3 m long saber that slides transverse to the axis of the seals, simulating the scrubbing motion anticipated in these engines. The capabilities of this text fixture along with preliminary data showing the dependence of seal leakage performance on high temperature cycling are covered.

Testing of the NASA Hypersonics Project Combined Cycle Engine Large Scale Inlet Mode Transition Experiment (CCE LlMX)

NASA Technical Reports Server (NTRS)

Saunders, J. D.; Stueber, T. J.; Thomas, S. R.; Suder, K. L.; Weir, L. J.; Sanders, B. W.

2012-01-01

Status on an effort to develop Turbine Based Combined Cycle (TBCC) propulsion is described. This propulsion technology can enable reliable and reusable space launch systems. TBCC propulsion offers improved performance and safety over rocket propulsion. The potential to realize aircraft-like operations and reduced maintenance are additional benefits. Among most the critical TBCC enabling technologies are: 1) mode transition from turbine to scramjet propulsion, 2) high Mach turbine engines and 3) TBCC integration. To address these TBCC challenges, the effort is centered on a propulsion mode transition experiment and includes analytical research. The test program, the Combined-Cycle Engine Large Scale Inlet Mode Transition Experiment (CCE LIMX), was conceived to integrate TBCC propulsion with proposed hypersonic vehicles. The goals address: (1) dual inlet operability and performance, (2) mode-transition sequences enabling a switch between turbine and scramjet flow paths, and (3) turbine engine transients during transition. Four test phases are planned from which a database can be used to both validate design and analysis codes and characterize operability and integration issues for TBCC propulsion. In this paper we discuss the research objectives, features of the CCE hardware and test plans, and status of the parametric inlet characterization testing which began in 2011. This effort is sponsored by the NASA Fundamental Aeronautics Hypersonics project

Hypersonic MHD Propulsion System Integration for the Mercury Lightcraft

NASA Astrophysics Data System (ADS)

Myrabo, L. N.; Rosa, R. J.

2004-03-01

Introduced herein are the design, systems integration, and performance analysis of an exotic magnetohydrodynamic (MHD) slipstream accelerator engine for a single-occupant ``Mercury'' lightcraft. This ultra-energetic, laser-boosted vehicle is designed to ride a `tractor beam' into space, transmitted from a future orbital network of satellite solar power stations. The lightcraft's airbreathing combined-cycle engine employs a rotary pulsed detonation thruster mode for lift-off & landing, and an MHD slipstream accelerator mode at hypersonic speeds. The latter engine transforms the transatmospheric acceleration path into a virtual electromagnetic `mass-driver' channel; the hypersonic momentum exchange process (with the atmosphere) enables engine specific impulses in the range of 6000 to 16,000 seconds, and propellant mass fractions as low as 10%. The single-stage-to-orbit, highly reusable lightcraft can accelerate at 3 Gs into low Earth orbit with its throttle just barely beyond `idle' power, or virtually `disappear' at 30 G's and beyond. The objective of this advanced lightcraft design is to lay the technological foundations for a safe, very low cost (e.g., 1000X below chemical rockets) air and space transportation for human life in the mid-21st Century - a system that will be completely `green' and independent of Earth's limited fossil fuel reserves.

An Overview of Brazilian Developments in Beamed Energy Aerospace Propulsion and Vehicle Performance Control

DOE Office of Scientific and Technical Information (OSTI.GOV)

Minucci, M. A. S.

Beamed energy propulsion and beamed energy vehicle performance control concepts are equally promising and challenging. In Brazil, the two concepts are being currently investigated at the Prof Henry T Nagamatsu Laboratory of Aerothermodynamics and Hypersonics, of the Institute for Advanced Studies--IEAv, in collaboration with the Rensselaer Polytechnic Institute--RPI, Troy, NY, and the United States Air force Research Laboratory-AFRL. Until recently, only laser energy addition for hypersonic flow control was being investigated at the Laboratory using a 0.3 m nozzle exit diameter hypersonic shock tunnel, T2, and two 7 joule CO{sub 2} TEA lasers. Flow visualization, model pressure and heat fluxmore » measurements of the laser energy addition perturbed flow around a model were produced as a result of this joint IEAv-RPI investigation. Presently, with the participation of AFRL and the newly commissioned 0.6 m. nozzle exit diameter hypersonic shock tunnel, T3, a more ambitious project is underway. Two 400 Joule Lumonics 620 CO{sub 2} TEA lasers will deliver a 20 cm X 25 cm propulsive laser beam to a complete laser propelled air breather/rocket hypersonic engine, located inside T3 test section. Schlieren photographs of the flow inside de engine as well as surface and heat flux measurements will be performed for free stream Mach numbers ranging from 6 to 25. The present paper discusses past, present and future Brazilian activities on beamed energy propulsion and related technologies.« less

An Overview of Brazilian Developments in Beamed Energy Aerospace Propulsion and Vehicle Performance Control

NASA Astrophysics Data System (ADS)

Minucci, M. A. S.

2008-04-01

Beamed energy propulsion and beamed energy vehicle performance control concepts are equally promising and challenging. In Brazil, the two concepts are being currently investigated at the Prof Henry T Nagamatsu Laboratory of Aerothermodynamics and Hypersonics, of the Institute for Advanced Studies—IEAv, in collaboration with the Rensselaer Polytechnic Institute—RPI, Troy, NY, and the United States Air force Research Laboratory-AFRL. Until recently, only laser energy addition for hypersonic flow control was being investigated at the Laboratory using a 0.3 m nozzle exit diameter hypersonic shock tunnel, T2, and two 7 joule CO2 TEA lasers. Flow visualization, model pressure and heat flux measurements of the laser energy addition perturbed flow around a model were produced as a result of this joint IEAv-RPI investigation. Presently, with the participation of AFRL and the newly commissioned 0.6 m. nozzle exit diameter hypersonic shock tunnel, T3, a more ambitious project is underway. Two 400 Joule Lumonics 620 CO2 TEA lasers will deliver a 20 cm X 25 cm propulsive laser beam to a complete laser propelled air breather/rocket hypersonic engine, located inside T3 test section. Schlieren photographs of the flow inside de engine as well as surface and heat flux measurements will be performed for free stream Mach numbers ranging from 6 to 25. The present paper discusses past, present and future Brazilian activities on beamed energy propulsion and related technologies.

High temperature dynamic engine seal technology development

NASA Technical Reports Server (NTRS)

Steinetz, Bruce M.; Dellacorte, Christopher; Machinchick, Michael; Mutharasan, Rajakkannu; Du, Guang-Wu; Ko, Frank; Sirocky, Paul J.; Miller, Jeffrey H.

1992-01-01

Combined cycle ramjet/scramjet engines being designed for advanced hypersonic vehicles, including the National Aerospace Plane (NASP), require innovative high temperature dynamic seals to seal the sliding interfaces of the articulated engine panels. New seals are required that will operate hot (1200 to 2000 F), seal pressures ranging from 0 to 100 psi, remain flexible to accommodate significant sidewall distortions, and resist abrasion over the engine's operational life. This report reviews the recent high temperature durability screening assessments of a new braided rope seal concept, braided of emerging high temperature materials, that shows promise of meeting many of the seal demands of hypersonic engines. The paper presents durability data for: (1) the fundamental seal building blocks, a range of candidate ceramic fiber tows; and for (2) braided rope seal subelements scrubbed under engine simulated sliding, temperature, and preload conditions. Seal material/architecture attributes and limitations are identified through the investigations performed. The paper summarizes the current seal technology development status and presents areas in which future work will be performed.

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.

Aerodynamic Design of a Dual-Flow Mach 7 Hypersonic Inlet System for a Turbine-Based Combined-Cycle Hypersonic Propulsion System

NASA Technical Reports Server (NTRS)

Sanders, Bobby W.; Weir, Lois J.

2008-01-01

A new hypersonic inlet for a turbine-based combined-cycle (TBCC) engine has been designed. This split-flow inlet is designed to provide flow to an over-under propulsion system with turbofan and dual-mode scramjet engines for flight from takeoff to Mach 7. It utilizes a variable-geometry ramp, high-speed cowl lip rotation, and a rotating low-speed cowl that serves as a splitter to divide the flow between the low-speed turbofan and the high-speed scramjet and to isolate the turbofan at high Mach numbers. The low-speed inlet was designed for Mach 4, the maximum mode transition Mach number. Integration of the Mach 4 inlet into the Mach 7 inlet imposed significant constraints on the low-speed inlet design, including a large amount of internal compression. The inlet design was used to develop mechanical designs for two inlet mode transition test models: small-scale (IMX) and large-scale (LIMX) research models. The large-scale model is designed to facilitate multi-phase testing including inlet mode transition and inlet performance assessment, controls development, and integrated systems testing with turbofan and scramjet engines.

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.

A Hot Dynamic Seal Rig for Measuring Hypersonic Engine Seal Durability and Flow Performance

NASA Technical Reports Server (NTRS)

Miller, Jeffrey H.; Steinetz, Bruce M.; Sirocky, Paul J.; Kren, Lawrence A.

1993-01-01

A test fixture for measuring the dynamic performance of candidate high-temperature engine seal concepts was installed at NASA Lewis Research Center. The test fixture was designed to evaluate seal concepts under development for advanced hypersonic engines, such as those being considered for the National Aerospace Plane (NASP). The fixture can measure dynamic seal leakage performance from room temperature up to 840 C (1550 F) and air pressure differentials up to 690 kPa (100 psi). Performance of the seals can be measured while sealing against flat or distorted walls. In the fixture two seals are preloaded against the sides of a 30 cm (1 ft) long saber that slides transverse to the axis of the seals, simulating the scrubbing motion anticipated in these engines. The capabilities of this test fixture along with preliminary data showing the dependence of seal leakage performance on high temperature cycling are addressed.

A hot dynamic seal rig for measuring hypersonic engine seal durability and flow performance

NASA Technical Reports Server (NTRS)

Miller, Jeffrey H.; Steinetz, Bruce M.; Sirocky, Paul J.; Kren, Lawrence A.

1993-01-01

A test fixture for measuring the dynamic performance of candidate high-temperature engine seal concepts has been installed at NASA Lewis Research Center. The test fixture has been designed to evaluate seal concepts under development for advanced hypersonic engines, such as those being considered for the National Aerospace Plane (NASP). The fixture can measure dynamic seal leakage performance from room temperature up to 840 C (1550 F) and air pressure differentials up to 690 kPa (100 psi). Performance of the seals can be measured while sealing against flat or distorted walls. In the fixture two seals are preloaded against the sides of a 30 cm (1 ft) long saber that slides transverse to the axis of the seals, simulating the scrubbing motion anticipated in these engines. This report covers the capabilities of this test fixture along with preliminary data showing the dependence of seal leakage performance on high temperature cycling.

Robust Adaptive Flight Control Design of Air-breathing Hypersonic Vehicles

DTIC Science & Technology

2016-12-07

dynamic inversion controller design for a non -minimum phase hypersonic vehicle is derived by Kuipers et al. [2008]. Moreover, integrated guidance and...stabilization time for inner loop variables is lesser than the intermediate loop variables because of the three-loop-control design methodology . The control...adaptive design . Control Engineering Practice, 2016. Michael A Bolender and David B Doman. A non -linear model for the longitudinal dynamics of a

Inlet and Propulsion Integration of Scram Propelled Vehicles

NASA Technical Reports Server (NTRS)

Povinelli, Louis A.

1996-01-01

The material to be presented in these two lectures begins with cycle considerations of the turbojet engine combined with a ramjet engine to provide thrust over the range of Mach 0 to 5. We will then examine in some detail the aerodynamic behavior that occurs in the inlet operating near the peak speed. Following that, we shall view a numerical simulation through a baseline scramjet engine, starting at the entrance to the inlet, proceeding into the combustor and through the nozzle. In the next segment, we examine a combined rocket and ramjet propulsion system. Analysis and test results will be examined with a view toward evaluation of the concept as a practical device. Two other inlets will then be reviewed: a Mach 12 inlet and a Mach 18 configuration. Finally, we close our lectures with a discussion of the Detonation Wave engine, and inspect the physical and chemical behavior obtained from numerical simulation. A few final remarks will be made regarding the application of CFD for hypersonic propulsion components.

Toward a CFD nose-to-tail capability - Hypersonic unsteady Navier-Stokes code validation

NASA Technical Reports Server (NTRS)

Edwards, Thomas A.; Flores, Jolen

1989-01-01

Computational fluid dynamics (CFD) research for hypersonic flows presents new problems in code validation because of the added complexity of the physical models. This paper surveys code validation procedures applicable to hypersonic flow models that include real gas effects. The current status of hypersonic CFD flow analysis is assessed with the Compressible Navier-Stokes (CNS) code as a case study. The methods of code validation discussed to beyond comparison with experimental data to include comparisons with other codes and formulations, component analyses, and estimation of numerical errors. Current results indicate that predicting hypersonic flows of perfect gases and equilibrium air are well in hand. Pressure, shock location, and integrated quantities are relatively easy to predict accurately, while surface quantities such as heat transfer are more sensitive to the solution procedure. Modeling transition to turbulence needs refinement, though preliminary results are promising.

Flight simulator for hypersonic vehicle and a study of NASP handling qualities

NASA Technical Reports Server (NTRS)

Ntuen, Celestine A.; Park, Eui H.; Deeb, Joseph M.; Kim, Jung H.

1992-01-01

The research goal of the Human-Machine Systems Engineering Group was to study the existing handling quality studies in aircraft with sonic to supersonic speeds and power in order to understand information requirements needed for a hypersonic vehicle flight simulator. This goal falls within the NASA task statements: (1) develop flight simulator for hypersonic vehicle; (2) study NASP handling qualities; and (3) study effects of flexibility on handling qualities and on control system performance. Following the above statement of work, the group has developed three research strategies. These are: (1) to study existing handling quality studies and the associated aircraft and develop flight simulation data characterization; (2) to develop a profile for flight simulation data acquisition based on objective statement no. 1 above; and (3) to develop a simulator and an embedded expert system platform which can be used in handling quality experiments for hypersonic aircraft/flight simulation training.

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.

Uncertainty Quantification of CFD Data Generated for a Model Scramjet Isolator Flowfield

NASA Technical Reports Server (NTRS)

Baurle, R. A.; Axdahl, E. L.

2017-01-01

Computational fluid dynamics is now considered to be an indispensable tool for the design and development of scramjet engine components. Unfortunately, the quantification of uncertainties is rarely addressed with anything other than sensitivity studies, so the degree of confidence associated with the numerical results remains exclusively with the subject matter expert that generated them. This practice must be replaced with a formal uncertainty quantification process for computational fluid dynamics to play an expanded role in the system design, development, and flight certification process. Given the limitations of current hypersonic ground test facilities, this expanded role is believed to be a requirement by some in the hypersonics community if scramjet engines are to be given serious consideration as a viable propulsion system. The present effort describes a simple, relatively low cost, nonintrusive approach to uncertainty quantification that includes the basic ingredients required to handle both aleatoric (random) and epistemic (lack of knowledge) sources of uncertainty. The nonintrusive nature of the approach allows the computational fluid dynamicist to perform the uncertainty quantification with the flow solver treated as a "black box". Moreover, a large fraction of the process can be automated, allowing the uncertainty assessment to be readily adapted into the engineering design and development workflow. In the present work, the approach is applied to a model scramjet isolator problem where the desire is to validate turbulence closure models in the presence of uncertainty. In this context, the relevant uncertainty sources are determined and accounted for to allow the analyst to delineate turbulence model-form errors from other sources of uncertainty associated with the simulation of the facility flow.

Optimal trajectories for an aerospace plane. Part 1: Formulation, results, and analysis

NASA Technical Reports Server (NTRS)

Miele, Angelo; Lee, W. Y.; Wu, G. D.

1990-01-01

The optimization of the trajectories of an aerospace plane is discussed. This is a hypervelocity vehicle capable of achieving orbital speed, while taking off horizontally. The vehicle is propelled by four types of engines: turbojet engines for flight at subsonic speeds/low supersonic speeds; ramjet engines for flight at moderate supersonic speeds/low hypersonic speeds; scramjet engines for flight at hypersonic speeds; and rocket engines for flight at near-orbital speeds. A single-stage-to-orbit (SSTO) configuration is considered, and the transition from low supersonic speeds to orbital speeds is studied under the following assumptions: the turbojet portion of the trajectory has been completed; the aerospace plane is controlled via the angle of attack and the power setting; the aerodynamic model is the generic hypersonic aerodynamics model example (GHAME). Concerning the engine model, three options are considered: (EM1), a ramjet/scramjet combination in which the scramjet specific impulse tends to a nearly-constant value at large Mach numbers; (EM2), a ramjet/scramjet combination in which the scramjet specific impulse decreases monotonically at large Mach numbers; and (EM3), a ramjet/scramjet/rocket combination in which, owing to stagnation temperature limitations, the scramjet operates only at M approx. less than 15; at higher Mach numbers, the scramjet is shut off and the aerospace plane is driven only by the rocket engines. Under the above assumptions, four optimization problems are solved using the sequential gradient-restoration algorithm for optimal control problems: (P1) minimization of the weight of fuel consumed; (P2) minimization of the peak dynamic pressure; (P3) minimization of the peak heating rate; and (P4) minimization of the peak tangential acceleration.

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.

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.

Configuration development study of the X-24C hypersonic research airplane, phase 1

NASA Technical Reports Server (NTRS)

Combs, H. G.

1976-01-01

Four hypersonic research airplane configurations found to be the most cost effective were selected for further refinement. The selection was based on a systematic analysis and evaluation of realistic designs, involving nine different configurations, evolving from three different structural/thermal concepts, coupled with existing rocket and sustainer engines. All configurations were constrained by the mission profiles, research requirements, aerodynamic envelope and maximum launch weight established by NASA.

Configuration development study of the X-24C hypersonic research airplane, phase 2

NASA Technical Reports Server (NTRS)

Combs, H. G.

1977-01-01

The X-24C Hypersonic Research Vehicle, configured with a heat-sink structure, a launch mass limit of 31.75 Mg and powered by an LR-105 Rocket Engine plus 12 LR-101 Sustainer Engines, was found to be the more cost effective of the candidate configurations. In addition, the configuration provides the maximum off design growth potential capability and subsequently, was selected as the candidate configuration to be subjected to the design refinement study in the remaining segment of the study. Selection of this configuration was based on the analytical study conducted on the performance growth capabilities of the candidate configurations selected from the Phase 1 Study.

A hypersonic research vehicle to develop scramjet engines

NASA Technical Reports Server (NTRS)

Gregorek, G. M.; Reuss, R. L.

1990-01-01

Four student design teams produced conceptual designs for a research vehicle to develop the supersonic combustion ramjet (scramjet) engines necessary for efficient hypersonic flight. This research aircraft would provide flight test data for prototype scramjets that is not available in groundbased test facilities. The design specifications call for a research aircraft to be launched from a carrier aircraft at 40,000 feet and a Mach number of 0.8. The aircraft must accelerate to Mach 6 while climbing to a 100,000 foot altitude and then ignite the experimental scramjet engines for acceleration to Mach 10. The research vehicle must then be recovered for another flight. The students responded with four different designs, two piloted waverider configurations, and two unmanned vehicles, one with a blended body-wing configuration, the other with a delta wing shape. All aircraft made use of an engine database provided by the General Electric Aircraft Engine Group; both turbofan ramjet and scramjet engine performance using liquid hydrogen fuel was available. Explained here are the students' conceptual designs and the aerodynamic and propulsion concepts that made their designs feasible.

Development and Evaluation of TiAl Sheet Structures for Hypersonic Applications

NASA Technical Reports Server (NTRS)

Draper, S. L.; Krause, D.; Lerch, B.; Locci, I. E.; Doehnert, B.; Nigam, R.; Das, G.; Sickles, P.; Tabernig, B.; Reger, N.;

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 exhaust flow from the engine by converting electrical current back into flow enthalpy to increase thrust. Though there has been considerable research into the use of MHD generators to produce electricity for industrial power plants, interest in the technology for flight-weight aerospace applications has developed only recently. In this research, electromagnetic fields coupled with weakly ionzed gases to slow hypersonic airflow were investigated within the confines of an MHD energy-bypass system with the goal of showing that it is possible for an air-breathing engine to transition from takeoff to Mach 7 without carrying a rocket propulsion system along with it. The MHD energy-bypass system was modeled for use on a supersonic turbojet engine. The model included all components envisioned for an MHD energy-bypass system; two preionizers, an MHD generator, and an MHD accelerator. A thermodynamic cycle analysis of the hypothesized MHD energy-bypass system on an existing supersonic turbojet engine was completed. In addition, a detailed thermodynamic, plasmadynamic, and electromagnetic analysis was combined to offer a single, comprehensive model to describe more fully the proper plasma flows and magnetic fields required for successful operation of the MHD energy bypass system. The unique contribution of this research involved modeling the current density, temperature, velocity, pressure, electric field, Hall parameter, and electrical power throughout an annular MHD generator and an annular MHD accelerator taking into account an external magnetic field within a moving flow field, collisions of electrons with neutral particles in an ionized flow field, and collisions of ions with neutral particles in an ionized flow field (ion slip). In previous research, the ion slip term has not been considered. The MHD energy-bypass system model showed that it is possible to expand the operating range of a supersonic jet engine from a maximum of Mach 3.5 to a maximum of Mach 7. The inclusion of ion slip within the analysis further showed that it is possible to 'drive' this system with maximum magnetic fields of 3 T and with maximum conductivity levels of 11 mhos/m. These operating parameters better the previous findings of 5 T and 10 mhos/m, and reveal that taking into account collisions between ions and neutral particles within a weakly ionized flow provides a more realistic model with added benefits of lower magnetic fields and conductivity levels especially at the higher Mach numbers. (Abstract shortened by UMI.).

Design, calibration and testing of a force balance for a hypersonic shock tunnel

NASA Astrophysics Data System (ADS)

Vadassery, Pravin

The forces acting on a flight vehicle are critical for determining its performance. Of particular interest is the hypersonic regime. Force measurements are much more complex in hypersonic flows, where those speeds are simulated in shock tunnels. A force balance for such facilities contains sensitive gages that measure stress waves and ultimately determine the different components of force acting on the model. An external force balance was designed and fabricated for the UTA Hypersonic shock tunnel to measure drag at Mach 10. Static and dynamic calibrations were performed to find the transfer function of the system. Forces were recovered using a deconvolution procedure. To validate the force balance, experiments were conducted on a blunt cone. The measured forces were compared to Newtonian theory.

Applicability of the Hypersonic Similarity Rule to Pressure Distributions which Include the Effects of Rotation for Bodies of Revolution at Zero Angle of Attack

NASA Technical Reports Server (NTRS)

Rossow, Vernon J

1951-01-01

The analysis of Technical Note 2250, 1950, is extended to include the effects of flow rotation. It is found that the theoretical pressure distributions over drive cylinders can be related by the hypersonic similarity rule with sufficient accuracy for most engineering purposes. The error introduced into pressure distributions and drag effective cylinders by ignoring the rotation term in the characteristic equations is investigated.

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 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 the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

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.

Emission calculations for a scramjet powered hypersonic transport

NASA Technical Reports Server (NTRS)

Lezberg, E. A.

1973-01-01

Calculations of exhaust emissions from a scramjet powered hypersonic transport burning hydrogen fuel were performed over a range of Mach numbers of 5 to 12 to provide input data for wake mixing calculations and forecasts of future levels of pollutants in the stratosphere. The calculations were performed utilizing a one-dimensional chemical kinetics computer program for the combustor and exhaust nozzle of a fixed geometry dual-mode scramjet engine. Inlet conditions to the combustor and engine size was based on a vehicle of 227,000 kg (500,000 lb) gross take of weight with engines sized for Mach 8 cruise. Nitric oxide emissions were very high for stoichiometric engine operation but for Mach 6 cruise at reduced equivalence ratio are in the range predicted for an advanced supersonic transport. Combustor designs which utilize fuel staging and rapid expansion to minimize residence time at high combustion temperatures were found to be effective in preventing nitric oxide formation from reaching equilibrium concentrations.

Further Investigations of Hypersonic Engine Seals

NASA Technical Reports Server (NTRS)

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

2004-01-01

Durable, flexible sliding seals are required in advanced hypersonic engines to seal the perimeters of movable engine ramps for efficient, safe operation in high heat flux environments at temperatures of 2000 to 2500 F. Current seal designs do not meet the demanding requirements for future engines, so NASA's Glenn Research Center is developing advanced seals and preloading devices to overcome these shortfalls. An advanced ceramic wafer seal design and two silicon nitride compression spring designs were evaluated in a series of compression, scrub, and flow tests. Silicon nitride wafer seals survived 2000 in. (50.8 m) of scrubbing at 2000 F against a silicon carbide rub surface with no chips or signs of damage. Flow rates measured for the wafers before and after scrubbing were almost identical and were up to 32 times lower than those recorded for the best braided rope seal flow blockers. Silicon nitride compression springs showed promise conceptually as potential seal preload devices to help maintain seal resiliency.

Hyper-X Vehicle Model - Side View

NASA Technical Reports Server (NTRS)

1996-01-01

A side-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 vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

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 vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

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 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 the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

Hyper-X Vehicle Model - Side View

NASA Technical Reports Server (NTRS)

1996-01-01

Sleek lines are apparent in this side-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 vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

Hyper-X Research Vehicle - Artist Concept in Flight

NASA Technical Reports Server (NTRS)

1997-01-01

An artist's conception of the X-43A Hypersonic Experimental Vehicle, or 'Hyper-X' in flight. 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, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

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, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

Hyper-X Vehicle Model - Top Front View

NASA Technical Reports Server (NTRS)

1996-01-01

A top front view of an early desk-top model of NASA's X-43A 'Hyper-X,' or Hypersonic Experimental Vehicle, 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 vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

X-43A Undergoing Controlled Radio Frequency Testing in the Benefield Anechoic Facility at Edwards Ai

NASA Technical Reports Server (NTRS)

2000-01-01

The X-43A Hypersonic Experimental (Hyper-X) Vehicle hangs suspended in the cavernous Benefield Aenechoic Facility at Edwards Air Force Base during radio frequency tests in January 2000. 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 the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

On the dependence of the domain of values of functionals of hypersonic aerodynamics on controls

NASA Astrophysics Data System (ADS)

Bilchenko, Grigory; Bilchenko, Nataly

2018-05-01

The properties of mathematical model of control of heat and mass transfer in laminar boundary layer on permeable cylindrical and spherical surfaces of the hypersonic aircraft are considered. Dependences of hypersonic aerodynamics functionals (the total heat flow and the total Newton friction force) on controls (the blowing into boundary layer, the temperature factor, the magnetic field) are investigated. The domains of allowed values of functionals of hypersonic aerodynamics are obtained. The results of the computational experiments are presented: the dependences of total heat flow on controls; the dependences of total Newton friction force on controls; the mutual dependences of functionals (as the domains of allowed values "Heat and Friction"); the dependences of blowing system power on controls. The influences of magnetic field and dissociation on the domain of "Heat and Friction" allowed values are studied. It is proved that for any fixed constant value of magnetic field the blowing system power is a symmetric function of constant dimensionless controls (the blowing into boundary layer and the temperature factor). It is shown that the obtained domain of allowed values of functionals of hypersonic aerodynamics depending on permissible range of controls may be used in engineering.

Advanced Technology Inlet Design, NRA 8-21 Cycle II: DRACO Flowpath Hypersonic Inlet Design

NASA Technical Reports Server (NTRS)

Sanders, Bobby W.; Weir, Lois J.

1999-01-01

The report outlines work performed in support of the flowpath development for the DRACO engine program. The design process initiated to develop a hypersonic axisymmetric inlet for a Mach 6 rocket-based combined cycle (RBCC) engine is discussed. Various design parametrics were investigated, including design shock-on-lip Mach number, cone angle, throat Mach number, throat angle. length of distributed compression, and subsonic diffuser contours. Conceptual mechanical designs consistent with installation into the D-21 vehicle were developed. Additionally, program planning for an intensive inlet development program to support a Critical Design Review in three years was performed. This development program included both analytical and experimental elements and support for a flight-capable inlet mechanical design.

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.

Fitting aerodynamics and propulsion into the puzzle

NASA Technical Reports Server (NTRS)

Johnston, Patrick J.; Whitehead, Allen H., Jr.; Chapman, Gary T.

1987-01-01

The development of an airbreathing single-stage-to-orbit vehicle, in particular the problems of aerodynamics and propulsion integration, is examined. The boundary layer transition on constant pressure surfaces at hypersonic velocities, and the effects of noise on the transition are investigated. The importance of viscosity, real-gas effects, and drag at hypersonic speeds is discussed. A propulsion system with sufficient propulsive lift to enhance the performance of the vehicle is being developed. The difficulties of engine-airframe integration are analyzed.

Development of a Test to Evaluate Aerothermal Response of Materials to Hypersonic Flow Using a Scramjet Wind Tunnel (Postprint)

DTIC Science & Technology

2010-05-01

SCRAMJET WIND TUNNEL (POSTPRINT) 5a. CONTRACT NUMBER FA8650-10-D-5226-0002 5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 62102F 6. AUTHOR(S...prototype scramjet engine as a wind tunnel . A sample holder was designed using combustion fluid dynamics results as inputs into structural models. The...Z39-18 Development of a Test to Evaluate Aerothermal Response of Materials to Hypersonic Flow Using a Scramjet Wind Tunnel Triplicane A

Approximate convective heating equations for hypersonic flows

NASA Technical Reports Server (NTRS)

Zoby, E. V.; Moss, J. N.; Sutton, K.

1979-01-01

Laminar and turbulent heating-rate equations appropriate for engineering predictions of the convective heating rates about blunt reentry spacecraft at hypersonic conditions are developed. The approximate methods are applicable to both nonreacting and reacting gas mixtures for either constant or variable-entropy edge conditions. A procedure which accounts for variable-entropy effects and is not based on mass balancing is presented. Results of the approximate heating methods are in good agreement with existing experimental results as well as boundary-layer and viscous-shock-layer solutions.

Engine sizing and integration requirements for hypersonic airbreathing missile applications

NASA Astrophysics Data System (ADS)

Waltrup, P. J.; Billig, F. S.; Stockbridge, R. D.

1982-03-01

A procedure that provides a rational means for selecting an inlet/combustor configuration for a hypersonic airbreathing missile is presented. The particular problem that is addressed is the design of the sustained engine of a two stage missile that is constrained to be launched from a stowage volume that is either square or circular in cross section. The sustainer engine accelerates from a low altitude separation at Mach 4 and climbs to high altitude for cruise at Mach 8. The results show that a missile with an axisymmetric nose inlet provides a somewhat higher thrust capability and slightly better fuel efficiency than a chin type inlet. Aft entry inlets are shown to have a substantially lower thrust potential and lower engine efficiency. A criterion for determining the maximum contraction ratio of a fixed geometry inlet is established and applied to the exemplary missile designs. Combustor area ratio is examined and found to have a relatively small effect on engine performance for area ratios equal to or larger than that required to obtain maximum thrust at the take-over Mach number.

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.

Flow-Tagging Velocimetry for Hypersonic Flows Using Fluorescence of Nitric Oxide

NASA Technical Reports Server (NTRS)

Danehy, P. M.; OByrne, S.; Houwing, A. F. P.

2001-01-01

We investigate a new type of flow-tagging velocimetry technique for hypersonic flows. The technique involves exciting a thin line of nitric oxide molecules with a laser beam and then, after some delay, acquiring an image of the displaced line. One component of velocity is determined from the time of flight. This method is applied to measure the velocity profile in a Mach 8.5 laminar, hypersonic boundary layer in the Australian National Universities T2 free-piston shock tunnel. The velocity is measured with an uncertainty of approximately 2%. Comparison with a CFD simulation of the flow shows reasonable agreement.

Trajectory optimization and guidance for an aerospace plane

NASA Technical Reports Server (NTRS)

Mease, Kenneth D.; Vanburen, Mark A.

1989-01-01

The first step in the approach to developing guidance laws for a horizontal take-off, air breathing single-stage-to-orbit vehicle is to characterize the minimum-fuel ascent trajectories. The capability to generate constrained, minimum fuel ascent trajectories for a single-stage-to-orbit vehicle was developed. A key component of this capability is the general purpose trajectory optimization program OTIS. The pre-production version, OTIS 0.96 was installed and run on a Convex C-1. A propulsion model was developed covering the entire flight envelope of a single-stage-to-orbit vehicle. Three separate propulsion modes, corresponding to an after burning turbojet, a ramjet and a scramjet, are used in the air breathing propulsion phase. The Generic Hypersonic Aerodynamic Model Example aerodynamic model of a hypersonic air breathing single-stage-to-orbit vehicle was obtained and implemented. Preliminary results pertaining to the effects of variations in acceleration constraints, available thrust level and fuel specific impulse on the shape of the minimum-fuel ascent trajectories were obtained. The results show that, if the air breathing engines are sized for acceleration to orbital velocity, it is the acceleration constraint rather than the dynamic pressure constraint that is active during ascent.

Hypersonic propulsion: Status and challenge

NASA Technical Reports Server (NTRS)

Guy, R. Wayne

1990-01-01

Scientists in the U.S. are again focusing on the challenge of hypersonic flight with the proposed National Aerospace Plane (NASP). This renewed interest has led to an expansion of research related to high speed airbreathing propulsion, in particular, the supersonic combustion ramjet, or scramjet. The history is briefly traced of scramjet research in the U.S., with emphasis on NASA sponsored efforts, from the Hypersonic Research Engine (HRE) to the current status of today's airframe integrated scramjets. The challenges of scramjet technology development from takeover to orbital speeds are outlined. Existing scramjet test facilities such as NASA Langley's Scramjet Test Complex as well as new high Mach number pulse facilities are discussed. The important partnership role of experimental methods and computational fluid dynamics is emphasized for the successful design of single stage to orbit vehicles.

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.

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.

Diagnostics for Hypersonic Engine Control

DTIC Science & Technology

2015-02-01

modeling efforts and/or lead to the development of sensors that can be used as part of scramjet engine control strategies. Activities included work on...of a model scramjet engine cannot rely on the presence of water. Instead, light sources operating at wavelengths resonant with molecular oxygen are...transmitted beam amplitude fluctuations (scintillation). Frequency axis is normalized. Figure 3. Oxygen absorption feature recorded using direct

X-43C Flight Demonstrator Project Overview

NASA Technical Reports Server (NTRS)

Moses, Paul L.

2003-01-01

The X-43C Flight Demonstrator Project is a joint NASA-USAF hypersonic propulsion technology flight demonstration project that will expand the hypersonic flight envelope for air-breathing engines. The Project will demonstrate sustained accelerating flight through three flights of expendable X-43C Demonstrator Vehicles (DVs). The approximately 16-foot long X-43C DV will be boosted to the starting test conditions, separate from the booster, and accelerate from Mach 5 to Mach 7 under its own power and autonomous control. The DVs will be powered by a liquid hydrocarbon-fueled, fuel-cooled, dual-mode, airframe integrated scramjet engine system developed under the USAF HyTech Program. The Project is managed by NASA Langley Research Center as part of NASA's Next Generation Launch Technology Program. Flight tests will be conducted by NASA Dryden Flight Research Center off the coast of California over water in the Pacific Test Range. The NASA/USAF/industry project is a natural extension of the Hyper-X Program (X-43A), which will demonstrate short duration (approximately 10 seconds) gaseous hydrogen-fueled scramjet powered flight at Mach 7 and Mach 10 using a heavy-weight, largely heat sink construction, experimental engine. The X-43C Project will demonstrate sustained accelerating flight from Mach 5 to Mach 7 (approximately 4 minutes) using a flight-weight, fuel-cooled, scramjet engine powered by much denser liquid hydrocarbon fuel. The X-43C DV design flows from integrating USAF HyTech developed engine technologies with a NASA Air-Breathing Launch Vehicle accelerator-class configuration and Hyper-X heritage vehicle systems designs. This paper describes the X-43C Project and provides the background for NASA's current hypersonic flight demonstration efforts.

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.

Hypersonic force measurements using internal balance based on optical micromachined Fabry-Perot interferometry.

PubMed

Qiu, Huacheng; Min, Fu; Zhong, Shaolong; Song, Xin; Yang, Yanguang

2018-03-01

Force measurements using wind tunnel balance are necessary for determining a variety of aerodynamic performance parameters, while the harsh environment in hypersonic flows requires that the measurement instrument should be reliable and robust, in against strong electromagnetic interference, high vacuum, or metal (oxide) dusts. In this paper, we demonstrated a three-component internal balance for hypersonic aerodynamic force measurements, using novel optical micromachined Fabry-Perot interferometric (FPI) strain gauges as sensing elements. The FPI gauges were fabricated using Micro-Opto-Electro-Mechanical Systems (MOEMS) surface and bulk fabrication techniques. High-reflectivity coatings are used to form a high-finesse Fabry-Perot cavity, which benefits a high resolution. Antireflective and passivation coatings are used to reduce unwanted interferences. The FPI strain gauge based balance has been calibrated and evaluated in a Mach 5 hypersonic flow. The results are compared with the traditional technique using the foil resistive strain gauge balance, indicating that the proposed balance based on the MOEMS FPI strain gauge is reliable and robust and is potentially suitable for the hypersonic wind tunnel harsh environment.

Hypersonic force measurements using internal balance based on optical micromachined Fabry-Perot interferometry

NASA Astrophysics Data System (ADS)

Qiu, Huacheng; Min, Fu; Zhong, Shaolong; Song, Xin; Yang, Yanguang

2018-03-01

Force measurements using wind tunnel balance are necessary for determining a variety of aerodynamic performance parameters, while the harsh environment in hypersonic flows requires that the measurement instrument should be reliable and robust, in against strong electromagnetic interference, high vacuum, or metal (oxide) dusts. In this paper, we demonstrated a three-component internal balance for hypersonic aerodynamic force measurements, using novel optical micromachined Fabry-Perot interferometric (FPI) strain gauges as sensing elements. The FPI gauges were fabricated using Micro-Opto-Electro-Mechanical Systems (MOEMS) surface and bulk fabrication techniques. High-reflectivity coatings are used to form a high-finesse Fabry-Perot cavity, which benefits a high resolution. Antireflective and passivation coatings are used to reduce unwanted interferences. The FPI strain gauge based balance has been calibrated and evaluated in a Mach 5 hypersonic flow. The results are compared with the traditional technique using the foil resistive strain gauge balance, indicating that the proposed balance based on the MOEMS FPI strain gauge is reliable and robust and is potentially suitable for the hypersonic wind tunnel harsh environment.

Laser-energized MHD generator for hypersonic electric air-turborockets

NASA Technical Reports Server (NTRS)

Myrabo, L. N.; Rosa, R. J.; Moder, J. P.; Blandino, J. S.; Frazier, S. R.

1987-01-01

The analysis and design of an open cycle Faraday MHD generator suitable for use in an electric air-turborocket cycle, the MHD-fanjet, is presented. The working fluid for the generators is unseeded high temperature hydrogen, generated by a standing, laser-supported combustion wave. This study also examines the performance of an advanced combined-cycle engine, powered by beamed energy, proposed for use in future SSTO aerospacecraft. This innovative engine incorporates the MHD-fanjet for the acceleration role within the hypersonic flight regime, from about Mach 11 to above Mach 25. Performance results indicate that specific impulses could fall in the range of 10,000 to 16,000 seconds. This would enable propellant mass fractions as low as 6 percent to 9 percent for such advanced shuttlecraft flying SSTO missions to low earth orbit.

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

NASA Astrophysics Data System (ADS)

Dalle, Derek J.

The development and application of a first-principles-derived reduced-order model called MASIV (Michigan/AFRL Scramjet In Vehicle) for an air-breathing hypersonic vehicle is discussed. Several significant and previously unreported aspects of hypersonic flight are investigated. A fortunate coupling between increasing Mach number and decreasing angle of attack is shown to extend the range of operating conditions for a class of supersonic inlets. Detailed maps of isolator unstart and ram-to-scram transition are shown on the flight corridor map for the first time. In scram mode the airflow remains supersonic throughout the engine, while in ram mode there is a region of subsonic flow. Accurately predicting the transition between these two modes requires models for complex shock interactions, finite-rate chemistry, fuel-air mixing, pre-combustion shock trains, and thermal choking, which are incorporated into a unified framework here. Isolator unstart occurs when the pre-combustion shock train is longer than the isolator, which blocks airflow from entering the engine. Finally, cooptimization of the vehicle design and trajectory is discussed. An optimal control technique is introduced that greatly reduces the number of computations required to optimize the simulated trajectory.

Mach 4 free-jet tunnel starting experiments for a hypersonic research engine model causing high blockage

NASA Technical Reports Server (NTRS)

Carson, G. T., Jr.; Midden, R. E.

1976-01-01

Tests of a full scale hypersonic research engine (HRE) were conducted in the hypersonic tunnel facility at Mach numbers of 5, 6, and 7. Since the HRE would cause a rather high blockage (48.83 percent of the nozzle area), subscale tests were conducted in various available small wind tunnels prior to the full scale tests to study the effects of model blockage on tunnel starting. The results of the Mach 4 subscale tests which utilized a model system at 0.0952 scale which simulated the HRE in the test section of the tunnel are presented. A satisfactory tunnel starting could not be achieved by varying the free jet length or diffuser size nor by inserting the model into the test stream after tunnel starting. However, the installation of a shroud around the HRE model allowed the tunnel to start with the model preset in the tunnel at a tunnel stagnation pressure to atmospheric exit pressure ratio of 13.4. The simulation of the discharge of instrumentation cooling water and the addition of test hardware at the aft end of the HRE model did not have a significant effect on the tunnel starting.

Studies of scramjet/airframe integration techniques for hypersonic aircraft

NASA Technical Reports Server (NTRS)

Edwards, C. L. W.; Small, W. J.; Weidner, J. P.; Johnston, P. J.

1975-01-01

New design and analysis techniques for engine-airframe integration were applied in a recent hypersonic vehicle design study. A new technique was developed to design the vehicle's forebody so that uniform precompressed flow was produced at the inlet entrance. Results are verified with three-dimensional characteristic calculations. Results from a new three-dimensional method for calculating nozzle flows show that the entire lower afterbody of the vehicle can be used as a scramjet exhaust nozzle to achieve efficient, controlled, and stable flight over a wide range of flight conditions.

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 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 the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

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

NASA Technical Reports Server (NTRS)

1999-01-01

The X-43A Hypersonic Experimental Vehicle, or 'Hyper-X,' carefully packed in a protective shipping framework, is unloaded from a container after its arrival at NASA's 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 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 the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

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-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 the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

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 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 the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

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

NASA Technical Reports Server (NTRS)

1999-01-01

A head-on view 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 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 the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

Hyper-X Engine Design and Ground Test Program

NASA Technical Reports Server (NTRS)

Voland, R. T.; Rock, K. E.; Huebner, L. D.; Witte, D. W.; Fischer, K. E.; McClinton, C. R.

1998-01-01

The Hyper-X Program, NASA's focused hypersonic technology program jointly run by NASA Langley and Dryden, is designed to move hypersonic, air-breathing vehicle technology from the laboratory environment to the flight environment, the last stage preceding prototype development. The Hyper-X research vehicle will provide the first ever opportunity to obtain data on an airframe integrated supersonic combustion ramjet propulsion system in flight, providing the first flight validation of wind tunnel, numerical and analytical methods used for design of these vehicles. A substantial portion of the integrated vehicle/engine flowpath development, engine systems verification and validation and flight test risk reduction efforts are experimentally based, including vehicle aeropropulsive force and moment database generation for flight control law development, and integrated vehicle/engine performance validation. The Mach 7 engine flowpath development tests have been completed, and effort is now shifting to engine controls, systems and performance verification and validation tests, as well as, additional flight test risk reduction tests. The engine wind tunnel tests required for these efforts range from tests of partial width engines in both small and large scramjet test facilities, to tests of the full flight engine on a vehicle simulator and tests of a complete flight vehicle in the Langley 8-Ft. High Temperature Tunnel. These tests will begin in the summer of 1998 and continue through 1999. The first flight test is planned for early 2000.

Combustor kinetic energy efficiency analysis of the hypersonic research engine data

NASA Astrophysics Data System (ADS)

Hoose, K. V.

1993-11-01

A one-dimensional method for measuring combustor performance is needed to facilitate design and development scramjet engines. A one-dimensional kinetic energy efficiency method is used for measuring inlet and nozzle performance. The objective of this investigation was to assess the use of kinetic energy efficiency as an indicator for scramjet combustor performance. A combustor kinetic energy efficiency analysis was performed on the Hypersonic Research Engine (HRE) data. The HRE data was chosen for this analysis due to its thorough documentation and availability. The combustor, inlet, and nozzle kinetic energy efficiency values were utilized to determine an overall engine kinetic energy efficiency. Finally, a kinetic energy effectiveness method was developed to eliminate thermochemical losses from the combustion of fuel and air. All calculated values exhibit consistency over the flight speed range. Effects from fuel injection, altitude, angle of attack, subsonic-supersonic combustion transition, and inlet spike position are shown and discussed. The results of analyzing the HRE data indicate that the kinetic energy efficiency method is effective as a measure of scramjet combustor performance.

In-stream measurements of combustion during Mach 5 to 7 tests of the Hypersonic Research Engine (HRE)

NASA Technical Reports Server (NTRS)

Lezberg, Erwin A.; Metzler, Allen J.; Pack, William D.

1993-01-01

Results of in-stream combustion measurements taken during Mach 5 to 7 true simulation testing of the Hypersonic Research Engine/Aerothermodynamic Integration Model (HRE/AIM) are presented. These results, the instrumentation techniques, and configuration changes to the engine installation that were required to test this model are described. In test runs at facility Mach numbers of 5 to 7, an exhaust instrumentation ring which formed an extension of the engine exhaust nozzle shroud provided diagnostic measurements at 10 circumferential locations in the HRE combustor exit plane. The measurements included static and pitot pressures using conventional conical probes, combustion gas temperatures from cooled-gas pyrometer probes, and species concentration from analysis of combustion gas samples. Results showed considerable circumferential variation, indicating that efficiency losses were due to nonuniform fuel distribution or incomplete mixing. Results using the Mach 7 facility nozzle but with Mach 6 temperature simulation, 1590 to 1670 K, showed indications of incomplete combustion. Nitric oxide measurements at the combustor exit peaked at 2000 ppmv for stoichiometric combustion at Mach 6.

Analytical and experimental investigations of the oblique detonation wave engine concept

NASA Technical Reports Server (NTRS)

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

1990-01-01

Wave combustors, which include the oblique detonation wave engine (ODWE), are attractive propulsion concepts for hypersonic flight. These engines utilize oblique shock or detonation waves to rapidly mix, ignite, and combust the air-fuel mixture in thin zones in the combustion chamber. Benefits of these combustion systems include shorter and lighter engines which require less cooling and can provide thrust at higher Mach numbers than conventional scramjets. The wave combustor's ability to operate at lower combustor inlet pressures may allow the vehicle to operate at lower dynamic pressures which could lessen the heating loads on the airframe. The research program at NASA-Ames includes analytical studies of the ODWE combustor using Computational Fluid Dynamics (CFD) codes which fully couple finite rate chemistry with fluid dynamics. In addition, experimental proof-of-concept studies are being performed in an arc heated hypersonic wind tunnel. Several fuel injection design were studied analytically and experimentally. In-stream strut fuel injectors were chosen to provide good mixing with minimal stagnation pressure losses. Measurements of flow field properties behind the oblique wave are compared to analytical predictions.

Analytical and experimental investigations of the oblique detonation wave engine concept

NASA Technical Reports Server (NTRS)

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

1991-01-01

Wave combustors, which include the Oblique Detonation Wave Engine (ODWE), are attractive propulsion concepts for hypersonic flight. These engines utilize oblique shock or detonation waves to rapidly mix, ignite, and combust the air-fuel mixture in thin zones in the combustion chamber. Benefits of these combustion systems include shorter and lighter engines which will require less cooling and can provide thrust at higher Mach numbers than conventional scramjets. The wave combustor's ability to operate at lower combustor inlet pressures may allow the vehicle to operate at lower dynamic pressures which could lessen the heating loads on the airframe. The research program at NASA-Ames includes analytical studies of the ODWE combustor using CFD codes which fully couple finite rate chemistry with fluid dynamics. In addition, experimental proof-of-concept studies are being carried out in an arc heated hypersonic wind tunnel. Several fuel injection designs were studied analytically and experimentally. In-stream strut fuel injectors were chosen to provide good mixing with minimal stagnation pressure losses. Measurements of flow field properties behind the oblique wave are compared to analytical predictions.

Conjugate Heat Transfer and Thermo-Structural Analysis of the Actively Cooled Multi-Stage Conical Nozzle and Hypersonic Low-Reynolds Diffuser of the New Arc-Heated Wind Tunnel (AWHT-II) of the University of Texas at Arlington

NASA Astrophysics Data System (ADS)

Campbell, David R.

Arc-heated wind tunnels are the primary test facility for screening and qualification of candidate materials for hypersonic thermal protection systems (TPS). Via an electric arc that largely augments the enthalpy (by tens of MJ/kg) of the working fluid (Air, Nitrogen, CO2 in case of Mars-entry studies) passed through a converging-diverging nozzle at specific stagnation conditions, different regimes encountered in entry and re-entry hypersonic aerothermodynamics can be simulated. Because of the high-enthalpies (and associated temperatures that generally exceed the limits required by the thermo-structural integrity of the facility) the active cooling of the arc-heated wind tunnel's parts exposed to the working gas is critical. This criticality is particularly severe in these facilities due to the time scales associated with their continuous operation capabilities (order of minutes). This research focuses on the design and the conjugate heat transfer and resultant thermo-structural analysis of a multi-segment nozzle and low-Reynolds, hypersonic diffuser for the new arc-heated wind tunnel (AHWT-II) of the University of Texas at Arlington. Nozzles and hypersonic diffusers are critical components that experience highly complex flows (non-equilibrium aerothermochemistry) and high (local and distributed) heat-flux loads which significantly augment the complexity of the problems associated with their thermal management. The proper design and thermo-mechanical analysis of these components are crucial elements for the operability of the new facility. This work is centered on the design considerations, methodologies and the detailed analysis of the aforementioned components which resulted in the definition of final parts and assemblies that are under manufacturing at this writing. The project is jointly sponsored by the Office of Naval Research (ONR) and the Defense Advanced Research Project Agency (DARPA).

Response Surface Modeling of Combined-Cycle Propulsion Components using Computational Fluid Dynamics

NASA Technical Reports Server (NTRS)

Steffen, C. J., Jr.

2002-01-01

Three examples of response surface modeling with CFD are presented for combined cycle propulsion components. The examples include a mixed-compression-inlet during hypersonic flight, a hydrogen-fueled scramjet combustor during hypersonic flight, and a ducted-rocket nozzle during all-rocket flight. Three different experimental strategies were examined, including full factorial, fractionated central-composite, and D-optimal with embedded Plackett-Burman designs. The response variables have been confined to integral data extracted from multidimensional CFD results. Careful attention to uncertainty assessment and modeling bias has been addressed. The importance of automating experimental setup and effectively communicating statistical results are emphasized.

Finite element elastic-plastic-creep and cyclic life analysis of a cowl lip

NASA Technical Reports Server (NTRS)

Arya, Vinod K.; Melis, Matthew E.; Halford, Gary R.

1990-01-01

Results are presented of elastic, elastic-plastic, and elastic-plastic-creep analyses of a test-rig component of an actively cooled cowl lip. A cowl lip is part of the leading edge of an engine inlet of proposed hypersonic aircraft and is subject to severe thermal loadings and gradients during flight. Values of stresses calculated by elastic analysis are well above the yield strength of the cowl lip material. Such values are highly unrealistic, and thus elastic stress analyses are inappropriate. The inelastic (elastic-plastic and elastic-plastic-creep) analyses produce more reasonable and acceptable stress and strain distributions in the component. Finally, using the results from these analyses, predictions are made for the cyclic crack initiation life of a cowl lip. A comparison of predicted cyclic lives shows the cyclic life prediction from the elastic-plastic-creep analysis to be the lowest and, hence, most realistic.

Transatmospheric vehicle research

NASA Technical Reports Server (NTRS)

Adelman, Henry G.; Cambier, Jean-Luc

1990-01-01

Research was conducted into the alternatives to the supersonic combustion ramjet (scramjet) engine for hypersonic flight. A new engine concept, the Oblique Detonation Wave Engine (ODWE) was proposed and explored analytically and experimentally. Codes were developed which can couple the fluid dynamics of supersonic flow with strong shock waves, with the finite rate chemistry necessary to model the detonation process. An additional study was conducted which compared the performance of a hypersonic vehicle powered by a scramjet or an ODWE. Engineering models of the overall performances of the two engines are included. This information was fed into a trajectory program which optimized the flight path to orbit. A third code calculated the vehicle size, weight, and aerodynamic characteristics. The experimental work was carried out in the Ames 20MW arc-jet wind tunnel, focusing on mixing and combustion of fuel injected into a supersonic airstream. Several injector designs were evaluated by sampling the stream behind the injectors and analyzing the mixture with an on-line mass spectrometer. In addition, an attempt was made to create a standing oblique detonation wave in the wind tunnel using hydrogen fuel. It appeared that the conditions in the test chamber were marginal for the generation of oblique detonation waves.

Overview of Transonic to Hypersonic Stage Separation Tool Development for Multi-Stage-to-Orbit Concepts

NASA Technical Reports Server (NTRS)

Murphy, Kelly J.; Bunning, Pieter G.; Pamadi, Bandu N.; Scallion, William I.; Jones, Kenneth M.

2004-01-01

An overview of research efforts at NASA in support of the stage separation and ascent aerothermodynamics research program is presented. The objective of this work is to develop a synergistic suite of experimental, computational, and engineering tools and methods to apply to vehicle separation across the transonic to hypersonic speed regimes. Proximity testing of a generic bimese wing-body configuration is on-going in the transonic (Mach numbers 0.6, 1.05, and 1.1), supersonic (Mach numbers 2.3, 3.0, and 4.5) and hypersonic (Mach numbers 6 and 10) speed regimes in four wind tunnel facilities at the NASA Langley Research Center. An overset grid, Navier-Stokes flow solver has been enhanced and demonstrated on a matrix of proximity cases and on a dynamic separation simulation of the bimese configuration. Steady-state predictions with this solver were in excellent agreement with wind tunnel data at Mach 3 as were predictions via a Cartesian-grid Euler solver. Experimental and computational data have been used to evaluate multi-body enhancements to the widely-used Aerodynamic Preliminary Analysis System, an engineering methodology, and to develop a new software package, SepSim, for the simulation and visualization of vehicle motions in a stage separation scenario. Web-based software will be used for archiving information generated from this research program into a database accessible to the user community. Thus, a framework has been established to study stage separation problems using coordinated experimental, computational, and engineering tools.

Overview of the Turbine Based Combined Cycle Discipline

NASA Technical Reports Server (NTRS)

Thomas, Scott R.; Walker, James F.; Pittman, James L.

2009-01-01

The NASA Fundamental Aeronautics Hypersonics project is focused on technologies for combined cycle, airbreathing propulsions systems to enable reusable launch systems for access to space. Turbine Based Combined Cycle (TBCC) propulsion systems offer specific impulse (Isp) improvements over rocket-based propulsion systems in the subsonic takeoff and return mission segments and offer improved safety. The potential to realize more aircraft-like operations with expanded launch site capability and reduced system maintenance are additional benefits. The most critical TBCC enabling technologies as identified in the National Aeronautics Institute (NAI) study were: 1) mode transition from the low speed propulsion system to the high speed propulsion system, 2) high Mach turbine engine development, 3) transonic aero-propulsion performance, 4) low-Mach-number dual-mode scramjet operation, 5) innovative 3-D flowpath concepts and 6) innovative turbine based combined cycle integration. To address several of these key TBCC challenges, NASA s Hypersonics project (TBCC Discipline) initiated an experimental mode transition task that includes an analytic research endeavor to assess the state-of-the-art of propulsion system performance and design codes. This initiative includes inlet fluid and turbine performance codes and engineering-level algorithms. This effort has been focused on the Combined Cycle Engine Large-Scale Inlet Mode Transition Experiment (CCE LIMX) which is a fully integrated TBCC propulsion system with flow path sizing consistent with previous NASA and DoD proposed Hypersonic experimental flight test plans. This experiment is being tested in the NASA-GRC 10 x 10 Supersonic Wind Tunnel (SWT) Facility. The goal of this activity is to address key hypersonic combined-cycle-engine issues: (1) dual integrated inlet operability and performance issues unstart constraints, distortion constraints, bleed requirements, controls, and operability margins, (2) mode-transition constraints imposed by the turbine and the ramjet/scramjet flow paths (imposed variable geometry requirements), (3) turbine engine transients (and associated time scales) during transition, (4) high-altitude turbine engine re-light, and (5) the operating constraints of a Mach 3-7 combustor (specific to the TBCC). The model will be tested in several test phases to develop a unique TBCC database to assess and validate design and analysis tools and address operability, integration, and interaction issues for this class of advanced propulsion systems. The test article and all support equipment is complete and available at the facility. The test article installation and facility build-up in preparation for the inlet performance and operability characterization is near completion and testing is planned to commence in FY11.

Recent Advances in Structures for Hypersonic Flight, part 2

NASA Technical Reports Server (NTRS)

1978-01-01

The papers at this symposium were presented by 24 speakers representing airframe, missile, and engine manufacturers, the U.S. Air Force, and two NASA Research Centers. The papers cover a variety of topics including engine structures, cooled airframe structures, hot structures, thermal protection systems, cryogenic tankage structures, cryogenic insulations, and analysis methods for thermal/structures.

Highlights from a Mach 4 Experimental Demonstration of Inlet Mode Transition for Turbine-Based Combined Cycle Hypersonic Propulsion

NASA Technical Reports Server (NTRS)

Foster, Lancert E.; Saunders, John D., Jr.; Sanders, Bobby W.; Weir, Lois J.

2012-01-01

NASA is focused on technologies for combined cycle, air-breathing propulsion systems to enable reusable launch systems for access to space. Turbine Based Combined Cycle (TBCC) propulsion systems offer specific impulse (Isp) improvements over rocket-based propulsion systems in the subsonic takeoff and return mission segments along with improved safety. Among the most critical TBCC enabling technologies are: 1) mode transition from the low speed propulsion system to the high speed propulsion system, 2) high Mach turbine engine development and 3) innovative turbine based combined cycle integration. To address these challenges, NASA initiated an experimental mode transition task including analytical methods to assess the state-of-the-art of propulsion system performance and design codes. One effort has been the Combined-Cycle Engine Large Scale Inlet Mode Transition Experiment (CCE-LIMX) which is a fully integrated TBCC propulsion system with flowpath sizing consistent with previous NASA and DoD proposed Hypersonic experimental flight test plans. This experiment was tested in the NASA GRC 10 by 10-Foot Supersonic Wind Tunnel (SWT) Facility. The goal of this activity is to address key hypersonic combined-cycle engine issues including: (1) dual integrated inlet operability and performance issues-unstart constraints, distortion constraints, bleed requirements, and controls, (2) mode-transition sequence elements caused by switching between the turbine and the ramjet/scramjet flowpaths (imposed variable geometry requirements), and (3) turbine engine transients (and associated time scales) during transition. Testing of the initial inlet and dynamic characterization phases were completed and smooth mode transition was demonstrated. A database focused on a Mach 4 transition speed with limited off-design elements was developed and will serve to guide future TBCC system studies and to validate higher level analyses.

Parametric Study of Cantilever Plates Exposed to Supersonic and Hypersonic Flows

NASA Astrophysics Data System (ADS)

Sri Harsha, A.; Rizwan, M.; Kuldeep, S.; Giridhara Prasad, A.; Akhil, J.; Nagaraja, S. R.

2017-08-01

Analysis of hypersonic flows associated with re-entry vehicles has gained a lot of significance due to the advancements in Aerospace Engineering. An area that is studied extensively by researchers is the simultaneous reduction aerodynamic drag and aero heating in re-entry vehicles. Out of the many strategies being studied, the use of aerospikes at the stagnation point of the vehicle is found to give favourable results. The structural stability of the aerospike becomes important as it is exposed to very high pressures and temperatures. Keeping this in view, the deflection and vibration of an inclined cantilever plate in hypersonic flow is carried out using ANSYS. Steady state pressure distribution obtained from Fluent is applied as load to the transient structural module for analysis. After due validation of the methods, the effects of parameters like flow Mach number, plate inclination and plate thickness on the deflection and vibration are studied.

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 desired capability.

High temperature NASP engine seals: A technology review

NASA Technical Reports Server (NTRS)

Steinetz, Bruce M.; Dellacorte, Christopher; Tong, Mike

1991-01-01

Progress in developing advanced high temperature engine seal concepts and related sealing technologies for advanced hypersonic engines are reviewed. Design attributes and issues requiring further development for both the ceramic wafer seal and the braided ceramic rope seal are examined. Leakage data are presented for these seals for engine simulated pressure and temperature conditions and compared to a target leakage limit. Basic elements of leakage flow models to predict leakage rates for each of these seals over the wide range of pressure and temperature conditions anticipated in the engine are also presented.

X-43C Plans and Status

NASA Technical Reports Server (NTRS)

Moses, Paul L.

2003-01-01

X-43C Project is a hypersonic flight demonstration being executed as a collaboration between the National Aeronautics and Space Administration (NASA) and the United States Air Force (USAF). X-43C will expand the hypersonic flight envelope for air breathing engines beyond the history making efforts of the Hyper-X Program (X-43A). X-43C will demonstrate sustained accelerating flight during three flight tests of expendable X-43C Demonstrator Vehicles (DVs). The approximately 16-foot long X-43C DV will be boosted to the starting test conditions, separate from the booster, and accelerate from Mach 5 to Mach 7 under its own power and autonomous control. The DVs are to be powered by a liquid hydrocarbon-fueled, fuel-cooled, dual-mode, airframe integrated scramjet engine system developed under the USAF HyTech Program. The Project is managed by NASA Langley Research Center as part of NASA s Next Generation Launch Technology Program. Flight tests will be conducted by NASA Dryden Flight Research Center over water off the coast of California in the Pacific Test Range. The NASA/USAF/industry project is a natural extension of the Hyper-X Program (X-43A), which will demonstrate short duration ( 10 seconds) gaseous hydrogen-fueled scramjet powered flight at Mach 7 and Mach 10 using a heavyweight, largely heat sink construction, experimental engine. The X-43C Project will demonstrate sustained accelerating flight from Mach 5 to Mach 7 ( 4 minutes) using a flight-weight, fuel-cooled, scramjet engine powered by much denser liquid hydrocarbon fuel. The X-43C DV design flows from integrating USAF HyTech developed engine technologies with a NASA Air Breathing Launch Vehicle accelerator-class configuration and Hyper-X heritage vehicle systems designs. This paper describes the X-43C Project and provides background for NASA s current hypersonic flight demonstration efforts.

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.

Numerical simulation of aerothermal loads in hypersonic engine inlets due to shock impingement

NASA Technical Reports Server (NTRS)

Ramakrishnan, R.

1992-01-01

The effect of shock impingement on an axial corner simulating the inlet of a hypersonic vehicle engine is modeled using a finite-difference procedure. A three-dimensional dynamic grid adaptation procedure is utilized to move the grids to regions with strong flow gradients. The adaptation procedure uses a grid relocation stencil that is valid at both the interior and boundary points of the finite-difference grid. A linear combination of spatial derivatives of specific flow variables, calculated with finite-element interpolation functions, are used as adaptation measures. This computational procedure is used to study laminar and turbulent Mach 6 flows in the axial corner. The description of flow physics and qualitative measures of heat transfer distributions on cowl and strut surfaces obtained from the analysis are compared with experimental observations. Conclusions are drawn regarding the capability of the numerical scheme for enhanced modeling of high-speed compressible flows.

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.

Multidisciplinary Analysis of a Hypersonic Engine

NASA Technical Reports Server (NTRS)

Stewart, M. E. M.; Suresh, A.; Liou, M. S.; Owen, A. K.; Messitt, D. G.

2002-01-01

This paper describes implementation of a technique used to obtain a high fidelity fluid-thermal-structural solution of a combined cycle engine at its scram design point. Single-discipline simulations are insufficient here since interactions from other disciplines are significant. Using off-the-shelf, validated solvers for the fluid, chemistry, thermal, and structural solutions, this approach couples together their results to obtain consistent solutions.

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.

Numerical Propulsion System Simulation

NASA Technical Reports Server (NTRS)

Naiman, Cynthia

2006-01-01

The NASA Glenn Research Center, in partnership with the aerospace industry, other government agencies, and academia, is leading the effort to develop an advanced multidisciplinary analysis environment for aerospace propulsion systems called the Numerical Propulsion System Simulation (NPSS). NPSS is a framework for performing analysis of complex systems. The initial development of NPSS focused on the analysis and design of airbreathing aircraft engines, but the resulting NPSS framework may be applied to any system, for example: aerospace, rockets, hypersonics, power and propulsion, fuel cells, ground based power, and even human system modeling. NPSS provides increased flexibility for the user, which reduces the total development time and cost. It is currently being extended to support the NASA Aeronautics Research Mission Directorate Fundamental Aeronautics Program and the Advanced Virtual Engine Test Cell (AVETeC). NPSS focuses on the integration of multiple disciplines such as aerodynamics, structure, and heat transfer with numerical zooming on component codes. Zooming is the coupling of analyses at various levels of detail. NPSS development includes capabilities to facilitate collaborative engineering. The NPSS will provide improved tools to develop custom components and to use capability for zooming to higher fidelity codes, coupling to multidiscipline codes, transmitting secure data, and distributing simulations across different platforms. These powerful capabilities extend NPSS from a zero-dimensional simulation tool to a multi-fidelity, multidiscipline system-level simulation tool for the full development life cycle.

Experimental Supersonic Combustion Research at NASA Langley

NASA Technical Reports Server (NTRS)

Rogers, R. Clayton; Capriotti, Diego P.; Guy, R. Wayne

1998-01-01

Experimental supersonic combustion research related to hypersonic airbreathing propulsion has been actively underway at NASA Langley Research Center (LaRC) since the mid-1960's. This research involved experimental investigations of fuel injection, mixing, and combustion in supersonic flows and numerous tests of scramjet engine flowpaths in LaRC test facilities simulating flight from Mach 4 to 8. Out of this research effort has come scramjet combustor design methodologies, ground test techniques, and data analysis procedures. These technologies have progressed steadily in support of the National Aero-Space Plane (NASP) program and the current Hyper-X flight demonstration program. During NASP nearly 2500 tests of 15 scramjet engine models were conducted in LaRC facilities. In addition, research supporting the engine flowpath design investigated ways to enhance mixing, improve and apply nonintrusive diagnostics, and address facility operation. Tests of scramjet combustor operation at conditions simulating hypersonic flight at Mach numbers up to 17 also have been performed in an expansion tube pulse facility. This paper presents a review of the LaRC experimental supersonic combustion research efforts since the late 1980's, during the NASP program, and into the Hyper-X Program.

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.

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.

Electro-chemical grinding

NASA Technical Reports Server (NTRS)

Feagans, P. L.

1972-01-01

Electro-chemical grinding technique has rotation speed control, constant feed rates, and contour control. Hypersonic engine parts of nickel alloys can be almost 100% machined, keeping tool pressure at virtual zero. Technique eliminates galling and permits constant surface finish and burr-free interrupted cutting.

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 system developed with the data obtained from the first two phases. Plans for a fourth phase include mode transition experiments with a turbine engine. This paper, focusing on the first two phases of experiments, presents developed operational and analysis tools for streamlined testing and data reduction procedures.

Theory and Design of Flight-Vehicle Engines

NASA Technical Reports Server (NTRS)

Zhdanov, V. T. (Editor); Kurziner, R. I. (Editor)

1987-01-01

Papers are presented on such topics as the testing of aircraft engines, errors in the experimental determination of the parameters of scramjet engines, the effect of the nonuniformity of supersonic flow with shocks on friction and heat transfer in the channel of a hypersonic ramjet engine, and the selection of the basic parameters of cooled GTE turbines. Consideration is also given to the choice of optimal total wedge angle for the acceleration of aerospace vehicles, the theory of an electromagnetic-resonator engine, the dynamic characteristics of the pumps and turbines of liquid propellant rocket engines in transition regimes, and a hierarchy of mathematical models for spacecraft control engines.

Hypersonic engine seal development at NASA Lewis Research Center

NASA Technical Reports Server (NTRS)

Steinetz, Bruce M.

1994-01-01

NASA Lewis Research Center is developing advanced seal concepts and sealing technology for advanced combined cycle ramjet/scramjet engines being designed for the National Aerospace Plane (NASP). Technologies are being developed for both the dynamic seals that seal the sliding interfaces between articulating engine panels and sidewalls, and for the static seals that seal the heat exchanger to back-up structure interfaces. This viewgraph presentation provides an overview of the candidate engine seal concepts, seal material assessments, and unique test facilities used to assess the leakage and thermal performance of the seal concepts.

Subsonic longitudinal aerodynamic characteristics and engine pressure distributions for an aircraft with an integrated scramjet designed for Mach 6 cruise. [conducted in Langley 7 by 10 foot high speed tunnel

NASA Technical Reports Server (NTRS)

Huffman, J. K.; Fox, C. H., Jr.; Johnston, P. J.

1977-01-01

A 1/10-scale model of a proposed hypersonic aircraft with an integrated scramjet was tested. The investigation took place over a Mach number range from 0.2 to 0.7 and an angle of attack range from 2 deg to approximately 17 deg at a sideslip angle of 0 deg. The primary configuration variables studied were engine location, internal engine geometry, and external engine geometry. The results are presented without analysis.

Hypersonic engine seal development at NASA Lewis Research Center

NASA Astrophysics Data System (ADS)

Steinetz, Bruce M.

1994-07-01

NASA Lewis Research Center is developing advanced seal concepts and sealing technology for advanced combined cycle ramjet/scramjet engines being designed for the National Aerospace Plane (NASP). Technologies are being developed for both the dynamic seals that seal the sliding interfaces between articulating engine panels and sidewalls, and for the static seals that seal the heat exchanger to back-up structure interfaces. This viewgraph presentation provides an overview of the candidate engine seal concepts, seal material assessments, and unique test facilities used to assess the leakage and thermal performance of the seal concepts.

Thrust modeling for hypersonic engines

NASA Technical Reports Server (NTRS)

Riggins, D. W.; Mcclinton, C. R.

1995-01-01

Expressions for the thrust losses of a scramjet engine are developed in terms of irreversible entropy increases and the degree of incomplete combustion. A method is developed which allows the calculation of the lost vehicle thrust due to different loss mechanisms within a given flow-field. This analysis demonstrates clearly the trade-off between mixing enhancement and resultant increased flow losses in scramjet combustors. An engine effectiveness parameter is defined in terms of thrust loss. Exergy and the thrust-potential method are related and compared.

Innovative Approaches to Fuel-Air Mixing and Combustion in Airbreathing Hypersonic Engines

NASA Astrophysics Data System (ADS)

MacLeod, C.

This paper describes some innovative methods for achieving enhanced fuel-air mixing and combustion in Scramjet-like spaceplane engines. A multimodal approach to the problem is discussed; this involves using several concurrent methods of forced mixing. The paper concentrates on Electromagnetic Activation (EMA) and Electrostatic Attraction as suitable techniques for this purpose - although several other potential methods are also discussed. Previously published empirical data is used to draw conclusions about the likely effectiveness of the system and possible engine topologies are outlined.

Hypersonic Engine Leading Edge Experiments in a High Heat Flux, Supersonic Flow Environment

NASA Technical Reports Server (NTRS)

Gladden, Herbert J.; Melis, Matthew E.

1994-01-01

A major concern in advancing the state-of-the-art technologies for hypersonic vehicles is the development of an aeropropulsion system capable of withstanding the sustained high thermal loads expected during hypersonic flight. Three aerothermal load related concerns are the boundary layer transition from laminar to turbulent flow, articulating panel seals in high temperature environments, and strut (or cowl) leading edges with shock-on-shock interactions. A multidisciplinary approach is required to address these technical concerns. A hydrogen/oxygen rocket engine heat source has been developed at the NASA Lewis Research Center as one element in a series of facilities at national laboratories designed to experimentally evaluate the heat transfer and structural response of the strut (or cowl) leading edge. A recent experimental program conducted in this facility is discussed and related to cooling technology capability. The specific objective of the experiment discussed is to evaluate the erosion and oxidation characteristics of a coating on a cowl leading edge (or strut leading edge) in a supersonic, high heat flux environment. Heat transfer analyses of a similar leading edge concept cooled with gaseous hydrogen is included to demonstrate the complexity of the problem resulting from plastic deformation of the structures. Macro-photographic data from a coated leading edge model show progressive degradation over several thermal cycles at aerothermal conditions representative of high Mach number flight.

Advanced ceramic matrix composite materials for current and future propulsion technology applications

NASA Astrophysics Data System (ADS)

Schmidt, S.; Beyer, S.; Knabe, H.; Immich, H.; Meistring, R.; Gessler, A.

2004-08-01

Current rocket engines, due to their method of construction, the materials used and the extreme loads to which they are subjected, feature a limited number of load cycles. Various technology programmes in Europe are concerned, besides developing reliable and rugged, low cost, throwaway equipment, with preparing for future reusable propulsion technologies. One of the key roles for realizing reusable engine components is the use of modern and innovative materials. One of the key technologies which concern various engine manufacturers worldwide is the development of fibre-reinforced ceramics—ceramic matrix composites. The advantages for the developers are obvious—the low specific weight, the high specific strength over a large temperature range, and their great damage tolerance compared to monolithic ceramics make this material class extremely interesting as a construction material. Over the past years, the Astrium company (formerly DASA) has, together with various partners, worked intensively on developing components for hypersonic engines and liquid rocket propulsion systems. In the year 2000, various hot-firing tests with subscale (scale 1:5) and full-scale nozzle extensions were conducted. In this year, a further decisive milestone was achieved in the sector of small thrusters, and long-term tests served to demonstrate the extraordinary stability of the C/SiC material. Besides developing and testing radiation-cooled nozzle components and small-thruster combustion chambers, Astrium worked on the preliminary development of actively cooled structures for future reusable propulsion systems. In order to get one step nearer to this objective, the development of a new fibre composite was commenced within the framework of a regionally sponsored programme. The objective here is to create multidirectional (3D) textile structures combined with a cost-effective infiltration process. Besides material and process development, the project also encompasses the development of special metal/ceramic and ceramic/ceramic joining techniques as well as studying and verifying non destructive investigation processes for the purpose of testing components.

An approximate method for calculating three-dimensional inviscid hypersonic flow fields

NASA Technical Reports Server (NTRS)

Riley, Christopher J.; Dejarnette, Fred R.

1990-01-01

An approximate solution technique was developed for 3-D inviscid, hypersonic flows. The method employs Maslen's explicit pressure equation in addition to the assumption of approximate stream surfaces in the shock layer. This approximation represents a simplification to Maslen's asymmetric method. The present method presents a tractable procedure for computing the inviscid flow over 3-D surfaces at angle of attack. The solution procedure involves iteratively changing the shock shape in the subsonic-transonic region until the correct body shape is obtained. Beyond this region, the shock surface is determined using a marching procedure. Results are presented for a spherically blunted cone, paraboloid, and elliptic cone at angle of attack. The calculated surface pressures are compared with experimental data and finite difference solutions of the Euler equations. Shock shapes and profiles of pressure are also examined. Comparisons indicate the method adequately predicts shock layer properties on blunt bodies in hypersonic flow. The speed of the calculations makes the procedure attractive for engineering design applications.

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 to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

Invariance of Hypersonic Normal Force Coefficients with Reynolds Number and Determination of Inviscid Wave Drag from Laminar Experimental Results

NASA Technical Reports Server (NTRS)

Hawkins, Richard; Penland, Jim A.

1997-01-01

Observations have been made and reported that the experimental normal force coefficients at a constant angle of attack were constant with a variation of more than 2 orders of magnitude of Reynolds number at a free-stream Mach number M(sub infinity) of 8.00 and more than 1 order of magnitude variation at M(sub infinity) = 6.00 on the same body-wing hypersonic cruise configuration. These data were recorded under laminar, transitional, and turbulent boundary layer conditions with both hot-wall and cold-wall models. This report presents experimental data on 25 configurations of 17 models of both simple and complex geometry taken at M(sub infinity) = 6.00, 6.86, and 8.00 in 4 different hypersonic facilities. Aerodynamic calculations were made by computational fluid dynamics (CID) and engineering methods to analyze these data. The conclusions were that the normal force coefficients at a given altitude are constant with Reynolds numbers at hypersonic speeds and that the axial force coefficients recorded under laminar boundary-layer conditions at several Reynolds numbers may be plotted against the laminar parameter (the reciprocal of the Reynolds number to the one-half power) and extrapolated to the ordinate axis to determine the inviscid-wave-drag coefficient at the intercept.

Recent advances in convectively cooled engine and airframe structures for hypersonic flight

NASA Technical Reports Server (NTRS)

Kelly, H. N.; Wieting, A. R.; Shore, C. P.; Nowak, R. J.

1978-01-01

A hydrogen-cooled structure for a fixed-geometry, airframe-integrated scramjet is described. The thermal/structural problems, concepts, design features, and technological advances are applicable to a broad range of engines. Convectively cooled airframe structural concepts that have evolved from an extensive series of investigations, the technology developments that have led to these concepts, and the benefits that accrue from their use are discussed.

Development of braided rope seals for hypersonic engine applications: Flow modeling

NASA Technical Reports Server (NTRS)

Mutharasan, Rajakkannu; Steinetz, Bruce M.; Tao, Xiaoming; Du, Guang-Wu; Ko, Frank

1992-01-01

A new type of engine seal is being developed to meet the needs of advanced hypersonic engines. A seal braided of emerging high temperature ceramic fibers comprised of a sheath-core construction was selected for study based on its low leakage rates. Flexible, low-leakage, high temperature seals are required to seal the movable engine panels of advanced ramjet-scramjet engines either preventing potentially dangerous leakage into backside engine cavities or limiting the purge coolant flow rates through the seals. To predict the leakage through these flexible, porous seal structures new analytical flow models are required. Two such models based on the Kozeny-Carman equations are developed herein and are compared to experimental leakage measurements for simulated pressure and seal gap conditions. The models developed allow prediction of the gas leakage rate as a function of fiber diameter, fiber packing density, gas properties, and pressure drop across the seal. The first model treats the seal as a homogeneous fiber bed. The second model divides the seal into two homogeneous fiber beds identified as the core and the sheath of the seal. Flow resistances of each of the main seal elements are combined to determine the total flow resistance. Comparisons between measured leakage rates and model predictions for seal structures covering a wide range of braid architectures show good agreement. Within the experimental range, the second model provides a prediction within 6 to 13 percent of the flow for many of the cases examined. Areas where future model refinements are required are identified.

Canard-Elevon Interactions on a Hypersonic Vehicle

DTIC Science & Technology

2010-01-01

AIAA. †Professor, Department of Aerospace Engineering, 1320 Beal Avenue, 3024 Francois-Xavier Bagnoud Building; cesnik@umich.edu. Associate Fellow...high number of grid points on each airfoil face, the CFD squares are clustered close together and, therefore, appear to be a solid line. V . Parametric

A Pre-Mixed Shock-Induced-Combustion Approach to Inlet and Combustor Design for Hypersonic Applications

NASA Technical Reports Server (NTRS)

Weidner, John P.

1996-01-01

The need for efficient access to space has created interest in airbreathing propulsion as a means of achieving that goal. The NASP program explored a single-stage-to-orbit approach which could require scramjet airbreathing propulsion out to Mach 16 to 20. Recent interest in global access could require hypersonic cruise engines operating efficiently in the Mach 10 to 12 speed range. A common requirement of both these types of propulsion systems is that they would have to be fully integrated with the aero configuration so that the forebody becomes a part of the external compression inlet and the nozzle expansion is completed on the vehicle aftbody.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Froning, H. David Jr

Although Australia has no Beamed Energy Propulsion programs at the present time, it is accomplishing significant scientific and technological activity that is of potential relevance to Beamed Energy Propulsion (BEP). These activities include: continual upgrading and enhancement of the Woomera Test Facility, Which is ideal for development and test of high power laser or microwave systems and the flight vehicles they would propel; collaborative development and test, with the US and UK of hypersonic missiles that embody many features needed by beam-propelled flight vehicles; hypersonic air breathing propulsion systems that embody inlet-engine-nozzle features needed for beam-riding agility by air breathingmore » craft; and research on specially conditioned EM fields that could reduce beamed energy lost during atmospheric propagation.« less

Cavity-type hypersonic phononic crystals

NASA Astrophysics Data System (ADS)

Sato, A.; Pennec, Y.; Yanagishita, T.; Masuda, H.; Knoll, W.; Djafari-Rouhani, B.; Fytas, G.

2012-11-01

We report on the engineering of the phonon dispersion diagram in monodomain anodic porous alumina (APA) films through the porosity and physical state of the material residing in the nanopores. Lattice symmetry and inclusion materials are theoretically identified to be the main factors which control the hypersonic acoustic wave propagation. This involves the interaction between the longitudinal and the transverse modes in the effective medium and a flat band characteristic of the material residing in the cavities. Air and filled nanopores, therefore, display markedly different dispersion relations and the inclusion materials lead to a locally resonant structural behavior uniquely determining their properties under confinement. APA films emerge as a new platform to investigate the rich acoustic phenomena of structured composite matter.

Shock-wave flow regimes at entry into the diffuser of a hypersonic ramjet engine: Influence of physical properties of the gas medium

NASA Astrophysics Data System (ADS)

Tarnavskii, G. A.

2006-07-01

The physical aspects of the effective-adiabatic-exponent model making it possible to decompose the total problem on modeling of high-velocity gas flows into individual subproblems (“physicochemical processes” and “ aeromechanics”), which ensures the creation of a universal and efficient computer complex divided into a number of independent units, have been analyzed. Shock-wave structures appearing at entry into the duct of a hypersonic aircraft have been investigated based on this methodology, and the influence of the physical properties of the gas medium in a wide range of variations of the effective adiabatic exponent has been studied.

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.

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.

Effects of reaction control system jet simulation on the stability and control characteristics of a 0.015-scale space shuttle orbiter model in the Ames Research Center 3.5-foot hypersonic wind tunnel

NASA Technical Reports Server (NTRS)

Dziubala, T. J.; Marroquin, J.; Cleary, J. W.; Mellenthin, J. A.

1973-01-01

An experimental investigation was performed in the Ames Research Center 3.5-Foot Hypersonic Wind Tunnel to obtain detailed effects which interactions between the RCS jet flow field and the local orbiter flow field have on orbiter hypersonic stability and control characteristics. Six-component force data were obtained through an angle-of-attack range of 15 to 35 deg with 0 deg angle of sideslip. The test was conducted with yaw, pitch and roll jet simulation at a free-stream Mach number of 10.3. These data simulate two SSV reentry flight conditions at Mach numbers of 28.3 and 10.3. Fuselage base pressures and pressures on the nonmetric RCS pods were obtained in addition to the basic force measurements. Model 42-0 was used for these tests.

An extended supersonic combustion model for the dynamic analysis of hypersonic vehicles

NASA Technical Reports Server (NTRS)

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

1993-01-01

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.

Critical Design Parameters for Pylon-Aided Gaseous Fuel Injection Upstream of a Flameholding Cavity

DTIC Science & Technology

2009-03-01

Spencer Bowen and Neil Rogers was critical to not only the successful completion of this research but my sanity as well. The care, support, and motivation...missile presents a relatively inexpensive alternative that while complex is much more feasible to achieve in the near term. The hypersonic bomber and...scramjet engine operates supersonically throughout. Though aerodynamically complex the scramjet engine is relatively simple, consisting of an inlet

Rocket-Based Combined-Cycle (RBCC) Propulsion Technology Workshop. Tutorial session

NASA Technical Reports Server (NTRS)

1992-01-01

The goal of this workshop was to illuminate the nation's space transportation and propulsion engineering community on the potential of hypersonic combined cycle (airbreathing/rocket) propulsion systems for future space transportation applications. Four general topics were examined: (1) selections from the expansive advanced propulsion archival resource; (2) related propulsion systems technical backgrounds; (3) RBCC engine multimode operations related subsystem background; and (4) focused review of propulsion aspects of current related programs.

Prediction of forces and moments for flight vehicle control effectors: Workplan

NASA Technical Reports Server (NTRS)

Maughmer, Mark D.

1989-01-01

Two research activities directed at hypersonic vehicle configurations are currently underway. The first involves the validation of a number of classical local surface inclination methods commonly employed in preliminary design studies of hypersonic flight vehicles. Unlike studies aimed at validating such methods for predicting overall vehicle aerodynamics, this effort emphasizes validating the prediction of forces and moments for flight control studies. Specifically, several vehicle configurations for which experimental or flight-test data are available are being examined. By comparing the theoretical predictions with these data, the strengths and weaknesses of the local surface inclination methods can be ascertained and possible improvements suggested. The second research effort, of significance to control during take-off and landing of most proposed hypersonic vehicle configurations, is aimed at determining the change due to ground effect in control effectiveness of highly swept delta planforms. Central to this research is the development of a vortex-lattice computer program which incorporates an unforced trailing vortex sheet and an image ground plane. With this program, the change in pitching moment of the basic vehicle due to ground proximity, and whether or not there is sufficient control power available to trim, can be determined. In addition to the current work, two different research directions are suggested for future study. The first is aimed at developing an interactive computer program to assist the flight controls engineer in determining the forces and moments generated by different types of control effectors that might be used on hypersonic vehicles. The first phase of this work would deal in the subsonic portion of the flight envelope, while later efforts would explore the supersonic/hypersonic flight regimes. The second proposed research direction would explore methods for determining the aerodynamic trim drag of a generic hypersonic flight vehicle and ways in which it can be minimized through vehicle design and trajectory optimization.

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.

Advanced Computational Techniques for Hypersonic Propulsion

NASA Technical Reports Server (NTRS)

Povinelli, Louis A.

1996-01-01

CFD has played a major role in the resurgence of hypersonic flight, on the premise that numerical methods will allow us to perform simulations at conditions for which no ground test capability exists. Validation of CFD methods is being established using the experimental data base available, which is below Mach 8. It is important, however, to realize the limitations involved in the extrapolation process as well as the deficiencies that exist in numerical methods at the present time. Current features of CFD codes are examined for application to propulsion system components. The shortcomings in simulation and modeling are identified and discussed.

Performance evaluations of oxidation-resistant carbon-carbon composites in simulated hypersonic vehicle environments

NASA Technical Reports Server (NTRS)

Barrett, D. M.; Maahs, H. G.; Ohlhorst, C. W.; Vaughn, W. L.; Martin, R. H.

1989-01-01

An evaluation is made of the oxidation-protection requirements of carbon-carbon composite (CCC) structural components in a hypersonic vehicle aerothermodynamic environment, where maximum test temperatures in air are of the order of 2800 F, and pressures range from 0.03 to 1.0 atm. The specimens were exposed to high humidity between tests. Attention was given to the effects of coating composition and thickness, and of substrate architecture and surface preparation, on the oxidation resistance of CCCs. Both surface preparation and coating chemistry have a profound effect on coating adherence and longevity.

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 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 the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

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.

Combustion Efficiency, Flameout Operability Limits and General Design Optimization for Integrated Ramjet-Scramjet Hypersonic Vehicles

NASA Astrophysics Data System (ADS)

Mbagwu, Chukwuka Chijindu

High speed, air-breathing hypersonic vehicles encounter a varied range of engine and operating conditions traveling along cruise/ascent missions at high altitudes and dynamic pressures. Variations of ambient pressure, temperature, Mach number, and dynamic pressure can affect the combustion conditions in conflicting ways. Computations were performed to understand propulsion tradeoffs that occur when a hypersonic vehicle travels along an ascent trajectory. Proper Orthogonal Decomposition methods were applied for the reduction of flamelet chemistry data in an improved combustor model. Two operability limits are set by requirements that combustion efficiency exceed selected minima and flameout be avoided. A method for flameout prediction based on empirical Damkohler number measurements is presented. Operability limits are plotted that define allowable flight corridors on an altitude versus flight Mach number performance map; fixed-acceleration ascent trajectories were considered for this study. Several design rules are also presented for a hypersonic waverider with a dual-mode scramjet engine. Focus is placed on ''vehicle integration" design, differing from previous ''propulsion-oriented" design optimization. The well-designed waverider falls between that of an aircraft (high lift-to-drag ratio) and a rocket (high thrust-to-drag ratio). 84 variations of an X-43-like vehicle were run using the MASIV scramjet reduced order model to examine performance tradeoffs. Informed by the vehicle design study, variable-acceleration trajectory optimization was performed for three constant dynamic pressures ascents. Computed flameout operability limits were implemented as additional constraints to the optimization problem. The Michigan-AFRL Scramjet In-Vehicle (MASIV) waverider model includes finite-rate chemistry, applied scaling laws for 3-D turbulent mixing, ram-scram transition and an empirical value of the flameout Damkohler number. A reduced-order modeling approach is justified (in lieu of higher-fidelity computational simulations) because all vehicle forces are computed multiple thousands of times to generate multi-dimensional performance maps. The findings of this thesis work present a number of compelling conclusions. It is found that the ideal operating conditions of a scramjet engine are heavily dependent on the ambient and combustor pressure (and less strongly on temperature). Combustor pressures of approximately 1.0 bar or greater achieve the highest combustion efficiency, in line with industry standards of more than 0.5 bar. Ascent trajectory analysis of combustion efficiency and lean-limit flameout dictate best operation at higher dynamic pressures and lower altitudes, but these goals are traded off by current structural limitations whereby dynamic pressures must remain below 100 kPa. Hypersonic waverider designs varied between an "airplane" and a "rocket" are found to have better performance with the latter design, with controllability and minimum elevon/rudder surface area as a stability constraint for the vehicle trim. Ultimately, these findings are beneficial and contribute to the overall understanding of dynamically stable waverider vehicles at hypersonic speeds. These types of vehicles have a range of applications from technology demonstration, to earth-to-low orbit payload transit, to most compellingly another step in the development and realization of viable supersonic commercial transport.

2007 Research and Engineering Annual Report

NASA Technical Reports Server (NTRS)

Stoliker, Patrick; Bowers, Albion; Cruciani, Everlyn

2008-01-01

Selected research and technology activities at NASA Dryden Flight Research Center are summarized. These following activities exemplify the Center's varied and productive research efforts: Developing a Requirements Development Guide for an Automatic Ground Collision Avoidance System; Digital Terrain Data Compression and Rendering for Automatic Ground Collision Avoidance Systems; Nonlinear Flutter/Limit Cycle Oscillations Prediction Tool; Nonlinear System Identification Using Orthonormal Bases: Application to Aeroelastic/Aeroservoelastic Systems; Critical Aerodynamic Flow Feature Indicators: Towards Application with the Aerostructures Test Wing; Multidisciplinary Design, Analysis, and Optimization Tool Development Using a Genetic Algorithm; Structural Model Tuning Capability in an Object-Oriented Multidisciplinary Design, Analysis, and Optimization Tool; Extension of Ko Straight-Beam Displacement Theory to the Deformed Shape Predictions of Curved Structures; F-15B with Phoenix Missile and Pylon Assembly--Drag Force Estimation; Mass Property Testing of Phoenix Missile Hypersonic Testbed Hardware; ARMD Hypersonics Project Materials and Structures: Testing of Scramjet Thermal Protection System Concepts; High-Temperature Modal Survey of the Ruddervator Subcomponent Test Article; ARMD Hypersonics Project Materials and Structures: C/SiC Ruddervator Subcomponent Test and Analysis Task; Ground Vibration Testing and Model Correlation of the Phoenix Missile Hypersonic Testbed; Phoenix Missile Hypersonic Testbed: Performance Design and Analysis; Crew Exploration Vehicle Launch Abort System-Pad Abort-1 (PA-1) Flight Test; Testing the Orion (Crew Exploration Vehicle) Launch Abort System-Ascent Abort-1 (AA-1) Flight Test; SOFIA Flight-Test Flutter Prediction Methodology; SOFIA Closed-Door Aerodynamic Analyses; SOFIA Handling Qualities Evaluation for Closed-Door Operations; C-17 Support of IRAC Engine Model Development; Current Capabilities and Future Upgrade Plans of the C-17 Data Rack; Intelligent Data Mining Capabilities as Applied to Integrated Vehicle Health Management; STARS Flight Demonstration No. 2 IP Data Formatter; Space-Based Telemetry and Range Safety (STARS) Flight Demonstration No. 2 Range User Flight Test Results; Aerodynamic Effects of the Quiet Spike(tm) on an F-15B Aircraft; F-15 Intelligent Flight Controls-Increased Destabilization Failure; F-15 Integrated Resilient Aircraft Control (IRAC) Improved Adaptive Controller; Aeroelastic Analysis of the Ikhana/Fire Pod System; Ikhana: Western States Fire Missions Utilizing the Ames Research Center Fire Sensor; Ikhana: Fiber-Optic Wing Shape Sensors; Ikhana: ARTS III; SOFIA Closed-Door Flutter Envelope Flight Testing; F-15B Quiet Spike(TM) Aeroservoelastic Flight Test Data Analysis; and UAVSAR Platform Precision Autopilot Flight Results.

ASUR

NASA Technical Reports Server (NTRS)

1991-01-01

The Ecole Polytechnique Feminine (EPF) is a French engineering school for women. The students who presented the project at the Summer Conference are in the fourth year of a five year program. For the second time, the EPF worked on an aeronautical project with the Ohio State University. This year, the theme was to design a hypersonic carrier aimed to launch an orbiter at Mach 6, a range of 375 miles and an altitude of 95,000 ft. The plane was called ASUR which in French means the blue sky, the same sky that links our countries across the ocean. Moreover, ASUR is an anagram of USRA. This work benefits from work on reusable hypersonic aircraft in Europe, and especially on two of them: STAR-H and Sanger. STAR-H is a French project. This hypersonic aircraft would replace Ariane 5 in launching a shuttle smaller than Hermes. Sanger is a German project. Its objective is to launch a manned shuttle called HORUS, but Ariane 5 would be kept for heavy cargo launches. These two projects are in competition in Europe to be a launcher of the European Space Agency.

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

NASA Astrophysics Data System (ADS)

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

1994-03-01

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.

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.

The success of the X-33 depends on its technology—an overview

NASA Astrophysics Data System (ADS)

Bunting, Jackie O.; Sasso, Steven E.

1996-03-01

The success of the X-33, and therefore the Reusable Launch Vehicle (RLV) program, is highly dependent on the maturity of the components and subsystems selected and the ability to verify their performance, cost, and operability goals. The success of the technology that will be developed to support these components and subsystems will be critical to developing an operationally efficient X-33 that is traceable to a full-scale RLV system. This paper will delineate the key objectives of each technology demonstration area and provide an assessment of its ability to meet the X-33/RLV requirements. It is our intent to focus on these key technology areas to achieve the ambitious but achievable goals of the RLV and X-33 programs. Based on our assessment of the X-33 and RLV systems, we have focused on the performance verification and validation of the linear aerospike engine. This engine, first developed in the mid-1960s, shows promise in achieving the RLV objectives. Equally critical to the engine selection is the development of cryogenic composite tanks and the associated health management system required to meet the operability goals. We are also developing a highly reusable form of thermal protection system based on years of hypersonic research and Space Shuttle experience. To meet the mass fraction goals, reduction in engine component weights will also be developed. Due to the high degree of operability required, we will investigate the use of real-time integrated system health management and propulsion systems diagnostics, and mature the use of electromechanical actuators for highly reusable systems. The rapid turn-around requirements will require an adaptive guidance, navigation, and control algorithm toolset, which is well underway. We envision our X-33 and RLV to use mature, low-risk technologies that will allow truly low-cost access to space (Lockheed Martin Internal Document, 1995).

Space shuttle: Static stability characteristics and control surface effectiveness of the Boeing .00435 scale model space shuttle booster H-32

NASA Technical Reports Server (NTRS)

Houser, J. F.; Runciman, W. H.

1971-01-01

Experimental aerodynamic investigations were made in the Grumman 36-inch hypersonic wind tunnel on a .00435 scale model of the H-32 reusable space shuttle booster. The objectives of the test were to determine the static stability characteristics and control surface effectiveness at hypersonic speeds. Data were taken at M = 8.12 over a range of angles of attack between -5 and 85 deg at beta = 0 deg and over a range of side slip angles between -10 and 10 deg at alpha = 0 and 70 deg. Six component balance data and base-cavity pressure data were recorded.

A Review of Hypersonics Aerodynamics, Aerothermodynamics and Plasmadynamics Activities within NASA's Fundamental Aeronautics Program

NASA Technical Reports Server (NTRS)

Salas, Manuel D.

2007-01-01

The research program of the aerodynamics, aerothermodynamics and plasmadynamics discipline of NASA's Hypersonic Project is reviewed. Details are provided for each of its three components: 1) development of physics-based models of non-equilibrium chemistry, surface catalytic effects, turbulence, transition and radiation; 2) development of advanced simulation tools to enable increased spatial and time accuracy, increased geometrical complexity, grid adaptation, increased physical-processes complexity, uncertainty quantification and error control; and 3) establishment of experimental databases from ground and flight experiments to develop better understanding of high-speed flows and to provide data to validate and guide the development of simulation tools.

Multidisciplinary Analysis of a Hypersonic Engine

NASA Technical Reports Server (NTRS)

Suresh, Ambady; Stewart, Mark

2003-01-01

The objective is to develop high fidelity tools that can influence ISTAR design In particular, tools for coupling Fluid-Thermal-Structural simulations RBCC/TBCC designers carefully balance aerodynamic, thermal, weight, & structural considerations; consistent multidisciplinary solutions reveal details (at modest cost) At Scram mode design point, simulations give details of inlet & combustor performance, thermal loads, structural deflections.

Cooled Ceramic Matrix Composite Propulsion Structures Demonstrated

NASA Technical Reports Server (NTRS)

Jaskowiak, Martha H.; Dickens, Kevin W.

2005-01-01

NASA's Next Generation Launch Technology (NGLT) Program has successfully demonstrated cooled ceramic matrix composite (CMC) technology in a scramjet engine test. This demonstration represented the world s largest cooled nonmetallic matrix composite panel fabricated for a scramjet engine and the first cooled nonmetallic composite to be tested in a scramjet facility. Lightweight, high-temperature, actively cooled structures have been identified as a key technology for enabling reliable and low-cost space access. Tradeoff studies have shown this to be the case for a variety of launch platforms, including rockets and hypersonic cruise vehicles. Actively cooled carbon and CMC structures may meet high-performance goals at significantly lower weight, while improving safety by operating with a higher margin between the design temperature and material upper-use temperature. Studies have shown that using actively cooled CMCs can reduce the weight of the cooled flow-path component from 4.5 to 1.6 lb/sq ft and the weight of the propulsion system s cooled surface area by more than 50 percent. This weight savings enables advanced concepts, increased payload, and increased range. The ability of the cooled CMC flow-path components to operate over 1000 F hotter than the state-of-the-art metallic concept adds system design flexibility to space-access vehicle concepts. Other potential system-level benefits include smaller fuel pumps, lower part count, lower cost, and increased operating margin.

Hypersonic technology-approach to an expanded program

NASA Technical Reports Server (NTRS)

Hearth, D. P.; Preyss, A. E.

1976-01-01

An overview of research, testing, and technology in the hypersonic range. Military and civilian hypersonic flight systems envisaged, ground testing facilities under development, methods for cooling the heated airframe, and use of hydrogen as fuel and coolant are discussed extensively. Air-breathing hypersonic cruise systems are emphasized, the airframe-integrated scramjet configuration is discussed and illustrated, materials proposed for hypersonic vehicles are reviewed, and test results on hypersonic flight (X-15 research aircraft) are indicated. Major advances and major problems in hypersonic flight and hypersonic technology are outlined, and the need for a hypersonic flying-laboratory research craft is stressed.

Modification to the Langley 8-foot high temperature tunnel for hypersonic propulsion testing

NASA Technical Reports Server (NTRS)

Reubush, D. E.; Puster, R. L.; Kelly, H. N.

1987-01-01

Described are the modifications currently under way to the Langley 8-Foot High Temperature Tunnel to produce a new, unique national resource for testing hypersonic air-breathing propulsion systems. The current tunnel, which has been used for aerothermal loads and structures research since its inception, is being modified with the addition of a LOX system to bring the oxygen content of the test medium up to that of air, the addition of alternate Mach number capability (4 and 5) to augment the current M=7 capability, improvements to the tunnel hardware to reduce maintenance downtime, the addition of a hydrogen system to allow the testing of hydrogen powered engines, and a new data system to increase both the quantity and quality of the data obtained.

Laboratory Facilities and Measurement Techniques for Beamed-Energy-Propulsion Experiments in Brazil

NASA Astrophysics Data System (ADS)

de Oliveira, Antonio Carlos; Chanes Júnior, José Brosler; Cordeiro Marcos, Thiago Victor; Pinto, David Romanelli; Santos Vilela, Renan Guilherme; Barros Galvão, Victor Alves; Mantovani, Arthur Freire; da Costa, Felipe Jean; dos Santos Assenção, José Adeildo; dos Santos, Alberto Monteiro; de Paula Toro, Paulo Gilberto; Sala Minucci, Marco Antonio; da Silveira Rêgo, Israel; Salvador, Israel Irone; Myrabo, Leik N.

2011-11-01

Laser propulsion is an innovative concept of accessing the space easier and cheaper where the propulsive energy is beamed to the aerospace vehicle in flight from ground—or even satellite-based high-power laser sources. In order to be realistic about laser propulsion, the Institute for Advanced Studies of the Brazilian Air Force in cooperation with the United States Air Force and the Rensselaer Polytechnic Institute are seriously investigating its basic physics mechanisms and engineering aspects at the Henry T. Hamamatsu Laboratory of Hypersonic and Aerothermodynamics in São José dos Campos, Brazil. This paper describes in details the existing facilities and measuring systems such as high-power laser devices, pulsed-hypersonic wind tunnels and high-speed flow visualization system currently utilized in the laboratory for experimentation on laser propulsion.

Evaluation of a CFD Method for Aerodynamic Database Development using the Hyper-X Stack Configuration

NASA Technical Reports Server (NTRS)

Parikh, Paresh; Engelund, Walter; Armand, Sasan; Bittner, Robert

2004-01-01

A computational fluid dynamic (CFD) study is performed on the Hyper-X (X-43A) Launch Vehicle stack configuration in support of the aerodynamic database generation in the transonic to hypersonic flow regime. The main aim of the study is the evaluation of a CFD method that can be used to support aerodynamic database development for similar future configurations. The CFD method uses the NASA Langley Research Center developed TetrUSS software, which is based on tetrahedral, unstructured grids. The Navier-Stokes computational method is first evaluated against a set of wind tunnel test data to gain confidence in the code s application to hypersonic Mach number flows. The evaluation includes comparison of the longitudinal stability derivatives on the complete stack configuration (which includes the X-43A/Hyper-X Research Vehicle, the launch vehicle and an adapter connecting the two), detailed surface pressure distributions at selected locations on the stack body and component (rudder, elevons) forces and moments. The CFD method is further used to predict the stack aerodynamic performance at flow conditions where no experimental data is available as well as for component loads for mechanical design and aero-elastic analyses. An excellent match between the computed and the test data over a range of flow conditions provides a computational tool that may be used for future similar hypersonic configurations with confidence.

Hypersonics. Volume 1 - Defining the hypersonic environment; Proceedings of the First Joint Europe/U.S. Short Course on Hypersonics, Paris, France, Dec. 7-11, 1987

NASA Astrophysics Data System (ADS)

Bertin, John J.; Glowinski, Roland; Periaux, Jacques

1989-05-01

The present work discusses the general characterization of hypersonic flows, the hypersonic phenomena to be encountered by the Hermes spacecraft, industrial methodologies for the design of hypersonic vehicles, the definition of aerodynamic methodology, and hypersonic airbreathing-propulsion vehicle design practices applicable to the U.S. National Aerospace Plane. Also discussed are real gas effects in the hypersonic regime, the influence of thermochemistry and of nonequilibrium and surface catalysis on hypersonic vehicle design, the modelling of nonequilibrium effects in high speed flows, air-dissociation thermochemistry, and rarefied gas dynamics effects for spacecraft.

Turbojet-type engines for the airbreathing propulsion of reusable winged launchers

NASA Astrophysics Data System (ADS)

Duparcq, J. L.; Hermant, E.; Scherrer, D.

Combined propulsion systems for hypersonic application have become new challenges for industrial and research organizations. In France, SNECMA and SEP, which have just joined together for a common effort on hypersonics within Hyperspace, and ONERA have been involved, under CNES (French space agency) contracts, in the assessment of new propulsion concepts for reusable winged launchers (SSTO or TSTO). As potential solutions for the airbreathing propulsion, some turbojet-type engines are presented: —the twin spool turbojet or turbofan with reheat —the turbojet with reheat —the twin-duct turbojet ramjet —the precooled turbojet with reheat. All these engines have been sized for a flight Mach number under seven with a cryogenic fuel (liquid hydrogen). Mainly due to total temperature and pressure encountered along the trajectory, the systems will have to withstand severe physical constraints. Coupled with performance and size requirements, like specific thrust and maximum air capture area, these operating conditions have been taken into account in order to select each engine cycle and technical arrangement. Performance and mass criteria make it possible to compare these systems and to emphasize their distinctive features among the propulsion concepts envisioned for the future reusable winged launchers (including airbreathing combined engines under study in France). The first step of the final selection, leading to the best adaptation between the engine and the vehicle, will then be tackled. This will be particularly enhanced by the analysis of potential advantages or technical difficulties, like thrust-to-weight ratio or needs of variable geometry and heat exchangers. The twin-duct turbojet ramjet, for example, is probably one of the best candidates for the first stages of propulsion of a reusable winged launcher.

Damage accumulation in titanium matrix composites under generic hypersonic vehicle flight simulation and sustained loads

DOE Office of Scientific and Technical Information (OSTI.GOV)

Johnson, W.S.; Mirdamadi, M.; Bakuckas, J.G. Jr.

1996-12-31

Titanium matrix composites (TMC), such as Ti-15V-3Cr-3Al-3Sn (Ti-15-3) reinforced with continuous silicon-carbide fibers (SCS-6), are being evaluated for use in hypersonic vehicles and advanced gas turbine engines where high strength-to-weight and high stiffness-to-weight ratios at elevated temperatures are critical. Such applications expose the composite to mechanical fatigue loading as well as thermally induced cycles. The damage accumulation behavior of a [0/90]2s laminate made of Ti-15V-3Cr-3Al-3Sn (Ti-15-3) reinforced with continuous silicon-carbide fibers (SCS-6) subjected to a simulated generic hypersonic flight profile, portions of the flight profile, and sustained loads was evaluated experimentally. Portions of the flight profile were used separately tomore » isolate combinations of load and time at temperature that influenced the fatigue behavior of the composite. Sustained load tests were also conducted and the results were compared with the fatigue results under the flight profile and its portions. The test results indicated that the fatigue strength of this materials system is considerably reduced by a combination of load and time at temperature.« less

The X-43A hypersonic research aircraft and its modified Pegasus booster rocket recently underwent c

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, Calif. The combined systems test was one of the last major milestones in the Hyper-X research program before the first X-43A flight. 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, or five times the speed of sound). The 12-foot, unpiloted research vehicle was developed and built by MicroCraft Inc., Tullahoma, Tenn., under NASA contract. The booster was built by Orbital Sciences Corp., Dulles, Va.,After being air-launched from NASA's venerable NB-52 mothership, the booster will accelerate the X-43A to test speed and altitude. The X-43A will then separate from the rocket 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.

The X-43A hypersonic research aircraft and its modified Pegasus booster rocket nestled under the wi

NASA Technical Reports Server (NTRS)

2001-01-01

The X-43A hypersonic research aircraft and its modified Pegasus booster rocket are nestled under the wing of NASA's NB-52B carrier aircraft during pre-flight systems testing at the Dryden Flight Research Center, Edwards, Calif. The combined systems test was one of the last major milestones in the Hyper-X research program before the first X-43A flight. 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, or five times the speed of sound). The 12-foot, unpiloted research vehicle was developed and built by MicroCraft Inc., Tullahoma, Tenn., under NASA contract. The booster was built by Orbital Sciences Corp., Dulles, Va. After being air-launched from NASA's venerable NB-52 mothership, the booster will accelerate the X-43A to test speed and altitude. The X-43A will then separate from the rocket 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.

The Road to Mach 10: A History of the X-43A Hypersonic Flight Test Program at NASA Dryden...Toward the Future. Part II

NASA Technical Reports Server (NTRS)

Peebles, Curtis

2007-01-01

In terms of technology, the X-43A/Hyper-X represented a singular milestone. After nearly a half century of high hopes, studies, wind tunnel tests, proposals, and canceled projects, a scramjet-powered vehicle had flown. The performance of the engine qualified the scramjet design tools and scaling laws. In turn, the theoretical calculations and ground testing could be used to design more advanced engine concepts. Just as important, both the scramjet and vehicle systems had successfully operated in the variable temperatures and densities of the atmosphere. The X-43A systems were able to maintain the exact flight conditions necessary for the scramjet to operate properly. Control deflections to correct the engine-induced moments were close to pre-flight predictions. When the unexpected occurred, such as when the vehicle pitched up during the cowl opening on the second flight, the control system was sufficiently designed to correct the situation. The airframe and wing structure, the thermal protection material, and the internal conditions of the X-43A performed largely as predicted. The HXLV thermal anomaly during the ascent on the third flight and "the Mach 8 unpleasantness" during the descent indicated that the HXLV and X-43A were not as resilient to aerodynamic heating as expected. The X-43A 's airframe drag and lift both were slightly higher than predicted, but still within preflight uncertainty predictions. The stability and control were as predicted, as was the boundary layer transition. The biggest aerodynamic worry before the flight was the separation of the HXLV and the X- 43A. After all was said and done, this went exactly as predicted, proving that non-symmetrical/high-dynamic pressure stage separations could be performed. This in turn meant that two-stage-to-orbit vehicles employing this technology were feasible. The Hyper-X program also served as a training ground for a new generation of scramjet and hypersonic researchers. This included both NASA and contractor personnel, providing them with experience in ground testing and component development; vehicle design, construction, integration, system checkout, and, ultimately, flight testing and data analysis. Additionally, researchers learned the practical details of running a project within finite budget and time limits, about the ambiguousness of risk assessment, and about the need to spend a significant amount of time and effort dealing with engineering problems, such as those with the FAS, that have nothing to do with the project's research goals. Finally, all those who worked on the X-43A project now know what it is like to spend years transforming an idea into a functional vehicle, only for it to be lost in a matter of seconds. And then to go through years of work to correct the problems, to face the possibility that still more might exist, and finally to savor the triumph of two successful flights. For those who will work on the hypersonic projects that emerge in coming years, these experiences may prove to be the most valuable of all.

Energy Conversion and Combustion Sciences

DTIC Science & Technology

2013-03-08

Property issues Flameholding (flammability limit) Flame propagation (turbulent-flame speed) combustion-Mixing interaction shock Cavity Based Scramjet ...focusing: • “Very-high” speed (space access) region • Overlapping interests and close coordination with AF programs ( scramjet , rockets etc.). • NSF...and Relevant Conditions Hypersonics Gas Turbines Rockets M > 0.1 Re ? Da ? wrinkled flame ball laminar flame Auto Engines PGC (1

Aerothermo-Structural Analysis of Low Cost Composite Nozzle/Inlet Components

NASA Technical Reports Server (NTRS)

Shivakumar, Kuwigai; Challa, Preeli; Sree, Dave; Reddy, D.

1999-01-01

This research is a cooperative effort among the Turbomachinery and Propulsion Division of NASA Glenn, CCMR of NC A&T State University, and the Tuskegee University. The NC A&T is the lead center and Tuskegee University is the participating institution. Objectives of the research were to develop an integrated aerodynamic, thermal and structural analysis code for design of aircraft engine components, such as, nozzles and inlets made of textile composites; conduct design studies on typical inlets for hypersonic transportation vehicles and setup standards test examples and finally manufacture a scaled down composite inlet. These objectives are accomplished through the following seven tasks: (1) identify the relevant public domain codes for all three types of analysis; (2) evaluate the codes for the accuracy of results and computational efficiency; (3) develop aero-thermal and thermal structural mapping algorithms; (4) integrate all the codes into one single code; (5) write a graphical user interface to improve the user friendliness of the code; (6) conduct test studies for rocket based combined-cycle engine inlet; and finally (7) fabricate a demonstration inlet model using textile preform composites. Tasks one, two and six are being pursued. Selected and evaluated NPARC for flow field analysis, CSTEM for in-depth thermal analysis of inlets and nozzles and FRAC3D for stress analysis. These codes have been independently verified for accuracy and performance. In addition, graphical user interface based on micromechanics analysis for laminated as well as textile composites was developed. Demonstration of this code will be made at the conference. A rocket based combined cycle engine was selected for test studies. Flow field analysis of various inlet geometries were studied. Integration of codes is being continued. The codes developed are being applied to a candidate example of trailblazer engine proposed for space transportation. A successful development of the code will provide a simpler, faster and user-friendly tool for conducting design studies of aircraft and spacecraft engines, applicable in high speed civil transport and space missions.

Results of an investigation of hypersonic viscous interaction effects of the space shuttle orbiter using a 0.010 scale model (51-0) in the AEDC-VKF tunnel F (OA160)

NASA Technical Reports Server (NTRS)

Elder, D. J.

1975-01-01

An experimental aerodynamic investigation was conducted in the AEDC-VKF Hypervelocity Wind Tunnel (Tunnel F) at a nomial Mach number of 19 to determine hypersonic viscous interaction effects on the space shuttle orbiter. The tests were conducted at an angle of attack of 30 degrees over a free-stream Reynolds number (based on fuselage length) variation from 0.1 to 0.4 million. Viscous interaction parameter was varied from 0.02 to 0.06. Six component static stability force and moment data were measured by an internally compensated internal strain gage balance. Resulting data are presented.

Asynchronous, macrotasked relaxation strategies for the solution of viscous, hypersonic flows

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.

1991-01-01

A point-implicit, asynchronous macrotasked relaxation of the steady, thin-layer, Navier-Stokes equations is presented. The method employs multidirectional, single-level storage Gauss-Seidel relaxation sweeps, which effectively communicate perturbations across the entire domain in 2n sweeps, where n is the dimension of the domain. In order to enhance convergence the application of relaxation factors to specific components of the Jacobian is examined using a stability analysis of the advection and diffusion equations. Attention is also given to the complications associated with asynchronous multitasking. Solutions are generated for hypersonic flows over blunt bodies in two and three dimensions with chemical reactions, utilizing single-tasked and multitasked relaxation strategies.

Langley Aerothermodynamic Facilities Complex: Enhancements and Testing Capabilities

NASA Technical Reports Server (NTRS)

Micol, J. R.

1998-01-01

Description, capabilities, recent upgrades, and utilization of the NASA Langley Research Center (LaRC) Aerothermodynamic Facilities Complex (AFC) are presented. The AFC consists of five hypersonic, blow-down-to-vacuum wind tunnels that collectively provide a range of Mach number from 6 to 20, unit Reynolds number from 0.04 to 22 million per foot and, most importantly for blunt configurations, normal shock density ratio from 4 to 12. These wide ranges of hypersonic simulation parameters are due, in part, to the use of three different test gases (air, helium, and tetrafluoromethane), thereby making several of the facilities unique. The Complex represents nearly three-fourths of the conventional (as opposed to impulse)-type hypersonic wind tunnels operational in this country. AFC facilities are used to assess and optimize the hypersonic aerodynamic performance and aeroheating characteristics of aerospace vehicle concepts and to provide benchmark aerodynamic/aeroheating data fr generating the flight aerodynamic databook and final design of the thermal protection system (TPS) (e.g., establishment of flight limitations not to exceed TPS design limits). Modifications and enhancements of AFC hardware components and instrumentation have been pursued to increase capability, reliability, and productivity in support of programmatic goals. Examples illustrating facility utilization in recent years to generate essentially all of the experimental hypersonic aerodynamic and aeroheating information for high-priority, fast-paced Agency programs are presented. These programs include Phase I of the Reusable Launch Vehicle (RLV) Advanced Technology Demonstrator, X-33 program, PHase II of the X-33 program, X-34 program, the Hyper-X program ( a Mach 5,7, and 10 airbreathing propulsion flight experiment), and the X-38 program (Experimental Crew Return Vehicle, X-CRV). Current upgrades/enchancements and future plans for the AFC are discussed.

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

Motion in Interactive Flows," MS Thesis , Aerospace Engineering and Engineering Mechanics Department, The University of Texas at Austin, Dec. 1986. 19...8217Cone drag in the transition from continuum to free molecular flow.’ PhD Thesis , Jesus College, Oxford (1984) 10-9 6i ______3&14 445 105.47 rn,=2.93...of Scramjets," Ph.D. Thesis , University of Sheffield (GB), October 1965. 88. Swithenbank, J., and Parsons, R.J., "Experimental Techniques for

Uncertainty Quantification and Statistical Engineering for Hypersonic Entry Applications

NASA Technical Reports Server (NTRS)

Cozmuta, Ioana

2011-01-01

NASA has invested significant resources in developing and validating a mathematical construct for TPS margin management: a) Tailorable for low/high reliability missions; b) Tailorable for ablative/reusable TPS; c) Uncertainty Quantification and Statistical Engineering are valuable tools not exploited enough; and d) Need to define strategies combining both Theoretical Tools and Experimental Methods. The main reason for this lecture is to give a flavor of where UQ and SE could contribute and hope that the broader community will work with us to improve in these areas.

Evaluation of a large capacity heat pump concept for active cooling of hypersonic aircraft structure

NASA Technical Reports Server (NTRS)

Pagel, L. L.; Herring, R. L.

1978-01-01

Results of engineering analyses assessing the conceptual feasibility of a large capacity heat pump for enhancing active cooling of hypersonic aircraft structure are presented. A unique heat pump arrangement which permits cooling the structure of a Mach 6 transport to aluminum temperatures without the aid of thermal shielding is described. The selected concept is compatible with the use of conventional refrigerants, with Freon R-11 selected as the preferred refrigerant. Condenser temperatures were limited to levels compatible with the use of conventional refrigerants by incorporating a unique multipass condenser design, which extracts mechanical energy from the hydrogen fuel, prior to each subsequent pass through the condenser. Results show that it is technically feasible to use a large capacity heat pump in lieu of external shielding. Additional analyses are required to optimally apply this concept.

Assessment Of The Aerodynamic And Aerothermodynamic Performance Of The USV-3 High-Lift Re-Entry Vehicle

NASA Astrophysics Data System (ADS)

Pezzella, Giuseppe; Richiello, Camillo; Russo, Gennaro

2011-05-01

This paper deals with the aerodynamic and aerothermodynamic trade-off analysis carried out with the aim to design a hypersonic flying test bed (FTB), namely USV3. Such vehicle will have to be launched with a small expendable launcher and shall re-enter the Earth atmosphere allowing to perform several experiments on critical re-entry phenomena. The demonstrator under study is a re-entry space glider characterized by a relatively simple vehicle architecture able to validate hypersonic aerothermodynamic design database and passenger experiments, including thermal shield and hot structures. Then, a summary review of the aerodynamic characteristics of two FTB concepts, compliant with a phase-A design level, has been provided hereinafter. Indeed, several design results, based both on engineering approach and computational fluid dynamics, are reported and discussed in the paper.

Improved Finite-Volume Method for Radiative Hydrodynamics

NASA Technical Reports Server (NTRS)

Wray, Alan

2012-01-01

Fully coupled simulations of hydrodynamics and radiative transfer are essential to a number of fields ranging from astrophysics to engineering applications. Of particular interest in this work are hypersonic atmospheric entries and associated experimental apparatus, e.g., shock tubes and high enthalpy testing facilities. The radiative transfer calculations must supply to the CFD a heating term in the energy equation in the form of the divergence of the radiative heat flux and the radiative heat fluxes to bounding surfaces. It is most efficient to solve the radiative transfer equation on the same grid as the CFD solution, and this work presents an algorithm with improved accuracy for such simulations on structured and unstructured grids compared to more conventional approaches. Results will be shown for shock radiation during hypersonic reentry. Issues of parallelization within a radiation sweep will also be discussed.

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.

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.

The X-43A hypersonic research aircraft and its modified Pegasus® booster rocket mounted to NASA's NB-52B carrier aircraft at the Dryden Flight Research Center, Edwards, California

NASA Image and Video Library

2001-03-13

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.

The effect of exhaust plume/afterbody interaction on installed Scramjet performance

NASA Technical Reports Server (NTRS)

Edwards, Thomas Alan

1988-01-01

Newly emerging aerospace technology points to the feasibility of sustained hypersonic flight. Designing a propulsion system capable of generating the necessary thrust is now the major obstacle. First-generation vehicles will be driven by air-breathing scramjet (supersonic combustion ramjet) engines. Because of engine size limitations, the exhaust gas leaving the nozzle will be highly underexpanded. Consequently, a significant amount of thrust and lift can be extracted by allowing the exhaust gases to expand along the underbody of the vehicle. Predicting how these forces influence overall vehicle thrust, lift, and moment is essential to a successful design. This work represents an important first step toward that objective. The UWIN code, an upwind, implicit Navier-Stokes computer program, has been applied to hypersonic exhaust plume/afterbody flow fields. The capability to solve entire vehicle geometries at hypersonic speeds, including an interacting exhaust plume, has been demonstrated for the first time. Comparison of the numerical results with available experimental data shows good agreement in all cases investigated. For moderately underexpanded jets, afterbody forces were found to vary linearly with the nozzle exit pressure, and increasing the exit pressure produced additional nose-down pitching moment. Coupling a species continuity equation to the UWIN code enabled calculations indicating that exhaust gases with low isentropic exponents (gamma) contribute larger afterbody forces than high-gamma exhaust gases. Moderately underexpanded jets, which remain attached to unswept afterbodies, underwent streamwise separation on upswept afterbodies. Highly underexpanded jets produced altogether different flow patterns, however. The highly underexpanded jet creates a strong plume shock, and the interaction of this shock with the afterbody was found to produce complicated patterns of crossflow separation. Finally, the effect of thrust vectoring on vehicle balance has been shown to alter dramatically the vehicle pitching moment.

Ceramic and coating applications in the hostile environment of a high temperature hypersonic wind tunnel. [Langley 8-foot high temperature structures tunnel

NASA Technical Reports Server (NTRS)

Puster, R. L.; Karns, J. R.; Vasquez, P.; Kelliher, W. C.

1981-01-01

A Mach 7, blowdown wind tunnel was used to investigate aerothermal structural phenomena on large to full scale high speed vehicle components. The high energy test medium, which provided a true temperature simulation of hypersonic flow at 24 to 40 km altitude, was generated by the combustion of methane with air at high pressures. Since the wind tunnel, as well as the models, must be protected from thermally induced damage, ceramics and coatings were used extensively. Coatings were used both to protect various wind tunnel components and to improve the quality of the test stream. Planned modifications for the wind tunnel included more extensive use of ceramics in order to minimize the number of active cooling systems and thus minimize the inherent operational unreliability and cost that accompanies such systems. Use of nonintrusive data acquisition techniques, such as infrared radiometry, allowed more widespread use of ceramics for models to be tested in high energy wind tunnels.

Fluid dynamic modeling and numerical simulation of low-density hypersonic flow

NASA Astrophysics Data System (ADS)

Cheng, H. K.; Wong, Eric Y.

1988-06-01

The concept of a viscous shock-layer and several related versions of continuum theories/methods are examined for their adequacy as a viable framework to study flow physics and aerothermodynamics of relevance to sustained hypersonic flights. Considering the flat plate at angle of attack, or the wedge, as a generic example for the major aerodynamic component of a hypersonic vehicle, the relative importance of the molecular-transport effects behind the shock (in the form of the 'shock slip') and the wall-slip effects are studied. In the flow regime where the shock-transition-zone thickness remains small compared to the shock radius of curvature, a quasi-one-dimensional shock structure under the Burnett/thirteen-moment approximation, as well as particulate/collisional models, can be consistently developed. The fully viscous version of the shock-layer model is shown to provide the crucial boundary condition downstream the shock in this case. The gas-kinetic basis of the continuum description for the flow behind the bow shock, and certain features affecting the non-equilibrium flow chemistry, are also discussed.

Wind-Tunnel Balance Characterization for Hypersonic Research Applications

NASA Technical Reports Server (NTRS)

Lynn, Keith C.; Commo, Sean A.; Parker, Peter A.

2012-01-01

Wind-tunnel research was recently conducted at the NASA Langley Research Center s 31-Inch Mach 10 Hypersonic Facility in support of the Mars Science Laboratory s aerodynamic program. Researchers were interested in understanding the interaction between the freestream flow and the reaction control system onboard the entry vehicle. A five-component balance, designed for hypersonic testing with pressurized flow-through capability, was used. In addition to the aerodynamic forces, the balance was exposed to both thermal gradients and varying internal cavity pressures. Historically, the effect of these environmental conditions on the response of the balance have not been fully characterized due to the limitations in the calibration facilities. Through statistical design of experiments, thermal and pressure effects were strategically and efficiently integrated into the calibration of the balance. As a result of this new approach, researchers were able to use the balance continuously throughout the wide range of temperatures and pressures and obtain real-time results. Although this work focused on a specific application, the methodology shown can be applied more generally to any force measurement system calibration.

Method for obtaining aerodynamic data on hypersonic configurations with scramjet exhaust flow simulation

NASA Technical Reports Server (NTRS)

Hartill, W. R.

1977-01-01

A hypersonic wind tunnel test method for obtaining credible aerodynamic data on a complete hypersonic vehicle (generic X-24c) with scramjet exhaust flow simulation is described. The general problems of simulating the scramjet exhaust as well as accounting for scramjet inlet flow and vehicle forces are analyzed, and candidate test methods are described and compared. The method selected as most useful makes use of a thrust-minus-drag flow-through balance with a completely metric model. Inlet flow is diverted by a fairing. The incremental effect of the fairing is determined in the testing of two reference models. The net thrust of the scramjet module is an input to be determined in large-scale module tests with scramjet combustion. Force accounting is described, and examples of force component levels are predicted. Compatibility of the test method with candidate wind tunnel facilities is described, and a preliminary model mechanical arrangement drawing is presented. The balance design and performance requirements are described in a detailed specification. Calibration procedures, model instrumentation, and a test plan for the model are outlined.

Fluid dynamic modeling and numerical simulation of low-density hypersonic flow

NASA Technical Reports Server (NTRS)

Cheng, H. K.; Wong, Eric Y.

1988-01-01

The concept of a viscous shock-layer and several related versions of continuum theories/methods are examined for their adequacy as a viable framework to study flow physics and aerothermodynamics of relevance to sustained hypersonic flights. Considering the flat plate at angle of attack, or the wedge, as a generic example for the major aerodynamic component of a hypersonic vehicle, the relative importance of the molecular-transport effects behind the shock (in the form of the 'shock slip') and the wall-slip effects are studied. In the flow regime where the shock-transition-zone thickness remains small compared to the shock radius of curvature, a quasi-one-dimensional shock structure under the Burnett/thirteen-moment approximation, as well as particulate/collisional models, can be consistently developed. The fully viscous version of the shock-layer model is shown to provide the crucial boundary condition downstream the shock in this case. The gas-kinetic basis of the continuum description for the flow behind the bow shock, and certain features affecting the non-equilibrium flow chemistry, are also discussed.

Anisotropic power spectrum of refractive-index fluctuation in hypersonic turbulence.

PubMed

Li, Jiangting; Yang, Shaofei; Guo, Lixin; Cheng, Mingjian

2016-11-10

An anisotropic power spectrum of the refractive-index fluctuation in hypersonic turbulence was obtained by processing the experimental image of the hypersonic plasma sheath and transforming the generalized anisotropic von Kármán spectrum. The power spectrum suggested here can provide as good a fit to measured spectrum data for hypersonic turbulence as that recorded from the nano-planar laser scattering image. Based on the newfound anisotropic hypersonic turbulence power spectrum, Rytov approximation was employed to establish the wave structure function and the spatial coherence radius model of electromagnetic beam propagation in hypersonic turbulence. Enhancing the anisotropy characteristics of the hypersonic turbulence led to a significant improvement in the propagation performance of electromagnetic beams in hypersonic plasma sheath. The influence of hypersonic turbulence on electromagnetic beams increases with the increase of variance of the refractive-index fluctuation and the decrease of turbulence outer scale and anisotropy parameters. The spatial coherence radius was much smaller than that in atmospheric turbulence. These results are fundamental to understanding electromagnetic wave propagation in hypersonic turbulence.

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

22 Figure 23. Flow Type and the reference empirical model ............................................................ 24 Figure 24. Baseline...Trajectory ...................................................................................................... 25 Figure 25. Flow Features Important...94 viii GLOSSARY ACCTE Advanced Ceramic Composites for Turbine Engines AFRL Air Force Research Laboratory AoA Angle of Attack ASE

System studies on space plane powered by scram/LACE propulsion system

NASA Astrophysics Data System (ADS)

Maita, Masataka; Miyajima, Hiroshi; Mori, Takashige

1992-12-01

Japan's NAL has undertaken concept-development studies for hypersonic technologies-integrating SSTO spaceplane configurations. Attention is presently given to the scramjet/liquefied air cycle engine (LACE). While the scramjet powers the vehicle begining at Mach 5, the LACE is used above Mach 12 on the basis of excess hydrogen fuel consumption; the Mach 20 orbital speed is thereby gained.

Hypersonic Technology Developments with EU Co-Funded Projects

DTIC Science & Technology

2010-09-01

metal or even high performance alloys . The hollow sphere technology allows high degrees of porosities, reproducible properties and fair process control...sandwich structure configuration will be investigated. Titanium alloys and Ti-aluminides exhibit excellent mechanical properties for applications where...cooling techniques, new alloys , improved thermodynamic cycles by increased pressure ratios and TIT, etc… As the Olympus 593 engine was based on the

A hypersonic aeroheating calculation method based on inviscid outer edge of boundary layer parameters

NASA Astrophysics Data System (ADS)

Meng, ZhuXuan; Fan, Hu; Peng, Ke; Zhang, WeiHua; Yang, HuiXin

2016-12-01

This article presents a rapid and accurate aeroheating calculation method for hypersonic vehicles. The main innovation is combining accurate of numerical method with efficient of engineering method, which makes aeroheating simulation more precise and faster. Based on the Prandtl boundary layer theory, the entire flow field is divided into inviscid and viscid flow at the outer edge of the boundary layer. The parameters at the outer edge of the boundary layer are numerically calculated from assuming inviscid flow. The thermodynamic parameters of constant-volume specific heat, constant-pressure specific heat and the specific heat ratio are calculated, the streamlines on the vehicle surface are derived and the heat flux is then obtained. The results of the double cone show that at the 0° and 10° angle of attack, the method of aeroheating calculation based on inviscid outer edge of boundary layer parameters reproduces the experimental data better than the engineering method. Also the proposed simulation results of the flight vehicle reproduce the viscid numerical results well. Hence, this method provides a promising way to overcome the high cost of numerical calculation and improves the precision.

STARS: An Integrated, Multidisciplinary, Finite-Element, Structural, Fluids, Aeroelastic, and Aeroservoelastic Analysis Computer Program

NASA Technical Reports Server (NTRS)

Gupta, K. K.

1997-01-01

A multidisciplinary, finite element-based, highly graphics-oriented, linear and nonlinear analysis capability that includes such disciplines as structures, heat transfer, linear aerodynamics, computational fluid dynamics, and controls engineering has been achieved by integrating several new modules in the original STARS (STructural Analysis RoutineS) computer program. Each individual analysis module is general-purpose in nature and is effectively integrated to yield aeroelastic and aeroservoelastic solutions of complex engineering problems. Examples of advanced NASA Dryden Flight Research Center projects analyzed by the code in recent years include the X-29A, F-18 High Alpha Research Vehicle/Thrust Vectoring Control System, B-52/Pegasus Generic Hypersonics, National AeroSpace Plane (NASP), SR-71/Hypersonic Launch Vehicle, and High Speed Civil Transport (HSCT) projects. Extensive graphics capabilities exist for convenient model development and postprocessing of analysis results. The program is written in modular form in standard FORTRAN language to run on a variety of computers, such as the IBM RISC/6000, SGI, DEC, Cray, and personal computer; associated graphics codes use OpenGL and IBM/graPHIGS language for color depiction. This program is available from COSMIC, the NASA agency for distribution of computer programs.

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.

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.

Langley Aerospace Research Summer Scholars (LARSS) Scholars Pres

NASA Image and Video Library

2013-08-07

250 students participated in the Langley Aerospace Research Summer Scholars (LARSS) Presentations focused on 3D modeling of STARBUKS calibration components in the National Transonic Facility, hypersonic aerodynamic inflatable decelerator, and optimization of a microphone-based array for flight testing. Reid Center LaRC Hampton, VA

Restoration of the Hypersonic Tunnel Facility at NASA Glenn Research Center, Plum Brook Station

NASA Technical Reports Server (NTRS)

Woodling, Mark A.

2000-01-01

The NASA Glenn Research Center's Hypersonic Tunnel Facility (HTF), located at the Plum Brook Station in Sandusky, Ohio, is a non-vitiated, free-jet facility, capable of testing large-scale propulsion systems at Mach Numbers from 5 to 7. As a result of a component failure in September of 1996, a restoration project was initiated in mid- 1997 to repair the damage to the facility. Following the 2-1/2 year effort, the HTF has been returned to an operational condition. Significant repairs and operational improvements have been implemented in order to ensure facility reliability and personnel safety. As of January 2000, this unique, state-of-the-art facility was ready for integrated systems testing.

Ultra High Temperature (UHT) SiC Fiber (Phase 2)

NASA Technical Reports Server (NTRS)

Dicarlo, James A.; Jacobson, Nathan S.; Lizcano, Maricela; Bhatt, Ramakrishna T.

2015-01-01

Silicon-carbide fiber-reinforced silicon-carbide ceramic matrix composites (SiCSiC CMC) are emerginglightweight re-usable structural materials not only for hot section components in gas turbine engines, but also for controlsurfaces and leading edges of reusable hypersonic vehicles as well as for nuclear propulsion and reactor components. Ithas been shown that when these CMC are employed in engine hot-section components, the higher the upper usetemperature (UUT) of the SiC fiber, the more performance benefits are accrued, such as higher operating temperatures,reduced component cooling air, reduced fuel consumption, and reduced emissions. The first generation of SiCSiC CMC with a temperature capability of 2200-2400F are on the verge of being introduced into the hot-section components ofcommercial and military gas turbine engines.Today the SiC fiber type currently recognized as the worlds best in terms ofthermo-mechanical performance is the Sylramic-iBN fiber. This fiber was previously developed by the PI at NASA GRC using patented processes to improve the high-cost commercial Sylramic fiber, which in turn was derived from anotherlow-cost low-performance commercial fiber. Although the Sylramic-iBN fiber shows state-of-the art creep and rupture resistance for use temperatures above 2550oF, NASA has shown by fundamental creep studies and model developmentthat its microstructure and creep resistance could theoretically be significantly improved to produce an Ultra HighTemperature (UHT) SiC fiber.This Phase II Seedling Fund effort has been focused on the key objective of effectively repeating the similar processes used for producing the Sylramic-iBN fiber using a design of experiments approach to first understand the cause of the less than optimum Sylramic-iBN microstructure and then attempting to develop processconditions that eliminate or minimize these key microstructural issues. In so doing, it is predicted that that theseadvanced process could result in an UHT SiC fiber with 20 times more creep resistance than the Sylramic-iBN fiber,which in turn would allow SiCSiC CMC to operate up to 2700oF and above, thereby further enhancing the performancebenefits of SiCSiC components in aero-propulsion engines. It was also envisioned that the fiber processes developedduring Phase II efforts would not only reduce production costs for the UHT fiber by using low-cost precursor fibers andcombined processes, but also allow the UHT fibers to be directly produced in preforms of the precursor fibers, possibly atthe facilities of the CMC fabricator.

The 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

NASA Image and Video Library

2001-03-15

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, Calif. The combined systems test was one of the last major milestones in the Hyper-X research program before the first X-43A flight. 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, or five times the speed of sound). The 12-foot, unpiloted research vehicle was developed and built by MicroCraft Inc., Tullahoma, Tenn., under NASA contract. The booster was built by Orbital Sciences Corp., Dulles, Va.,After being air-launched from NASA's venerable NB-52 mothership, the booster will accelerate the X-43A to test speed and altitude. The X-43A will then separate from the rocket 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.

The X-43A hypersonic research aircraft and its modified Pegasus® booster rocket nestled under the wing of NASA's NB-52B carrier aircraft during pre-flight systems testing

NASA Image and Video Library

2001-03-15

The X-43A hypersonic research aircraft and its modified Pegasus® booster rocket are nestled under the wing of NASA's NB-52B carrier aircraft during pre-flight systems testing at the Dryden Flight Research Center, Edwards, Calif. The combined systems test was one of the last major milestones in the Hyper-X research program before the first X-43A flight. 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, or five times the speed of sound). The 12-foot, unpiloted research vehicle was developed and built by MicroCraft Inc., Tullahoma, Tenn., under NASA contract. The booster was built by Orbital Sciences Corp., Dulles, Va. After being air-launched from NASA's venerable NB-52 mothership, the booster will accelerate the X-43A to test speed and altitude. The X-43A will then separate from the rocket 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.

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.

High-Temperature Modal Survey of a Hot-Structure Control Surface

NASA Technical Reports Server (NTRS)

Spivey, Natalie Dawn

2010-01-01

Ground vibration tests or modal surveys are routinely conducted for supporting flutter analysis for subsonic and supersonic vehicles; however, for hypersonic vehicle applications, thermoelastic vibration testing techniques are not well established and are not routinely performed for supporting hypersonic flutter analysis. New high-temperature material systems, fabrication technologies and high-temperature sensors expand the opportunities to develop advanced techniques for performing ground vibration tests at elevated temperatures. High-temperature materials have the unique property of increasing in stiffness when heated. When these materials are incorporated into a hot-structure, which includes metallic components that decrease in stiffness with increasing temperature, the interaction between the two materials systems needs to be understood because that interaction could ultimately affect the hypersonic flutter analysis. Performing a high-temperature modal survey will expand the research database for hypersonics and will help build upon the understanding of the dual material interaction. This paper will discuss the vibration testing of the Carbon-Silicon Carbide Ruddervator Subcomponent Test Article which is a truncated version of the full-scale X-37 hot-structure control surface. In order to define the modal characteristics of the test article during the elevated-temperature modal survey, two series of room-temperature modal test configurations had to be performed. The room-temperature test series included one with the test article suspended from a bungee cord (free-free) and the second with it mounted on the strongback (fixed boundary condition) in NASA Dryden's Flight Loads Lab large nitrogen test chamber.

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

DTIC Science & Technology

2013-01-01

coupled with combustor – Combustor, component for subsonic or supersonic combustion – Nozzle , expands flow for high thrust and may provide lift... supersonic solution method that is used for both the inlet and nozzle components. The supersonic model SAMURI is a substantial improvement over previous models...purely supersonic inviscid flow. As a result, the model is also appropriate for other applications, including the nozzle , which is important 19 Figure

2-D and 3-D mixing flow analyses of a scramjet-afterbody configuration

NASA Technical Reports Server (NTRS)

Baysal, Oktay; Eleshaky, Mohamed E.; Engelund, Walter C.

1989-01-01

A cold simulant gas study of propulsion/airframe integration for a hypersonic vehicle powered by a scramjet engine is presented. The specific heat ratio of the hot exhaust gases are matched by utilizing a cold mixture of argon and Freon-12. Solutions are obtained for a hypersonic corner flow and a supersonic rectangular flow in order to provide the upstream boundary conditions. The computational test examples also provide a comparison of this flow with that of air as the expanding supersonic jet, where the specific heats are assumed to be constant. It is shown that the three-dimensional computational fluid capabilities developed for these types of flow may be utilized to augment the conventional wind tunnel studies of scramjet afterbody flows using cold simulant exhaust gases, which in turn can help in the design of a scramjet internal-external nozzle.

System technology analysis of aeroassisted orbital transfer vehicles - Moderate lift/drag

NASA Technical Reports Server (NTRS)

Florence, D. E.; Fischer, G.

1983-01-01

The utilization of procedures involving aerodynamic braking and/or aerodynamic maneuvering on return from higher altitude orbits to low-earth orbit makes it possible to realize significant performance benefits. The present study is concerned with a number of mission scenarios for Aeroassisted Orbital Transfer Vehicles (AOTV) and the impact of potential technology advances in the performance enhancement of the class of AOTV's having a hypersonic lift to drag ratio (L/D) of 0.75 to 1.5. It is found that the synergistic combination of a hypersonic L/D of 1.2, an advanced cryopropelled engine, and an LH2 drop tank (1-1/2 stage) leads to a single 65,000 pound shuttle, two-man geosynchronous mission with 2100 pounds of useful paylod. Additional payload enhancement is possible with AOTV dry weight reductions due to technology advances in the areas of vehicle structures and thermal protection systems and other subsystems.

Development of a High Accuracy Angular Measurement System for Langley Research Center Hypersonic Wind Tunnel Facilities

NASA Technical Reports Server (NTRS)

Newman, Brett; Yu, Si-bok; Rhew, Ray D. (Technical Monitor)

2003-01-01

Modern experimental and test activities demand innovative and adaptable procedures to maximize data content and quality while working within severely constrained budgetary and facility resource environments. This report describes development of a high accuracy angular measurement capability for NASA Langley Research Center hypersonic wind tunnel facilities to overcome these deficiencies. Specifically, utilization of micro-electro-mechanical sensors including accelerometers and gyros, coupled with software driven data acquisition hardware, integrated within a prototype measurement system, is considered. Development methodology addresses basic design requirements formulated from wind tunnel facility constraints and current operating procedures, as well as engineering and scientific test objectives. Description of the analytical framework governing relationships between time dependent multi-axis acceleration and angular rate sensor data and the desired three dimensional Eulerian angular state of the test model is given. Calibration procedures for identifying and estimating critical parameters in the sensor hardware is also addressed.

Diagnostics for Hypersonic Engine Control

DTIC Science & Technology

2013-02-01

weredirected across the flow at the entrance to the isolator just downstream of the facility nozzle . The near-infrared beams were frequency tuned across...the facility nozzle were used to study the dynamics of the shock train structure during these transient combustor events. They revealed the...entropy fluctuations in supersonic boundary layers can be quite short in time – on the order of tens of microseconds. We therefore sought data

Hearings Before the Committee on Aeronautical and Space Sciences. United States Senate, Ninety-Third Congress, Second Session

DTIC Science & Technology

1974-07-18

concepts which, when de - veloped, will provide for the combination of new structural design concepts with the new developments in airframe fabri- cation...19-20 9. Cargo aircraft design for special applications ------ 20 10. Advanced fighter concepts------------------------ 21 11-12. Hypersonic...Research, De - velopment and Engineering. U.S. Army Aviation Systems Command -------------------------------------------- 172 Appendix: Statement on behalf

X-15: Extending the Frontiers of Flight

NASA Technical Reports Server (NTRS)

Jenkins, Dennis R.

2007-01-01

A history of the design and achievements of the high-speed, 1950s-era X-15 airplane is presented. The following chapters are included: A New Science; A Hypersonic Research Airplane; Conflict and Innovation; The Million-Horsepower Engine; High Range and Dry Lakes; Preparations; The Flight Program; and the Research Program. Selected biographies, flight logs and physical characteristics of the X-15 Airplane are included in the appendices.

Hypersonic research engine project. Phase 2: Aerothermodynamic integration model development

NASA Technical Reports Server (NTRS)

Jilly, L. F. (Editor)

1970-01-01

The analytical effort was directed towards (1) completing the design of the combustor exit instrumentation assembly, (2) analyzing the coolant flow distribution of the cowl leading edge tip section, (3) determining effects of purge gas pressure on AIM performance analysis, and (4) analyzing heat transfer and associated stress problems related to the cowl leading edge tip section and the nozzle shroud assembly for test conditions.

2017 Air Force Global Strike Command Innovation and Technology Symposium

DTIC Science & Technology

2017-11-15

and technological changes that may have occurred within American society and therefore may have different expectations, motivations and values...Participate in a discussion on hypersonic technology as a national imperative and what are government, industry and academia doing to accelerate this...USAF Deputy Chief, Weapons Requirements Division HAF/A5RW Mr. Robert B. Addis Defense Technologies Engineering Division, Lawrence Livermore National

International Hypersonic Waverider Symposium, 1st, University of Maryland, College Park, MD, Oct. 17-19, 1990, Proceedings

NASA Technical Reports Server (NTRS)

Anderson, John D., Jr. (Editor); Lewis, Mark J. (Editor); Corda, Stephen (Editor); Blankson, Isaiah M. (Editor)

1990-01-01

The papers presented in this volume provide an overview of current theoretical and experimental research in the field of hypersonic waveriders. In particular, attention is given to efficient waveriders from known axisymmetric flow fields, hypersonic waverider design from given shock waves, limitations of waveriders, and aerodynamic stability theory of hypersonic waveriders. The discussion also covers momentum analysis of waverider flow fields, tethered aerothermodynamic research for hypersonic waveriders, simulation of hypersonic waveriders, and an idealized tip-to-tail waverider model.

NASP - Waveriders in a hypersonic sky.

NASA Astrophysics Data System (ADS)

Baker, David

1993-01-01

A development history is presented for the hydrogen-fueled, airbreathing (scramjet) engine-propelled National Aerospace Plane (NASP), which will be able to cruise endoatmospherically at hypersopnic speeds or rise exoatmospherically, by converting to rocket power, to LEO. Attention is given to the technology-development and configuration-validation services that the X-30 project will render the far larger NASP vehicle; the configurational and propulsion system factors in question encompass the use of 'slush' hydrogen fuel, the integration of engine inlets into the aircraft forebody and exhaust nozzles into the afterbody, and the conversion from turbojet or rocket propulsion to scramjet mode and back.

An Investigation of Acoustic Wave Propagation in Mach 2 Flow

NASA Astrophysics Data System (ADS)

Nieberding, Zachary J.

Hypersonic technology is the next advancement to enter the aerospace community; it is defined as the study of flight at speeds Mach 5 and higher where intense aerodynamic heating is prevalent. Hypersonic flight is achieved through use of scramjet engines, which intake air and compress it by means of shock waves and geometry design. The airflow is then directed through an isolator where it is further compressed, it is then delivered to the combustor at supersonic speeds. The combusted airflow and fuel mixture is then accelerated through a nozzle to achieve the hypersonic speeds. Unfortunately, scramjet engines can experience a phenomenon known as an inlet unstart, where the combustor produces pressures large enough to force the incoming airflow out of the inlet of the engine, resulting in a loss of acceleration and power. There have been several government-funded programs that look to prove the concept of the scramjet engine and also tackle this inlet unstart issue. The research conducted in this thesis is a fundamental approach towards controlling the unstart problem: it looks at the basic concept of sending a signal upstream through the boundary layer of a supersonic flow and being able to detect a characterizeable signal. Since conditions within and near the combustor are very harsh, hardware is unable to be installed in that area, so this testing will determine if a signal can be sent and if so, how far upstream can the signal be detected. This experimental approach utilizes several acoustic and mass injection sources to be evaluated over three test series in a Mach 2 continuous flow wind tunnel that will determine the success of the objective. The test series vary in that the conditions of the flow and the test objectives change. The research shows that a characterizeable signal can be transmitted upstream roughly 12 inches through the subsonic boundary layer of a supersonic cross flow. It is also shown that the signal attenuates as the distance between the source and sensors increases. Individual studies including detection sensor and source comparison, material selection, transfer rates, and shadowgraph imagery are also investigated. The acoustic signal is affected by the boundary layer, which is impacted by the shock train and its location. With the capability to characterize an acoustic signal within a scramjet engine to detect the shock train location, any disturbance in the acoustic signals can be linked to shock train displacement that could lead to an inlet unstart. With these results in mind, it is possible that acoustic hardware can be designed to be implemented into the scramjet engine to detect an inlet unstart before it should happen.

Diffusion coefficient of hydrogen in a cast gamma titanium aluminide

DOE Office of Scientific and Technical Information (OSTI.GOV)

Sundaram, P.A.; Wessel, E.; Ennis, P.J.

1999-06-04

Gamma titanium aluminides have the potential for high temperature applications because of their high specific strength and specific modulus. Their oxidation resistance is good, especially at intermediate temperatures and with suitable alloying additions, good oxidation resistance can be obtained up to 800 C. One critical area of application is in combustion engines in aero-space vehicles such as hypersonic airplanes and high speed civil transport airplanes. This entails the use of hydrogen as a fuel component and hence the effect of hydrogen on the mechanical properties of gamma titanium aluminides is of significant scientific and technological utility. The purpose of thismore » short investigation is to use an electrochemical method under galvanostatic conditions to determine the diffusion coefficient of hydrogen in a cast gamma titanium aluminide, a typical technical alloy with potential application in gas turbines under creep conditions. This result will be then compared with that obtained by microhardness profiling of electrolytically hydrogen precharged material.« less

The HYTHIRM Project: Flight Thermography of the Space Shuttle During the Hypersonic Re-entry

NASA Technical Reports Server (NTRS)

Horvath, Thomas J.; Tomek, Deborah M.; Berger, Karen T.; Zalameda, Joseph N.; Splinter, Scott C.; Krasa, Paul W.; Schwartz, Richard J.; Gibson, David M.; Tietjen, Alan B.; Tack, Steve

2010-01-01

This report describes a NASA Langley led endeavor sponsored by the NASA Engineering Safety Center, the Space Shuttle Program Office and the NASA Aeronautics Research Mission Directorate to demonstrate a quantitative thermal imaging capability. A background and an overview of several multidisciplinary efforts that culminated in the acquisition of high resolution calibrated infrared imagery of the Space Shuttle during hypervelocity atmospheric entry is presented. The successful collection of thermal data has demonstrated the feasibility of obtaining remote high-resolution infrared imagery during hypersonic flight for the accurate measurement of surface temperature. To maximize science and engineering return, the acquisition of quantitative thermal imagery and capability demonstration was targeted towards three recent Shuttle flights - two of which involved flight experiments flown on Discovery. In coordination with these two Shuttle flight experiments, a US Navy NP-3D aircraft was flown between 26-41 nautical miles below Discovery and remotely monitored surface temperature of the Orbiter at Mach 8.4 (STS-119) and Mach 14.7 (STS-128) using a long-range infrared optical package referred to as Cast Glance. This same Navy aircraft successfully monitored the Orbiter Atlantis traveling at approximately Mach 14.3 during its return from the successful Hubble repair mission (STS-125). The purpose of this paper is to describe the systematic approach used by the Hypersonic Thermodynamic Infrared Measurements team to develop and implement a set of mission planning tools designed to establish confidence in the ability of an imaging platform to reliably acquire, track and return global quantitative surface temperatures of the Shuttle during entry. The mission planning tools included a pre-flight capability to predict the infrared signature of the Shuttle. Such tools permitted optimization of the hardware configuration to increase signal-to-noise and to maximize the available dynamic range while mitigating the potential for saturation. Post flight, analysis tools were used to assess atmospheric effects and to convert the 2-D intensity images to 3-D temperature maps of the windward surface. Comparison of the spatially resolved global thermal measurements to surface thermocouples and CFD prediction is made. Successful demonstration of a quantitative, spatially resolved, global temperature measurement on the Shuttle suggests future applications towards hypersonic flight test programs within NASA, DoD and DARPA along with flight test opportunities supporting NASA's project Constellation.

STB-White

NASA Technical Reports Server (NTRS)

Molnar, Dan; Ammon, Rob; Gallagher, Todd; Gouhin, Pat; Hermann, Steve; Roos, John Bryan; Saurer, Craig; White, Heather

1990-01-01

The final design of a hypersonic, SCRAMjet research aircraft, which is to be dropped from a carrier plane, is considered. Topics such as propulsion systems, aerodynamics, component weight analysis, and aircraft design with waverider analyses are stressed with smaller emphasis placed on aircraft systems such as cockpit design and landing gear configurations. Propulsion systems include analysis of the turbofanramjet for acceleration to low hypersonic speed (Mach 6.0) and analysis of the SCRAMjets themselves to carry the aircraft to Mach 10.0. Both analyses include the use of liquid hydrogen as fuel. Inlet design for both propulsion systems is analyzed as well. Aerodynamic properties are found using empirical and theoretical formulas for lift and drag on delta-wing aircraft. The aircraft design involves the integration of all preliminary studies into a modified waverider configuration.

Flow-Tagging Velocimetry for Hypersonic Flows Using Fluorescence of Nitric Oxide

NASA Technical Reports Server (NTRS)

Danehy, Paul M.; OByrne, Sean; Houwing, A. Frank P.; Fox, Jodie S.; Smith, Daniel R.

2003-01-01

We demonstrate a new variation of molecular-tagging velocimetry for hypersonic flows based on laser-induced fluorescence. A thin line of nitric-oxide molecules is excited with a laser beam and then, after a time delay, a fluorescence image of the displaced line is acquired. One component of velocity is determined from the time of flight. This method is applied to measure the velocity profile in a Mach 8.5 laminar, hypersonic boundary layer in the Australian National University s T2 free-piston shock tunnel. The single-shot velocity measurement uncertainty in the freestream was found to be 3.5%, based on 90% confidence. The method is also demonstrated in the separated flow region forward of a blunt fin attached to a flat plate in a Mach 7.4 flow produced by the Australian National University s T3 free-piston shock tunnel. The measurement uncertainty in the blunt fin experiment is approximately 30%, owing mainly to low fluorescence intensities, which could be improved significantly in future experiments. This velocimetry method is applicable to very high-speed flows that have low collisional quenching of the fluorescing species. It is particularly convenient in facilities where planar laser-induced fluorescence is already being performed.

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.

Hybrid Solution-Adaptive Unstructured Cartesian Method for Large-Eddy Simulation of Detonation in Multi-Phase Turbulent Reactive Mixtures

DTIC Science & Technology

2012-03-27

pulse- detonation engines ( PDE ), stage separation, supersonic cav- ity oscillations, hypersonic aerodynamics, detonation induced structural...ADAPTIVE UNSTRUCTURED CARTESIAN METHOD FOR LARGE-EDDY SIMULATION OF DETONATION IN MULTI-PHASE TURBULENT REACTIVE MIXTURES 5b. GRANT NUMBER FA9550...CCL Report TR-2012-03-03 Hybrid Solution-Adaptive Unstructured Cartesian Method for Large-Eddy Simulation of Detonation in Multi-Phase Turbulent

A high-frequency servosystem for fuel control in hypersonic engines

NASA Technical Reports Server (NTRS)

Simon, Donald L.

1991-01-01

A hydrogen fuel-flow valve with an electrohydraulic servosystem is described. An analysis of the servosystem is presented along with a discussion of the limitations imposed on system performance by nonlinearities. The response of the valve to swept-frequency inputs is experimentally determined and compared with analytical results obtained from a computer model. The valve is found to perform favorably for frequencies up to 200 Hz.

Advanced Ceramic Materials for Future Aerospace Applications

NASA Technical Reports Server (NTRS)

Misra, Ajay

2015-01-01

With growing trend toward higher temperature capabilities, lightweight, and multifunctionality, significant advances in ceramic matrix composites (CMCs) will be required for future aerospace applications. The presentation will provide an overview of material requirements for future aerospace missions, and the role of ceramics and CMCs in meeting those requirements. Aerospace applications will include gas turbine engines, aircraft structure, hypersonic and access to space vehicles, space power and propulsion, and space communication.

Overview of CMC Research at NASA Glenn Research Center

NASA Technical Reports Server (NTRS)

Grady, Joseph E.

2011-01-01

CMC technology development in the Ceramics Branch at NASA Glenn Research Center addresses Aeronautics propulsion goals across subsonic, supersonic and hypersonic flight regimes. Combustor, turbine and exhaust nozzle applications of CMC materials will enable NASA to demonstrate reduced fuel consumption, emissions, and noise in advanced gas turbine engines. Applications ranging from basic Fundamental Aeronautics research activities to technology demonstrations in the new Integrated Systems Research Program will be discussed.

Numerical Solutions for a Cylindrical Laser Diffuser Flowfield

DTIC Science & Technology

1990-06-01

exhaust conditions with minimum losses to optimize performance of the system. Thus, the handling of the system of shock waves to decelerate the flow...requirement for exhaustive experimental work will result in significant savings of both time and resources. As more advanced computers are developed, the...Mach number (ɚ.5) flows. Recent interest in hypersonic engine inlet performance has resulted in an extension of the methodology to high Mach number

NASA X-34 Technology in Motion

NASA Technical Reports Server (NTRS)

Beech, Geoffrey; Chandler, Kristie

1997-01-01

The X-34 technology development program is a joint industry/government project to develop, test, and operate a small, fully-reusable hypersonic flight vehicle. The objective is to demonstrate key technologies and operating concepts applicable to future reusable launch vehicles. Integrated in the vehicle are various systems to assure successful completion of mission objectives, including the Main Propulsion System (MPS). NASA-Marshall Space Flight Center (MSFC) is responsible for developing the X-34's MPS including the design and complete build package for the propulsion system components. The X-34 will be powered by the Fastrac Engine, which is currently in design and development at NASA-MSFC. Fastrac is a single-stage main engine, which burns a mixture of liquid oxygen (LOX) and kerosene(RP-1). The interface between the MPS and Fastrac engine are critical for proper system operation and technologies applicable to future reusable launch vehicles. Deneb's IGRIP software package with the Dynamic analysis option provided a key tool for conducting studies critical to this interface as well as a mechanism to drive the design of the LOX and RP-1 feedlines. Kinematic models were created for the Fastrac Engine and the feedlines for various design concepts. Based on the kinematic simulation within Envision, design and joint limits were verified and system interference controlled. It was also critical to the program to evaluate the effect of dynamic loads visually, providing a verification tool for dynamic analysis and in some cases uncovering areas that had not been considered. Deneb's software put the X-34 technology in motion and has been a key factor in facilitating the strenuous design schedule.

Anomaly detection of microstructural defects in continuous fiber reinforced composites

NASA Astrophysics Data System (ADS)

Bricker, Stephen; Simmons, J. P.; Przybyla, Craig; Hardie, Russell

2015-03-01

Ceramic matrix composites (CMC) with continuous fiber reinforcements have the potential to enable the next generation of high speed hypersonic vehicles and/or significant improvements in gas turbine engine performance due to their exhibited toughness when subjected to high mechanical loads at extreme temperatures (2200F+). Reinforced fiber composites (RFC) provide increased fracture toughness, crack growth resistance, and strength, though little is known about how stochastic variation and imperfections in the material effect material properties. In this work, tools are developed for quantifying anomalies within the microstructure at several scales. The detection and characterization of anomalous microstructure is a critical step in linking production techniques to properties, as well as in accurate material simulation and property prediction for the integrated computation materials engineering (ICME) of RFC based components. It is desired to find statistical outliers for any number of material characteristics such as fibers, fiber coatings, and pores. Here, fiber orientation, or `velocity', and `velocity' gradient are developed and examined for anomalous behavior. Categorizing anomalous behavior in the CMC is approached by multivariate Gaussian mixture modeling. A Gaussian mixture is employed to estimate the probability density function (PDF) of the features in question, and anomalies are classified by their likelihood of belonging to the statistical normal behavior for that feature.

X-ray based extensometry

NASA Technical Reports Server (NTRS)

Jordan, E. H.; Pease, D. M.

1988-01-01

A totally new method of extensometry using an X-ray beam was proposed. The intent of the method is to provide a non-contacting technique that is immune to problems associated with density variations in gaseous environments that plague optical methods. X-rays are virtually unrefractable even by solids. The new method utilizes X-ray induced X-ray fluorescence or X-ray induced optical fluorescence of targets that have melting temperatures of over 3000 F. Many different variations of the basic approaches are possible. In the year completed, preliminary experiments were completed which strongly suggest that the method is feasible. The X-ray induced optical fluorescence method appears to be limited to temperatures below roughly 1600 F because of the overwhelming thermal optical radiation. The X-ray induced X-ray fluorescence scheme appears feasible up to very high temperatures. In this system there will be an unknown tradeoff between frequency response, cost, and accuracy. The exact tradeoff can only be estimated. It appears that for thermomechanical tests with cycle times on the order of minutes a very reasonable system may be feasible. The intended applications involve very high temperatures in both materials testing and monitoring component testing. Gas turbine engines, rocket engines, and hypersonic vehicles (NASP) all involve measurement needs that could partially be met by the proposed technology.

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.

Life assessment of structural components using inelastic finite element analyses

NASA Technical Reports Server (NTRS)

Arya, Vinod K.; Halford, Gary R.

1993-01-01

The need for enhanced and improved performance of structural components subject to severe cyclic thermal/mechanical loadings, such as in the aerospace industry, requires development of appropriate solution technologies involving time-dependent inelastic analyses. Such analyses are mandatory to predict local stress-strain response and to assess more accurately the cyclic life time of structural components. The NASA-Lewis Research Center is cognizant of this need. As a result of concerted efforts at Lewis during the last few years, several such finite element solution technologies (in conjunction with the finite element program MARC) were developed and successfully applied to numerous uniaxial and multiaxial problems. These solution technologies, although developed for use with MARC program, are general in nature and can easily be extended for adaptation with other finite element programs such as ABAQUS, ANSYS, etc. The description and results obtained from two such inelastic finite element solution technologies are presented. The first employs a classical (non-unified) creep-plasticity model. An application of this technology is presented for a hypersonic inlet cowl-lip problem. The second of these technologies uses a unified creep-plasticity model put forth by Freed. The structural component for which this finite element solution technology is illustrated, is a cylindrical rocket engine thrust chamber. The advantages of employing a viscoplastic model for nonlinear time-dependent structural analyses are demonstrated. The life analyses for cowl-lip and cylindrical thrust chambers are presented. These analyses are conducted by using the stress-strain response of these components obtained from the corresponding finite element analyses.

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.

NASA's NB-52B carrier aircraft rolls down a taxiway with the X-43A hypersonic research aircraft and its modified Pegasus® booster rocket attached to a pylon under its right wing.

NASA Image and Video Library

2001-03-15

As part of a combined systems test conducted by NASA Dryden Flight Research Center, NASA's NB-52B carrier aircraft rolls down a taxiway at Edwards Air Force Base with the X-43A hypersonic research aircraft and its modified Pegasus® booster rocket attached to a pylon under its right wing. The taxi test was one of the last major milestones in the Hyper-X research program before the first X-43A flight. 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, or five times the speed of sound). The 12-foot, unpiloted research vehicle was developed and built by MicroCraft Inc., Tullahoma, Tenn., under NASA contract. The booster was built by Orbital Sciences Corp., Dulles, Va. After being air-launched from NASA's venerable NB-52 mothership, the booster will accelerate the X-43A to test speed and altitude. The X-43A will then separate from the rocket 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.

The 1 × 1 m hypersonic wind tunnel Kochel/Tullahoma 1940-1960

NASA Astrophysics Data System (ADS)

Eckardt, Dietrich

2015-03-01

Peenemünde and Cape Canaveral mark cornerstones of space history. Kochel in Southern Germany and Tullahoma in Tennessee, USA also belong in this category. The technically unique Kochel wind tunnel was part of the German long-distance missile development strategy, planned and prepared in secret before the beginning of World War II. A 57 MW closed-circuit wind tunnel facility with 1 × 1 m measuring section was planned for continuous-flow simulation at high Mach numbers Ma 7-10. In the early 1940 s a site beside the Walchensee Power Station at Kochel am See in Upper Bavaria, Germany was chosen to provide the required altitude difference of 200 m for the hydraulic turbine drives. The preparatory activities for the erection of this impressive hypersonic wind tunnel facility were pushed ahead until an enforced temporary pause in September 1944. In early May 1945 US troops occupied the area and, in due course, scientists of General Arnold's Scientific Advisory Group, the `von Kármán team', ordered the transfer to the USA of available equipment, design materials and other paperwork. Here, at the Arnold Engineering Development Center (AEDC) Tullahoma, TN this `Tunnel A' was built to begin operation around 1957. The testing was conducted on the Mach 7 experimental aircraft X-15, space shuttle developments and still secret investigations on unmanned hypersonic vehicles.

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.

The Effect of Magnetohydrodynamic (MHD) Energy Bypass on Specific Thrust for a Supersonic Turbojet Engine

NASA Technical Reports Server (NTRS)

Benyo, Theresa L.

2010-01-01

This paper describes the preliminary results of a thermodynamic cycle analysis of a supersonic turbojet engine with a magnetohydrodynamic (MHD) energy bypass system that explores a wide range of MHD enthalpy extraction parameters. Through the analysis described here, it is shown that applying a magnetic field to a flow path in the Mach 2.0 to 3.5 range can increase the specific thrust of the turbojet engine up to as much as 420 N/(kg/s) provided that the magnitude of the magnetic field is in the range of 1 to 5 Tesla. The MHD energy bypass can also increase the operating Mach number range for a supersonic turbojet engine into the hypersonic flight regime. In this case, the Mach number range is shown to be extended to Mach 7.0.

Modeling of Inhomogeneous Compressible Turbulence Using a Two-Scale Statistical Theory

NASA Technical Reports Server (NTRS)

Hamba, Fujihiro

1996-01-01

Turbulence modeling plays an important role in the study of high-speed flows in engineering and aerodynamic problems; they include flows in supersonic combustion engines and over hypersonic transport aircraft. The enhancement of the kinetic energy dissipation by the dilatational terms is one of the typical compressibility effects. Zeman (1990) and Sarkar et al. (1991) proposed that the dilatation dissipation is proportional to the solenoidal dissipation and is a function of the turbulent Mach number. Sarkar (1992) also modeled the pressure-dilatation correlation using the turbulent Mach number. Zeman (1991) related the correlation to the rate of change of the pressure variance.

Evaluation of an innovative high temperature ceramic wafer seal for hypersonic engine applications. Ph.D. Thesis, 1991

NASA Technical Reports Server (NTRS)

Steinetz, Bruce M.

1992-01-01

A critical mechanical system in advanced hypersonic engines is the panel-edge seal system that seals gaps between the articulating engine panels and the adjacent engine splitter walls. Significant advancements in seal technology are required to meet the extreme demands placed on the seals, including the simultaneous requirements of low leakage, conformable, high temperature, high pressure, sliding operation. In this investigation, the design, development, analytical and experimental evaluation of a new ceramic wafer seal that shows promise of meeting these demands will be addressed. A high temperature seal test fixture was designed and fabricated to measure static seal leakage performance under engine simulated conditions. Ceramic wafer seal leakage rates are presented for engine-simulated air pressure differentials (up to 100 psi), and temperature (up to 1350 F), sealing both flat and distorted wall conditions, where distortions can be as large as 0.15 inches in only an 18 inch span. Seal leakage rates are low, meeting an industry-established tentative leakage limit for all combinations of temperature, pressure and wall conditions considered. A seal leakage model developed from externally-pressurized gas film bearing theory is also presented. Predicted leakage rates agree favorably with the measured data for nearly all conditions of temperature and pressure. Discrepancies noted at high engine pressure and temperature are attributed to thermally-induced, non-uniform changes in the size and shape of the leakage gap condition. The challenging thermal environment the seal must operate in places considerable demands on the seal concept and material selection. Of the many high temperature materials considered in the design, ceramics were the only materials that met the many challenging seal material design requirements. Of the aluminum oxide, silicon carbide, and silicon nitride ceramics considered in the material ranking scheme developed herein, the silicon nitride class of ceramics ranked the highest because of their high temperature strength; resistance to the intense heating rates; resistance to hydrogen damage; and good structural properties. Baseline seal feasibility has been established through the research conducted in this investigation. Recommendations for future work are also discussed.

Computational mechanics and physics at NASA Langley Research Center

NASA Technical Reports Server (NTRS)

South, Jerry C., Jr.

1987-01-01

An overview is given of computational mechanics and physics at NASA Langley Research Center. Computational analysis is a major component and tool in many of Langley's diverse research disciplines, as well as in the interdisciplinary research. Examples are given for algorithm development and advanced applications in aerodynamics, transition to turbulence and turbulence simulation, hypersonics, structures, and interdisciplinary optimization.

Hypersonic ramjet experiment project. Phase 1: Computer program description, ramjet and scramjet cycle performance

NASA Technical Reports Server (NTRS)

Jackson, R. J.; Wang, T. T.

1974-01-01

A computer program was developed to describe the performance of ramjet and scramjet cycles. The program performs one dimensional calculations of the equilibrium, real-gas internal flow properties of the engine. The program can be used for the following: (1) preliminary design calculation and (2) design analysis of internal flow properties corresponding to stipulated flow areas. Only the combustion of hydrogen in air is considered in this case.

Reduced-Order Aerothermoelastic Analysis of Hypersonic Vehicle Structures

DTIC Science & Technology

2012-01-01

concentrates on performance of the scramjet powered X-43 vehicle and flight data relative to pretest predictions and some limited posttest analysis...for his dedication to my development both as a student and as a person. He has given me the freedom to pursue my specific research interests and has...to grow as a researcher and engineer. I am especially grateful for the opportunities he has provided for me to attend technical conferences which have

Hydrogen-fueled scramjets: Potential for detailed combustor analysis

NASA Technical Reports Server (NTRS)

Beach, H. L., Jr.

1976-01-01

Combustion research related to hypersonic scramjet (supersonic combustion ramjet) propulsion is discussed from the analytical point of view. Because the fuel is gaseous hydrogen, mixing is single phase and the chemical kinetics are well known; therefore, the potential for analysis is good relative to hydro-carbon fueled engines. Recent progress in applying two and three dimensional analytical techniques to mixing and reacting flows indicates cause for optimism, and identifies several areas for continuing effort.

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.

Center of Excellence for Hypersonics Research

DTIC Science & Technology

2012-01-25

detailed simulations of actual combustor configurations, and ultimately for the optimization of hypersonic air - breathing propulsion system flow paths... vehicle development programs. The Center engaged leading experts in experimental and computational analysis of hypersonic flows to provide research...advanced hypersonic vehicles and space access systems will require significant advances in the design methods and ground testing techniques to ensure

CFD for hypersonic airbreathing aircraft

NASA Technical Reports Server (NTRS)

Kumar, Ajay

1989-01-01

A general discussion is given on the use of advanced computational fluid dynamics (CFD) in analyzing the hypersonic flow field around an airbreathing aircraft. Unique features of the hypersonic flow physics are presented and an assessment is given of the current algorithms in terms of their capability to model hypersonic flows. Several examples of advanced CFD applications are then presented.

The second X-43A hypersonic research aircraft, shown here in its protective shipping jig, arrives at NASA's Dryden Flight Research Center

NASA Image and Video Library

2001-01-31

The second of three X-43A hypersonic research aircraft, shown here in its protective shipping jig, arrived at NASA's Dryden Flight Research Center, Edwards, California, on January 31, 2001. The arrival of the second X-43A from its manufacturer, MicroCraft, Inc., of Tullahoma, Tenn., followed by only a few days the mating of the first X-43A and its specially-designed adapter to the first stage of a modified Pegasus® booster rocket. The booster, built by Orbital Sciences Corp., Dulles, Va., will accelerate the 12-foot-long, unpiloted research aircraft to a predetermined altitude and speed after the X-43A/booster "stack" is air-launched from NASA's venerable NB-52 mothership. The X-43A will then separate from the rocket 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, 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.

Results of the NASP Ames Integrated Mixing Hypersonic Engine (AIMHYE) Scramjet Test Program

NASA Technical Reports Server (NTRS)

Cavolowsky, John A.; Loomis, Mark P.; Deiwert, George S.

1995-01-01

This paper describes the test techniques and results from the National Aerospace Plane Government Work Package 53, the Ames Integrated Mixing Hypersonic Engine (AIMHYE) Scramjet Test program conducted in the NASA Ames 16-Inch Combustion Driven Shock Tunnel. This was a series of near full-scale scramjet combustor tests with the objective to obtain high speed combustor and nozzle data from an engine with injector configurations similar to the NASP E21 and E22a designs. The experimental test approach was to use a large combustor model (80-100% throat height) designed and fabricated for testing in the semi-free jet mode. The conditions tested were similar to the "blue book" conditions at Mach 12, 14, and 16. GWP 53 validated use of large, long test time impulse facilities, specifically the Ames 16-Inch Shock Tunnel, for high Mach number scramjet propulsion testing an integrated test rig (inlet, combustor, and nozzle). Discussion of key features of the test program will include: effects of the 2-D combustor inlet pressure profile; performance of large injectors' fueling system that included nozzlettes, base injection, and film cooling; and heat transfer measurements to the combustor. Significant instrumentation development and application efforts include the following: combustor force balance application for measurement of combustor drag for comparison with integrated point measurements of skin friction; nozzle metric strip for measuring thrust with comparison to integrated pressure measurements; and nonintrusive optical fiber-based diode laser absorption measurements of combustion products for determination of combustor performance. Direct measurements will be reported for specific test article configurations and compared with CFD solutions.

Flight Control Laws for NASA's Hyper-X Research Vehicle

NASA Technical Reports Server (NTRS)

Davidson, J.; Lallman, F.; McMinn, J. D.; Martin, J.; Pahle, J.; Stephenson, M.; Selmon, J.; Bose, D.

1999-01-01

The goal of the Hyper-X program is to demonstrate and validate technology for design and performance predictions of hypersonic aircraft with an airframe-integrated supersonic-combustion ramjet propulsion system. Accomplishing this goal requires flight demonstration of a hydrogen-fueled scramjet powered hypersonic aircraft. A key enabling technology for this flight demonstration is flight controls. Closed-loop flight control is required to enable a successful stage separation, to achieve and maintain the design condition during the engine test, and to provide a controlled descent. Before the contract award, NASA developed preliminary flight control laws for the Hyper-X to evaluate the feasibility of the proposed scramjet test sequence and descent trajectory. After the contract award, a Boeing/NASA partnership worked to develop the current control laws. This paper presents a description of the Hyper-X Research Vehicle control law architectures with performance and robustness analyses. Assessments of simulated flight trajectories and stability margin analyses demonstrate that these control laws meet the flight test requirements.

Generic Hypersonic Inlet Module Analysis

NASA Technical Reports Server (NTRS)

Cockrell, Chares E., Jr.; Huebner, Lawrence D.

2004-01-01

A computational study associated with an internal inlet drag analysis was performed for a generic hypersonic inlet module. The purpose of this study was to determine the feasibility of computing the internal drag force for a generic scramjet engine module using computational methods. The computational study consisted of obtaining two-dimensional (2D) and three-dimensional (3D) computational fluid dynamics (CFD) solutions using the Euler and parabolized Navier-Stokes (PNS) equations. The solution accuracy was assessed by comparisons with experimental pitot pressure data. The CFD analysis indicates that the 3D PNS solutions show the best agreement with experimental pitot pressure data. The internal inlet drag analysis consisted of obtaining drag force predictions based on experimental data and 3D CFD solutions. A comparative assessment of each of the drag prediction methods is made and the sensitivity of CFD drag values to computational procedures is documented. The analysis indicates that the CFD drag predictions are highly sensitive to the computational procedure used.

Design applications for supercomputers

NASA Technical Reports Server (NTRS)

Studerus, C. J.

1987-01-01

The complexity of codes for solutions of real aerodynamic problems has progressed from simple two-dimensional models to three-dimensional inviscid and viscous models. As the algorithms used in the codes increased in accuracy, speed and robustness, the codes were steadily incorporated into standard design processes. The highly sophisticated codes, which provide solutions to the truly complex flows, require computers with large memory and high computational speed. The advent of high-speed supercomputers, such that the solutions of these complex flows become more practical, permits the introduction of the codes into the design system at an earlier stage. The results of several codes which either were already introduced into the design process or are rapidly in the process of becoming so, are presented. The codes fall into the area of turbomachinery aerodynamics and hypersonic propulsion. In the former category, results are presented for three-dimensional inviscid and viscous flows through nozzle and unducted fan bladerows. In the latter category, results are presented for two-dimensional inviscid and viscous flows for hypersonic vehicle forebodies and engine inlets.

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.

Hypersonic Flight Test Windows for Technology Development Testing

DTIC Science & Technology

2013-11-01

used. 2.1 Propulsion and Controls Test Window The technologies dealing with scramjet propulsion (inlets, fuel injection, etc.) and hypersonic ...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

High-Temperature Modal Survey of a Hot-Structure Control Surface

NASA Technical Reports Server (NTRS)

Spivey, Natalie D.

2011-01-01

Ground vibration tests are routinely conducted for supporting flutter analysis for subsonic and supersonic vehicles; however, for hypersonic vehicles, thermoelastic vibration testing techniques are neither well established nor routinely performed. New high-temperature material systems, fabrication technologies and high-temperature sensors expand the opportunities to develop advanced techniques for performing ground vibration tests at elevated temperatures. When high-temperature materials, which increase in stiffness when heated, are incorporated into a hot-structure that contains metallic components that decrease in stiffness when heated, the interaction between those materials can affect the hypersonic flutter analysis. A high-temperature modal survey will expand the research database for hypersonics and improve the understanding of this dual-material interaction. This report discusses the vibration testing of the carbon-silicon carbide Ruddervator Subcomponent Test Article, which is a truncated version of a full-scale hot-structure control surface. Two series of room-temperature modal test configurations were performed in order to define the modal characteristics of the test article during the elevated-temperature modal survey: one with the test article suspended from a bungee cord (free-free) and the second with it mounted on the strongback (fixed boundary). Testing was performed in the NASA Dryden Flight Research Center Flight Loads Laboratory Large Nitrogen Test Chamber.

Proceedings of the Seventh Annual Summer Conference. NASA/USRA: University Advanced Design Program

NASA Technical Reports Server (NTRS)

1991-01-01

The Advanced Design Program (ADP) is a unique program that brings together students and faculty from U.S. engineering schools with engineers from the NASA centers through integration of current and future NASA space and aeronautics projects into university engineering design curriculum. The Advanced Space Design Program study topics cover a broad range of projects that could be undertaken during a 20-30 year period beginning with the deployment of the Space Station Freedom. The Advanced Aeronautics Design Program study topics typically focus on nearer-term projects of interest to NASA, covering from small, slow-speed vehicles through large, supersonic passenger transports and on through hypersonic research vehicles. Student work accomplished during the 1990-91 academic year and reported at the 7th Annual Summer Conference is presented.

Heat transfer in aerospace propulsion

NASA Technical Reports Server (NTRS)

Simoneau, Robert J.; Hendricks, Robert C.; Gladden, Herbert J.

1988-01-01

Presented is an overview of heat transfer related research in support of aerospace propulsion, particularly as seen from the perspective of the NASA Lewis Research Center. Aerospace propulsion is defined to cover the full spectrum from conventional aircraft power plants through the Aerospace Plane to space propulsion. The conventional subsonic/supersonic aircraft arena, whether commercial or military, relies on the turbine engine. A key characteristic of turbine engines is that they involve fundamentally unsteady flows which must be properly treated. Space propulsion is characterized by very demanding performance requirements which frequently push systems to their limits and demand tailored designs. The hypersonic flight propulsion systems are subject to severe heat loads and the engine and airframe are truly one entity. The impact of the special demands of each of these aerospace propulsion systems on heat transfer is explored.

Global strike hypersonic weapons

NASA Astrophysics Data System (ADS)

Lewis, Mark J.

2017-11-01

Beginning in the 1940's, the United States has pursued the development of hypersonic technologies, enabling atmospheric flight in excess of five times the speed of sound. Hypersonic flight has application to a range of military and civilian applications, including commercial transport, space access, and various weapons and sensing platforms. A number of flight tests of hypersonic vehicles have been conducted by countries around the world, including the United States, Russia, and China, that could lead the way to future hypersonic global strike weapon systems. These weapons would be especially effective at penetrating conventional defenses, and could pose a significant risk to national security.

Shock Tunnel Studies of the Hypersonic Flowfield around the Hypervelocity Ballistic Models with Aerospikes

NASA Astrophysics Data System (ADS)

Balakalyani, G.; Saravanan, S.; Jagadeesh, G.

Reduced drag and aerodynamic heating are the two basic design requirements for any hypersonic vehicle [1]. The flowfield around an axisymmetric blunt body is characterized by a bow shockwave standing ahead of its nose. The pressure and temperature behind this shock wave are very high. This increased pressure and temperature are responsible for the high levels of drag and aerodynamic heating over the body. In the past, there have been many investigations on the use of aerospikes as a drag reduction tool. These studies on spiked bodies aim at reducing both the drag and aerodynamic heating by modifying the hypersonic flowfield ahead of the nose of the body [2]. However, most of them used very simple configurations to experimentally study the drag reduction using spikes at hypersonic speeds [3] and therefore very little experimental data is available for a realistic geometric configuration. In the present study, the standard AGARD Hypervelocity Ballistic model 1 is used as the test model. The addition of the spike to the blunt body significantly alters the flowfield ahead of the nose, leading to the formation of a low pressure conical recirculation region, thus causing a reduction in drag and wall heat flux [4]. In the present investigation, aerodynamic drag force is measured over the Hypervelocity Ballistic model-1, with and without spike, at a flow enthalpy of 1.7 MJ/kg. The experiments are carried out at a Mach number of 8 and at zero angle of attack. An internally mountable accelerometer based 3-component force balance system is used to measure the aerodynamic forces on the model. Also computational studies are carried out to complement the experiments.

Thermal coatings for titanium-aluminum alloys

NASA Technical Reports Server (NTRS)

Cunnington, George R.; Clark, Ronald K.; Robinson, John C.

1993-01-01

Titanium aluminides and titanium alloys are candidate materials for use in hot structure and heat-shield components of hypersonic vehicles because of their good strength-to-weight characteristics at elevated temperature. However, in order to utilize their maximum temperature capability, they must be coated to resist oxidation and to have a high total remittance. Also, surface catalysis for recombination of dissociated species in the aerodynamic boundary layer must be minimized. Very thin chemical vapor deposition (CVD) coatings are attractive candidates for this application because of durability and very light weight. To demonstrate this concept, coatings of boron-silicon and aluminum-boron-silicon compositions were applied to the titanium-aluminides alpha2 (Ti-14Al-21Nb), super-alpha2 (Ti-14Al-23-Nb-2V), and gamma (Ti-33Al-6Nb-1Ta) and to the titanium alloy beta-21S (Ti-15Mo-3Al-3Nb-0.2Si). Coated specimens of each alloy were subjected to a set of simulated hypersonic vehicle environmental tests to determine their properties of oxidation resistance, surface catalysis, radiative emittance, and thermal shock resistance. Surface catalysis results should be viewed as relative performance only of the several coating-alloy combinations tested under the specific environmental conditions of the LaRC Hypersonic Materials Environmental Test System (HYMETS) arc-plasma-heated hypersonic wind tunnel. Tests were also conducted to evaluate the hydrogen transport properties of the coatings and any effects of the coating processing itself on fatigue life of the base alloys. Results are presented for three types of coatings, which are as follows: (1) a single layer boron silicon coating, (2) a single layer aluminum-boron-silicon coating, and (3) a multilayer coating consisting of an aluminum-boron-silicon sublayer with a boron-silicon outer layer.

Hypersonic Flight Vehicle and Ramjet Engine. Considerations of Coordination and Configuration,

DTIC Science & Technology

1980-04-01

oxidizing agent tank of the rocket and the initial weight is reduced. Analysis shows that the weight reduced by liquifying oxygen from the air not opnly...6fvaVThL- tE ~ dtid J ~v~Vf / Prick t D~ri, y? 1~ O 9 21 22 Hight lv’b ,~,i~ ~ 1cvV/L R~e~ viol ~s. lht Plug Pf-05ram,, Ae,+rO"O(ih- 1961. ’- *, s

Approximate method for calculating convective heat flux on the surface of bodies of simple geometric shapes

NASA Astrophysics Data System (ADS)

Kuzenov, V. V.; Ryzhkov, S. V.

2017-02-01

The paper formulated engineering and physical mathematical model for aerothermodynamics hypersonic flight vehicle (HFV) in laminar and turbulent boundary layers (model designed for an approximate estimate of the convective heat flow in the range of speeds M = 6-28, and height H = 20-80 km). 2D versions of calculations of convective heat flows for bodies of simple geometric forms (individual elements of the design HFV) are presented.

Propulsion/airframe integration issues for waverider aircraft

NASA Technical Reports Server (NTRS)

Blankson, Isaiah M.; Hagseth, Paul

1993-01-01

While many propulsion concepts and technologies developed for nonwaverider-type hypersonic vehicles may apply to waveriders, some aspects of these configurations require unique technological approaches. An evaluation is made of such distinctive opportunities in the cases of engine cycle selection, inlets, nozzle designs and integration, longitudinal stability, and thermal management. Also discussed are waverider requirements for control surface effectiveness, inlet boundary layer ingestion effects, and structural/configurational optimization, giving attention to trades in volumetric/structural efficiency and vehicle L/D.

OVRhyp, Scramjet Test Aircraft

NASA Technical Reports Server (NTRS)

Aslan, J.; Bisard, T.; Dallinga, S.; Draper, K.; Hufford, G.; Peters, W.; Rogers, J.

1990-01-01

A preliminary design for an unmanned hypersonic research vehicle to test scramjet engines is presented. The aircraft will be launched from a carrier aircraft at an altitude of 40,000 feet at Mach 0.8. The vehicle will then accelerate to Mach 6 at an altitude of 100,000 feet. At this stage the prototype scramjet will be employed to accelerate the vehicle to Mach 10 and maintain Mach 10 flight for 2 minutes. The aircraft will then decelerate and safely land.

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.

Conceptual Design of a 50--100 MW Electron Beam Accelerator System for the National Hypersonic Wind Tunnel Program

DOE Office of Scientific and Technical Information (OSTI.GOV)

SCHNEIDER,LARRY X.

2000-06-01

The National Hypersonic Wind Tunnel program requires an unprecedented electron beam source capable of 1--2 MeV at a beam power level of 50--100 MW. Direct-current electron accelerator technology can readily generate high average power beams to approximately 5 MeV at output efficiencies greater than 90%. However, due to the nature of research and industrial applications, there has never been a requirement for a single module with an output power exceeding approximately 500 kW. Although a 50--100 MW module is a two-order extrapolation from demonstrated power levels, the scaling of accelerator components appears reasonable. This paper presents an evaluation of componentmore » and system issues involved in the design of a 50--100 MW electron beam accelerator system with precision beam transport into a high pressure flowing air environment.« less

Modeling of aerodynamic heat flux and thermoelastic behavior of nose caps of hypersonic vehicles

NASA Astrophysics Data System (ADS)

Persova, Marina G.; Soloveichik, Yury G.; Belov, Vasiliy K.; Kiselev, Dmitry S.; Vagin, Denis V.; Domnikov, Petr A.; Patrushev, Ilya I.; Kurskiy, Denis N.

2017-07-01

In this paper, the problem of numerical modeling of thermoelastic behavior of nose caps of hypersonic vehicles at different angles of attack is considered. 3D finite element modeling is performed by solving the coupled heat and elastic problems taking into account thermal and mechanical properties variations with temperature. A special method for calculating the aerodynamic heat flux entering the nose cap from its surface is proposed. This method is characterized by very low computational costs and allows calculating the aerodynamic heat flux at different values of the Mach number and angles of attack which may vary during the aerodynamic heating. The numerical results obtained by the proposed approach are compared with the numerical results and experimental data obtained by other authors. The developed approach has been used for studying the impact of the angle of attack on the thermoelastic behavior of nose caps main components.

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.

Description and preliminary calibration results for the Langley hypersonic CF4 tunnel

NASA Technical Reports Server (NTRS)

Midden, R. E.; Miller, C. G.

1978-01-01

A detailed description of the hypersonic CF4 tunnel is presented along with discussion of the basic components, instrumentation, and operating procedure. Operational experience with the CF4 reclaimer and lead-bath heater is discussed. Comparison of measured and predicted shock detachment distance on a sphere and pressure distributions measured on a sharp leading-edge flat plate revealed the absence of significant flow nonuniformity and lent creditability to predicted free stream flow conditions. The economic operation of this facility centers about the CF4 reclaimer, which was designed to operate at an efficiency of 90 to 95 percent. A number of modifications were made to the reclaimer system to improve its performance, and presently, the system reclaims approximately 75 percent of the test gas. Even with current budgetary constraints, this efficiency permits the CF4 tunnel to be operated as a viable research wind tunnel.

The Additional Error of Inertial Sensors Induced by Hypersonic Flight Conditions

PubMed Central

Karachun, Volodimir; Mel’nick, Viktorij; Korobiichuk, Igor; Nowicki, Michał; Szewczyk, Roman; Kobzar, Svitlana

2016-01-01

The emergence of hypersonic technology pose a new challenge for inertial navigation sensors, widely used in aerospace industry. The main problems are: extremely high temperatures, vibration of the fuselage, penetrating acoustic radiation and shock N-waves. The nature of the additional errors of the gyroscopic inertial sensor with hydrostatic suspension components under operating conditions generated by forced precession of the movable part of the suspension due to diffraction phenomena in acoustic fields is explained. The cause of the disturbing moments in the form of the Coriolis inertia forces during the transition of the suspension surface into the category of impedance is revealed. The boundaries of occurrence of the features on the resonance wave match are described. The values of the “false” angular velocity as a result of the elastic-stress state of suspension in the acoustic fields are determined. PMID:26927122

NASA's hypersonic flight research program

NASA Technical Reports Server (NTRS)

Blankson, Isaiah; Pyle, Jon

1993-01-01

The NASA hypersonic flight research program is reviewed focusing on program history, philosophy, and rationale. Flight research in the high Mach numbers, high dynamic pressure flight regime is considered to be essential to the development of future operational hypersonic systems. The piggy-back experiments which are to be carried out on the Pegasus will develop instrumentation packages for hypersonic data acquisition and will provide unique data of high value to designers and researchers.

NASA's NB-52B carrier aircraft rolls down a taxiway with the X-43A hypersonic research aircraft and its modified Pegasus® booster rocket slung from a pylon under its right wing

NASA Image and Video Library

2001-03-15

NASA's NB-52B carrier aircraft rolls down a taxiway at Edwards Air Force Base with the X-43A hypersonic research aircraft and its modified Pegasus® booster rocket slung from a pylon under its right wing. Part of a combined systems test conducted by NASA's Dryden Flight Research Center at Edwards, the taxi test was one of the last major milestones in the Hyper-X research program before the first X-43A flight. 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, or five times the speed of sound). The 12-foot, unpiloted research vehicle was developed and built by MicroCraft Inc., Tullahoma, Tenn., under NASA contract. The booster was built by Orbital Sciences Corp., Dulles, Va.,After being air-launched from NASA's venerable NB-52 mothership, the booster will accelerate the X-43A to test speed and altitude. The X-43A will then separate from the rocket 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, with the first tentatively scheduled for late spring to early summer, 2001.

Design Study of Wafer Seals for Future Hypersonic Vehicles

NASA Technical Reports Server (NTRS)

Dunlap, Patrick H.; Finkbeiner, Joshua R.; Steinetz, Bruce M.; DeMange, Jeffrey J.

2005-01-01

Future hypersonic vehicles require high temperature, dynamic seals in advanced hypersonic engines and on the vehicle airframe to seal the perimeters of movable panels, flaps, and doors. Current seals do not meet the demanding requirements of these applications, so NASA Glenn Research Center is developing improved designs to overcome these shortfalls. An advanced ceramic wafer seal design has shown promise in meeting these needs. Results from a design of experiments study performed on this seal revealed that several installation variables played a role in determining the amount of leakage past the seals. Lower leakage rates were achieved by using a tighter groove width around the seals, a higher seal preload, a tighter wafer height tolerance, and a looser groove length. During flow testing, a seal activating pressure acting behind the wafers combined with simulated vibrations to seat the seals more effectively against the sealing surface and produce lower leakage rates. A seal geometry study revealed comparable leakage for full-scale wafers with 0.125 and 0.25 in. thicknesses. For applications in which lower part counts are desired, fewer 0.25-in.-thick wafers may be able to be used in place of 0.125-in.-thick wafers while achieving similar performance. Tests performed on wafers with a rounded edge (0.5 in. radius) in contact with the sealing surface resulted in flow rates twice as high as those for wafers with a flat edge. Half-size wafers had leakage rates approximately three times higher than those for full-size wafers.

Integrated Design and Engineering Analysis (IDEA) Environment - Propulsion Related Module Development and Vehicle Integration

NASA Technical Reports Server (NTRS)

Kamhawi, Hilmi N.

2013-01-01

This report documents the work performed during the period from May 2011 - October 2012 on the Integrated Design and Engineering Analysis (IDEA) environment. IDEA is a collaborative environment based on an object-oriented, multidisciplinary, distributed framework using the Adaptive Modeling Language (AML). This report will focus on describing the work done in the areas of: (1) Integrating propulsion data (turbines, rockets, and scramjets) in the system, and using the data to perform trajectory analysis; (2) Developing a parametric packaging strategy for a hypersonic air breathing vehicles allowing for tank resizing when multiple fuels and/or oxidizer are part of the configuration; and (3) Vehicle scaling and closure strategies.

International Symposium on Numerical Methods in Engineering, 5th, Ecole Polytechnique Federale de Lausanne, Switzerland, Sept. 11-15, 1989, Proceedings. Volumes 1 & 2

NASA Astrophysics Data System (ADS)

Gruber, Ralph; Periaux, Jaques; Shaw, Richard Paul

Recent advances in computational mechanics are discussed in reviews and reports. Topics addressed include spectral superpositions on finite elements for shear banding problems, strain-based finite plasticity, numerical simulation of hypersonic viscous continuum flow, constitutive laws in solid mechanics, dynamics problems, fracture mechanics and damage tolerance, composite plates and shells, contact and friction, metal forming and solidification, coupling problems, and adaptive FEMs. Consideration is given to chemical flows, convection problems, free boundaries and artificial boundary conditions, domain-decomposition and multigrid methods, combustion and thermal analysis, wave propagation, mixed and hybrid FEMs, integral-equation methods, optimization, software engineering, and vector and parallel computing.

JPRS report: Science and technology. Central Eurasia: Engineering and equipment

NASA Astrophysics Data System (ADS)

1993-10-01

Translated articles cover the following topics: transient gas dynamic processes in ramjet engines; aerodynamic characteristics of delta wings with detached leading edge shock wave at hypersonic flight velocities; effect of atmospheric density gradient on aerodynamic stabilization; measurement of target radar scattering characteristics using frequency synthesized signals; assessing survivability and ensuring safety of large axial-flow compressor blades; procedure for experimentally determining transient aerodynamic forces caused by flat vane cascade; analysis of aerodynamic interaction of profile and vortex; laser machine for balancing dynamically adjusted gyros; use of heat pumps in solar heat supply systems; numerical simulation of deflagration transition to detonation in homogeneous combustible fuel mixture; and investigation of chemically nonequilibrium flow about bodies allowing for vibrational relaxation.

800 C Silicon Carbide (SiC) Pressure Sensors for Engine Ground Testing

NASA Technical Reports Server (NTRS)

Okojie, Robert S.

2016-01-01

MEMS-based 4H-SiC piezoresistive pressure sensors have been demonstrated at 800 C, leading to the discovery of strain sensitivity recovery with increasing temperatures above 400 C, eventually achieving up to, or near, 100 recovery of the room temperature values at 800 C. This result will allow the insertion of highly sensitive pressure sensors closer to jet, rocket, and hypersonic engine combustion chambers to improve the quantification accuracy of combustor dynamics, performance, and increase safety margin. Also, by operating at higher temperature and locating closer to the combustion chamber, reduction of the length (weight) of pressure tubes that are currently used will be achieved. This will result in reduced costlb to access space.

A Study of Premixed, Shock-Induced Combustion With Application to Hypervelocity Flight

NASA Technical Reports Server (NTRS)

Axdahl, Erik L.

2013-01-01

One of the current goals of research in hypersonic, airbreathing propulsion is access to higher Mach numbers. A strong driver of this goal is the desire to integrate a scramjet engine into a transatmospheric vehicle airframe in order to improve performance to low Earth orbit (LEO) or the performance of a semiglobal transport. An engine concept designed to access hypervelocity speeds in excess of Mach 10 is the shock-induced combustion ramjet (i.e. shcramjet). This dissertation presents numerical studies simulating the physics of a shcramjet vehicle traveling at hypervelocity speeds with the goal of understanding the physics of fuel injection, wall autoignition mitigation, and combustion instability in this flow regime.

Rocket-Based Combined Cycle Engine Technology Development: Inlet CFD Validation and Application

NASA Technical Reports Server (NTRS)

DeBonis, J. R.; Yungster, S.

1996-01-01

A CFD methodology has been developed for inlet analyses of Rocket-Based Combined Cycle (RBCC) Engines. A full Navier-Stokes analysis code, NPARC, was used in conjunction with pre- and post-processing tools to obtain a complete description of the flow field and integrated inlet performance. This methodology was developed and validated using results from a subscale test of the inlet to a RBCC 'Strut-Jet' engine performed in the NASA Lewis 1 x 1 ft. supersonic wind tunnel. Results obtained from this study include analyses at flight Mach numbers of 5 and 6 for super-critical operating conditions. These results showed excellent agreement with experimental data. The analysis tools were also used to obtain pre-test performance and operability predictions for the RBCC demonstrator engine planned for testing in the NASA Lewis Hypersonic Test Facility. This analysis calculated the baseline fuel-off internal force of the engine which is needed to determine the net thrust with fuel on.

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

DOE Office of Scientific and Technical Information (OSTI.GOV)

Toro, P. G. P.; Minucci, M. A. S.; Chanes, J. B. Jr

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 producedmore » 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.« less

Remote Thermal Analysis Through the Internet

NASA Astrophysics Data System (ADS)

Malroy, Eric T.

2002-07-01

The Heater of the Hypersonic Tunnel Facility (HTF) was modeled using SINDA/FLUINT thermal software. A description of the model is given. The project presented the opportunity of interfacing the thermal model with the Internet and was a demonstration that complex analysis is possible through the Internet. Some of the issues that need to be addressed related to interfacing software with the Internet are the following: justification for using the Internet, selection of the web server, choice of the CGI language, security of the system, communication among the parties, maintenance of state between web pages, and simultaneous users on the Internet system. The opportunities available for using the Internet for analysis are many and can present a significant jump in technology. This paper presents a vision how interfacing with the Internet could develop in the future. Using a separate Optical Internet (OI) for analysis, coupled with virtual reality analysis rooms (VRAR), could provide a synergistic environment to couple together engineering analysis within industry, academia, and government. The process of analysis could be broken down into sub-components so that specialization could occur resulting in superior quality, minimized cost and reduced time for engineering analysis and manufacturing. Some possible subcomponents of the system are solver routines, databases, Graphical User Interfaces, engineering design software, VRARs, computer processing, CAD systems, manufacturing, and a plethora of other options only limited by ones imagination. On a larger scope, the specialization of companies on the optical network would allow companies to rapidly construct and reconstruct their infrastructure based on changing economic conditions. This could transform business.

Structural Design of Dual-Mode Ramjets and Associated System Issues (Conception structurale des statoreacteurs mixtes et defis systeme associes)

DTIC Science & Technology

2010-09-01

Institute PAO Protection Against Oxidation RBCC Rocket Based Combined Cycle RSL Reusable Space Launchers SSTO Single Stage to Orbit TOW Take-Off Weight...Orbit) or Single Stage To Orbit ( SSTO ) vehicles, or other kind of hypersonic vehicles. For example, in the scope of the French PREPHA program, the...study of a generic SSTO vehicle led to conclusion that the best type of airbreathing engine could be the dual-mode ramjet (subsonic then supersonic

Optimization of Tangential Mass Injection for Minimizing Flow Separation in a Scramjet Inlet

DTIC Science & Technology

1991-12-01

34 Aerospace EnQineering, Vol. 11. No. 8, August 1991, p.23. 26. Heppenheimer , Thomas A . Lecture notes from Hypersonic Technologies seminar. University...AFIT/GAE/ENY,/9 lD-2 ( /~ AD-A243 868 "DTIC OPTIMIZATION OF TANGENTIAL MASS INJECTION FOR MINIMIZING FLOW SEPARATION IN A SC.R-.MJET INLET THESIS...OF TANGENTIAL MASS INJECTION FOR MINIMIZING FLOW SEPARATION IN A SCRAMJET INLEr THESIS Presented to the Faculty of the School of E.ngineering of the

Thermal and Oxidation Response of UHTC Leading Edge Samples Exposed to Simulated Hypersonic Flight Conditions (Postprint)

DTIC Science & Technology

2013-03-01

such that the oxygen mole fraction of the test gas matches that of clean air. A supersonic nozzle then acceler- ates the test gas to the proper Mach...25 km. Its key limitation is that the gas chemistry is different from air, with lower oxygen and higher moisture and carbon-dioxide levels.5 Among the...Refs. 33,34 Briefly, it is a scramjet engine built to study supersonic combustion over a range of simulated flight J. Smialek—contributing editor

Nonlinear Modeling and Initial Condition Estimation for Identifying the Aerothermodynamic Environment of the Space Shuttle Orbiter.

DTIC Science & Technology

1983-12-01

8217°%. .. o..’% - * 2’ . *. -o- . *o.oo o ,o ;j ’:-’ List of Figures Figure Page 1. System Identification of the Aerothermodynamic Environment of... System (STS) has of fered the engineering community a unique opportunity to flight test a reentry, hypersonic vehicle. The key 4 to the Shuttle’s...of the system (Refs. 7,8,9,10). Although the initial test flights have now been completed, data analysis and expansion of the existing data base

The present status and the future of missile aerodynamics

NASA Technical Reports Server (NTRS)

Nielsen, Jack N.

1989-01-01

Recent developments in the state of the art in missile aerodynamics are reviewed. Among the subjects covered are: (1) Tri-service/NASA data base, (2) wing-body interference, (3) nonlinear controls, (4) hypersonic transition, (5) vortex interference, (6) airbreathers, supersonic inlets, (7) store separation problems, (8) correlation of missile data, (9) CFD codes for complete configurations, (10) engineering prediction methods, and (11) future configurations. Suggestions are made for future research and development to advance the state of the art of missile aerodynamics.

The present status and the future of missile aerodynamics

NASA Technical Reports Server (NTRS)

Nielsen, Jack N.

1988-01-01

Some recent developments in the state of the art in missile aerodynamics are reviewed. Among the subjects covered are: (1) tri-service/NASA data base, (2) wing-body interference, (3) nonlinear controls, (4) hypersonic transition, (5) vortex interference, (6) airbreathers, supersonic inlets, (7) store separation problems, (8) correlation of missile data, (9) CFD codes for complete configurations, (10) engineering prediction methods, and (11) future configurations. Suggestions are made for future research and development to advance the state of the art of missile aerodynamics.

Human Planetary Landing System (HPLS) Capability Roadmap NRC Progress Review

NASA Technical Reports Server (NTRS)

Manning, Rob; Schmitt, Harrison H.; Graves, Claude

2005-01-01

Capability Roadmap Team. Capability Description, Scope and Capability Breakdown Structure. Benefits of the HPLS. Roadmap Process and Approach. Current State-of-the-Art, Assumptions and Key Requirements. Top Level HPLS Roadmap. Capability Presentations by Leads. Mission Drivers Requirements. "AEDL" System Engineering. Communication & Navigation Systems. Hypersonic Systems. Super to Subsonic Decelerator Systems. Terminal Descent and Landing Systems. A Priori In-Situ Mars Observations. AEDL Analysis, Test and Validation Infrastructure. Capability Technical Challenges. Capability Connection Points to other Roadmaps/Crosswalks. Summary of Top Level Capability. Forward Work.

Shuttle Tethered Aerothermodynamics Research Facility (STARFAC) Instrumentation Requirements

NASA Technical Reports Server (NTRS)

Wood, George M.; Siemers, Paul M.; Carlomagno, Giovanni M.; Hoffman, John

1986-01-01

The instrumentation requirements for the Shuttle Tethered Aerothermodynamic Research Facility (STARFAC) are presented. The typical physical properties of the terrestrial atmosphere are given along with representative atmospheric daytime ion concentrations and the equilibrium and nonequilibrium gas property comparison from a point away from a wall. STARFAC science and engineering measurements are given as are the TSS free stream gas analysis. The potential nonintrusive measurement techniques for hypersonic boundary layer research are outlined along with the quantitative physical measurement methods for aerothermodynamic studies.

Mechanism analysis of Magnetohydrodynamic heat shield system and optimization of externally applied magnetic field

NASA Astrophysics Data System (ADS)

Li, Kai; Liu, Jun; Liu, Weiqiang

2017-04-01

As a novel thermal protection technique for hypersonic vehicles, Magnetohydrodynamic (MHD) heat shield system has been proved to be of great intrinsic value in the hypersonic field. In order to analyze the thermal protection mechanisms of such a system, a physical model is constructed for analyzing the effect of the Lorentz force components in the counter and normal directions. With a series of numerical simulations, the dominating Lorentz force components are analyzed for the MHD heat flux mitigation in different regions of a typical reentry vehicle. Then, a novel magnetic field with variable included angle between magnetic induction line and streamline is designed, which significantly improves the performance of MHD thermal protection in the stagnation and shoulder areas. After that, the relationships between MHD shock control and MHD thermal protection are investigated, based on which the magnetic field above is secondarily optimized obtaining better performances of both shock control and thermal protection. Results show that the MHD thermal protection is mainly determined by the Lorentz force's effect on the boundary layer. From the stagnation to the shoulder region, the flow deceleration effect of the counter-flow component is weakened while the flow deflection effect of the normal component is enhanced. Moreover, there is no obviously positive correlation between the MHD shock control and thermal protection. But once a good Lorentz force's effect on the boundary layer is guaranteed, the thermal protection performance can be further improved with an enlarged shock stand-off distance by strengthening the counter-flow Lorentz force right after shock.

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.

Phoenix Missile Hypersonic Testbed (PMHT): Project Concept Overview

NASA Technical Reports Server (NTRS)

Jones, Thomas P.

2007-01-01

An over view of research into a low cost hypersonic research flight test capability to increase the amount of hypersonic flight data to help bridge the large developmental gap between ground testing/analysis and major flight demonstrator Xplanes is provided. The major objectives included: develop an air launched missile booster research testbed; accurately deliver research payloads through programmable guidance to hypersonic test conditions; low cost; a high flight rate minimum of two flights per year and utilize surplus air launched missiles and NASA aircraft.

Perspectives on hypersonic viscous flow research

NASA Technical Reports Server (NTRS)

Cheng, H. K.

1993-01-01

Issues and advances in current hypersonic flow research perceived to be of interest in theoretical fluid/gas dynamics are reviewed. Particular attention is given to the hypersonic aircraft as waverider, computational methods and theoretical development in the study of viscous interaction, and boundary-layer instability and transition studies. In the present framework the study of viscous hypersonic flow faces transition problems of two kinds which represent the two major areas of current research: the turbulence transition in the high Re range and the transition to the free-molecule limit.

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.

Modeling and life prediction methodology for Titanium Matrix Composites subjected to mission profiles

NASA Technical Reports Server (NTRS)

Mirdamadi, M.; Johnson, W. S.

1994-01-01

Titanium matrix composites (TMC) are being evaluated as structural materials for elevated temperature applications in future generation hypersonic vehicles. In such applications, TMC components are subjected to complex thermomechanical loading profiles at various elevated temperatures. Therefore, thermomechanical fatigue (TMF) testing, using a simulated mission profile, is essential for evaluation and development of life prediction methodologies. The objective of the research presented in this paper was to evaluate the TMF response of the (0/90)2s SCS-6/Timetal-21S subjected to a generic hypersonic flight profile and its portions with a temperature ranging from -130 C to 816 C. It was found that the composite modulus, prior to rapid degradation, had consistent values for all the profiles tested. A micromechanics based analysis was used to predict the stress-strain response of the laminate and of the constituents in each ply during thermomechanical loading conditions by using only constituent properties as input. The fiber was modeled as elastic with transverse orthotropic and temperature dependent properties. The matrix was modeled using a thermoviscoplastic constitutive relation. In the analysis, the composite modulus degradation was assumed to result from matrix cracking and was modeled by reducing the matrix modulus. Fatigue lives of the composite subjected to the complex generic hypersonic flight profile were well correlated using the predicted stress in 0 degree fibers.

Improving the Flow

NASA Technical Reports Server (NTRS)

2004-01-01

In early 1995, NASA s Glenn Research Center (then Lewis Research Center) formed an industry-government team with several jet engine companies to develop the National Combustion Code (NCC), which would help aerospace engineers solve complex aerodynamics and combustion problems in gas turbine, rocket, and hypersonic engines. The original development team consisted of Allison Engine Company (now Rolls-Royce Allison), CFD Research Corporation, GE Aircraft Engines, Pratt and Whitney, and NASA. After the baseline beta version was established in July 1998, the team focused its efforts on consolidation, streamlining, and integration, as well as enhancement, evaluation, validation, and application. These activities, mainly conducted at NASA Glenn, led to the completion of NCC version 1.0 in October 2000. NCC version 1.0 features high-fidelity representation of complex geometry, advanced models for two-phase turbulent combustion, and massively parallel computing. Researchers and engineers at Glenn have been using NCC to provide analysis and design support for various aerospace propulsion technology development projects. NASA transfers NCC technology to external customers using non- exclusive Space Act Agreements. Glenn researchers also communicate research and development results derived from NCC's further development through publications and special sessions at technical conferences.

Parametric scramjet analysis

NASA Astrophysics Data System (ADS)

Choi, Jongseong

The performance of a hypersonic flight vehicle will depend on existing materials and fuels; this work presents the performance of the ideal scramjet engine for three different combustion chamber materials and three different candidate fuels. Engine performance is explored by parametric cycle analysis for the ideal scramjet as a function of material maximum service temperature and the lower heating value of jet engine fuels. The thermodynamic analysis is based on the Brayton cycle as similarly employed in describing the performance of the ramjet, turbojet, and fanjet ideal engines. The objective of this work is to explore material operating temperatures and fuel possibilities for the combustion chamber of a scramjet propulsion system to show how they relate to scramjet performance and the seven scramjet engine parameters: specific thrust, fuel-to-air ratio, thrust-specific fuel consumption, thermal efficiency, propulsive efficiency, overall efficiency, and thrust flux. The information presented in this work has not been done by others in the scientific literature. This work yields simple algebraic equations for scramjet performance which are similar to that of the ideal ramjet, ideal turbojet and ideal turbofan engines.

Classifier utility modeling and analysis of hypersonic inlet start/unstart considering training data costs

NASA Astrophysics Data System (ADS)

Chang, Juntao; Hu, Qinghua; Yu, Daren; Bao, Wen

2011-11-01

Start/unstart detection is one of the most important issues of hypersonic inlets and is also the foundation of protection control of scramjet. The inlet start/unstart detection can be attributed to a standard pattern classification problem, and the training sample costs have to be considered for the classifier modeling as the CFD numerical simulations and wind tunnel experiments of hypersonic inlets both cost time and money. To solve this problem, the CFD simulation of inlet is studied at first step, and the simulation results could provide the training data for pattern classification of hypersonic inlet start/unstart. Then the classifier modeling technology and maximum classifier utility theories are introduced to analyze the effect of training data cost on classifier utility. In conclusion, it is useful to introduce support vector machine algorithms to acquire the classifier model of hypersonic inlet start/unstart, and the minimum total cost of hypersonic inlet start/unstart classifier can be obtained by the maximum classifier utility theories.

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 matrix of flush ports is placed in the vicinity of the aircraft nose, and the airdata are inferred indirectly from the measured pressures.

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 application of the suite. The current models are accurate at low supersonic speed and reasonable for engineering approximation at hypersonic speeds. Improvements to extend the models fully into the hypersonic regime are given in the Recommendations section.

Mission and Design Sensitivities for Human Mars Landers Using Hypersonic Inflatable Aerodynamic Decelerators

NASA Technical Reports Server (NTRS)

Polsgrove, Tara P.; Thomas, Herbert D.; Dwyer Ciancio, Alicia; Collins, Tim; Samareh, Jamshid

2017-01-01

Landing humans on Mars is one of NASA's long term goals. NASA's Evolvable Mars Campaign (EMC) is focused on evaluating architectural trade options to define the capabilities and elements needed to sustain human presence on the surface of Mars. The EMC study teams have considered a variety of in-space propulsion options and surface mission options. Understanding how these choices affect the performance of the lander will allow a balanced optimization of this complex system of systems problem. This paper presents the effects of mission and vehicle design options on lander mass and performance. Beginning with Earth launch, options include fairing size assumptions, co-manifesting elements with the lander, and Earth-Moon vicinity operations. Capturing into Mars orbit using either aerocapture or propulsive capture is assessed. For entry, descent, and landing both storable as well as oxygen and methane propellant combinations are considered, engine thrust level is assessed, and sensitivity to landed payload mass is presented. This paper focuses on lander designs using the Hypersonic Inflatable Aerodynamic Decelerators, one of several entry system technologies currently considered for human missions.

Aerodynamic characteristics at Mach numbers from 0.33 to 1.20 of a wing-body design concept for a hypersonic research airplane

NASA Technical Reports Server (NTRS)

Dillon, J. L.; Pittman, J. L.

1977-01-01

An experimental investigation of the static aerodynamic characteristics of a model of one design concept for the proposed National Hypersonic Flight Research Facility was conducted in the Langley 8 foot transonic pressure tunnel. The experiment consisted of configuration buildup from the basic body by adding a wing, center vertical tail, and a three module or six module scramjet engine. The freestream test Mach numbers were 0.33, 0.80, 0.90, 0.95, 0.98, 1.10, and 1.20 at Reynolds numbers per meter ranging from 4.8 x 1 million to 10.4 x 1 million. The test angle of attack range was approximately -4 deg to 22 deg at constant angles of sideslip of 0 deg and 4 deg; the angle of sideslip ranged from about -6 deg to 6 deg at constant angles of attack of 0 deg and 17 deg. The elevons were deflected 0 deg, -10 deg, and -20 deg with rudder deflections of 0 deg and 15.6 deg.

Overview of hypersonic CFD code calibration studies

NASA Technical Reports Server (NTRS)

Miller, Charles G.

1987-01-01

The topics are presented in viewgraph form and include the following: definitions of computational fluid dynamics (CFD) code validation; climate in hypersonics and LaRC when first 'designed' CFD code calibration studied was initiated; methodology from the experimentalist's perspective; hypersonic facilities; measurement techniques; and CFD code calibration studies.

Strain sensing technology for high temperature applications

NASA Technical Reports Server (NTRS)

Williams, W. Dan

1993-01-01

This review discusses the status of strain sensing technology for high temperature applications. Technologies covered are those supported by NASA such as required for applications in hypersonic vehicles and engines, advanced subsonic engines, as well as material and structure development. The applications may be at temperatures of 540 C (1000 F) to temperatures in excess of 1400 C (2500 F). The most promising technologies at present are the resistance strain gage and remote sensing schemes. Resistance strain gages discussed include the BCL gage, the LaRC compensated gage, and the PdCr gage. Remote sensing schemes such as laser based speckle strain measurement, phase-shifling interferometry, and x-ray extensometry are discussed. Present status and limitations of these technologies are presented.

The dual combustor ramjet: A versatile propulsion system for hypersonic tactical missile applications

NASA Astrophysics Data System (ADS)

Waltrup, Paul J.

1992-09-01

Procedures for designing and maximizing the performance of Dual Combustor Ramjet (DCR) engines and vehicles powered by this engine are presented. Comparisons of DCR powered vehicles with scramjet powered vehicles for Mach 4 to 8 flight show that the DCR provides better performance at the Mach 4 flight condition, while the scramjet is better at Mach 8. Comparisons of the DCR with a ramjet for Mach 3 to 6 flight, with both having the same, but low, thrust level at Mach 3, show that the DCR exhibits better performance at and near the cruise condition at Mach 6 and similar performance during acceleration. Suggested additional comparisons to broaden the scope of the conclusions are also given.

Application of computational fluid dynamics to the design of the Space Transportation Main Engine subscale nozzle

NASA Technical Reports Server (NTRS)

Garrett, J. L.; Syed, S. A.

1992-01-01

CFD analyses of the Space Transportation Main Engine film/dump cooled subscale nozzle are presented, with an emphasis on the timely impact of CFD in the design of the subscale nozzle secondary coolant system. Calculations were performed with the Generalized Aerodynamic Simulation Program (GASP), using a Baldwin-Lomas Turbulence model, and finite rate hydrogen-oxygen chemistry. Design iterations for both the secondary coolant cavity passage and the secondary coolant lip are presented. In addition, validation of the GASP chemistry and turbulence models by comparison with data and other CFD codes are presented for a hypersonic laminar separation corner, a backward facing step, and a 2D scramjet nozzle with hydrogen-oxygen kinetics.

Aeroheating Design Issues for Reusable Launch Vehicles: A Perspective

NASA Technical Reports Server (NTRS)

Zoby, E. Vincent; Thompson, Richard A.; Wurster, Kathryn E.

2004-01-01

An overview of basic aeroheating design issues for Reusable Launch Vehicles (RLV), which addresses the application of hypersonic ground-based testing, and computational fluid dynamic (CFD) and engineering codes, is presented. Challenges inherent to the prediction of aeroheating environments required for the successful design of the RLV Thermal Protection System (TPS) are discussed in conjunction with the importance of employing appropriate experimental/computational tools. The impact of the information garnered by using these tools in the resulting analyses, ultimately enhancing the RLV TPS design is illustrated. A wide range of topics is presented in this overview; e.g. the impact of flow physics issues such as boundary-layer transition, including effects of distributed and discrete roughness, shock-shock interactions, and flow separation/reattachment. Also, the benefit of integrating experimental and computational studies to gain an improved understanding of flow phenomena is illustrated. From computational studies, the effect of low-density conditions and of uncertainties in material surface properties on the computed heating rates a r e highlighted as well as the significant role of CFD in improving the Outer Mold Line (OML) definition to reduce aeroheating while maintaining aerodynamic performance. Appropriate selection of the TPS design trajectories and trajectory shaping to mitigate aeroheating levels and loads are discussed. Lastly, an illustration of an aeroheating design process is presented whereby data from hypersonic wind-tunnel tests are integrated with predictions from CFD codes and engineering methods to provide heating environments along an entry trajectory as required for TPS design.

Aeroheating Design Issues for Reusable Launch Vehicles: A Perspective

NASA Technical Reports Server (NTRS)

Zoby, E. Vincent; Thompson, Richard A.; Wurster, Kathryn E.

2004-01-01

An overview of basic aeroheating design issues for Reusable Launch Vehicles (RLV), which addresses the application of hypersonic ground-based testing, and computational fluid dynamic (CFD) and engineering codes, is presented. Challenges inherent to the prediction of aeroheating environments required for the successful design of the RLV Thermal Protection System (TPS) are discussed in conjunction with the importance of employing appropriate experimental/computational tools. The impact of the information garnered by using these tools in the resulting analyses, ultimately enhancing the RLV TPS design is illustrated. A wide range of topics is presented in this overview; e.g. the impact of flow physics issues such as boundary-layer transition, including effects of distributed and discrete roughness, shockshock interactions, and flow separation/reattachment. Also, the benefit of integrating experimental and computational studies to gain an improved understanding of flow phenomena is illustrated. From computational studies, the effect of low-density conditions and of uncertainties in material surface properties on the computed heating rates are highlighted as well as the significant role of CFD in improving the Outer Mold Line (OML) definition to reduce aeroheating while maintaining aerodynamic performance. Appropriate selection of the TPS design trajectories and trajectory shaping to mitigate aeroheating levels and loads are discussed. Lastly, an illustration of an aeroheating design process is presented whereby data from hypersonic wind-tunnel tests are integrated with predictions from CFD codes and engineering methods to provide heating environments along an entry trajectory as required for TPS design.

Aerothermodynamics and Turbulence

DTIC Science & Technology

2013-03-08

Surface Heat Transfer and Detailed Flow Structure Fuel Injection in a Scramjet Combustor Reduced Uncertainty in Complex Flows Addressing... hypersonic flight data to capture shock interaction unsteadiness National Hypersonic Foundational Research Plan Joint Technology Office... Hypersonics Basic Science Roadmap Assessment of SOA and Future Research Directions Ongoing Basic Research for Understanding and Controlling Noise

Numerical method of carbon-based material ablation effects on aero-heating for half-sphere

NASA Astrophysics Data System (ADS)

Wang, Jiang-Feng; Li, Jia-Wei; Zhao, Fa-Ming; Fan, Xiao-Feng

2018-05-01

A numerical method of aerodynamic heating with material thermal ablation effects for hypersonic half-sphere is presented. A surface material ablation model is provided to analyze the ablation effects on aero-thermal properties and structural heat conduction for thermal protection system (TPS) of hypersonic vehicles. To demonstrate its capability, applications for thermal analysis of hypersonic vehicles using carbonaceous ceramic ablators are performed and discussed. The numerical results show the high efficiency and validation of the method developed in thermal characteristics analysis of hypersonic aerodynamic heating.

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.

Development of an aerodynamic measurement system for hypersonic rarefied flows

NASA Astrophysics Data System (ADS)

Ozawa, T.; Fujita, K.; Suzuki, T.

2015-01-01

A hypersonic rarefied wind tunnel (HRWT) has lately been developed at Japan Aerospace Exploration Agency in order to improve the prediction of rarefied aerodynamics. Flow characteristics of hypersonic rarefied flows have been investigated experimentally and numerically. By conducting dynamic pressure measurements with pendulous models and pitot pressure measurements, we have probed flow characteristics in the test section. We have also improved understandings of hypersonic rarefied flows by integrating a numerical approach with the HRWT measurement. The development of the integration scheme between HRWT and numerical approach enables us to estimate the hypersonic rarefied flow characteristics as well as the direct measurement of rarefied aerodynamics. Consequently, this wind tunnel is capable of generating 25 mm-core flows with the free stream Mach number greater than 10 and Knudsen number greater than 0.1.

Controllers for Flow-Field Survey Apparatus

NASA Technical Reports Server (NTRS)

Ashby George C., JR.; Vaccarelli, M. D.

1986-01-01

Control systems of flow-field survey apparatuses of 22-inch (56centimeter) Hypersonic Helium Facility (two-dimensional) and 20-inch (51centimeter) Mach 6 Tunnel (three-dimensional) at Langley Research Center equipped with single-chip microcomputer and single-board microcomputer, respectively, to drive probes at selected speeds and perform other functions automatically. Various modes of operation programed as need arises. Both of these control systems fabricated relatively inexpensively from commercially available stock components.

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.

Computational method to predict thermodynamic, transport, and flow properties for the modified Langley 8-foot high-temperature tunnel

NASA Technical Reports Server (NTRS)

Venkateswaran, S.; Hunt, L. Roane; Prabhu, Ramadas K.

1992-01-01

The Langley 8 foot high temperature tunnel (8 ft HTT) is used to test components of hypersonic vehicles for aerothermal loads definition and structural component verification. The test medium of the 8 ft HTT is obtained by burning a mixture of methane and air under high pressure; the combustion products are expanded through an axisymmetric conical contoured nozzle to simulate atmospheric flight at Mach 7. This facility was modified to raise the oxygen content of the test medium to match that of air and to include Mach 4 and Mach 5 capabilities. These modifications will facilitate the testing of hypersonic air breathing propulsion systems for a wide range of flight conditions. A computational method to predict the thermodynamic, transport, and flow properties of the equilibrium chemically reacting oxygen enriched methane-air combustion products was implemented in a computer code. This code calculates the fuel, air, and oxygen mass flow rates and test section flow properties for Mach 7, 5, and 4 nozzle configurations for given combustor and mixer conditions. Salient features of the 8 ft HTT are described, and some of the predicted tunnel operational characteristics are presented in the carpet plots to assist users in preparing test plans.

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.

Hyper-X Mach 7 Scramjet Design, Ground Test and Flight Results

NASA Technical Reports Server (NTRS)

Ferlemann, Shelly M.; McClinton, Charles R.; Rock, Ken E.; Voland, Randy T.

2005-01-01

The successful Mach 7 flight test of the Hyper-X (X-43) research vehicle has provided the major, essential demonstration of the capability of the airframe integrated scramjet engine. This flight was a crucial first step toward realizing the potential for airbreathing hypersonic propulsion for application to space launch vehicles. However, it is not sufficient to have just achieved a successful flight. The more useful knowledge gained from the flight is how well the prediction methods matched the actual test results in order to have confidence that these methods can be applied to the design of other scramjet engines and powered vehicles. The propulsion predictions for the Mach 7 flight test were calculated using the computer code, SRGULL, with input from computational fluid dynamics (CFD) and wind tunnel tests. This paper will discuss the evolution of the Mach 7 Hyper-X engine, ground wind tunnel experiments, propulsion prediction methodology, flight results and validation of design methods.

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.

Pulse Detonation Engines for High Speed Flight

NASA Technical Reports Server (NTRS)

Povinelli, Louis A.

2002-01-01

Revolutionary concepts in propulsion are required in order to achieve high-speed cruise capability in the atmosphere and for low cost reliable systems for earth to orbit missions. One of the advanced concepts under study is the air-breathing pulse detonation engine. Additional work remains in order to establish the role and performance of a PDE in flight applications, either as a stand-alone device or as part of a combined cycle system. In this paper, we shall offer a few remarks on some of these remaining issues, i.e., combined cycle systems, nozzles and exhaust systems and thrust per unit frontal area limitations. Currently, an intensive experimental and numerical effort is underway in order to quantify the propulsion performance characteristics of this device. In this paper, we shall highlight our recent efforts to elucidate the propulsion potential of pulse detonation engines and their possible application to high-speed or hypersonic systems.

Toward an Improved Hypersonic Engine Seal

NASA Technical Reports Server (NTRS)

Dunlap, Patrick H., Jr.; Steinetz, Bruce M.; DeMange,Jeffrey J.; Taylor, Shawn C.

2003-01-01

High temperature, dynamic seals are required in advanced engines to seal the perimeters of movable engine ramps for efficient, safe operation in high heat flux environments at temperatures from 2000 to 2500 F. Current seal designs do not meet the demanding requirements for future engines, so NASA s Glenn Research Center (GRC) is developing advanced seals to overcome these shortfalls. Two seal designs and two types of seal preloading devices were evaluated in a series of compression tests at room temperature and 2000 F and flow tests at room temperature. Both seals lost resiliency with repeated load cycling at room temperature and 2000 F, but seals with braided cores were significantly more flexible than those with cores composed of uniaxial ceramic fibers. Flow rates for the seals with cores of uniaxial fibers were lower than those for the seals with braided cores. Canted coil springs and silicon nitride compression springs showed promise conceptually as potential seal preloading devices to help maintain seal resiliency.

PIC Simulations of Hypersonic Plasma Instabilities

NASA Astrophysics Data System (ADS)

Niehoff, D.; Ashour-Abdalla, M.; Niemann, C.; Decyk, V.; Schriver, D.; Clark, E.

2013-12-01

The plasma sheaths formed around hypersonic aircraft (Mach number, M > 10) are relatively unexplored and of interest today to both further the development of new technologies and solve long-standing engineering problems. Both laboratory experiments and analytical/numerical modeling are required to advance the understanding of these systems; it is advantageous to perform these tasks in tandem. There has already been some work done to study these plasmas by experiments that create a rapidly expanding plasma through ablation of a target with a laser. In combination with a preformed magnetic field, this configuration leads to a magnetic "bubble" formed behind the front as particles travel at about Mach 30 away from the target. Furthermore, the experiment was able to show the generation of fast electrons which could be due to instabilities on electron scales. To explore this, future experiments will have more accurate diagnostics capable of observing time- and length-scales below typical ion scales, but simulations are a useful tool to explore these plasma conditions theoretically. Particle in Cell (PIC) simulations are necessary when phenomena are expected to be observed at these scales, and also have the advantage of being fully kinetic with no fluid approximations. However, if the scales of the problem are not significantly below the ion scales, then the initialization of the PIC simulation must be very carefully engineered to avoid unnecessary computation and to select the minimum window where structures of interest can be studied. One method of doing this is to seed the simulation with either experiment or ion-scale simulation results. Previous experiments suggest that a useful configuration for studying hypersonic plasma configurations is a ring of particles rapidly expanding transverse to an external magnetic field, which has been simulated on the ion scale with an ion-hybrid code. This suggests that the PIC simulation should have an equivalent configuration; however, modeling a plasma expanding radially in every direction is computationally expensive. In order to reduce the computational expense, we use a radial density profile from the hybrid simulation results to seed a self-consistent PIC simulation in one direction (x), while creating a current in the direction (y) transverse to both the drift velocity and the magnetic field (z) to create the magnetic bubble observed in experiment. The simulation will be run in two spatial dimensions but retain three velocity dimensions, and the results will be used to explore the growth of micro-instabilities present in hypersonic plasmas in the high-density region as it moves through the simulation box. This will still require a significantly large box in order to compare with experiment, as the experiments are being performed over distances of 104 λDe and durations of 105 ωpe-1.

Mathematical modeling and characteristic analysis for over-under turbine based combined cycle engine

NASA Astrophysics Data System (ADS)

Ma, Jingxue; Chang, Juntao; Ma, Jicheng; Bao, Wen; Yu, Daren

2018-07-01

The turbine based combined cycle engine has become the most promising hypersonic airbreathing propulsion system for its superiority of ground self-starting, wide flight envelop and reusability. The simulation model of the turbine based combined cycle engine plays an important role in the research of performance analysis and control system design. In this paper, a turbine based combined cycle engine mathematical model is built on the Simulink platform, including a dual-channel air intake system, a turbojet engine and a ramjet. It should be noted that the model of the air intake system is built based on computational fluid dynamics calculation, which provides valuable raw data for modeling of the turbine based combined cycle engine. The aerodynamic characteristics of turbine based combined cycle engine in turbojet mode, ramjet mode and mode transition process are studied by the mathematical model, and the influence of dominant variables on performance and safety of the turbine based combined cycle engine is analyzed. According to the stability requirement of thrust output and the safety in the working process of turbine based combined cycle engine, a control law is proposed that could guarantee the steady output of thrust by controlling the control variables of the turbine based combined cycle engine in the whole working process.

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

NASA Astrophysics Data System (ADS)

Salvador, Israel Irone

The present research campaign centered on static and hypersonic experiments performed with a two-dimensional, repetitively-pulsed (RP) laser Lightcraft model. The future application of interest for this basic research endeavor is the laser launch of nano- and micro-satellites (i.e., 1-100 kg payloads) into Low Earth Orbit (LEO), at low-cost and "on-demand". This research began with an international collaboration on Beamed Energy Propulsion between the United States Air Force and Brazilian Air Force to conduct experiments at the Henry T. Nagamatsu Laboratory of Aerothermodynamics and Hypersonics (HTN-LAH). The laser propulsion (LP) experiments employed the T3 Hypersonic Shock Tunnel (HST), integrated with twin gigawatt pulsed Lumonics 620-TEA CO2 lasers to produce the required test conditions. Following an introduction of the pulsed laser thermal propulsion concept and a state-of-the-art review of the topic, the principal physical processes are outlined starting from the onset of the laser pulse and subsequent laser-induced air-breakdown, to the expansion and exhaust of the resulting blast wave. After installation of the 254 mm wide, 2D Lightcraft model into the T3 tunnel, static LP tests were performed under quiescent (no-flow) conditions at ambient pressures of 0.06, 0.15, 0.3 and 1 bar, using the T3 test-section/dump-tank as a vacuum chamber. Time-dependent surface pressure distributions were measured over the engine thrust-generating surfaces following laser energy deposition; the delivered impulse and momentum coupling coefficients (Cm) were calculated from that pressure data. A Schlieren visualization system (using a high-speed Cordin digital camera) captured the laser breakdown and blast wave expansion process. The 2D model's Cm performance of 600 to 3000 N/MW was 2.5-5x higher than theoretical projections available in the literature, but indeed in the realm of feasibility for static conditions. Also, these Cm values exceed that for smaller Lightcraft models (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.

Infrared measurement and composite tracking algorithm for air-breathing hypersonic vehicles

NASA Astrophysics Data System (ADS)

Zhang, Zhao; Gao, Changsheng; Jing, Wuxing

2018-03-01

Air-breathing hypersonic vehicles have capabilities of hypersonic speed and strong maneuvering, and thus pose a significant challenge to conventional tracking methodologies. To achieve desirable tracking performance for hypersonic targets, this paper investigates the problems related to measurement model design and tracking model mismatching. First, owing to the severe aerothermal effect of hypersonic motion, an infrared measurement model in near space is designed and analyzed based on target infrared radiation and an atmospheric model. Second, using information from infrared sensors, a composite tracking algorithm is proposed via a combination of the interactive multiple models (IMM) algorithm, fitting dynamics model, and strong tracking filter. During the procedure, the IMMs algorithm generates tracking data to establish a fitting dynamics model of the target. Then, the strong tracking unscented Kalman filter is employed to estimate the target states for suppressing the impact of target maneuvers. Simulations are performed to verify the feasibility of the presented composite tracking algorithm. The results demonstrate that the designed infrared measurement model effectively and continuously observes hypersonic vehicles, and the proposed composite tracking algorithm accurately and stably tracks these targets.

Dynamics Evolution Investigation of Mack Mode Instability in a Hypersonic Boundary Layer by Bicoherence Spectrum Analysis

NASA Astrophysics Data System (ADS)

Han, Jian; Jiang, Nan

2012-07-01

The instability of a hypersonic boundary layer on a cone is investigated by bicoherence spectrum analysis. The experiment is conducted at Mach number 6 in a hypersonic wind tunnel. The time series signals of instantaneous fluctuating surface-thermal-flux are measured by Pt-thin-film thermocouple temperature sensors mounted at 28 stations on the cone surface along streamwise direction to investigate the development of the unstable disturbances. The bicoherence spectrum analysis based on wavelet transform is employed to investigate the nonlinear interactions of the instability of Mack modes in hypersonic laminar boundary layer transition. The results show that wavelet bicoherence is a powerful tool in studying the unstable mode nonlinear interaction of hypersonic laminar-turbulent transition. The first mode instability gives rise to frequency shifts to higher unstable modes at the early stage of hypersonic laminar-turbulent transition. The modulations subsequently lead to the second mode instability occurrence. The second mode instability governs the last stage of instability and final breakdown to turbulence with multi-scale disturbances growth.

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.

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.

Results of investigations on a 0.010-scale model of the configuration 3 space shuttle orbiter and external tank in the NASA/Ames Research Center 3.5-foot hypersonic wind tunnel (IA15)

NASA Technical Reports Server (NTRS)

Petrozzi, M. T.; Milam, M. D.; Mellenthin, J. A.

1974-01-01

Experimental aerodynamic investigations were conducted in a 3.5-foot hypersonic wind tunnel. The model used for this test was a 0.010-scale of the Configuration 2 Space Shuttle Orbiter and the External Tank. Six-component aerodynamic force and moment data were recorded over an angle of attack range from -8 deg to +30 deg at 0 deg and 5 deg angles of sideslip. Data was also recorded during beta sweeps of -8 deg to +10 deg at angles of attack of -10 deg, 0 deg, and 30 deg. All testing was done at Mach 7.3. Various elevon, rudder and orbiter to external tank attaching structures and fairings were tested to determine longitudinal and lateral-directional stability characteristics. Non-metric exhaust plumes were installed during a portion of the testing to determine the effects of the main propulsion system rocket plumes.

Pitot pressure analyses in CO2 condensing rarefied hypersonic flows

NASA Astrophysics Data System (ADS)

Ozawa, T.; Suzuki, T.; Fujita, K.

2016-11-01

In order to improve the accuracy of rarefied aerodynamic prediction, a hypersonic rarefied wind tunnel (HRWT) was developed at Japan Aerospace Exploration Agency. While this wind tunnel has been limited to inert gases, such as nitrogen or argon, we recently extended the capability of HRWT to CO2 hypersonic flows for several Mars missions. Compared to our previous N2 cases, the condensation effect may not be negligible for CO2 rarefied aerodynamic measurements. Thus, in this work, we have utilized both experimental and numerical approaches to investigate the condensation and rarefaction effects in CO2 hypersonic nozzle flows.

The Center of Excellence for Hypersonics Training and Research at the University of Texas at Austin

NASA Technical Reports Server (NTRS)

Dolling, David S.

1993-01-01

Over the period of this grant (1986-92), 23 graduate students were supported by the Center and received education and training in hypersonics through MS and Ph.D. programs. An additional 8 Ph.D. candidates and 2 MS candidates, with their own fellowship support, were attracted to The University of Texas and were recruited into the hypersonics program because of the Center. Their research, supervised by the 10 faculty involved in the Center, resulted in approximately 50 publications and presentations in journals and at national and international technical conferences. To provide broad-based training, a new hypersonics curriculum was created, enabling students to take 8 core classes in theoretical, computational, and experimental hypersonics, and other option classes over a two to four semester period. The Center also developed an active continuing education program. The Hypersonics Short Course was taught 3 times, twice in the USA and once in Europe. Approximately 300 persons were attracted to hear lectures by more than 25 of the leading experts in the field. In addition, a hypersonic aerodynamics short course was offered through AIAA, as well as short courses on computational fluid dynamics (CFD) and advanced CFD. The existence of the Center also enabled faculty to leverage a substantial volume of additional funds from other agencies, for research and graduate student training. Overall, this was a highly successful and highly visible program.

The Kinetic Nonequilibrium Processes in the Internal Flow and in the Plume of Subsonic and Supersonic Aircrafts

NASA Technical Reports Server (NTRS)

Starik, Alexander M.

1997-01-01

(1) Our results show that under combustion of thermal destruction products of n-C8H18, and other hydrocarbon fuels with air at the equivalent ratio -0.5 and less the chemical equilibrium is not realized at the exit plane of combustion chamber and in the gas turbine and nozzle for most of small components such as NO2, NO3, HNO, HNO2, HNO3, N(x)H(y), HO2, OH. The chemical equilibrium is not realized in the internal flow of ramjet hydrogen combustion engine too. So at the nozzle exit plane both of gas-turbine hydrocarbon combustion engine and of ramjet hydrogen combustion engine the relatively large values of concentration of such small components as NO3, HNO2, N2O, HNO3, HNO, NH, N2H, HO2, H2O2 may be realized. The exact definition of these component concentration as well as concentration of NO(x), OH, SO2, O, H, H2, H2O at the nozzle exit plane is very important for plume chemistry. (2) The results which were obtained for subsonic and hypersonic aircrafts indicate on the considerable change of the composition of the gas mixture along the plume. This change can be caused not only by the mixture of combustion products with the atmosphere air but by proceeding of whole complex of nonequilibrium photochemical reactions. The photodissociation processes begin to influence on the formation of the free atoms and radicals at flight altitude H greater than or equal to 18 km. Neglect of these processes can result in essential (up to 10(exp 4) times) mistakes of values gamma(sub OH), gamma(sub O), gamma(sub H), gamma(sub HSO3) and some products of CFC's disintegration. It was found that penetration of Cl-containing species from the atmosphere into the exhaust flow and its interaction with nitrogen oxides leads to essential increasing of the concentration of Cl, Cl2, ClO2, ClNO3, CH3Cl and sometimes HCl and the decreasing of ClO concentration by comparison with background values. The results of our analysis show that the plume aircraft with both hydrocarbon and hydrogen combustion engine may be source of various pollutant components such as HNO, HNO4,ClO2, CH3NO2, CH3NO3, CH2O, Cl, H2O2, but not only NO, NO2, HNO2, HNO3, N2O5, SO2, SO3, H2SO4 as it was supposed before.

Special Course on Engineering Methods in Aerodynamic Analysis and Design of Aircraft (Les Methodes d’Ingenierie Employee lors de l’Analyse et de la Conception des Aeronefs)

DTIC Science & Technology

1992-01-01

study assumed to the VLM-SA method (solid carve), obtained by combininglie between 15 and 75% of the local wing chord) The ex - the potential-flow...the Consultant and Lxt hange Programme and the Aeroslace Apphations Studies Programme The results oa AGARD work are reported it) the member rations...Testing AGARD CP.429, July 1988 Aerodynamics of Hypersonic, Lifting Vehicles AGARD CP-428, Novemnber 1987 Aerodynamic and Related Hydr dynamic Studies

Experimental Investigation of Nozzle/Plume Aerodynamics at Hypersonic Speeds

NASA Technical Reports Server (NTRS)

Heinemann, K.; Bogdanoff, David W.; Cambier, Jean-Luc

1992-01-01

The work performed by D. W. Bogdanoff and J.-L. Cambier during the period of 1 Feb. - 31 Oct. 1992 is presented. The following topics are discussed: (1) improvement in the operation of the facility; (2) the wedge model; (3) calibration of the new test section; (4) combustor model; (5) hydrogen fuel system for combustor model; (6) three inch calibration/development tunnel; (7) shock tunnel unsteady flow; (8) pulse detonation wave engine; (9) DCAF flow simulation; (10) high temperature shock layer simulation; and (11) the one dimensional Godunov CFD code.

Dynamic, High-Temperature, Flexible Seal

NASA Technical Reports Server (NTRS)

Steinetz, Bruce M.; Sirocky, Paul J.

1989-01-01

New seal consists of multiple plies of braided ceramic sleeves filled with small ceramic balls. Innermost braided sleeve supported by high-temperature-wire-mesh sleeve that provides both springback and preload capabilities. Ceramic balls reduce effect of relatively high porosity of braided ceramic sleeves by acting as labyrinth flow path for gases and thereby greatly increasing pressure gradient seal can sustain. Dynamic, high-temperature, flexible seal employed in hypersonic engines, two-dimensional convergent/divergent and vectorized-thrust exhaust nozzles, reentry vehicle airframes, rocket-motor casings, high-temperature furnaces, and any application requiring non-asbestos high-temperature gaskets.

Design of three-dimensional scramjet inlets for hypersonic propulsion

NASA Technical Reports Server (NTRS)

Simmons, J. M.; Weidner, E. H.

1986-01-01

The paper outlines an approach to the design of three-dimensional inlets for scramjet engines. The basis of the techniques used is the method of streamline tracing through an inviscid axisymmetric flow field. A technique is described for making a smooth change of cross-section shape from rectangular to circular. A feature is the considerable use of computer-graphics to provide a 'user-oriented' procedure which can produce promising design configurations for subsequent analysis with CFD codes. An example is given to demonstrate the capabilities of the design techniques.

Propellant Management and Conditioning within the X-34 Main Propulsion System

NASA Technical Reports Server (NTRS)

Brown, T. M.; McDonald, J. P.; Hedayat, A.; Knight, K. C.; Champion, R. H., Jr.

1998-01-01

The X-34 hypersonic flight vehicle is currently under development by Orbital Sciences Corporation (Orbital). The Main Propulsion ystem as been designed around the liquid propellant Fastrac rocket engine currently under development at NASA Marshall Space Flight Center. This paper presents analyses of the MPS subsystems used to manage the liquid propellants. These subsystems include the propellant tanks, the tank vent/relief subsystem, and the dump/fill/drain subsystem. Analyses include LOX tank chill and fill time estimates, LOX boil-off estimates, propellant conditioning simulations, and transient propellant dump simulations.

Effect of Sidewall Configurations on Hypersonic Intake Performance

NASA Astrophysics Data System (ADS)

Kim, Seihwan; Park, Ji Hyun; Jeung, In-Seuck; Lee, Hyoung Jin

For reusable space launchers and hypersonic flight vehicles, use of an air-breathing propulsion system with supersonic combustion is the most promising option in terms of cost effectiveness. At this point, only the scramjet propulsion system provides a real alternative to expensive rocket driven systems, which currently are the only way to reach a hypersonics speeds.

AFRL Research in Plasma-Assisted Combustion

DTIC Science & Technology

2013-10-23

Scramjet propulsion Non-equilibrium flows Diagnostics for scramjet controls  Boundary-layer transition  Structural sciences for...hypersonic vehicles Computational sciences for hypersonic flight 3 DISTRIBUTION STATEMENT A – Unclassified, Unlimited Distribution Overview Research...within My Division HIFiRE-5 Vehicle Launched 23 April 2012 can payload transition section Orion S-30 Focus on hypersonic flight: scalability

CFD validation experiments for hypersonic flows

NASA Technical Reports Server (NTRS)

Marvin, Joseph G.

1992-01-01

A roadmap for CFD code validation is introduced. 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 could provide new validation data.

Experimental aerothermodynamic research of hypersonic aircraft

NASA Technical Reports Server (NTRS)

Cleary, Joseph W.

1987-01-01

The 2-D and 3-D advance computer codes being developed for use in the design of such hypersonic aircraft as the National Aero-Space Plane require comparison of the computational results with a broad spectrum of experimental data to fully assess the validity of the codes. This is particularly true for complex flow fields with control surfaces present and for flows with separation, such as leeside flow. Therefore, the objective is to provide a hypersonic experimental data base required for validation of advanced computational fluid dynamics (CFD) computer codes and for development of more thorough understanding of the flow physics necessary for these codes. This is being done by implementing a comprehensive test program for a generic all-body hypersonic aircraft model in the NASA/Ames 3.5 foot Hypersonic Wind Tunnel over a broad range of test conditions to obtain pertinent surface and flowfield data. Results from the flow visualization portion of the investigation are presented.

Experimental and finite element investigation of the buckling characteristics of a beaded skin panel for a hypersonic aircraft. Ph.D. Thesis. Final Report

NASA Technical Reports Server (NTRS)

Siegel, W. H.

1978-01-01

As part of NASA's continuing research into hypersonics and 85 square foot hypersonic wing test section of a proposed hypersonic research airplane was laboratory tested. The project reported on in this paper has carried the hypersonic wing test structure project one step further by testing a single beaded panel to failure. The primary interest was focused upon the buckling characteristics of the panel under pure compression with boundary conditions similar to those found in a wing mounted condition. Three primary phases of analysis are included in the report. These phases include: experimental testing of the beaded panel to failure; finite element structural analysis of the beaded panel with the computer program NASTRAN; a summary of the semiclassical buckling equations for the beaded panel under purely compressive loads. Comparisons between each of the analysis methods are also included.

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 engine model. HYFIM performs the aerodynamic analysis of forebodies and inlet characteristics of RBCC powered SSTO launch vehicles. HYFIM is applicable to the analysis of the ramjet/scramjet engine operations modes (Mach 3-12), and provides estimates of parameters such as air capture area, shock-on-lip Mach number, design Mach number, compression ratio, etc., based on a basic geometry routine for modeling axisymmetric cones, 2-D wedge geometries. HYFIM also estimates the variation of shock layer properties normal to the forebody surface. The thermal protection system (TPS) is directly linked to determination of the vehicle moldline and the shaping of the trajectory. Thermal protection systems to maintain the structural integrity of the vehicle must be able to mitigate the heat transfer to the structure and be lightweight. Herein lies the interdependency, in that as the vehicle's speed increases, the TPS requirements are increased. And as TPS masses increase the effect on the propulsion system and all other systems is compounded. The need to analyze vehicle forebody and engine inlet is critical to be able to design the RBCC vehicle. To adequately determine insulation masses for an RBCC vehicle, the hypersonic aerodynamic environment and aeroheating loads must be calculated and the TPS thicknesses must be calculated for the entire vehicle. To accomplish this an ascent or reentry trajectory is obtained using the computer code Program to Optimize Simulated Trajectories (POST). The trajectory is then used to calculate the convective heat rates on several locations on the vehicles using the Miniature Version of the JA70 Aerodynamic Heating Computer Program (MINIVER). Once the heat rates are defined for each body point on the vehicle, then insulation thicknesses that are required to maintain the vehicle within structural limits are calculated using Systems Improved Numerical Differencing Analyzer (SINDA) models. If the TPS masses are too heavy for the performance of the vehicle the process may be repeated altering the trajectory or some other input to reduce the TPS mass. E-PSURBCC is an "engine performance" model and requires the specification of inlet air static temperature and pressure as well as Mach number (which it pulls from the HYFIM and POST trajectory files), and calculates the corresponding stagnation properties. The engine air flow path geometry includes inlet, a constant area section where the rocket is positioned, a subsonic diffuser, a constant area afterburner, and either a converging nozzle or a converging-diverging nozzle. The current capabilities of E-PSURBCC ejector and ramjet mode treatment indicated that various complex flow phenomena including multiple choking and internal shocks can occur for combinations of geometry/flow conditions. For a given input deck defining geometry/flow conditions, the program first goes through a series of checks to establish whether the input parameters are sound in terms of a solution path. If the vehicle/engine performance fails mission goals, the engineer is able to collaboratively alter the vehicle moldline to change aerodynamics, or trajectory, or some other input to achieve orbit. The problem described is an example of the need for collaborative design and analysis. RECIPE is a cross-platform application capable of hosting a number of engineers and designers across the Internet for distributed and collaborative engineering environments. Such integrated system design environments allow for collaborative team design analysis for performing individual or reduced team studies. To facilitate the larger number of potential runs that may need to be made, RECIPE connects the computer codes that calculate the trajectory data, aerodynamic data based on vehicle geometry, heat rate data, TPS masses, and vehicle and engine performance, so that the output from each tool is easily transferred to the model input files that need it.

Transition and turbulence measurements in hypersonic flows

NASA Technical Reports Server (NTRS)

Owen, F. K.

1990-01-01

This paper reviews techniques for transitional- and turbulent-flow measurements and describes current research in support of turbulence modeling. Special attention is given to the potential of applying hot wire and laser velocimeter to measuring turbulent fluctuations in hypersonic flow fields. The results of recent experiments conducted in two hypersonic wind tunnels are presented and compared with previous hot-wire turbulence measurements.

A comparison of upwind schemes for computation of three-dimensional hypersonic real-gas flows

NASA Technical Reports Server (NTRS)

Gerbsch, R. A.; Agarwal, R. K.

1992-01-01

The method of Suresh and Liou (1992) is extended, and the resulting explicit noniterative upwind finite-volume algorithm is applied to the integration of 3D parabolized Navier-Stokes equations to model 3D hypersonic real-gas flowfields. The solver is second-order accurate in the marching direction and employs flux-limiters to make the algorithm second-order accurate, with total variation diminishing in the cross-flow direction. The algorithm is used to compute hypersonic flow over a yawed cone and over the Ames All-Body Hypersonic Vehicle. The solutions obtained agree well with other computational results and with experimental data.

Power-on performance predictions for a complete generic hypersonic vehicle configuration

NASA Technical Reports Server (NTRS)

Bennett, Bradford C.

1991-01-01

The Compressible Navier-Stokes (CNS) code was developed to compute external hypersonic flow fields. It has been applied to various hypersonic external flow applications. Here, the CNS code was modified to compute hypersonic internal flow fields. Calculations were performed on a Mach 18 sidewall compression inlet and on the Lewis Mach 5 inlet. The use of the ARC3D diagonal algorithm was evaluated for internal flows on the Mach 5 inlet flow. The initial modifications to the CNS code involved generalization of the boundary conditions and the addition of viscous terms in the second crossflow direction and modifications to the Baldwin-Lomax turbulence model for corner flows.

Tomographic PIV investigation of roughness-induced transition in a hypersonic boundary layer

NASA Astrophysics Data System (ADS)

Avallone, F.; Ye, Q.; Schrijer, F. F. J.; Scarano, F.; Cardone, G.

2014-11-01

The disturbance generated by roughness elements in a hypersonic laminar boundary layer is investigated, with attention to its three-dimensional properties. The transition of the boundary layer is inspected with tomographic particle image velocimetry that is applied for the first time at Mach 7.5 inside a short duration hypersonic wind tunnel. A low aspect ratio cylindrical roughness element is installed on a flat plate, and experiments are conducted downstream of the element describing the mean velocity field and the turbulent fluctuations. Details of the experimental procedure needed to realize these measurements are discussed, along with the fluid dynamic behaviour of the perturbed hypersonic boundary layer.

Aerothermodynamic shape optimization of hypersonic blunt bodies

NASA Astrophysics Data System (ADS)

Eyi, Sinan; Yumuşak, Mine

2015-07-01

The aim of this study is to develop a reliable and efficient design tool that can be used in hypersonic flows. The flow analysis is based on the axisymmetric Euler/Navier-Stokes and finite-rate chemical reaction equations. The equations are coupled simultaneously and solved implicitly using Newton's method. The Jacobian matrix is evaluated analytically. A gradient-based numerical optimization is used. The adjoint method is utilized for sensitivity calculations. The objective of the design is to generate a hypersonic blunt geometry that produces the minimum drag with low aerodynamic heating. Bezier curves are used for geometry parameterization. The performances of the design optimization method are demonstrated for different hypersonic flow conditions.

Experimental and computational surface and flow-field results for an all-body hypersonic aircraft

NASA Technical Reports Server (NTRS)

Lockman, William K.; Lawrence, Scott L.; Cleary, Joseph W.

1990-01-01

The objective of the present investigation is to establish a benchmark experimental data base for a generic hypersonic vehicle shape for validation and/or calibration of advanced computational fluid dynamics computer codes. This paper includes results from the comprehensive test program conducted in the NASA/Ames 3.5-foot Hypersonic Wind Tunnel for a generic all-body hypersonic aircraft model. Experimental and computational results on flow visualization, surface pressures, surface convective heat transfer, and pitot-pressure flow-field surveys are presented. Comparisons of the experimental results with computational results from an upwind parabolized Navier-Stokes code developed at Ames demonstrate the capabilities of this code.

Pegasus hypersonic flight research

NASA Technical Reports Server (NTRS)

Curry, Robert E.; Meyer, Robert R., Jr.; Budd, Gerald D.

1992-01-01

Hypersonic aeronautics research using the Pegasus air-launched space booster is described. Two areas are discussed in the paper: previously obtained results from Pegasus flights 1 and 2, and plans for future programs. Proposed future research includes boundary-layer transition studies on the airplane-like first stage and also use of the complete Pegasus launch system to boost a research vehicle to hypersonic speeds. Pegasus flight 1 and 2 measurements were used to evaluate the results of several analytical aerodynamic design tools applied during the development of the vehicle as well as to develop hypersonic flight-test techniques. These data indicated that the aerodynamic design approach for Pegasus was adequate and showed that acceptable margins were available. Additionally, the correlations provide insight into the capabilities of these analytical tools for more complex vehicles in which design margins may be more stringent. Near-term plans to conduct hypersonic boundary-layer transition studies are discussed. These plans involve the use of a smooth metallic glove at about the mid-span of the wing. Longer-term opportunities are proposed which identify advantages of the Pegasus launch system to boost large-scale research vehicles to the real-gas hypersonic flight regime.

Fluid Mechanics, Drag Reduction and Advanced Configuration Aeronautics

NASA Technical Reports Server (NTRS)

Bushnell, Dennis M.

2000-01-01

This paper discusses Advanced Aircraft configurational approaches across the speed range, which are either enabled, or greatly enhanced, by clever Flow Control. Configurations considered include Channel Wings with circulation control for VTOL (but non-hovering) operation with high cruise speed, strut-braced CTOL transports with wingtip engines and extensive ('natural') laminar flow control, a midwing double fuselage CTOL approach utilizing several synergistic methods for drag-due-to-lift reduction, a supersonic strut-braced configuration with order of twice the L/D of current approaches and a very advanced, highly engine flow-path-integrated hypersonic cruise machine. This paper indicates both the promise of synergistic flow control approaches as enablers for 'Revolutions' in aircraft performance and fluid mechanic 'areas of ignorance' which impede their realization and provide 'target-rich' opportunities for Fluids Research.

Advanced fabrication techniques for hydrogen-cooled engine structures

NASA Technical Reports Server (NTRS)

Buchmann, O. A.; Arefian, V. V.; Warren, H. A.; Vuigner, A. A.; Pohlman, M. J.

1985-01-01

Described is a program for development of coolant passage geometries, material systems, and joining processes that will produce long-life hydrogen-cooled structures for scramjet applications. Tests were performed to establish basic material properties, and samples constructed and evaluated to substantiate fabrication processes and inspection techniques. Results of the study show that the basic goal of increasing the life of hydrogen-cooled structures two orders of magnitude relative to that of the Hypersonic Research Engine can be reached with available means. Estimated life is 19000 cycles for the channels and 16000 cycles for pin-fin coolant passage configurations using Nickel 201. Additional research is required to establish the fatigue characteristics of dissimilar-metal coolant passages (Nickel 201/Inconel 718) and to investigate the embrittling effects of the hydrogen coolant.

Numerical Simulation Of Shock Response To Wall Changes In High Speed Intakes

NASA Astrophysics Data System (ADS)

Fincham, J.; Taylor, N. V.

2011-05-01

Hypersonic flight presents a number of challenges to the designer, one of which is the intake behaviour. Minimising drag requires careful positioning of the intake shock structure, while accurate understanding of the dynamic behaviour is required to allow minimisation of margins. In this paper, a two shock external compression intake derived from the Reaction Engines Limited SABRE engine is examined using inviscid axisymmetric CFD analysis to determine the response of the normal shockwave to axial motion of the intake centrebody. An approximately linear relationship between centrebody position and both the normal shock position and additive drag in steady flow is demonstrated. Initial results from an unsteady analysis are also given, which show complex behaviours may be triggered by rapid motion of the centrebody in response to control input.

Seals Flow Code Development 1993

NASA Technical Reports Server (NTRS)

Liang, Anita D. (Compiler); Hendricks, Robert C. (Compiler)

1994-01-01

Seals Workshop of 1993 code releases include SPIRALI for spiral grooved cylindrical and face seal configurations; IFACE for face seals with pockets, steps, tapers, turbulence, and cavitation; GFACE for gas face seals with 'lift pad' configurations; and SCISEAL, a CFD code for research and design of seals of cylindrical configuration. GUI (graphical user interface) and code usage was discussed with hands on usage of the codes, discussions, comparisons, and industry feedback. Other highlights for the Seals Workshop-93 include environmental and customer driven seal requirements; 'what's coming'; and brush seal developments including flow visualization, numerical analysis, bench testing, T-700 engine testing, tribological pairing and ceramic configurations, and cryogenic and hot gas facility brush seal results. Also discussed are seals for hypersonic engines and dynamic results for spiral groove and smooth annular seals.

Fullerene data mining using bibliometrics and database tomography

PubMed

Kostoff; Braun; Schubert; Toothman; Humenik

2000-01-01

Database tomography (DT) is a textual database analysis system consisting of two major components: (1) algorithms for extracting multiword phrase frequencies and phrase proximities (physical closeness of the multiword technical phrases) from any type of large textual database, to augment (2) interpretative capabilities of the expert human analyst. DT was used to derive technical intelligence from a fullerenes database derived from the Science Citation Index and the Engineering Compendex. Phrase frequency analysis by the technical domain experts provided the pervasive technical themes of the fullerenes database, and phrase proximity analysis provided the relationships among the pervasive technical themes. Bibliometric analysis of the fullerenes literature supplemented the DT results with author/journal/institution publication and citation data. Comparisons of fullerenes results with past analyses of similarly structured near-earth space, chemistry, hypersonic/supersonic flow, aircraft, and ship hydrodynamics databases are made. One important finding is that many of the normalized bibliometric distribution functions are extremely consistent across these diverse technical domains and could reasonably be expected to apply to broader chemical topics than fullerenes that span multiple structural classes. Finally, lessons learned about integrating the technical domain experts with the data mining tools are presented.

State of the art in high-temperature fiber optic sensors

NASA Astrophysics Data System (ADS)

Fielder, Robert S.; Stinson-Bagby, Kelly L.; Palmer, Matthew E.

2004-12-01

The objective of the work presented was to develop a suite of sensors for use in high-temperature aerospace environments, including turbine engine monitoring, hypersonic vehicle skin friction measurements, and support ground and flight test operations. A fiber optic sensor platform was used to construct the sensor suite. Successful laboratory demonstrations include calibration of pressure sensors to 500psi at a gas temperature of 800°C. Additionally, pressure sensors were demonstrated at 800°C in combination with a high-speed (1.0MHz) fiber optic readout system enabling previously unobtainable dynamic measurements at high-temperatures. Temperature sensors have been field tested up to 1400°C and as low as -195°C. The key advancement that enabled the operation of these novel harsh environment sensors was a fiber optic packaging methodology that allowed the coupling of alumina and sapphire transducer components, optical fiber, and high-temperature alloy housing materials. The basic operation of the sensors and early experimental results are presented. Each of the sensors described here represent a quantifiable advancement in the state of the art in high-temperature physical sensors and will have a significant impact on the aerospace propulsion instrumentation industry.

Harsh-environment fiber optic sensors for structural monitoring applications

NASA Astrophysics Data System (ADS)

Fielder, Robert S.; Stinson-Bagby, Kelly L.; Palmer, Matthew E.

2004-07-01

The objective of the work presented was to develop a suite of sensors for use in high-temperature aerospace environments, including turbine engine monitoring, hypersonic vehicle skin friction measurements, and support ground and flight test operations. A fiber optic sensor platform was used to construct the sensor suite. Successful laboratory demonstrations include calibration of a pressure sensor to 100psi at a gas temperature of 800°C, calibration of an accelerometer to 2.5g at a substrate temperature of 850°C. Temperature sensors have been field tested up to 1400°C, and a skin friction sensor designed for 870°C operation has been constructed. The key advancement that enabled the operation of these novel harsh environment sensors was a fiber optic packaging methodology that allowed the coupling of alumina and sapphire transducer components, optical fiber, and high-temperature alloy housing materials. The basic operation of the sensors and early experimental results are presented. Each of the sensors described here represent a quantifiable advancement in the state of the art in high-temperature physical sensors and will have a significant impact on the aerospace propulsion instrumentation industry.

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.

Evaluation and ranking of candidate ceramic wafer engine seal materials

NASA Technical Reports Server (NTRS)

Steinetz, Bruce M.

1991-01-01

Modern engineered ceramics offer high temperature capabilities not found in even the best superalloy metals. The high temperature properties of several selected ceramics including aluminum oxide, silicon carbide, and silicon nitride are reviewed as they apply to hypersonic engine seal design. A ranking procedure is employed to objectively differentiate among four different monolithic ceramic materials considered, including: a cold-pressed and sintered aluminum oxide; a sintered alpha-phase silicon carbide; a hot-isostatically pressed silicon nitride; and a cold-pressed and sintered silicon nitride. This procedure is used to narrow the wide range of potential ceramics considered to an acceptable number for future detailed and costly analyses and tests. The materials are numerically scored according to their high temperature flexural strength; high temperature thermal conductivity; resistance to crack growth; resistance to high heating rates; fracture toughness; Weibull modulus; and finally according to their resistance to leakage flow, where materials having coefficients of thermal expansion closely matching the engine panel material resist leakage flow best. The cold-pressed and sintered material (Kyocera SN-251) ranked the highest in the overall ranking especially when implemented in engine panels made of low expansion rate materials being considered for the engine, including Incoloy and titanium alloys.

Rocketdyne RBCC Engine Concept Development

NASA Technical Reports Server (NTRS)

Ratckin, G.; Goldman, A.; Ortwerth, P.; Weisberg, S.

1999-01-01

Boeing Rocketdyne is pursuing the development of Rocket Based Combined Cycle (RBCC), propulsion systems as demonstrated by significant contract work in the hypersonic arena (ART, NASP, SCT, system studies) and over 12 years of steady company discretionary investment. The Rocketdyne concept is a fixed geometry integrated rocket, ramjet, scramjet which is hydrogen fueled and uses hydrogen regenerative cooling. The baseline engine structural configuration uses an integral structure that eliminates panel seals. seal purge gas, and closeout side attachments. Rocketdyne's experimental RBCC engine (Engine A5) was constructed under contract with the NASA Marshall Space Flight Center. Engine A5 models the complete flight engine flowpath consisting of an inlet, isolator, airbreathing combustor and nozzle. High performance rocket thrusters are integrated into the engine to enable both air-augmented rocket (AAR) and pure rocket operation. Engine A5 was tested in CASL's new FAST facility as an air-augmented rocket, a ramjet and a pure rocket. Measured performance demonstrated vision vehicle performance levels for Mach 3 AAR operation and ramjet operation from Mach 3 to 4. Rocket mode performance was above predictions. For the first time. testing also demonstrated transition from AAR operation to ramjet operation. This baseline configuration has also been shown, in previous testing, to perform well in the scramjet mode.

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.

A CFD validation roadmap for hypersonic flows

NASA Technical Reports Server (NTRS)

Marvin, Joseph G.

1993-01-01

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.

Lift, drag and thrust measurement in a hypersonic impulse facility

NASA Technical Reports Server (NTRS)

Tuttle, S. L.; Mee, D. J.; Simmons, J. M.

1995-01-01

This paper reports the extension of the stress wave force balance to the measurement of forces on models which are non-axisymmetric or which have non-axisymmetric load distributions. Recent results are presented which demonstrate the performance of the stress wave force balance for drag measurement, for three-component force measurement and preliminary results for thrust measurement on a two-dimensional scramjet nozzle. In all cases, the balances respond within a few hundred microseconds.

Aerothermodynamic flow phenomena of the airframe-integrated supersonic combustion ramjet

NASA Technical Reports Server (NTRS)

Walton, James T.

1992-01-01

The unique component flow phenomena is discussed of the airframe-integrated supersonic combustion ramjet (scramjet) in a format geared towards new players in the arena of hypersonic propulsion. After giving an overview of the scramjet aerothermodynamic cycle, the characteristics are then covered individually of the vehicle forebody, inlet, combustor, and vehicle afterbody/nozzle. Attention is given to phenomena such as inlet speeding, inlet starting, inlet spillage, fuel injection, thermal choking, and combustor-inlet interaction.

Results of tests of a 0.010- and 0.015-scale models of space shuttle orbiter configurations 3 and 3A in the Ames Research Center 3.5 foot hypersonic wind tunnel (OA23)

NASA Technical Reports Server (NTRS)

Dziubala, T. J.; Milam, M. D.; Cleary, J. W.; Mellenthin, J. A.

1974-01-01

Longitudinal and lateral-directional stability and control characteristics were evaluated at Mach numbers of 5.3, 7.3 and 10.3 at angles of attack up to 50 degrees with Beta = 0 degrees and, for a few cases, Beta = 5 degrees. Component force data, fuselage base pressures and shadowgraph patterns were recorded.

Experimental Study of Shock Wave Interference Heating on a Cylindrical Leading Edge. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Wieting, Allan R.

1987-01-01

An experimental study of shock wave interference heating on a cylindrical leading edge representative of the cowl of a rectangular hypersonic engine inlet at Mach numbers of 6.3, 6.5, and 8.0 is presented. Stream Reynolds numbers ranged from 0.5 x 106 to 4.9 x 106 per ft. and stream total temperature ranged from 2100 to 3400 R. The model consisted of a 3" dia. cylinder and a shock generation wedge articulated to angles of 10, 12.5, and 15 deg. A fundamental understanding was obtained of the fluid mechanics of shock wave interference induced flow impingement on a cylindrical leading edge and the attendant surface pressure and heat flux distributions. The first detailed heat transfer rate and pressure distributions for two dimensional shock wave interference on a cylinder was provided along with insight into the effects of specific heat variation with temperature on the phenomena. Results show that the flow around a body in hypersonic flow is altered significantly by the shock wave interference pattern that is created by an oblique shock wave from an external source intersecting the bow shock wave produced in front of the body.

Mission and Design Sensitivities for Human Mars Landers Using Hypersonic Inflatable Aerodynamic Decelerators

NASA Technical Reports Server (NTRS)

Polsgrove, Tara P.; Thomas, Herbert D.; Collins, Tim; Dwyer Cianciolo, Alicia; Samareh, Jamshid

2017-01-01

Landing humans on Mars is one of NASA's long term goals. The Evolvable Mars Campaign (EMC) is focused on evaluating architectural trade options to define the capabilities and elements needed for a sustainable human presence on the surface of Mars. The EMC study teams have considered a variety of in-space propulsion options and surface mission options. As we seek to better understand how these choices affect the performance of the lander, this work informs and influences requirements for transportation systems to deliver the landers to Mars and enable these missions. This paper presents the effects of mission and vehicle design options on lander mass and performance. Beginning with Earth launch, options include fairing size assumptions, co-manifesting other elements with the lander, and Earth-Moon vicinity operations. Capturing into Mars orbit using either aerocapture or propulsive capture is assessed. For entry, descent, and landing both storable as well as oxygen and methane propellant combinations are considered, engine thrust level is assessed, and sensitivity to landed payload mass is presented. This paper focuses on lander designs using the Hypersonic Inflatable Aerodynamic Decelerators (HIAD), one of several entry system technologies currently considered for human missions.

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. Both facility and engine starting were achieved. Further, the static pressure distributions compared well with the results previously obtained in a 40% subscale flowpath study conducted in the LERC 1X1 supersonic wind tunnel (SWT), as well as the results of CFD analysis. Fueled performance results were obtained for the engine at both simulated Mach 6 (1670 K) and Mach 6.6 (1950 K) conditions. For all these tests the primary fuel was liquid JP-10 with gaseous silane (a mixture of 20% SiH4 and 80% H2 by volume) as an ignitor/pilot. These tests verified performance of this engine flowpath in a freejet mode. High combustor pressures were reached and significant changes in axial force were achieved due to combustion. Future test plans include redistributing the fuel to improve mixing, and consequently performance, at higher equivalence ratios.

Thermal control/oxidation resistant coatings for titanium-based alloys

NASA Technical Reports Server (NTRS)

Clark, Ronald K.; Wallace, Terryl A.; Cunnington, George R.; Wiedemann, Karl E.

1992-01-01

Extensive research and development efforts have been expended toward development of thermal control and environmental protection coatings for NASP and generic hypersonic vehicle applications. The objective of the coatings development activities summarized here was to develop light-weight coatings for protecting advanced titanium alloys from oxidation in hypersonic vehicle applications. A number of new coating concepts have been evaluated. Coated samples were exposed to static oxidation tests at temperatures up to 1000 C using a thermogravimetric apparatus. Samples were also exposed to simulated hypersonic flight conditions for up to 10 hr to determine their thermal and chemical stability and catalytic efficiency. The emittance of samples was determined before and after exposure to simulated hypersonic flight conditions.

Acoustic phonon dispersion at hypersonic frequencies in Si and Ge

NASA Astrophysics Data System (ADS)

Kuok, M. H.; Ng, S. C.; Rang, Z. L.; Lockwood, D. J.

2000-11-01

Brillouin spectra of the longitudinal acoustic (LA) mode, traveling along the [001] direction, in silicon and germanium have been recorded in 180° backscattering using 457.9-514.5-nm laser lines. The wave velocity of the LA phonon propagating in the [001] direction was determined at hypersonic frequencies, from the measured acoustic phonon dispersion in silicon and germanium. The elastic modulus c11 of the two semiconductors has been calculated from the respective measured hypersonic wave velocities and the results are compared with values determined from lower-frequency ultrasonic and other measurements. Interestingly, the hypersonic velocities are consistently lower by ~1-2 % than the ultrasonic ones, but they generally agree within the present experimental accuracy.

N-S/DSMC hybrid simulation of hypersonic flow over blunt body including wakes

NASA Astrophysics Data System (ADS)

Li, Zhonghua; Li, Zhihui; Li, Haiyan; Yang, Yanguang; Jiang, Xinyu

2014-12-01

A hybrid N-S/DSMC method is presented and applied to solve the three-dimensional hypersonic transitional flows by employing the MPC (modular Particle-Continuum) technique based on the N-S and the DSMC method. A sub-relax technique is adopted to deal with information transfer between the N-S and the DSMC. The hypersonic flows over a 70-deg spherically blunted cone under different Kn numbers are simulated using the CFD, DSMC and hybrid N-S/DSMC method. The present computations are found in good agreement with DSMC and experimental results. The present method provides an efficient way to predict the hypersonic aerodynamics in near-continuum transitional flow regime.

Optimization of the Upper Surface of Hypersonic Vehicle Based on CFD Analysis

NASA Astrophysics Data System (ADS)

Gao, T. Y.; Cui, K.; Hu, S. C.; Wang, X. P.; Yang, G. W.

2011-09-01

For the hypersonic vehicle, the aerodynamic performance becomes more intensive. Therefore, it is a significant event to optimize the shape of the hypersonic vehicle to achieve the project demands. It is a key technology to promote the performance of the hypersonic vehicle with the method of shape optimization. Based on the existing vehicle, the optimization to the upper surface of the Simplified hypersonic vehicle was done to obtain a shape which suits the project demand. At the cruising condition, the upper surface was parameterized with the B-Spline curve method. The incremental parametric method and the reconstruction technology of the local mesh were applied here. The whole flow field was been calculated and the aerodynamic performance of the craft were obtained by the computational fluid dynamic (CFD) technology. Then the vehicle shape was optimized to achieve the maximum lift-drag ratio at attack angle 3°, 4° and 5°. The results will provide the reference for the practical design.

Trajectory Approaches for Launching Hypersonic Flight Tests (Preprint)

DTIC Science & Technology

2014-08-01

This paper presents some approaches toward designing trajectories for hypersonic testing at up to Mach 10 speed using a reusable rocket -powered first...Program to Optimize Simulated Trajectories (POST) code to look at different ways of flying to Mach 10 with a reusable first stage rocket . These trajectories...are good starting points for how to setup a trajectory simulation to meet hypersonic testing needs. 15. SUBJECT TERMS responsive and reusable rocket

Computation of hypersonic flows with finite rate condensation and evaporation of water

NASA Technical Reports Server (NTRS)

Perrell, Eric R.; Candler, Graham V.; Erickson, Wayne D.; Wieting, Alan R.

1993-01-01

A computer program for modelling 2D hypersonic flows of gases containing water vapor and liquid water droplets is presented. The effects of interphase mass, momentum and energy transfer are studied. Computations are compared with existing quasi-1D calculations on the nozzle of the NASA Langley Eight Foot High Temperature Tunnel, a hypersonic wind tunnel driven by combustion of natural gas in oxygen enriched air.

A Combined CFD/Characteristic Method for Prediction and Design of Hypersonic Inlet with Nose Bluntness

NASA Astrophysics Data System (ADS)

Gao, Wenzhi; Li, Zhufei; Yang, Jiming

Leading edge bluntness is widely used in hypersonic inlet design for thermal protection[1]. Detailed research of leading edge bluntness on hypersonic inlet has been concentrated on shock shape correlation[2], boundary layer flow[3], inlet performance[4], etc. It is well known that blunted noses cause detached bow shocks which generate subsonic regions around the noses and entropy layers in the flowfield.

Calculation of Supersonic Combustion Using Implicit Schemes

NASA Technical Reports Server (NTRS)

Yoon, Seokkwan; Kwak, Dochan (Technical Monitor)

2003-01-01

One of the technology goals of NASA for advanced space transportation is to develop highly efficient propulsion systems to reduce the cost of payload for space missions. Developments of rockets for the second generation Reusable Launch Vehicle (RLV) in the past several years have been focused on low-cost versions of conventional engines. However, recent changes in the Integrated Space Transportation Program to build a crew transportation vehicle to extend the life of the Space Shuttle fleet might suggest that air-breathing rockets could reemerge as a possible propulsion system for the third generation RLV to replace the Space Shuttle after 2015. The weight of the oxygen tank exceeds thirty percent of the total weight of the Space Shuttle at launch while the payload is only one percent of the total weight. The air-breathing rocket propulsion systems, which consume oxygen in the air, offer clear advantages by making vehicles lighter and more efficient. Experience in the National Aerospace Plane Program in the late 1980s indicates that scramjet engines can achieve high specific impulse for low hypersonic vehicle speeds. Whether taking a form of Rocket Based Combined Cycle (RBCC) or Turbine Based Combined Cycle (TBCC), the scramjet is an essential mode of operation for air-breathing rockets. It is well known that fuel-air mixing and rapid combustion are of crucial importance for the success of scramjet engines since the spreading rate of the supersonic mixing layer decreases as the Mach number increases. A factored form of the Gauss-Seidel relaxation method has been widely used in hypersonic flow research since its first application to non-equilibrium flows. However, difficulties in stability and convergence have been encountered when there is strong interaction between fluid motion and chemical reaction, such as multiple fuel injection problems. The present paper reports the results from investigation of the effect of modifications to the original algorithm on the performance for multiple injectors.

New Air-Launched Small Missile (ALSM) Flight Testbed for Hypersonic Systems

NASA Technical Reports Server (NTRS)

Bui, Trong T.; Lux, David P.; Stenger, Mike; Munson, Mike; Teate, George

2006-01-01

A new testbed for hypersonic flight research is proposed. Known as the Phoenix air-launched small missile (ALSM) flight testbed, it was conceived to help address the lack of quick-turnaround and cost-effective hypersonic flight research capabilities. The Phoenix ALSM testbed results from utilization of two unique and very capable flight assets: the United States Navy Phoenix AIM-54 long-range, guided air-to-air missile and the NASA Dryden F-15B testbed airplane. The U.S. Navy retirement of the Phoenix AIM-54 missiles from fleet operation has presented an excellent opportunity for converting this valuable flight asset into a new flight testbed. This cost-effective new platform will fill an existing gap in the test and evaluation of current and future hypersonic systems for flight Mach numbers ranging from 3 to 5. Preliminary studies indicate that the Phoenix missile is a highly capable platform. When launched from a high-performance airplane, the guided Phoenix missile can boost research payloads to low hypersonic Mach numbers, enabling flight research in the supersonic-to-hypersonic transitional flight envelope. Experience gained from developing and operating the Phoenix ALSM testbed will be valuable for the development and operation of future higher-performance ALSM flight testbeds as well as responsive microsatellite small-payload air-launched space boosters.

Numerical simulation for the influence of laser-induced plasmas addition on air mass capture of hypersonic inlet

NASA Astrophysics Data System (ADS)

Zhao, Wei; Dou, Zhiguo; Li, Qian

2012-03-01

The theory of laser-induced plasmas addition to hypersonic airflow off a vehicle to increase air mass capture and improve the performance of hypersonic inlets at Mach numbers below the design value is explored. For hypersonic vehicles, when flying at mach numbers lower than the design one, we can increase the mass capture ratio of inlet through laser-induced plasmas injection to the hypersonic flow upstream of cowl lip to form a virtual cowl. Based on the theory, the model of interaction between laser-induced plasmas and hypersonic flow was established. The influence on the effect of increasing mass capture ratio was studied at different positions of laser-induced plasmas region for the external compression hypersonic inlet at Mach 5 while the design value is 6, the power of plasmas was in the range of 1-8mJ. The main results are as follows: 1. the best location of the plasma addition region is near the intersection of the nose shock of the vehicle with the continuation of the cowl line, and slightly below that line. In that case, the shock generated by the heating is close to the shock that is a reflection of the vehicle nose shock off the imaginary solid surface-extension of the cowl. 2. Plasma addition does increase mass capture, and the effect becomes stronger as more energy is added, the peak value appeared when the power of plasma was about 4mJ, when the plasma energy continues to get stronger, the mass capture will decline slowly.

Numerical simulation for the influence of laser-induced plasmas addition on air mass capture of hypersonic inlet

NASA Astrophysics Data System (ADS)

Zhao, Wei; Dou, Zhiguo; Li, Qian

2011-11-01

The theory of laser-induced plasmas addition to hypersonic airflow off a vehicle to increase air mass capture and improve the performance of hypersonic inlets at Mach numbers below the design value is explored. For hypersonic vehicles, when flying at mach numbers lower than the design one, we can increase the mass capture ratio of inlet through laser-induced plasmas injection to the hypersonic flow upstream of cowl lip to form a virtual cowl. Based on the theory, the model of interaction between laser-induced plasmas and hypersonic flow was established. The influence on the effect of increasing mass capture ratio was studied at different positions of laser-induced plasmas region for the external compression hypersonic inlet at Mach 5 while the design value is 6, the power of plasmas was in the range of 1-8mJ. The main results are as follows: 1. the best location of the plasma addition region is near the intersection of the nose shock of the vehicle with the continuation of the cowl line, and slightly below that line. In that case, the shock generated by the heating is close to the shock that is a reflection of the vehicle nose shock off the imaginary solid surface-extension of the cowl. 2. Plasma addition does increase mass capture, and the effect becomes stronger as more energy is added, the peak value appeared when the power of plasma was about 4mJ, when the plasma energy continues to get stronger, the mass capture will decline slowly.

Optical Measurements at the Combustor Exit of the HIFiRE 2 Ground Test Engine

NASA Technical Reports Server (NTRS)

Brown, Michael S.; Herring, Gregory C.; Cabell, Karen; Hass, Neal; Barhorst, Todd F.; Gruber, Mark

2012-01-01

The development of optical techniques capable of measuring in-stream flow properties of air breathing hypersonic engines is a goal of the Aerospace Propulsion Division at AFRL. Of particular interest are techniques such as tunable diode laser absorption spectroscopy that can be implemented in both ground and flight test efforts. We recently executed a measurement campaign at the exit of the combustor of the HIFiRE 2 ground test engine during Phase II operation of the engine. Data was collected in anticipation of similar data sets to be collected during the flight experiment. The ground test optical data provides a means to evaluate signal processing algorithms particularly those associated with limited line of sight tomography. Equally important, this in-stream data was collected to compliment data acquired with surface-mounted instrumentation and the accompanying flowpath modeling efforts-both CFD and lower order modeling. Here we discuss the specifics of hardware and data collection along with a coarse-grained look at the acquired data and our approach to processing and analyzing it.

Thrust augmentation nozzle (TAN) concept for rocket engine booster applications

NASA Astrophysics Data System (ADS)

Forde, Scott; Bulman, Mel; Neill, Todd

2006-07-01

Aerojet used the patented thrust augmented nozzle (TAN) concept to validate a unique means of increasing sea-level thrust in a liquid rocket booster engine. We have used knowledge gained from hypersonic Scramjet research to inject propellants into the supersonic region of the rocket engine nozzle to significantly increase sea-level thrust without significantly impacting specific impulse. The TAN concept overcomes conventional engine limitations by injecting propellants and combusting in an annular region in the divergent section of the nozzle. This injection of propellants at moderate pressures allows for obtaining high thrust at takeoff without overexpansion thrust losses. The main chamber is operated at a constant pressure while maintaining a constant head rise and flow rate of the main propellant pumps. Recent hot-fire tests have validated the design approach and thrust augmentation ratios. Calculations of nozzle performance and wall pressures were made using computational fluid dynamics analyses with and without thrust augmentation flow, resulting in good agreement between calculated and measured quantities including augmentation thrust. This paper describes the TAN concept, the test setup, test results, and calculation results.

The X-43A Hyper-X Mach 7 Flight 2 Guidance, Navigation, and Control Overview and Flight Test Results

NASA Technical Reports Server (NTRS)

Bahm, Catherine; Baumann, Ethan; Martin, John; Bose, David; Beck, Roger E.; Strovers, Brian

2005-01-01

The objective of the Hyper-X program was to flight demonstrate an airframe-integrated hypersonic vehicle. On March 27, 2004, the Hyper-X program team successfully conducted flight 2 and achieved all of the research objectives. The Hyper-X research vehicle successfully separated from the Hyper-X launch vehicle and achieved the desired engine test conditions before the experiment began. The research vehicle rejected the disturbances caused by the cowl door opening and the fuel turning on and off and maintained the engine test conditions throughout the experiment. After the engine test was complete, the vehicle recovered and descended along a trajectory while performing research maneuvers. The last data acquired showed that the vehicle maintained control to the water. This report will provide an overview of the research vehicle guidance and control systems and the performance of the vehicle during the separation event and engine test. The research maneuvers were performed to collect data for aerodynamics and flight controls research. This report also will provide an overview of the flight controls related research and results.

Engineering Aerothermal Analysis for X-34 Thermal Protection System Design

NASA Technical Reports Server (NTRS)

Wurster, Kathryn E.; Riley, Christopher J.; Zoby, E. Vincent

1998-01-01

Design of the thermal protection system for any hypersonic flight vehicle requires determination of both the peak temperatures over the surface and the heating-rate history along the flight profile. In this paper, the process used to generate the aerothermal environments required for the X-34 Testbed Technology Demonstrator thermal protection system design is described as it has evolved from a relatively simplistic approach based on engineering methods applied to critical areas to one of detailed analyses over the entire vehicle. A brief description of the trajectory development leading to the selection of the thermal protection system design trajectory is included. Comparisons of engineering heating predictions with wind-tunnel test data and with results obtained using a Navier-Stokes flowfield code and an inviscid/boundary layer method are shown. Good agreement is demonstrated among all these methods for both the ground-test condition and the peak heating flight condition. Finally, the detailed analysis using engineering methods to interpolate the surface-heating-rate results from the inviscid/boundary layer method to predict the required thermal environments is described and results presented.

Engineering Aerothermal Analysis for X-34 Thermal Protection System Design

NASA Technical Reports Server (NTRS)

Wurster, Kathryn E.; Riley, Christopher J.; Zoby, E. Vincent

1998-01-01

Design of the thermal protection system for any hypersonic flight vehicle requires determination of both the peak temperatures over the surface and the heating-rate history along the flight profile. In this paper, the process used to generate the aerothermal environments required for the X-34 Testbed Technology Demonstrator thermal protection system design is described as it has evolved from a relatively simplistic approach based on engineering methods applied to critical areas to one of detailed analyses over the entire vehicle. A brief description of the trajectory development leading to the selection of the thermal protection system design trajectory is included. Comparisons of engineering heating predictions with wind-tunnel test data and with results obtained using a Navier- Stokes flowfield code and an inviscid/boundary layer method are shown. Good agreement is demonstrated among all these methods for both the ground-test condition and the peak heating flight condition. Finally, the detailed analysis using engineering methods to interpolate the surface-heating-rate results from the inviscid/boundary layer method to predict the required thermal environments is described and results presented.

CFD transient simulation of an isolator shock train in a scramjet engine

NASA Astrophysics Data System (ADS)

Hoeger, Troy Christopher

For hypersonic flight, the scramjet engine uses an isolator to contain the pre-combustion shock train formed by the pressure difference between the inlet and the combustion chamber. If this shock train were to reach the inlet, it would cause an engine unstart, disrupting the flow through the engine and leading to a loss of thrust and potential loss of the vehicle. Prior to this work, a Computational Fluid Dynamics (CFD) simulation of the isolator was needed for simulating and characterizing the isolator flow and for finding the relationship between back pressure and changes in the location of the leading edge of the shock train. In this work, the VULCAN code was employed with back pressure as an input to obtain the time history of the shock train leading location. Results were obtained for both transient and steady-state conditions. The simulation showed a relationship between back-to-inlet pressure ratios and final locations of the shock train. For the 2-D runs, locations were within one isolator duct height of experimental results while for 3-D runs, the results were within two isolator duct heights.

Roles of Engineering Correlations in Hypersonic Entry Boundary Layer Transition Prediction

NASA Technical Reports Server (NTRS)

Campbell, Charles H.; Anderson, Brian P.; King, Rudolph A.; Kegerise, Michael A.; Berry, Scott A.; Horvath, Thomas J.

2010-01-01

Efforts to design and operate hypersonic entry vehicles are constrained by many considerations that involve all aspects of an entry vehicle system. One of the more significant physical phenomenon that affect entry trajectory and thermal protection system design is the occurrence of boundary layer transition from a laminar to turbulent state. During the Space Shuttle Return To Flight activity following the loss of Columbia and her crew of seven, NASA's entry aerothermodynamics community implemented an engineering correlation based framework for the prediction of boundary layer transition on the Orbiter. The methodology for this implementation relies upon similar correlation techniques that have been is use for several decades. What makes the Orbiter boundary layer transition correlation implementation unique is that a statistically significant data set was acquired in multiple ground test facilities, flight data exists to assist in establishing a better correlation and the framework was founded upon state of the art chemical nonequilibrium Navier Stokes flow field simulations. Recent entry flight testing performed with the Orbiter Discovery now provides a means to validate this engineering correlation approach to higher confidence. These results only serve to reinforce the essential role that engineering correlations currently exercise in the design and operation of entry vehicles. The framework of information related to the Orbiter empirical boundary layer transition prediction capability will be utilized to establish a fresh perspective on this role, and to discuss the characteristics which are desirable in a next generation advancement. The details of the paper will review the experimental facilities and techniques that were utilized to perform the implementation of the Orbiter RTF BLT Vsn 2 prediction capability. Statistically significant results for multiple engineering correlations from a ground testing campaign will be reviewed in order to describe why only certain correlations were selected for complete implementation to support the Shuttle Program. Historical Orbiter flight data on early boundary layer transition due to protruding gap fillers will be described in relation to the selected empirical correlations. In addition, Orbiter entry flight testing results from the BLT Flight Experiment will be discussed in relation to these correlations. Applicability of such correlations to the entry design problem will be reviewed, and finally a perspective on the desirable characteristics for a next generation capability based on high fidelity physical models will be provided.

A Gas-Surface Interaction Model based on Accelerated Reactive Molecular Dynamics for Hypersonic Conditions including Thermal Conduction

DTIC Science & Technology

2012-02-28

Interaction Model based on Accelerated Reactive Molecular Dynamics for Hypersonic conditions including Thermal Conduction FA9550-09-1-0157 Schwartzentruber...Dynamics for Hypersonic Conditions including Thermal Conduction Grant/Contract Number: FA9550-09-1-0157 Program Manager: Dr. John Schmisseur PI...through the boundary layer and may chemically react with the vehicle’s thermal protection system (TPS). Many TPS materials act as a catalyst for the

On two special values of temperature factor in hypersonic flow stagnation point

NASA Astrophysics Data System (ADS)

Bilchenko, G. G.; Bilchenko, N. G.

2018-03-01

The hypersonic aircraft permeable cylindrical and spherical surfaces laminar boundary layer heat and mass transfer control mathematical model properties are investigated. The nonlinear algebraic equations systems are obtained for two special values of temperature factor in the hypersonic flow stagnation point. The mappings bijectivity between heat and mass transfer local parameters and controls is established. The computation experiments results are presented: the domains of allowed values “heat-friction” are obtained.

Hypersonic research at Stanford University

NASA Technical Reports Server (NTRS)

Candler, Graham; Maccormack, Robert

1988-01-01

The status of the hypersonic research program at Stanford University is discussed and recent results are highlighted. The main areas of interest in the program are the numerical simulation of radiating, reacting and thermally excited flows, the investigation and numerical solution of hypersonic shock wave physics, the extension of the continuum fluid dynamic equations to the transition regime between continuum and free-molecule flow, and the development of novel numerical algorithms for efficient particulate simulations of flowfields.

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

NASA Astrophysics Data System (ADS)

Mehta, U.

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.

Supersonic Combustion in Air-Breathing Propulsion Systems for Hypersonic Flight

NASA Astrophysics Data System (ADS)

Urzay, Javier

2018-01-01

Great efforts have been dedicated during the last decades to the research and development of hypersonic aircrafts that can fly at several times the speed of sound. These aerospace vehicles have revolutionary applications in national security as advanced hypersonic weapons, in space exploration as reusable stages for access to low Earth orbit, and in commercial aviation as fast long-range methods for air transportation of passengers around the globe. This review addresses the topic of supersonic combustion, which represents the central physical process that enables scramjet hypersonic propulsion systems to accelerate aircrafts to ultra-high speeds. The description focuses on recent experimental flights and ground-based research programs and highlights associated fundamental flow physics, subgrid-scale model development, and full-system numerical simulations.

Fusion Propulsion and Power for Future Flight

NASA Technical Reports Server (NTRS)

Froning, H. D., Jr.

1996-01-01

There are innovative magnetic and electric confinement fusion power and propulsion system designs with potential for: vacuum specific impulses of 1500-2000 seconds with rocket engine thrust/mass ratios of 5-10 g's; environmentally favorable exhaust emissions if aneutronic fusion propellants can be used; a 2 to 3-fold reduction in the mass of hypersonic airliners and SSTO aerospace planes; a 10 to 20 fold reduction in Mars expedition mass and cost (if propellant from planetary atmospheres is used); and feasibility or in-feasibility of these systems could be confirmed with a modest applied research and exploratory development cost.

Fabrication Materials for a Closed Cycle Brayton Turbine Wheel

NASA Technical Reports Server (NTRS)

Khandelwal, Suresh; Hah, Chunill; Powers, Lynn M.; Stewart, Mark E.; Suresh, Ambady; Owen, Albert K.

2006-01-01

A multidisciplinary analysis of a radial inflow turbine rotor is presented. This work couples high-fidelity fluid, structural, and thermal simulations in a seamless multidisciplinary analysis to investigate the consequences of material selection. This analysis extends multidisciplinary techniques previously demonstrated on rocket turbopumps and hypersonic engines. Since no design information is available for the anticipated Brayton rotating machinery, an existing rotor design (the Brayton Rotating Unit (BRU)) was used in the analysis. Steady state analysis results of a notional turbine rotor indicate that stress levels are easily manageable at the turbine inlet temperature, and stress levels anticipated using either superalloys or ceramics.

Aerodynamic Characteristics, Database Development and Flight Simulation of the X-34 Vehicle

NASA Technical Reports Server (NTRS)

Pamadi, Bandu N.; Brauckmann, Gregory J.; Ruth, Michael J.; Fuhrmann, Henri D.

2000-01-01

An overview of the aerodynamic characteristics, development of the preflight aerodynamic database and flight simulation of the NASA/Orbital X-34 vehicle is presented in this paper. To develop the aerodynamic database, wind tunnel tests from subsonic to hypersonic Mach numbers including ground effect tests at low subsonic speeds were conducted in various facilities at the NASA Langley Research Center. Where wind tunnel test data was not available, engineering level analysis is used to fill the gaps in the database. Using this aerodynamic data, simulations have been performed for typical design reference missions of the X-34 vehicle.

Nose-to-tail analysis of an airbreathing hypersonic vehicle using an in-house simplified tool

NASA Astrophysics Data System (ADS)

Piscitelli, Filomena; Cutrone, Luigi; Pezzella, Giuseppe; Roncioni, Pietro; Marini, Marco

2017-07-01

SPREAD (Scramjet PREliminary Aerothermodynamic Design) is a simplified, in-house method developed by CIRA (Italian Aerospace Research Centre), able to provide a preliminary estimation of the performance of engine/aeroshape for airbreathing configurations. It is especially useful for scramjet engines, for which the strong coupling between the aerothermodynamic (external) and propulsive (internal) flow fields requires real-time screening of several engine/aeroshape configurations and the identification of the most promising one/s with respect to user-defined constraints and requirements. The outcome of this tool defines the base-line configuration for further design analyses with more accurate tools, e.g., CFD simulations and wind tunnel testing. SPREAD tool has been used to perform the nose-to-tail analysis of the LAPCAT-II Mach 8 MR2.4 vehicle configuration. The numerical results demonstrate SPREAD capability to quickly predict reliable values of aero-propulsive balance (i.e., net-thrust) and aerodynamic efficiency in a pre-design phase.

National Aeronautics and Space Administration (NASA)/american Society for Engineering Education (ASEE) Summer Faculty Fellowship Program, 1991, Volume 2

NASA Technical Reports Server (NTRS)

Hyman, William A. (Editor); Goldstein, Stanley H. (Editor)

1991-01-01

The objectives of the program are: (1) to further the professional knowledge of qualified engineering and science faculty members; (2) to stimulate an exchange of ideas between participants and NASA; (3) to enrich and refresh the research and teaching activities of participant's institutions; and (4) to contribute to the research objectives of the NASA Centers. A compilation of the final reports on the research projects done by the faculty fellows during the summer of 1991 are presented. Some of the topics covered include: collision avoidance for rover vehicles, bioinstrumentation, neural nets, total quality management of flexible space structures, project scheduling, nondestructive tests, orthostatic intolerance to bedrest, hypersonic reentry simulation, measuring human energy expenditure, tribological models, trace element movement in Anarctic ice, gastrointestinal function, and computer assisted instruction.

Integrated Multidisciplinary Optimization Objects

NASA Technical Reports Server (NTRS)

Alston, Katherine

2014-01-01

OpenMDAO is an open-source MDAO framework. It is used to develop an integrated analysis and design environment for engineering challenges. This Phase II project integrated additional modules and design tools into OpenMDAO to perform discipline-specific analysis across multiple flight regimes at varying levels of fidelity. It also showcased a refined system architecture that allows the system to be less customized to a specific configuration (i.e., system and configuration separation). By delivering a capable and validated MDAO system along with a set of example applications to be used as a template for future users, this work greatly expands NASA's high-fidelity, physics-based MDAO capabilities and enables the design of revolutionary vehicles in a cost-effective manner. This proposed work complements M4 Engineering's expertise in developing modeling and simulation toolsets that solve relevant subsonic, supersonic, and hypersonic demonstration applications.

Computation of Hypersonic Shock Wave Flows of Multi-Component Reactive Gas Mixtures Using the Generalized Boltzmann Equation

DTIC Science & Technology

2009-03-27

ones like the Lennard - Jones potential with established parameters for each gas (e.g. N2 and 02), and for inelastic collisions DSMC method employs...solution of the collision integral. Lennard - Jones potential with two free parameters is used to obtain the elastic cross-section of the gas molecules...and the so called "combinatory relations" are used to obtain parameters of Lennard - Jones potential for an interaction of molecule A with molecule B

Hypersonic Wind-Tunnel Measurements of Boundary-Layer Pressure Fluctuations

DTIC Science & Technology

2009-08-01

experiments. The third and fourth sections of the cone are designed to hold the instrumentation. The model can be run as a 0.102-m base-diameter cone...using the third section only, or 28 Figure 3.10. Glow-perturber section the fourth section can be added to increase the cone base diameter to 0.127 m...the second sensor. The third sensor shows an increase in frequency components above 15 kHz as well as a rise in lower frequencies. As transition

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.

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.

Experimental Study of Hypersonic Wing/Fin Root Heating at Mach 8

DTIC Science & Technology

2012-12-27

at 700%, 200% and 60% for 45?, 55? and 65? of sweep respectively. 15. SUBJECT TERMS Hypersonics, Hypersonic Test Facilities , Shock Tunnels , wing...consisting of a flat plate and a cylinder with an adjustable sweep angle. The tests were conducted in the T4 shock tunnel at conditions simulating Mach 8...root experiment began with an assessment of design considerations for the experiment and the parameters of the T4 shock tunnel facilities. A CAD

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.

Phoenix Missile Hypersonic Testbed (PMHT): System Concept Overview

NASA Technical Reports Server (NTRS)

Jones, Thomas P.

2007-01-01

A viewgraph presentation of the Phoenix Missile Hypersonic Testbed (PMHT) is shown. The contents include: 1) Need and Goals; 2) Phoenix Missile Hypersonic Testbed; 3) PMHT Concept; 4) Development Objectives; 5) Possible Research Payloads; 6) Possible Research Program Participants; 7) PMHT Configuration; 8) AIM-54 Internal Hardware Schematic; 9) PMHT Configuration; 10) New Guidance and Armament Section Profiles; 11) Nomenclature; 12) PMHT Stack; 13) Systems Concept; 14) PMHT Preflight Activities; 15) Notional Ground Path; and 16) Sample Theoretical Trajectories.

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.

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.

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.

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.

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.

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.

Second-mode control in hypersonic boundary layers over assigned complex wall impedance

NASA Astrophysics Data System (ADS)

Sousa, Victor; Patel, Danish; Chapelier, Jean-Baptiste; Scalo, Carlo

2017-11-01

The durability and aerodynamic performance of hypersonic vehicles greatly relies on the ability to delay transition to turbulence. Passive aerodynamic flow control devices such as porous acoustic absorbers are a very attractive means to damp ultrasonic second-mode waves, which govern transition in hypersonic boundary layers under idealized flow conditions (smooth walls, slender geometries, small angles of attack). The talk will discuss numerical simulations modeling such absorbers via the time-domain impedance boundary condition (TD-IBC) approach by Scalo et al. in a hypersonic boundary layer flow over a 7-degree wedge at freestream Mach numbers M∞ = 7.3 and Reynolds numbers Rem = 1.46 .106 . A three-parameter impedance model tuned to the second-mode waves is tested first with varying resistance, R, and damping ratio, ζ, revealing complete mode attenuation for R < 20. A realistic IBC is then employed, derived via an inverse Helmholtz solver analysis of an ultrasonically absorbing carbon-fiber-reinforced carbon ceramic sample used in recent hypersonic transition experiments by Dr. Wagner and co-workers at DLR-Göttingen.

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.

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.

Aeroelastic Analysis of a Trimmed Generic Hypersonic Vehicle

NASA Technical Reports Server (NTRS)

Nydick, I.; Friedmann, P. P.

1999-01-01

The aeroelastic equations of motion governing a hypersonic vehicle in free flight are derived. The equations of motion for a translating and rotating flexible body using Lagrange's equations in terms of quasi-coordinates are presented. These equations are simplified for the case of a vehicle with pitch and plunge rigid body degrees of freedom and small elastic displacements. The displacements are approximated by a truncated series of the unrestrained mode shapes, which are obtained using equivalent plate theory. Subsequently, the nonlinear equations of motion are linearized about the trim state, which is obtained using a rigid body trim model and steady hypersonic aerodynamics. The appropriate flutter derivatives are calculated from piston theory. Results describing mode shapes, trim behavior, and aeroelastic stability of a generic hypersonic vehicle are presented.

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.

Computational effects of inlet representation on powered hypersonic, airbreathing models

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Tatum, Kenneth E.

1993-01-01

Computational results are presented to illustrate the powered aftbody effects of representing the scramjet inlet on a generic hypersonic vehicle with a fairing, to divert the external flow, as compared to an operating flow-through scramjet inlet. This study is pertinent to the ground testing of hypersonic, airbreathing models employing scramjet exhaust flow simulation in typical small-scale hypersonic wind tunnels. The comparison of aftbody effects due to inlet representation is well-suited for computational study, since small model size typically precludes the ability to ingest flow into the inlet and perform exhaust simulation at the same time. Two-dimensional analysis indicates that, although flowfield differences exist for the two types of inlet representations, little, if any, difference in surface aftbody characteristics is caused by fairing over the inlet.

Computational and experimental aftbody flow fields for hypersonic, airbreathing configurations with scramjet exhaust flow simulation

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Tatum, Kenneth E.

1991-01-01

Computational results are presented for three issues pertinent to hypersonic, airbreathing vehicles employing scramjet exhaust flow simulation. The first issue consists of a comparison of schlieren photographs obtained on the aftbody of a cruise missile configuration under powered conditions with two-dimensional computational solutions. The second issue presents the powered aftbody effects of modeling the inlet with a fairing to divert the external flow as compared to an operating flow-through inlet on a generic hypersonic vehicle. Finally, a comparison of solutions examining the potential of testing powered configurations in a wind-off, instead of a wind-on, environment, indicate that, depending on the extent of the three-dimensional plume, it may be possible to test aftbody powered hypersonic, airbreathing configurations in a wind-off environment.

Conjugate Heat Transfer Study in Hypersonic Flows

NASA Astrophysics Data System (ADS)

Sahoo, Niranjan; Kulkarni, Vinayak; Peetala, Ravi Kumar

2018-04-01

Coupled and decoupled conjugate heat transfer (CHT) studies are carried out to imitate experimental studies for heat transfer measurement in hypersonic flow regime. The finite volume based solvers are used for analyzing the heat interaction between fluid and solid domains. Temperature and surface heat flux signals are predicted by both coupled and decoupled CHT analysis techniques for hypersonic Mach numbers. These two methodologies are also used to study the effect of different wall materials on surface parameters. Effectiveness of these CHT solvers has been verified for the inverse problem of wall heat flux recovery using various techniques reported in the literature. Both coupled and decoupled CHT techniques are seen to be equally useful for prediction of local temperature and heat flux signals prior to the experiments in hypersonic flows.

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.

Multi Laser Pulse Investigation of the DEAS Concept in Hypersonic Flow

DOE Office of Scientific and Technical Information (OSTI.GOV)

Minucci, M.A.S.; Toro, P.G.P.; Oliveira, A.C.

2004-03-30

The present paper presents recent experimental results on the Laser-Supported Directed Energy 'Air Spike' - DEAS in hypersonic flow achieved by the Laboratory of Aerothermodynamics and Hypersonics - LAH, Brazil. Two CO2 TEA lasers, sharing the same optical cavity, have been used in conjunction with the IEAv 0.3m Hypersonic Shock Tunnel - HST to demonstrate the Laser-Supported DEAS concept. A single and double laser pulse, generated during the tunnel useful test time, were focused through a NaCl lens upstream of a Double Apollo Disc model fitted with seven piezoelectric pressure transducers and six platinum thin film heat transfer gauges. Themore » objective being to corroborate previous results as well as to obtain additional pressure and heat flux distributions information when two laser pulses are used.« less

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, and results of turbulence model assessments for the six models that have been extensively applied to the hypersonic validation database are compiled and presented in graphical form. While some of the turbulence models do provide reasonable predictions for the surface pressure, the predictions for surface heat flux are generally poor, and often in error by a factor of four or more. In the vast majority of the turbulence model validation studies we review, the authors fail to adequately address the numerical accuracy of the simulations (i.e., discretization and iterative error) and the sensitivities of the model predictions to freestream turbulence quantities or near-wall y+ mesh spacing. We recommend new hypersonic experiments be conducted which (1) measure not only surface quantities but also mean and fluctuating quantities in the interaction region and (2) provide careful estimates of both random experimental uncertainties and correlated bias errors for the measured quantities and freestream conditions. For the turbulence models, we recommend that a wide-range of turbulence models (including newer models) be re-examined on the current hypersonic experimental database, including the more recent experiments. Any future turbulence model validation efforts should carefully assess the numerical accuracy and model sensitivities. In addition, model corrections (e.g., compressibility corrections) should be carefully examined for their effects on a standard, low-speed validation database. Finally, as new experiments or direct numerical simulation data become available with information on mean and fluctuating quantities, they should be used to improve the turbulence models and thus increase their predictive capability.

Status of the Combined Cycle Engine Rig

NASA Technical Reports Server (NTRS)

Saunders, Dave; Slater, John; Dippold, Vance

2009-01-01

Status for the past year is provided of the turbine-based Combined-Cycle Engine (CCE) Rig for the hypersonic project. As part of the first stage propulsion of a two-stage-to-orbit vehicle concept, this engine rig is designed with a common inlet that supplies flow to a turbine engine and a dual-mode ramjet / scramjet engine in an over/under configuration. At Mach 4 the inlet has variable geometry to switch the airflow from the turbine to the ramjet / scramjet engine. This process is known as inlet mode-transition. In addition to investigating inlet aspects of mode transition, the rig will allow testing of turbine and scramjet systems later in the test series. Fully closing the splitter cowl "cocoons" the turbine engine and increases airflow to the scramjet duct. The CCE Rig will be a testbed to investigate integrated propulsion system and controls technology objectives. Four phases of testing are planned to 1) characterize the dual inlet database, 2) collect inlet dynamics using system identification techniques, 3) implement an inlet control to demonstrate mode-transition scenarios and 4) demonstrate integrated inlet/turbine engine operation through mode-transition. Status of the test planning and preparation activities is summarized with background on the inlet design and small-scale testing, analytical CFD predictions and some details of the large-scale hardware. The final stages of fabrication are underway.

High-Speed Schlieren and 10-Hz Kr PLIF for the new AFRL Mach-6 Ludwieg Tube Hypersonic Wind Tunnel

DTIC Science & Technology

2017-11-01

STATES AIR FORCE AFRL-RQ-WP-TP-2017-0167 HIGH -SPEED SCHLIEREN AND 10-HZ KR PLIF FOR THE NEW AFRL MACH 6 LUDWIEG TUBE HYPERSONIC WIND TUNNEL Roger...L. Kimmel and Campbell D. Carter Hypersonic Sciences Branch High Speed Systems Division Joshua D. Pickles and Venkateswaran Narayanaswamy North...Public Affairs Office (PAO) and is available to the general public, including foreign nationals. Copies may be obtained from the Defense Technical

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.

A new Lagrangian random choice method for steady two-dimensional supersonic/hypersonic flow

NASA Technical Reports Server (NTRS)

Loh, C. Y.; Hui, W. H.

1991-01-01

Glimm's (1965) random choice method has been successfully applied to compute steady two-dimensional supersonic/hypersonic flow using a new Lagrangian formulation. The method is easy to program, fast to execute, yet it is very accurate and robust. It requires no grid generation, resolves slipline and shock discontinuities crisply, can handle boundary conditions most easily, and is applicable to hypersonic as well as supersonic flow. It represents an accurate and fast alternative to the existing Eulerian methods. Many computed examples are given.