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.

A Photogrammetric System for Model Attitude Measurement in Hypersonic Wind Tunnels

NASA Technical Reports Server (NTRS)

Jones, Thomas W.; Lunsford, Charles B.

2007-01-01

A series of wind tunnel tests have been conducted to evaluate a multi-camera videogrammetric system designed to measure model attitude in hypersonic facilities. The technique utilizes processed video data and photogrammetric principles for point tracking to compute model position including pitch, roll and yaw. A discussion of the constraints encountered during the design, and a review of the measurement results obtained from the NASA Langley Research Center (LaRC) 31-Inch Mach 10 tunnel are presented.

An iodine hypersonic wind tunnel for the study of nonequilibrium reacting flows

NASA Technical Reports Server (NTRS)

Pham-Van-diep, G. C.; Muntz, E. P.; Weaver, D. P.; Dewitt, T. G.; Bradley, M. K.; Erwin, D. A.; Kunc, J. A.

1992-01-01

A pilot scale hypersonic wind tunnel operating on pure iodine vapor has been designed and tested. The wind tunnel operates intermittently with a run phase lasting approximately 20 minutes. Successful recirculation of the iodine used during the run phase has been achieved but can be improved. Relevant issues regarding the full scale facility's design and operation, and the use of iodine as a working gas are discussed. Continuous wave laser induced fluorescence was used to monitor number densities within the plume flowfield, while pulsed laser induced fluorescence was used in an initial attempt to measure vibrational energy state population distributions. Preliminary nozzle flow calculations based on finite rate chemistry are presented.

CAN-DO, CFD-based Aerodynamic Nozzle Design and Optimization program for supersonic/hypersonic wind tunnels

NASA Technical Reports Server (NTRS)

Korte, John J.; Kumar, Ajay; Singh, D. J.; White, J. A.

1992-01-01

A design program is developed which incorporates a modern approach to the design of supersonic/hypersonic wind-tunnel nozzles. The approach is obtained by the coupling of computational fluid dynamics (CFD) with design optimization. The program can be used to design a 2D or axisymmetric, supersonic or hypersonic, wind-tunnel nozzles that can be modeled with a calorically perfect gas. The nozzle design is obtained by solving a nonlinear least-squares optimization problem (LSOP). The LSOP is solved using an iterative procedure which requires intermediate flowfield solutions. The nozzle flowfield is simulated by solving the Navier-Stokes equations for the subsonic and transonic flow regions and the parabolized Navier-Stokes equations for the supersonic flow regions. The advantages of this method are that the design is based on the solution of the viscous equations eliminating the need to make separate corrections to a design contour, and the flexibility of applying the procedure to different types of nozzle design problems.

Least-squares/parabolized Navier-Stokes procedure for optimizing hypersonic wind tunnel nozzles

NASA Technical Reports Server (NTRS)

Korte, John J.; Kumar, Ajay; Singh, D. J.; Grossman, B.

1991-01-01

A new procedure is demonstrated for optimizing hypersonic wind-tunnel-nozzle contours. The procedure couples a CFD computer code to an optimization algorithm, and is applied to both conical and contoured hypersonic nozzles for the purpose of determining an optimal set of parameters to describe the surface geometry. A design-objective function is specified based on the deviation from the desired test-section flow-field conditions. The objective function is minimized by optimizing the parameters used to describe the nozzle contour based on the solution to a nonlinear least-squares problem. The effect of the changes in the nozzle wall parameters are evaluated by computing the nozzle flow using the parabolized Navier-Stokes equations. The advantage of the new procedure is that it directly takes into account the displacement effect of the boundary layer on the wall contour. The new procedure provides a method for optimizing hypersonic nozzles of high Mach numbers which have been designed by classical procedures, but are shown to produce poor flow quality due to the large boundary layers present in the test section. The procedure is demonstrated by finding the optimum design parameters for a Mach 10 conical nozzle and a Mach 6 and a Mach 15 contoured nozzle.

Direct numerical simulations of an arc-powered heater for used in a hypersonic wind tunnel

NASA Astrophysics Data System (ADS)

Kim, Pilbum; Panesi, Marco; Freund, Jonathan

2017-11-01

We study a model arc-heater using direct numerical simulations, in a configuration motivated by its used to generated inflow of a high-speed wind tunnel for hypersonics research. The flow is assumed to be in local thermal equilibrium (LTE) and is modeled with with 11 species (N2, O2, NO, N, O, N2+,O2+,NO+, N+, O+, e-). The flow equations are solved in conjunction with an electrostatic field solver and the gas electric conductivity in LTE. The flow rate and the mean arc power are set to be 50.42 g/s and 84.7 kW with 214.0 V of the mean arc voltage , respectively. We study the flow details, the heading and thrust mechanisms, and make general comparisons with a corresponding, though geometrically more complex, experimental configuration. We particularly interested in the radical species it produces and will potentially be present in the wind-tunnel test section. This material is based in part upon work supported by the Department of Energy, National Nuclear Security Administration, under Award Number DE-NA0002374.

Measurement of temperature and pressure on the surface of a blunt cone using FBG sensor in hypersonic wind tunnel

NASA Astrophysics Data System (ADS)

Prasad, A. S. Guru; Sharath, U.; Nagarjun, V.; Hegde, G. M.; Asokan, S.

2013-09-01

Measurement of temperature and pressure exerted on the leeward surface of a blunt cone specimen has been demonstrated in the present work in a hypersonic wind tunnel using fiber Bragg grating (FBG) sensors. The experiments were conducted on a 30° apex-angle blunt cone with 51 mm base diameter at wind flow speeds of Mach 6.5 and 8.35 in a 300 mm hypersonic wind tunnel of Indian Institute of Science, Bangalore. A special pressure insensitive temperature sensor probe along with the conventional bare FBG sensors was used for explicit temperature and aerodynamic pressure measurement respectively on the leeward surface of the specimen. computational fluid dynamics (CFD) simulation of the flow field around the blunt cone specimen has also been carried out to obtain the temperature and pressure at conditions analogous to experiments. The results obtained from FBG sensors and the CFD simulations are found to be in good agreement with each other.

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

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.

Wind-US Code Physical Modeling Improvements to Complement Hypersonic Testing and Evaluation

NASA Technical Reports Server (NTRS)

Georgiadis, Nicholas J.; Yoder, Dennis A.; Towne, Charles S.; Engblom, William A.; Bhagwandin, Vishal A.; Power, Greg D.; Lankford, Dennis W.; Nelson, Christopher C.

2009-01-01

This report gives an overview of physical modeling enhancements to the Wind-US flow solver which were made to improve the capabilities for simulation of hypersonic flows and the reliability of computations to complement hypersonic testing. The improvements include advanced turbulence models, a bypass transition model, a conjugate (or closely coupled to vehicle structure) conduction-convection heat transfer capability, and an upgraded high-speed combustion solver. A Mach 5 shock-wave boundary layer interaction problem is used to investigate the benefits of k- s and k-w based explicit algebraic stress turbulence models relative to linear two-equation models. The bypass transition model is validated using data from experiments for incompressible boundary layers and a Mach 7.9 cone flow. The conjugate heat transfer method is validated for a test case involving reacting H2-O2 rocket exhaust over cooled calorimeter panels. A dual-mode scramjet configuration is investigated using both a simplified 1-step kinetics mechanism and an 8-step mechanism. Additionally, variations in the turbulent Prandtl and Schmidt numbers are considered for this scramjet configuration.

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.

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.

A CFD-based aerodynamic design procedure for hypersonic wind-tunnel nozzles

NASA Technical Reports Server (NTRS)

Korte, John J.

1993-01-01

A new procedure which unifies the best of current classical design practices, computational fluid dynamics (CFD), and optimization procedures is demonstrated for designing the aerodynamic lines of hypersonic wind-tunnel nozzles. The new procedure can be used to design hypersonic wind tunnel nozzles with thick boundary layers where the classical design procedure has been shown to break down. An efficient CFD code, which solves the parabolized Navier-Stokes (PNS) equations using an explicit upwind algorithm, is coupled to a least-squares (LS) optimization procedure. A LS problem is formulated to minimize the difference between the computed flow field and the objective function, consisting of the centerline Mach number distribution and the exit Mach number and flow angle profiles. The aerodynamic lines of the nozzle are defined using a cubic spline, the slopes of which are optimized with the design procedure. The advantages of the new procedure are that it allows full use of powerful CFD codes in the design process, solves an optimization problem to determine the new contour, can be used to design new nozzles or improve sections of existing nozzles, and automatically compensates the nozzle contour for viscous effects as part of the unified design procedure. The new procedure is demonstrated by designing two Mach 15, a Mach 12, and a Mach 18 helium nozzles. The flexibility of the procedure is demonstrated by designing the two Mach 15 nozzles using different constraints, the first nozzle for a fixed length and exit diameter and the second nozzle for a fixed length and throat diameter. The computed flow field for the Mach 15 least squares parabolized Navier-Stokes (LS/PNS) designed nozzle is compared with the classically designed nozzle and demonstrates a significant improvement in the flow expansion process and uniform core region.

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.

Characterization of a hypersonic quiet wind tunnel nozzle

NASA Astrophysics Data System (ADS)

Sweeney, Cameron J.

The Boeing/AFOSR Mach-6 Quiet Tunnel at Purdue University has been able to achieve low-disturbance flows at high Reynolds numbers for approximately ten years. The flow in the nozzle was last characterized in 2010. However, researchers have noted that the performance of the nozzle has changed in the intervening years. Understand ing the tunnel characteristics is critical for the hypersonic boundary-layer transition research performed at the facility and any change in performance could have signif icant effects on research performed at the facility. Pitot probe measurements were made using Kulite and PCB pressure transducers to quantify the performance changes since characterization was last performed. Aspects of the nozzle that were investi gated include the radial uniformity of the flow, the effects that time and stagnation pressure have on the flow, and the Reynolds number limits of low-disturbance flows. Measurements showed that freestream noise levels are consistently around 0.01% to 0.02% for the majority of the quiet flow core, with quiet flow now achievable for Reynolds numbers up to Re = 13.0x10 6/m. Additionally, while pitot probes are a widely used measurement technique for quantifying freestream disturbances, pitot probes are not without drawbacks. In order to provide a more complete methodology for freestream noise measurement other researchers have started experimenting with alternate geometries, such as cones. Using a newly designed 30° half-angle cone model, measurements were performed to quantify the freestream noise in the BAM6QT and compare the performance with another hypersonic wind tunnel. Also, measurements were made with three newly designed pitot sleeves to study the effects of probe geometry on freestream noise measurements. The results were compared to recent DNS calculations.

Comparison of thin-film resistance heat-transfer gages with thin-skin transient calorimeter gages in conventional hypersonic wind tunnels

NASA Technical Reports Server (NTRS)

Miller, C. G., III

1981-01-01

Thin film gages deposited at the stagnation region of small (8.1-mm-diameter) hemispheres and gages mounted flush with the surface of a sharp-leading-edge flat plate were tested in the Langley continuous-flow hypersonic tunnel and in the Langley hypersonic CF4 tunnel. Two substrate materials were tested, quartz and a machinable glass-ceramic. Small hemispheres were also tested utilizing the thin-skin transient calorimeter technique usually employed in conventional tunnels. One transient calorimeter model was a thin shell of stainless steel, and the other was a thin-skin insert of stainless steel mounted into a hemisphere fabricated from a machinable-glass-ceramic. Measured heat-transfer rates from the various hemispheres were compared with one another and with predicted rates. The results demonstrate the feasibility and advantages of using-film resistance heat-transfer gages in conventional hypersonic wind tunnels over a wide range of conditions.

Magnetohydrodynamics Accelerator Research Into Advanced Hypersonics (MARIAH). Part 1

NASA Technical Reports Server (NTRS)

Micheletti, David A.; Baughman, Jack A.; Nelson, Gordon L.; Simmons, Gloyd A.

1997-01-01

This report documents the activities, results, conclusions and recommendations of the Magnetohydrodynamics Accelerator Research Into Advanced Hypersonics (MARIAH) Project in which the use of magnetohydrodynamics (MHD) technology is investigated for its applicability to augment hypersonic wind tunnels. The long range objective of this investigation is to advance the development of ground test facilities to support the development of hypervelocity flight vehicles. The MHD accelerator adds kinetic energy directly to the wind tunnel working fluid, thereby increasing its Mach number to hypervelocity levels. Several techniques for MHD augmentation, as well as other physical characteristics of the process are studied to enhance the overall performance of hypersonic wind tunnel design. Specific recommendations are presented to improve the effectiveness of ground test facilities. The work contained herein builds on nearly four decades of research and experimentation by the aeronautics ground test and evaluation community, both foreign and domestic.

Magnetohydrodynamics Accelerator Research into Advanced Hypersonics (MARIAH). Part 2

NASA Technical Reports Server (NTRS)

Baughman, Jack A.; Micheletti, David A.; Nelson, Gordon L.; Simmons, Gloyd A.

1997-01-01

This report documents the activities, results, conclusions and recommendations of the Magnetohydrodynamics Accelerator Research Into Advanced Hypersonics (MARIAH) Project in which the use of magnetohydrodynamics (MHD) technology is investigated for its applicability to augment hypersonic wind tunnels. The long range objective of this investigation is to advance the development of ground test facilities to support the development of hypervelocity flight vehicles. The MHD accelerator adds kinetic energy directly to the wind tunnel working fluid, thereby increasing its Mach number to hypervelocity levels. Several techniques for MHD augmentation, as well as other physical characteristics of the process are studied to enhance the overall performance of hypersonic wind tunnel design. Specific recommendations are presented to improve the effectiveness of ground test facilities. The work contained herein builds on nearly four decades of research and experimentation by the aeronautics ground test and evaluation community, both foreign and domestic.

Conceptual design for an electron-beam heated hypersonic wind tunnel

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

Lipinski, R.J.; Kensek, R.P.

1997-07-01

There is a need for hypersonic wind-tunnel testing at about mach 10 and above using natural air and simulating temperatures and pressures which are prototypic of flight at 50 km altitude or below. With traditional wind-tunnel techniques, gas cooling during expansion results in exit temperatures which are too low. Miles, et al., have proposed overcoming this difficulty by heating the air with a laser beam as it expands in the wind-tunnel nozzle. This report discusses an alternative option of using a high-power electron beam to heat the air as it expands. In the e-beam heating concept, the electron beam ismore » injected into the wind-tunnel nozzle near the exit and then is guided upstream toward the nozzle throat by a strong axial magnetic field. The beam deposits most of its power in the dense air near the throat where the expansion rate is greatest. A conceptual design is presented for a large-scale system which achieves Mach 14 for 0.1 seconds with an exit diameter of 2.8 meters. It requires 450 MW of electron beam power (5 MeV at 90 A). The guiding field is 500 G for most of the transport length and increases to 100 kG near the throat to converge the beam to a 1.0-cm diameter. The beam generator is a DC accelerator using a Marx bank (of capacitors) and a diode stack with a hot cathode. 14 refs. 38 figs., 9 tabs.« less

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.

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.

Investigation of Condensation/Clustering Effects on Rayleigh Scattering Measurements in a Hypersonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Tyler, Charles

1996-01-01

Rayleigh scattering, a nonintrusive measurement technique for the measurement of density in a hypersonic wind tunnel, is under investigation at Wright Laboratory's Mach 6 wind tunnel. Several adverse effects, i.e., extraneous scatter off walls and windows, hinder Rayleigh scattering measurements. Condensation and clustering of flow constituents also present formidable obstacles. Overcoming some of these difficulties, measurements have been achieved while the Mach 6 test section was pumped down to a vacuum, as well as for actual tunnel operation for various stagnation pressures at fixed stagnation temperatures. Stagnation pressures ranged from 0.69 MPa to 6.9 MPa at fixed stagnation temperatures of 511, 556, and 611 K. Rayleigh scatter results show signal levels much higher than expected for molecular scattering in the wind tunnel. Even with higher than expected signals, scattering measurements have been made in the flowfield of an 8-degree half-angle blunt nose cone with a nose radius of 1.5 cm.

Testing of Flexible Ballutes in Hypersonic Wind Tunnels for Planetary Aerocapture

NASA Technical Reports Server (NTRS)

Buck, Gregory M.

2007-01-01

Studies were conducted for the In-Space Propulsion (ISP) Ultralightweight Ballute Technology Development Program to increase the technical readiness level of inflatable decelerator systems for planetary aerocapture. The present experimental study was conducted to develop the capability for testing lightweight, flexible materials in hypersonic facilities. The primary objectives were to evaluate advanced polymer film materials in a high-temperature, high-speed flow environment and provide experimental data for comparisons with fluid-structure interaction modeling tools. Experimental testing was conducted in the Langley Aerothermodynamics Laboratory 20-Inch Hypersonic CF4 and 31-Inch Mach 10 Air blowdown wind tunnels. Quantitative flexure measurements were made for 60 degree half angle afterbody-attached ballutes, in which polyimide films (1-mil and 3- mil thick) were clamped between a 1/2-inch diameter disk and a base ring (4-inch and 6-inch diameters). Deflection measurements were made using a parallel light silhouette of the film surface through an existing schlieren optical system. The purpose of this paper is to discuss these results as well as free-flying testing techniques being developed for both an afterbody-attached and trailing toroidal ballute configuration to determine dynamic fluid-structural stability. Methods for measuring polymer film temperature were also explored using both temperature sensitive paints (for up to 370 C) and laser-etched thin-film gages.

Testing of Flexible Ballutes in Hypersonic Wind Tunnels for Planetary Aerocapture

NASA Technical Reports Server (NTRS)

Buck, Gregory M.

2006-01-01

Studies were conducted for the In-Space Propulsion (ISP) Ultralightweight Ballute Technology Development Program to increase the technical readiness level of inflatable decelerator systems for planetary aerocapture. The present experimental study was conducted to develop the capability for testing lightweight, flexible materials in hypersonic facilities. The primary objectives were to evaluate advanced polymer film materials in a high-temperature, high-speed flow environment and provide experimental data for comparisons with fluid-structure interaction modeling tools. Experimental testing was conducted in the Langley Aerothermodynamics Laboratory 20-Inch Hypersonic CF4 and 31-Inch Mach 10 Air blowdown wind tunnels. Quantitative flexure measurements were made for 60 degree half angle afterbody-attached ballutes, in which polyimide films (1-mil and 3-mil thick) were clamped between a 1/2-inch diameter disk and a base ring (4-inch and 6-inch diameters). Deflection measurements were made using a parallel light silhouette of the film surface through an existing schlieren optical system. The purpose of this paper is to discuss these results as well as free-flying testing techniques being developed for both an afterbody-attached and trailing toroidal ballute configuration to determine dynamic fluid-structural stability. Methods for measuring polymer film temperature were also explored using both temperature sensitive paints (for up to 370 C) and laser-etched thin-film gages.

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.

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.

Hypersonic Wind Tunnel Calibration Using the Modern Design of Experiments

NASA Technical Reports Server (NTRS)

Rhode, Matthew N.; DeLoach, Richard

2005-01-01

A calibration of a hypersonic wind tunnel has been conducted using formal experiment design techniques and response surface modeling. Data from a compact, highly efficient experiment was used to create a regression model of the pitot pressure as a function of the facility operating conditions as well as the longitudinal location within the test section. The new calibration utilized far fewer design points than prior experiments, but covered a wider range of the facility s operating envelope while revealing interactions between factors not captured in previous calibrations. A series of points chosen randomly within the design space was used to verify the accuracy of the response model. The development of the experiment design is discussed along with tactics used in the execution of the experiment to defend against systematic variation in the results. Trends in the data are illustrated, and comparisons are made to earlier findings.

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

NASA Astrophysics Data System (ADS)

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

2016-11-01

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

Assessment of a flow-through balance for hypersonic wind tunnel models with scramjet exhaust flow simulation

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Kniskern, Marc W.; Monta, William J.

1993-01-01

The purpose of this investigation were twofold: first, to determine whether accurate force and moment data could be obtained during hypersonic wind tunnel tests of a model with a scramjet exhaust flow simulation that uses a representative nonwatercooled, flow-through balance; second, to analyze temperature time histories on various parts of the balance to address thermal effects on force and moment data. The tests were conducted in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel at free-stream Reynolds numbers ranging from 0.5 to 7.4 x 10(exp 6)/ft and nominal angles of attack of -3.5 deg, 0 deg, and 5 deg. The simulant exhaust gases were cold air, hot air, and a mixture of 50 percent Argon and 50 percent Freon by volume, which reached stagnation temperatures within the balance of 111, 214, and 283 F, respectively. All force and moment values were unaffected by the balance thermal response from exhaust gas simulation and external aerodynamic heating except for axial-force measurements, which were significantly affected by balance heating. This investigation showed that for this model at the conditions tested, a nonwatercooled, flow-through balance is not suitable for axial-force measurements during scramjet exhaust flow simulation tests at hypersonic speeds. In general, heated exhaust gas may produce unacceptable force and moment uncertainties when used with thermally sensitive balances.

Design and optimization of resistance wire electric heater for hypersonic wind tunnel

NASA Astrophysics Data System (ADS)

Rehman, Khurram; Malik, Afzaal M.; Khan, I. J.; Hassan, Jehangir

2012-06-01

The range of flow velocities of high speed wind tunnels varies from Mach 1.0 to hypersonic order. In order to achieve such high speed flows, a high expansion nozzle is employed in the converging-diverging section of wind tunnel nozzle. The air for flow is compressed and stored in pressure vessels at temperatures close to ambient conditions. The stored air is dried and has minimum amount of moisture level. However, when this air is expanded rapidly, its temperature drops significantly and liquefaction conditions can be encountered. Air at near room temperature will liquefy due to expansion cooling at a flow velocity of more than Mach 4.0 in a wind tunnel test section. Such liquefaction may not only be hazardous to the model under test and wind tunnel structure; it may also affect the test results. In order to avoid liquefaction of air, a pre-heater is employed in between the pressure vessel and the converging-diverging section of a wind tunnel. A number of techniques are being used for heating the flow in high speed wind tunnels. Some of these include the electric arc heating, pebble bed electric heating, pebble bed natural gas fired heater, hydrogen burner heater, and the laser heater mechanisms. The most common are the pebble bed storage type heaters, which are inefficient, contaminating and time consuming. A well designed electrically heating system can be efficient, clean and simple in operation, for accelerating the wind tunnel flow up to Mach 10. This paper presents CFD analysis of electric preheater for different configurations to optimize its design. This analysis has been done using ANSYS 12.1 FLUENT package while geometry and meshing was done in GAMBIT.

MHz-Rate NO PLIF Imaging in a Mach 10 Hypersonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Jiang, N.; Webster, M.; Lempert, Walter R.; Miller, J. D.; Meyer, T. R.; Danehy, Paul M.

2010-01-01

NO PLIF imaging at repetition rates as high as 1 MHz is demonstrated in the NASA Langley 31 inch Mach 10 hypersonic wind tunnel. Approximately two hundred time correlated image sequences, of between ten and twenty individual frames, were obtained over eight days of wind tunnel testing spanning two entries in March and September of 2009. The majority of the image sequences were obtained from the boundary layer of a 20 flat plate model, in which transition was induced using a variety of cylindrical and triangular shaped protuberances. The high speed image sequences captured a variety of laminar and transitional flow phenomena, ranging from mostly laminar flow, typically at lower Reynolds number and/or in the near wall region of the model, to highly transitional flow in which the temporal evolution and progression of characteristic streak instabilities and/or corkscrew-shaped vortices could be clearly identified. A series of image sequences were also obtained from a 20 compression ramp at a 10 angle of attack in which the temporal dynamics of the characteristic separated flow was captured in a time correlated manner.

Electron-Beam Diagnostic Methods for Hypersonic Flow Diagnostics

NASA Technical Reports Server (NTRS)

1994-01-01

The purpose of this work was the evaluation of the use of electron-bean fluorescence for flow measurements during hypersonic flight. Both analytical and numerical models were developed in this investigation to evaluate quantitatively flow field imaging concepts based upon the electron beam fluorescence technique for use in flight research and wind tunnel applications. Specific models were developed for: (1) fluorescence excitation/emission for nitrogen, (2) rotational fluorescence spectrum for nitrogen, (3) single and multiple scattering of electrons in a variable density medium, (4) spatial and spectral distribution of fluorescence, (5) measurement of rotational temperature and density, (6) optical filter design for fluorescence imaging, and (7) temperature accuracy and signal acquisition time requirements. Application of these models to a typical hypersonic wind tunnel flow is presented. In particular, the capability of simulating the fluorescence resulting from electron impact ionization in a variable density nitrogen or air flow provides the capability to evaluate the design of imaging instruments for flow field mapping. The result of this analysis is a recommendation that quantitative measurements of hypersonic flow fields using electron-bean fluorescence is a tractable method with electron beam energies of 100 keV. With lower electron energies, electron scattering increases with significant beam divergence which makes quantitative imaging difficult. The potential application of the analytical and numerical models developed in this work is in the design of a flow field imaging instrument for use in hypersonic wind tunnels or onboard a flight research vehicle.

Review of Aeronautical Wind Tunnel Facilities

NASA Technical Reports Server (NTRS)

1988-01-01

The nation's aeronautical wind tunnel facilities constitute a valuable technological resource and make a significant contribution to the global supremacy of U.S. aircraft, both civil and military. At the request of NASA, the National Research Council's Aeronautics and Space Engineering Board organized a commitee to review the state of repair, adequacy, and future needs of major aeronautical wind tunnel facilities in meeting national goals. The comittee identified three main areas where actions are needed to sustain the capability of NASA's aeronautical wind tunnel facilities to support the national aeronautical research and development activities: tunnel maintenance and upgrading, productivity enhancement, and accommodation of new requirements (particularly in hypersonics). Each of these areas are addressed and the committee recommendations for appropriate actions presented.

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

Experimental results for a hypersonic nozzle/afterbody flow field

NASA Technical Reports Server (NTRS)

Spaid, Frank W.; Keener, Earl R.; Hui, Frank C. L.

1995-01-01

This study was conducted to experimentally characterize the flow field created by the interaction of a single-expansion ramp-nozzle (SERN) flow with a hypersonic external stream. Data were obtained from a generic nozzle/afterbody model in the 3.5 Foot Hypersonic Wind Tunnel at the NASA Ames Research Center, in a cooperative experimental program involving Ames and McDonnell Douglas Aerospace. The model design and test planning were performed in close cooperation with members of the Ames computational fluid dynamics (CFD) team for the National Aerospace Plane (NASP) program. This paper presents experimental results consisting of oil-flow and shadow graph flow-visualization photographs, afterbody surface-pressure distributions, rake boundary-layer measurements, Preston-tube skin-friction measurements, and flow field surveys with five-hole and thermocouple probes. The probe data consist of impact pressure, flow direction, and total temperature profiles in the interaction flow field.

Development and Characterization of a Low-Pressure Calibration System for Hypersonic Wind Tunnels

NASA Technical Reports Server (NTRS)

Green, Del L.; Everhart, Joel L.; Rhode, Matthew N.

2004-01-01

Minimization of uncertainty is essential for accurate ESP measurements at very low free-stream static pressures found in hypersonic wind tunnels. Statistical characterization of environmental error sources requires a well defined and controlled calibration method. A calibration system has been constructed and environmental control software developed to control experimentation to eliminate human induced error sources. The initial stability study of the calibration system shows a high degree of measurement accuracy and precision in temperature and pressure control. Control manometer drift and reference pressure instabilities induce uncertainty into the repeatability of voltage responses measured from the PSI System 8400 between calibrations. Methods of improving repeatability are possible through software programming and further experimentation.

Free-stream static pressure measurements in the Longshot hypersonic wind tunnel and sensitivity analysis

NASA Astrophysics Data System (ADS)

Grossir, Guillaume; Van Hove, Bart; Paris, Sébastien; Rambaud, Patrick; Chazot, Olivier

2016-05-01

The performance of fast-response slender static pressure probes is evaluated in the short-duration, cold-gas, VKI Longshot hypersonic wind tunnel. Free-stream Mach numbers range between 9.5 and 12, and unit Reynolds numbers are within 3-10 × 106/m. Absolute pressure sensors are fitted within the probes, and an inexpensive calibration method, suited to low static pressure environments (200-1000 Pa), is described. Transfer functions relating the probe measurements p w to the free-stream static pressure p ∞ are established for the Longshot flow conditions based on numerical simulations. The pressure ratios p w / p ∞ are found to be close to unity for both laminar and turbulent boundary layers. Weak viscous effects characterized by small viscous interaction parameters {bar{χ }}<1.5 are confirmed experimentally for probe aspect ratios of L/ D > 16.5 by installing multiple pressure sensors in a single probe. The effect of pressure orifice geometry is also evaluated experimentally and found to be negligible for either straight or chamfered holes, 0.6-1 mm in diameter. No sensitivity to probe angle of attack could be evidenced for α < 0.33°. Pressure measurements are compared to theoretical predictions assuming an isentropic nozzle flow expansion. Significant deviations from this ideal case and the Mach 14 contoured nozzle design are uncovered. Validation of the static pressure measurements is obtained by comparing shock wave locations on Schlieren photographs to numerical predictions using free-stream properties derived from the static pressure probes. While these results apply to the Longshot wind tunnel, the present methodology and sensitivity analysis can guide similar investigations for other hypersonic test facilities.

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.

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.

Portable Fluorescence Imaging System for Hypersonic Flow Facilities

NASA Technical Reports Server (NTRS)

Wilkes, J. A.; Alderfer, D. W.; Jones, S. B.; Danehy, P. M.

2003-01-01

A portable fluorescence imaging system has been developed for use in NASA Langley s hypersonic wind tunnels. The system has been applied to a small-scale free jet flow. Two-dimensional images were taken of the flow out of a nozzle into a low-pressure test section using the portable planar laser-induced fluorescence system. Images were taken from the center of the jet at various test section pressures, showing the formation of a barrel shock at low pressures, transitioning to a turbulent jet at high pressures. A spanwise scan through the jet at constant pressure reveals the three-dimensional structure of the flow. Future capabilities of the system for making measurements in large-scale hypersonic wind tunnel facilities are discussed.

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.

Design and Navier-Stokes analysis of hypersonic wind tunnel nozzles. M.S. Thesis

NASA Technical Reports Server (NTRS)

Benton, James R.

1989-01-01

Four hypersonic wind tunnel nozzles ranging in Mach number from 6 to 17 are designed with the method of characteristics and boundary layer approach (MOC/BL) and analyzed with a Navier-Stokes solver. Limitations of the MOC/BL approach when applied to thick high speed boundary layers with non-zero normal pressure gradients are investigated. Working gases include ideal air, thermally perfect nitrogen and virial CF4. Agreement between the design conditions and Navier-Stokes solutions for ideal air at Mach 6 is good. Thermally perfect nitrogen showed poor agreement at Mach 13.5 and Mach 17. Navier-Stokes solutions for CF4 are not obtained, but comparison of the effects of low gamma to those of high Mach number suggests that the Navier-Stokes solution would not compare well with design.

Heating Augmentation for Short Hypersonic Protuberances

NASA Technical Reports Server (NTRS)

Mazaheri, Ali R.; Wood, William A.

2008-01-01

Computational aeroheating analyses of the Space Shuttle Orbiter plug repair models are validated against data collected in the Calspan University of Buffalo Research Center (CUBRC) 48 inch shock tunnel. The comparison shows that the average difference between computed heat transfer results and the data is about 9.5%. Using CFD and Wind Tunnel (WT) data, an empirical correlation for estimating heating augmentation on short hypersonic protuberances (k/delta less than 0.3) is proposed. This proposed correlation is compared with several computed flight simulation cases and good agreement is achieved. Accordingly, this correlation is proposed for further investigation on other short hypersonic protuberances for estimating heating augmentation.

Research on hypersonic aircraft using pre-cooled turbojet engines

NASA Astrophysics Data System (ADS)

Taguchi, Hideyuki; Kobayashi, Hiroaki; Kojima, Takayuki; Ueno, Atsushi; Imamura, Shunsuke; Hongoh, Motoyuki; Harada, Kenya

2012-04-01

Systems analysis of a Mach 5 class hypersonic aircraft is performed. The aircraft can fly across the Pacific Ocean in 2 h. A multidisciplinary optimization program for aerodynamics, structure, propulsion, and trajectory is used in the analysis. The result of each element model is improved using higher accuracy analysis tools. The aerodynamic performance of the hypersonic aircraft is examined through hypersonic wind tunnel tests. A thermal management system based on the data of the wind tunnel tests is proposed. A pre-cooled turbojet engine is adopted as the propulsion system for the hypersonic aircraft. The engine can be operated continuously from take-off to Mach 5. This engine uses a pre-cooling cycle using cryogenic liquid hydrogen. The high temperature inlet air of hypersonic flight would be cooled by the same liquid hydrogen used as fuel. The engine is tested under sea level static conditions. The engine is installed on a flight test vehicle. Both liquid hydrogen fuel and gaseous hydrogen fuel are supplied to the engine from a tank and cylinders installed within the vehicle. The designed operation of major components of the engine is confirmed. A large amount of liquid hydrogen is supplied to the pre-cooler in order to make its performance sufficient for Mach 5 flight. Thus, fuel rich combustion is adopted at the afterburner. The experiments are carried out under the conditions that the engine is mounted upon an experimental airframe with both set up either horizontally or vertically. As a result, the operating procedure of the pre-cooled turbojet engine is demonstrated.

Combined free-stream disturbance measurements and receptivity studies in hypersonic wind tunnels by means of a slender wedge probe and direct numerical simulation

NASA Astrophysics Data System (ADS)

Wagner, Alexander; Schülein, Erich; Petervari, René; Hannemann, Klaus; Ali, Syed R. C.; Cerminara, Adriano; Sandham, Neil D.

2018-05-01

Combined free-stream disturbance measurements and receptivity studies in hypersonic wind tunnels were conducted by means of a slender wedge probe and direct numerical simulation. The study comprises comparative tunnel noise measurements at Mach 3, 6 and 7.4 in two Ludwieg tube facilities and a shock tunnel. Surface pressure fluctuations were measured over a wide range of frequencies and test conditions including harsh test environments not accessible to measurement techniques such as pitot probes and hot-wire anemometry. Quantitative results of the tunnel noise are provided in frequency ranges relevant for hypersonic boundary layer transition. In combination with the experimental studies, direct numerical simulations of the leading-edge receptivity to fast and slow acoustic waves were performed for the slender wedge probe at conditions corresponding to the experimental free-stream conditions. The receptivity to fast acoustic waves was found to be characterized by an early amplification of the induced fast mode. For slow acoustic waves an initial decay was found close to the leading edge. At all Mach numbers, and for all considered frequencies, the leading-edge receptivity to fast acoustic waves was found to be higher than the receptivity to slow acoustic waves. Further, the effect of inclination angles of the acoustic wave with respect to the flow direction was investigated. The combined numerical and experimental approach in the present study confirmed the previous suggestion that the slow acoustic wave is the dominant acoustic mode in noisy hypersonic wind tunnels.

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

NASA Technical Reports Server (NTRS)

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

1975-01-01

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

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.

Experimental And Numerical Investigation Of Aerothermal Characteristics Of The IXV Hypersonic Vehicle

NASA Astrophysics Data System (ADS)

Paris, S.; Charbonnier, D.; Tran, D.

2011-05-01

The main results of the aerothermodynamic hypersonic characterization of Intermediate eXperimental Vehicle (IXV), by means of both CFD simulations and wind tunnel measurements, have been reported and analyzed. In the framework of ESA FLPP Program, the VKI (Von Karman Institute) was in charge of an experimental test campaign for the consolidation of the aerothermal database in cold hypersonic regime. The tests campaign has been carried out at VKI Free Piston Longshot wind tunnel at mach 14. The numerical simulations have been performed for VKI wind tunnel conditions by CFSE with the in-house NSMB flow solver (Navier-Stokes Multi-Blocks 3D), the goal being to support the procedure of extrapolation-to-flight of the measurements and the general aerothermal characterization. Laminar, transitional and fully turbulent flows have been computed, with air considered as an ideal gas, for the wind tunnel tests numerical rebuilding. A detailed comparison of all measured and predicted hypersonic relevant phenomena and parameters (surface pressure and heat flux) is reported in the paper, together with a detailed description of configuration, freestream conditions, model attitude effects and flap deflection effect. The detailed analyze of the experimental and numerical data gives information on the nature of the flow on the body and on the flaps for the most critical configuration

Ultra-High Pressure Driver and Nozzle Survivability in the RDHWT/MARIAH II Hypersonic Wind Tunnel

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

Costantino, M.; Brown, G.; Raman, K.

2000-06-02

An ultra-high pressure device provides a high enthalpy (> 2500 kJ/kg), low entropy (< 5 kJ/kg-K) air source for the RDHWT/MARIAH II Program Medium Scale Hypersonic Wind Tunnel. The design uses stagnation conditions of 2300 MPa (330,000 Psi) and 750 K (900 F) in a radial configuration of intensifiers around an axial manifold to deliver pure air at 100 kg/s mass flow rates for run times suitable for aerodynamic, combustion, and test and evaluation applications. Helium injection upstream of the nozzle throat reduces the throat wall recovery temperature to about 1200 K and reduces the oxygen concentration at the nozzlemore » wall.« less

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.

MHz-rate nitric oxide planar laser-induced fluorescence imaging in a Mach 10 hypersonic wind tunnel.

PubMed

Jiang, Naibo; Webster, Matthew; Lempert, Walter R; Miller, Joseph D; Meyer, Terrence R; Ivey, Christopher B; Danehy, Paul M

2011-02-01

Nitric oxide planar laser-induced fluorescence (NO PLIF) imaging at repetition rates as high as 1 MHz is demonstrated in the NASA Langley 31 in. Mach 10 hypersonic wind tunnel. Approximately 200 time-correlated image sequences of between 10 and 20 individual frames were obtained over eight days of wind tunnel testing spanning two entries in March and September of 2009. The image sequences presented were obtained from the boundary layer of a 20° flat plate model, in which transition was induced using a variety of different shaped protuberances, including a cylinder and a triangle. The high-speed image sequences captured a variety of laminar and transitional flow phenomena, ranging from mostly laminar flow, typically at a lower Reynolds number and/or in the near wall region of the model, to highly transitional flow in which the temporal evolution and progression of characteristic streak instabilities and/or corkscrew-shaped vortices could be clearly identified.

Aerodynamic Characteristics of Two Waverider-Derived Hypersonic Cruise Configurations

NASA Technical Reports Server (NTRS)

Cockrell, Charles E., Jr.; Huebner, Lawrence D.; Finley, Dennis B.

1996-01-01

An evaluation was made on the effects of integrating the required aircraft components with hypersonic high-lift configurations known as waveriders to create hypersonic cruise vehicles. Previous studies suggest that waveriders offer advantages in aerodynamic performance and propulsion/airframe integration (PAI) characteristics over conventional non-waverider hypersonic shapes. A wind-tunnel model was developed that integrates vehicle components, including canopies, engine components, and control surfaces, with two pure waverider shapes, both conical-flow-derived waveriders for a design Mach number of 4.0. Experimental data and limited computational fluid dynamics (CFD) solutions were obtained over a Mach number range of 1.6 to 4.63. The experimental data show the component build-up effects and the aerodynamic characteristics of the fully integrated configurations, including control surface effectiveness. The aerodynamic performance of the fully integrated configurations is not comparable to that of the pure waverider shapes, but is comparable to previously tested hypersonic models. Both configurations exhibit good lateral-directional stability characteristics.

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.

Hypersonic research engine project. Phase 2: Aerothermodynamic Integration Model (AIM) test report

NASA Technical Reports Server (NTRS)

Andersen, W. L.; Kado, L.

1975-01-01

The Hypersonic Research Engine-Aerothermodynamic Integration Model (HRE-AIM) was designed, fabricated, and tested in the Hypersonic Tunnel Facility. The HRE-AIM is described along with its installation in the wind tunnel facility. Test conditions to which the HRE-AIM was subjected and observations made during the tests are discussed. The overall engine performance, component interaction, and ignition limits for the design are evaluated.

An overview of HyFIE Technical Research Project: cross-testing in main European hypersonic wind tunnels on EXPERT body

NASA Astrophysics Data System (ADS)

Brazier, Jean-Philippe; Martinez Schramm, Jan; Paris, Sébastien; Gawehn, Thomas; Reimann, Bodo

2016-09-01

HyFIE project aimed at improving the measurement techniques in hypersonic wind tunnels and comparing the experimental data provided by four major European facilities: DLR HEG and H2K, ONERA F4 and VKI Longshot. A common geometry of EXPERT body was chosen and four different models were used. A large amount of experimental data was collected and compared with the results of numerical simulations. Collapsing all the measured values showed a good agreement between the different facilities, as well as between experimental and computed data.

Planar laser-induced fluorescence (PLIF) investigation of hypersonic flowfields in a Mach 10 wind tunnel

NASA Technical Reports Server (NTRS)

Danehy, Paul M.; Wilkes, Jennifer A.; Aderfer, David W.; Jones, Stephen B.; Robbins, Anthony W.; Pantry, Danny P.; Schwartz, Richard J.

2006-01-01

Planar laser-induced fluorescence (PLIF) of nitric oxide (NO) was used to visualize four different hypersonic flowfields in the NASA Langley Research Center 31-Inch Mach 10 Air wind tunnel. The four configurations were: (1) the wake flowfield of a fuselage-only X-33 lifting body, (2) flow over a flat plate containing a rectangular cavity, (3) flow over a 70deg blunted cone with a cylindrical afterbody, formerly studied by an AGARD working group, and (4) an Apollo-geometry entry capsule - relevant to the Crew Exploration Vehicle currently being developed by NASA. In all cases, NO was seeded into the flowfield through tubes inside or attached to the model sting and strut. PLIF was used to visualize the NO in the flowfield. In some cases pure NO was seeded into the flow while in other cases a 5% NO, 95% N2 mix was injected. Several parameters were varied including seeding method and location, seeding mass flow rate, model angle of attack and tunnel stagnation pressure, which varies the unit Reynolds number. The location of the laser sheet was as also varied to provide three dimensional flow information. Virtual Diagnostics Interface (ViDI) technology developed at NASA Langley was used to visualize the data sets in post processing. The measurements demonstrate some of the capabilities of the PLIF method for studying hypersonic flows.

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.

CFD on hypersonic flow geometries with aeroheating

NASA Astrophysics Data System (ADS)

Sohail, Muhammad Amjad; Chao, Yan; Hui, Zhang Hui; Ullah, Rizwan

2012-11-01

The hypersonic flowfield around a blunted cone and cone-flare exhibits some of the major features of the flows around space vehicles, e.g. a detached bow shock in the stagnation region and the oblique shock wave/boundary layer interaction at the cone-flare junction. The shock wave/boundary layer interaction can produce a region of separated flow. This phenomenon may occur, for example, at the upstream-facing corner formed by a deflected control surface on a hypersonic entry vehicle, where the length of separation has implications for control effectiveness. Computational fluid-dynamics results are presented to show the flowfield around a blunted cone and cone-flare configurations in hypersonic flow with separation. This problem is of particular interest since it features most of the aspects of the hypersonic flow around planetary entry vehicles. The region between the cone and the flare is particularly critical with respect to the evaluation of the surface pressure and heat flux with aeroheating. Indeed, flow separation is induced by the shock wave boundary layer interaction, with subsequent flow reattachment, that can dramatically enhance the surface heat transfer. The exact determination of the extension of the recirculation zone is a particularly delicate task for numerical codes. Laminar flow and turbulent computations have been carried out using a full Navier-Stokes solver, with freestream conditions provided by the experimental data obtained at Mach 6, 8, and 16.34 wind tunnel. The numerical results are compared with the measured pressure and surface heat flux distributions in the wind tunnel and a good agreement is found, especially on the length of the recirculation region and location of shock waves. The critical physics of entropy layer, boundary layers, boundary layers and shock wave interaction and flow behind shock are properly captured and elaborated.. Hypersonic flows are characterized by high Mach number and high total enthalpy. An elevated

Supersonic investigation of two dimensional hypersonic exhaust nozzles

NASA Technical Reports Server (NTRS)

Carboni, Jeanne D.; Shyne, Rickey J.; Leavitt, Laurence D.; Taylor, John G.; Lamb, Milton

1992-01-01

An experimental investigation was conducted in the NASA Lewis 10 x 10 ft supersonic Wind Tunnel to determine the performance characteristics of 2D hypersonic exhaust nozzles/afterbodies at low supersonic conditions. Generally, this type of application requires a single expansion ramp nozzle (SERN) that is highly integrated with the airframe of the hypersonic vehicle. At design conditions (hypersonic speeds), the nozzle generally exhibits acceptable performance. At off-design conditions (transonic to mid-supersonic speeds), nozzle performance of a fixed geometry configuration is generally poor. Various 2-D nozzle configurations were tested at off-design conditions from Mach 2.0 to 3.5. Performance data is presented at nozzle pressure ratios from 1 to 35. Jet exhaust was simulated with high-pressure air. To study performance of different geometries, nozzle configurations were varied by interchanging the following model parts: internal upstream contour, expansion ramp, sidewalls, and cowl.

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.

DSMC Simulation of Separated Flows About Flared Bodies at Hypersonic Conditions

NASA Technical Reports Server (NTRS)

Moss, James N.

2000-01-01

This paper describes the results of a numerical study of interacting hypersonic flows at conditions that can be produced in ground-based test facilities. The computations are made with the direct simulation Monte Carlo (DSMC) method of Bird. The focus is on Mach 10 flows about flared axisymmetric configurations, both hollow cylinder flares and double cones. The flow conditions are those for which experiments have been or will be performed in the ONERA R5Ch low-density wind tunnel and the Calspan-University of Buffalo Research Center (CUBRC) Large Energy National Shock (LENS) tunnel. The range of flow conditions, model configurations, and model sizes provides a significant range of shock/shock and shock/boundary layer interactions at low Reynolds number conditions. Results presented will highlight the sensitivity of the calculations to grid resolution, contrast the differences in flow structure for hypersonic cold flows and those of more energetic but still low enthalpy flows, and compare the present results with experimental measurements for surface heating, pressure, and extent of separation.

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.

Quantifying Non-Equilibrium in Hypersonic Flows Using Entropy Generation

DTIC Science & Technology

2007-03-01

do this, two experimental cases performed at the Calspan- University of Buffalo Research Center ( CUBRC ) were modeled using Navier-Stokes based CFD...data provided by the CUBRC hypersonic wind tunnel facility (Holden and Wadhams, 2004). The wall data in Figure 9 and Figure 10 reveals some difference

CFD applications in hypersonic flight

NASA Technical Reports Server (NTRS)

Edwards, T. A.

1992-01-01

Design studies are underway for a variety of hypersonic flight vehicles. The National Aero-Space Plane will provide a reusable, single-stage-to-orbit capability for routine access to low earth orbit. Flight-capable satellites will dip into the atmosphere to maneuver to new orbits, while planetary probes will decelerate at their destination by atmospheric aerobraking. To supplement limited experimental capabilities in the hypersonic regime, CFD is being used to analyze the flow about these configurations. The governing equations include fluid dynamic as well as chemical species equations, which are solved with robust upwind differencing schemes. Examples of CFD applications to hypersonic vehicles suggest an important role this technology will play in the development of future aerospace systems. The computational resources needed to obtain solutions are large, but various strategies are being exploited to reduce the time required for complete vehicle simulations.

Wind Energy at NREL's National Wind Technology Center

ScienceCinema

None

2017-12-09

It is a pure, plentiful natural resource. Right now wind is in high demand and it holds the potential to transform the way we power our homes and businesses. NREL is at the forefront of wind energy research and development. NREL's National Wind Technology Center (NWTC) is a world-class facility dedicated to accelerating and deploying wind technology.

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.

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.

Center for Hypersonic Combined Cycle Flow Physics

DTIC Science & Technology

2015-03-24

team of expert experimentalists and numerical and chemical kinetic modelers. Flowfields were examined in the turbine /ramjet dual inlet mode transition...using data from the NASA Glenn IMX facility and RANS calculations. In the ramjet/scramjet mode regime a dual-mode combustion wind tunnel was developed...SUBJECT TERMS Hypersonic combined cycle propulsion, turbine /ram dual-inlet transition, ram/scram dual-mode transition, hypervelocity regime, RANS, Hybrid

The Wind Integration National Dataset (WIND) toolkit (Presentation)

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

Caroline Draxl: NREL

2014-01-01

Regional wind integration studies require detailed wind power output data at many locations to perform simulations of how the power system will operate under high penetration scenarios. The wind datasets that serve as inputs into the study must realistically reflect the ramping characteristics, spatial and temporal correlations, and capacity factors of the simulated wind plants, as well as being time synchronized with available load profiles.As described in this presentation, the WIND Toolkit fulfills these requirements by providing a state-of-the-art national (US) wind resource, power production and forecast dataset.

Hypersonic nozzle/afterbody CFD code validation. I - Experimental measurements

NASA Technical Reports Server (NTRS)

Spaid, Frank W.; Keener, Earl R.

1993-01-01

This study was conducted to obtain a detailed experimental description of the flow field created by the interaction of a single-expansion-ramp-nozzle flow with a hypersonic external stream. Data were obtained from a generic nozzle/afterbody model in the 3.5-Foot Hypersonic Wind Tunnel of the NASA Ames Research Center in a cooperative experimental program involving Ames and the McDonnell Douglas Research Laboratories. This paper presents experimental results consisting primarily of surveys obtained with a five-hole total-pressure/flow-direction probe and a total-temperature probe. These surveys were obtained in the flow field created by the interaction between the underexpanded jet plume and the external flow.

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

NASA Technical Reports Server (NTRS)

Cleary, Joseph W.

1989-01-01

A comprehensive test program is defined which is being implemented in the NASA/Ames 3.5 foot Hypersonic Wind Tunnel for obtaining data on a generic all-body hypersonic vehicle for computational fluid dynamics (CFD) code validation. Computational methods (approximate inviscid methods and an upwind parabolized Navier-Stokes code) currently being applied to the all-body model are outlined. Experimental and computational results on surface pressure distributions and Pitot-pressure surveys for the basic sharp-nose model (without control surfaces) at a free-stream Mach number of 7 are presented.

Direct Numerical Simulation of Acoustic Noise Generation from the Nozzle Wall of a Hypersonic Wind Tunnel

NASA Astrophysics Data System (ADS)

Huang, Junji; Duan, Lian; Choudhari, Meelan; Missouri Univ of Sci; Tech Team; NASA Langley Research Center Team

2017-11-01

Direct numerical simulations (DNS) are used to examine the acoustic noise generation from the turbulent boundary layer on the nozzle wall of a Mach 6 Ludwieg Tube. The emphasis is on characterizing the freestream acoustic pressure disturbances radiated from the nozzle-wall turbulent boundary layer and comparing it with acoustic noise generated from a single, flat wall in an unconfined setting at a similar freestream Mach number to assess the effects of noise reverberation. In particular, the numerical database is used to provide insights into the pressure disturbance spectrum and amplitude scaling with respect to the boundary-layer parameters as well as to understand the acoustic source mechanisms. Such information is important for characterizing the freestream disturbance environment in conventional (i.e., noisy) hypersonic wind tunnels. Air Force Office of Scientific Research Award No. FA9550-14-1-0170.

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).

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

NASA Technical Reports Server (NTRS)

Woike, Mark R.; Willis, Brian P.

2001-01-01

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

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.

An assessment of laser velocimetry in hypersonic flow

NASA Technical Reports Server (NTRS)

1992-01-01

Although extensive progress has been made in computational fluid mechanics, reliable flight vehicle designs and modifications still cannot be made without recourse to extensive wind tunnel testing. Future progress in the computation of hypersonic flow fields is restricted by the need for a reliable mean flow and turbulence modeling data base which could be used to aid in the development of improved empirical models for use in numerical codes. Currently, there are few compressible flow measurements which could be used for this purpose. In this report, the results of experiments designed to assess the potential for laser velocimeter measurements of mean flow and turbulent fluctuations in hypersonic flow fields are presented. Details of a new laser velocimeter system which was designed and built for this test program are described.

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.

Wind Integration National Dataset Toolkit | Grid Modernization | NREL

Science.gov Websites

information, share tips The WIND Toolkit includes meteorological conditions and turbine power for more than Integration National Dataset Toolkit Wind Integration National Dataset Toolkit The Wind Integration National Dataset (WIND) Toolkit is an update and expansion of the Eastern Wind Integration Data Set and

Thermal and Pressure Characterization of a Wind Tunnel Force Balance Using the Single Vector System. Experimental Design and Analysis Approach to Model Pressure and Temperature Effects in Hypersonic Wind Tunnel Research

NASA Technical Reports Server (NTRS)

Lynn, Keith C.; Commo, Sean A.; Johnson, Thomas H.; Parker, Peter A,

2011-01-01

Wind tunnel research at NASA Langley Research Center s 31-inch Mach 10 hypersonic facility utilized a 5-component force balance, which provided a pressurized flow-thru capability to the test article. The goal of the research was to determine the interaction effects between the free-stream flow and the exit flow from the reaction control system on the Mars Science Laboratory aeroshell during planetary entry. In the wind tunnel, the balance was exposed to aerodynamic forces and moments, steady-state and transient thermal gradients, and various internal balance cavity pressures. Historically, these effects on force measurement accuracy have not been fully characterized due to limitations in the calibration apparatus. A statistically designed experiment was developed to adequately characterize the behavior of the balance over the expected wind tunnel operating ranges (forces/moments, temperatures, and pressures). The experimental design was based on a Taylor-series expansion in the seven factors for the mathematical models. Model inversion was required to calculate the aerodynamic forces and moments as a function of the strain-gage readings. Details regarding transducer on-board compensation techniques, experimental design development, mathematical modeling, and wind tunnel data reduction are included in this paper.

Hypersonic Boundary Layer Stability Experiments in a Quiet Wind Tunnel with Bluntness Effects

NASA Technical Reports Server (NTRS)

Lachowicz, Jason T.; Chokani, Ndaona

1996-01-01

Hypersonic boundary layer measurements over a flared cone were conducted in a Mach 6 quiet wind tunnel at a freestream unit Reynolds number of 2.82 million/ft. This Reynolds number provided laminar-to-transitional flow over the cone model in a low-disturbance environment. Four interchangeable nose-tips, including a sharp-tip, were tested. Point measurements with a single hot-wire using a novel constant voltage anemometer were used to measure the boundary layer disturbances. Surface temperature and schlieren measurements were also conducted to characterize the transitional state of the boundary layer and to identify instability modes. Results suggest that second mode disturbances were the most unstable and scaled with the boundary layer thickness. The second mode integrated growth rates compared well with linear stability theory in the linear stability regime. The second mode is responsible for transition onset despite the existence of a second mode subharmonic. The subharmonic disturbance wavelength also scales with the boundary layer thickness. Furthermore, the existence of higher harmonics of the fundamental suggests that nonlinear disturbances are not associated with 'high' free stream disturbance levels. Nose-tip radii greater than 2.7% of the base radius completely stabilized the second mode.

Features of the Upgraded Imaging for Hypersonic Experimental Aeroheating Testing (IHEAT) Software

NASA Technical Reports Server (NTRS)

Mason, Michelle L.; Rufer, Shann J.

2016-01-01

The Imaging for Hypersonic Experimental Aeroheating Testing (IHEAT) software is used at the NASA Langley Research Center to analyze global aeroheating data on wind tunnel models tested in the Langley Aerothermodynamics Laboratory. One-dimensional, semi-infinite heating data derived from IHEAT are used in the design of thermal protection systems for hypersonic vehicles that are exposed to severe aeroheating loads, such as reentry vehicles during descent and landing procedures. This software program originally was written in the PV-WAVE(Registered Trademark) programming language to analyze phosphor thermography data from the two-color, relative-intensity system developed at Langley. To increase the efficiency, functionality, and reliability of IHEAT, the program was migrated to MATLAB(Registered Trademark) syntax and compiled as a stand-alone executable file labeled version 4.0. New features of IHEAT 4.0 include the options to perform diagnostic checks of the accuracy of the acquired data during a wind tunnel test, to extract data along a specified multi-segment line following a feature such as a leading edge or a streamline, and to batch process all of the temporal frame data from a wind tunnel run. Results from IHEAT 4.0 were compared on a pixel level to the output images from the legacy software to validate the program. The absolute differences between the heat transfer data output from the two programs were on the order of 10(exp -5) to 10(exp -7). IHEAT 4.0 replaces the PV-WAVE(Registered Trademark) version as the production software for aeroheating experiments conducted in the hypersonic facilities at NASA Langley.

Wind Integration National Dataset (WIND) Toolkit; NREL (National Renewable Energy Laboratory)

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

Draxl, Caroline; Hodge, Bri-Mathias

A webinar about the Wind Integration National Dataset (WIND) Toolkit was presented by Bri-Mathias Hodge and Caroline Draxl on July 14, 2015. It was hosted by the Southern Alliance for Clean Energy. The toolkit is a grid integration data set that contains meteorological and power data at a 5-minute resolution across the continental United States for 7 years and hourly power forecasts.

Application of unsteady aeroelastic analysis techniques on the national aerospace plane

NASA Technical Reports Server (NTRS)

Pototzky, Anthony S.; Spain, Charles V.; Soistmann, David L.; Noll, Thomas E.

1988-01-01

A presentation provided at the Fourth National Aerospace Plane Technology Symposium held in Monterey, California, in February 1988 is discussed. The objective is to provide current results of ongoing investigations to develop a methodology for predicting the aerothermoelastic characteristics of NASP-type (hypersonic) flight vehicles. Several existing subsonic and supersonic unsteady aerodynamic codes applicable to the hypersonic class of flight vehicles that are generally available to the aerospace industry are described. These codes were evaluated by comparing calculated results with measured wind-tunnel aeroelastic data. The agreement was quite good in the subsonic speed range but showed mixed agreement in the supersonic range. In addition, a future endeavor to extend the aeroelastic analysis capability to hypersonic speeds is outlined. An investigation to identify the critical parameters affecting the aeroelastic characteristics of a hypersonic vehicle, to define and understand the various flutter mechanisms, and to develop trends for the important parameters using a simplified finite element model of the vehicle is summarized. This study showed the value of performing inexpensive and timely aeroelastic wind-tunnel tests to expand the experimental data base required for code validation using simple to complex models that are representative of the NASP configurations and root boundary conditions are discussed.

Hyper-X: Flight Validation of Hypersonic Airbreathing Technology

NASA Technical Reports Server (NTRS)

Rausch, Vincent L.; McClinton, Charles R.; Crawford, J. Larry

1997-01-01

This paper provides an overview of NASA's focused hypersonic technology program, i.e. the Hyper-X program. This program is designed to move hypersonic, air breathing vehicle technology from the laboratory environment to the flight environment, the last stage preceding prototype development. This paper presents some history leading to the flight test program, research objectives, approach, schedule and status. Substantial experimental data base and concept validation have been completed. The program is concentrating on Mach 7 vehicle development, verification and validation in preparation for wind tunnel testing in 1998 and flight testing in 1999. It is also concentrating on finalization of the Mach 5 and 10 vehicle designs. Detailed evaluation of the Mach 7 vehicle at the flight conditions is nearing completion, and will provide a data base for validation of design methods once flight test data are available.

Results of pressure distribution tests of a 0.010-scale space shuttle orbiter model (61-0) in the NASA/ARC 3.5-foot hypersonic wind tunnel (test OH38), volume 1

NASA Technical Reports Server (NTRS)

Dye, W. H.; Polek, T.

1975-01-01

Test results are presented of hypersonic pressure distributions at simulated atmospheric entry conditions. Pressure data were obtained at Mach numbers of 7.4 and 10.4 and Reynolds numbers of 3.0 and 6.5 million per foot. Data are presented in both plotted and tabulated data form. Photographs of wind tunnel apparatus and test configurations are provided.

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.

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.

Adaptive computations of multispecies mixing between scramjet nozzle flows and hypersonic freestream

NASA Technical Reports Server (NTRS)

Baysa, Oktay; Engelund, Walter C.; Eleshaky, Mohamed E.; Pittman, James L.

1989-01-01

The objective of this paper is to compute the expansion of a supersonic flow through an internal-external nozzle and its viscous mixing with the hypersonic flow of air. The supersonic jet may be that of a multispecies gas other than air. Calculations are performed for one case where both flows are those of air, and another case where a mixture of freon-12 and argon is discharged supersonically to mix with the hypersonic airflow. Comparisons are made between these two cases with respect to gas compositions, and fixed versus flow-adaptive grids. All the computational results are compared successfully with the wind-tunnel tests results.

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.

Fluorescence Visualization of Hypersonic Flow over Rapid Prototype Wind-Tunnel Models

NASA Technical Reports Server (NTRS)

Alderfer, D. W.; Danehy, P. M.; Inma, J. A.; Berger, K. T.; Buck, G. M.; Schwartz, R J.

2007-01-01

Reentry models for use in hypersonic wind tunnel tests were fabricated using a stereolithography apparatus. These models were produced in one day or less, which is a significant time savings compared to the manufacture of ceramic or metal models. The models were tested in the NASA Langley Research Center 31-Inch Mach 10 Air Tunnel. Most of the models did not survive repeated tests in the tunnel, and several failure modes of the models were identified. Planar laser-induced fluorescence (PLIF) of nitric oxide (NO) was used to visualize the flowfields in the wakes of these models. Pure NO was either seeded through tubes plumbed into the model or via a tube attached to the strut holding the model, which provided localized addition of NO into the model s wake through a porous metal cylinder attached to the end of the tube. Models included several 2-inch diameter Inflatable Reentry Vehicle Experiment (IRVE) models and 5-inch diameter Crew Exploration Vehicle (CEV) models. Various configurations were studied including different sting placements relative to the models, different model orientations and attachment angles, and different NO seeding methods. The angle of attack of the models was also varied and the location of the laser sheet was scanned to provide three-dimensional flowfield information. Virtual Diagnostics Interface technology, developed at NASA Langley, was used to visualize the data sets in post processing. The use of calibration "dotcards" was investigated to correct for camera perspective and lens distortions in the PLIF images. Lessons learned and recommendations for future experiments are discussed.

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.

National Offshore Wind Energy Grid Interconnection Study

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

Daniel, John P.; Liu, Shu; Ibanez, Eduardo

2014-07-30

The National Offshore Wind Energy Grid Interconnection Study (NOWEGIS) considers the availability and potential impacts of interconnecting large amounts of offshore wind energy into the transmission system of the lower 48 contiguous United States. A total of 54GW of offshore wind was assumed to be the target for the analyses conducted. A variety of issues are considered including: the anticipated staging of offshore wind; the offshore wind resource availability; offshore wind energy power production profiles; offshore wind variability; present and potential technologies for collection and delivery of offshore wind energy to the onshore grid; potential impacts to existing utility systemsmore » most likely to receive large amounts of offshore wind; and regulatory influences on offshore wind development. The technologies considered the reliability of various high-voltage ac (HVAC) and high-voltage dc (HVDC) technology options and configurations. The utility system impacts of GW-scale integration of offshore wind are considered from an operational steady-state perspective and from a regional and national production cost perspective.« less

An Upgrade of the Imaging for Hypersonic Experimental Aeroheating Testing (IHEAT) Software

NASA Technical Reports Server (NTRS)

Mason, Michelle L.; Rufer, Shann J.

2015-01-01

The Imaging for Hypersonic Experimental Aeroheating Testing (IHEAT) code is used at NASA Langley Research Center to analyze global aeroheating data on wind tunnel models tested in the Langley Aerothermodynamics Laboratory. One-dimensional, semi-infinite heating data derived from IHEAT are used to design thermal protection systems to mitigate the risks due to the aeroheating loads on hypersonic vehicles, such as re-entry vehicles during descent and landing procedures. This code was originally written in the PV-WAVE programming language to analyze phosphor thermography data from the two-color, relativeintensity system developed at Langley. To increase the efficiency, functionality, and reliability of IHEAT, the code was migrated to MATLAB syntax and compiled as a stand-alone executable file labeled version 4.0. New features of IHEAT 4.0 include the options to batch process all of the data from a wind tunnel run, to map the two-dimensional heating distribution to a three-dimensional computer-aided design model of the vehicle to be viewed in Tecplot, and to extract data from a segmented line that follows an interesting feature in the data. Results from IHEAT 4.0 were compared on a pixel level to the output images from the legacy code to validate the program. The differences between the two codes were on the order of 10-5 to 10-7. IHEAT 4.0 replaces the PV-WAVE version as the production code for aeroheating experiments conducted in the hypersonic facilities at NASA Langley.

CFD analysis of hypersonic, chemically reacting flow fields

NASA Technical Reports Server (NTRS)

Edwards, T. A.

1993-01-01

Design studies are underway for a variety of hypersonic flight vehicles. The National Aero-Space Plane will provide a reusable, single-stage-to-orbit capability for routine access to low earth orbit. Flight-capable satellites will dip into the atmosphere to maneuver to new orbits, while planetary probes will decelerate at their destination by atmospheric aerobraking. To supplement limited experimental capabilities in the hypersonic regime, computational fluid dynamics (CFD) is being used to analyze the flow about these configurations. The governing equations include fluid dynamic as well as chemical species equations, which are being solved with new, robust numerical algorithms. Examples of CFD applications to hypersonic vehicles suggest an important role this technology will play in the development of future aerospace systems. The computational resources needed to obtain solutions are large, but solution adaptive grids, convergence acceleration, and parallel processing may make run times manageable.

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.

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.

75 FR 4417 - Wind Cave National Park, Custer County, SD

Federal Register 2010, 2011, 2012, 2013, 2014

2010-01-27

... DEPARTMENT OF THE INTERIOR Wind Cave National Park, Custer County, SD AGENCY: National Park... Final Environmental Impact Statement, Wind Cave National Park, Custer County, South Dakota. SUMMARY... Management Plan and Final Environmental Impact Statement (Plan), Wind Cave National Park, Custer County...

Hypersonic missile propulsion system

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

Kazmar, R.R.

1998-11-01

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

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.

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.

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.

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.

A Compact Tunable Diode Laser Absorption Spectrometer to Monitor CO2 at 2.7 μm Wavelength in Hypersonic Flows

PubMed Central

Vallon, Raphäel; Soutadé, Jacques; Vérant, Jean-Luc; Meyers, Jason; Paris, Sébastien; Mohamed, Ajmal

2010-01-01

Since the beginning of the Mars planet exploration, the characterization of carbon dioxide hypersonic flows to simulate a spaceship’s Mars atmosphere entry conditions has been an important issue. We have developed a Tunable Diode Laser Absorption Spectrometer with a new room-temperature operating antimony-based distributed feedback laser (DFB) diode laser to characterize the velocity, the temperature and the density of such flows. This instrument has been tested during two measurement campaigns in a free piston tunnel cold hypersonic facility and in a high enthalpy arc jet wind tunnel. These tests also demonstrate the feasibility of mid-infrared fiber optics coupling of the spectrometer to a wind tunnel for integrated or local flow characterization with an optical probe placed in the flow. PMID:22219703

A compact tunable diode laser absorption spectrometer to monitor CO2 at 2.7 μm wavelength in hypersonic flows.

PubMed

Vallon, Raphäel; Soutadé, Jacques; Vérant, Jean-Luc; Meyers, Jason; Paris, Sébastien; Mohamed, Ajmal

2010-01-01

Since the beginning of the Mars planet exploration, the characterization of carbon dioxide hypersonic flows to simulate a spaceship's Mars atmosphere entry conditions has been an important issue. We have developed a Tunable Diode Laser Absorption Spectrometer with a new room-temperature operating antimony-based distributed feedback laser (DFB) diode laser to characterize the velocity, the temperature and the density of such flows. This instrument has been tested during two measurement campaigns in a free piston tunnel cold hypersonic facility and in a high enthalpy arc jet wind tunnel. These tests also demonstrate the feasibility of mid-infrared fiber optics coupling of the spectrometer to a wind tunnel for integrated or local flow characterization with an optical probe placed in the flow.

Heating Augmentation for Short Hypersonic Protuberances

NASA Technical Reports Server (NTRS)

Mazaheri, Alireza R.; Wood, William A.

2008-01-01

Computational aeroheating analyses of the Space Shuttle Orbiter plug repair models are validated against data collected in the Calspan University of Buffalo Research Center (CUBRC) 48 inch shock tunnel. The comparison shows that the average difference between computed heat transfer results and the data is about 9:5%. Using CFD and Wind Tunnel (WT) data, an empirical correlation for estimating heating augmentation on short hyper- sonic protuberances (k/delta < 0.33) is proposed. This proposed correlation is compared with several computed flight simulation cases and good agreement is achieved. Accordingly, this correlation is proposed for further investigation on other short hypersonic protuberances for estimating heating augmentation.

Results of heat transfer tests of an 0.0175-scale space shuttle vehicle model 22 OTS in the NASA-Ames 3.5-foot hypersonic wind tunnel (IH3), volume 4

NASA Technical Reports Server (NTRS)

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

1975-01-01

Heat-transfer data for the 0.0175-scale Space Shuttle Vehicle 3 are presented. Interference heating effects were investigated by a model build-up technique of Orbiter alone, tank alone, second, and first stage configurations. The test program was conducted in the NASA-Ames 3.5-Foot Hypersonic Wind Tunnel at Mach 5.3 for nominal free-stream Reynolds number per foot values of 1.5 x 1,000,000 and 5.0 x 1,000,000.

Hypersonic aerodynamic characteristics of an all-body research aircraft configuration

NASA Technical Reports Server (NTRS)

Clark, L. E.

1973-01-01

An experimental investigation was conducted at Mach 6 to determine the hypersonic aerodynamic characteristics of an all-body, delta-planform, hypersonic research aircraft (HYFAC configuration). The aerodynamic characteristics were obtained at Reynolds numbers based on model length of 2.84 million and 10.5 million and over an angle-of-attack range from minus 4 deg to 20 deg. The experimental results show that the HYFAC configuration is longitudinally stable and can be trimmed over the range of test conditions. The configuration had a small degree of directional stability over the angle-of-attack range and positive effective dihedral at angles of attack greater than 2 deg. Addition of canards caused a decrease in longitudinal stability and an increase in directional stability. Oil-flow studies revealed extensive areas of separated and vortex flow on the fuselage lee surface. A limited comparison of wind-tunnel data with several hypersonic approximations indicated that, except for the directional stability, the tangent-cone method gave adequate agreement at control settings between 5 deg and minus 5 deg and positive lift coefficient. A limited comparison indicated that the HYFAC configuration had greater longitudinal stability than an elliptical-cross-section configuration, but a lower maximum lift-drag ratio.

Advancements in Wind Integration Study Data Modeling: The Wind Integration National Dataset (WIND) Toolkit; Preprint

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

Draxl, C.; Hodge, B. M.; Orwig, K.

2013-10-01

Regional wind integration studies in the United States require detailed wind power output data at many locations to perform simulations of how the power system will operate under high-penetration scenarios. The wind data sets that serve as inputs into the study must realistically reflect the ramping characteristics, spatial and temporal correlations, and capacity factors of the simulated wind plants, as well as be time synchronized with available load profiles. The Wind Integration National Dataset (WIND) Toolkit described in this paper fulfills these requirements. A wind resource dataset, wind power production time series, and simulated forecasts from a numerical weather predictionmore » model run on a nationwide 2-km grid at 5-min resolution will be made publicly available for more than 110,000 onshore and offshore wind power production sites.« less

Investigation of flow fields within large scale hypersonic inlet models

NASA Technical Reports Server (NTRS)

Gnos, A. V.; Watson, E. C.; Seebaugh, W. R.; Sanator, R. J.; Decarlo, J. P.

1973-01-01

Analytical and experimental investigations were conducted to determine the internal flow characteristics in model passages representative of hypersonic inlets for use at Mach numbers to about 12. The passages were large enough to permit measurements to be made in both the core flow and boundary layers. The analytical techniques for designing the internal contours and predicting the internal flow-field development accounted for coupling between the boundary layers and inviscid flow fields by means of a displacement-thickness correction. Three large-scale inlet models, each having a different internal compression ratio, were designed to provide high internal performance with an approximately uniform static-pressure distribution at the throat station. The models were tested in the Ames 3.5-Foot Hypersonic Wind Tunnel at a nominal free-stream Mach number of 7.4 and a unit free-stream Reynolds number of 8.86 X one million per meter.

Validation of WIND for a Series of Inlet Flows

NASA Technical Reports Server (NTRS)

Slater, John W.; Abbott, John M.; Cavicchi, Richard H.

2002-01-01

Validation assessments compare WIND CFD simulations to experimental data for a series of inlet flows ranging in Mach number from low subsonic to hypersonic. The validation procedures follow the guidelines of the AIAA. The WIND code performs well in matching the available experimental data. The assessments demonstrate the use of WIND and provide confidence in its use for the analysis of aircraft inlets.

Evaluation of candidate working fluid formulations for the electrothermal-chemical wind tunnel

NASA Technical Reports Server (NTRS)

Akyurtlu, Jale F.; Akyurtlu, Ates

1993-01-01

A new hypersonic test facility which can simulate conditions typical of atmospheric flight at Mach numbers up to 20 is currently under study at the NASA/LaRC Hypersonic Propulsion Branch. In the proposed research, it was suggested that a combustion augmented electrothermal wind tunnel concept may be applied to the planned hypersonic testing facility. The purpose of the current investigation is to evaluate some candidate working fluid formulations which may be used in the chemical-electrothermal wind. The efforts in the initial phase of this research were concentrated on acquiring the code used by GASL to model the electrothermal wind tunnel and testing it using the conditions of GASL simulation. The early version of the general chemical kinetics code (GCKP84) was obtained from NASA and the latest updated version of the code (LSENS) was obtained from the author Dr. Bittker. Both codes are installed on a personal computer with a 486 25 MHz processor and 16 Mbyte RAM. Since the available memory was not sufficient to debug LSENS, for the current work GCKP84 was used.

Hypersonic Navier Stokes Comparisons to Orbiter Flight Data

NASA Technical Reports Server (NTRS)

Campbell, Charles H.; Nompelis, Ioannis; Candler, Graham; Barnhart, Michael; Yoon, Seokkwan

2009-01-01

Hypersonic chemical nonequilibrium simulations of low earth orbit entry flow fields are becoming increasingly commonplace as software and computational capabilities become more capable. However, development of robust and accurate software to model these environments will always encounter a significant barrier in developing a suite of high quality calibration cases. The US3D hypersonic nonequilibrium Navier Stokes analysis capability has been favorably compared to a number of wind tunnel test cases. Extension of the calibration basis for this software to Orbiter flight conditions will provide an incremental increase in confidence. As part of the Orbiter Boundary Layer Transition Flight Experiment and the Hypersonic Thermodynamic Infrared Measurements project, NASA is performing entry flight testing on the Orbiter to provide valuable aerothermodynamic heating data. An increase in interest related to orbiter entry environments is resulting from this activity. With the advent of this new data, comparisons of the US3D software to the new flight testing data is warranted. This paper will provide information regarding the framework of analyses that will be applied with the US3D analysis tool. In addition, comparisons will be made to entry flight testing data provided by the Orbiter BLT Flight Experiment and HYTHIRM projects. If data from digital scans of the Orbiter windward surface become available, simulations will also be performed to characterize the difference in surface heating between the CAD reference OML and the digitized surface provided by the surface scans.

National Aero-Space Plane (NASP) program

NASA Technical Reports Server (NTRS)

Tank, Ming H.

1991-01-01

A program to develop the technology for reusable airbreathing hypersonic/transatmospheric vehicles is addressed. Information on the following topics is presented in viewgraph form: (1) the National Aerospace Plane (NASP) program schedule; (2) the NASP program organization; (3) competitive strategy; (4) propulsion options; (5) wind tunnel data available for NASP; (6) ground track of envelope expansion; and (7) altitude vs. Mach number. A NASP/Space Shuttle comparison, NASP configuration matrix, and the propulsion concept of a high speed scramjet are also briefly addressed.

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.

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.

Stability of hypersonic compression cones

NASA Astrophysics Data System (ADS)

Reed, Helen; Kuehl, Joseph; Perez, Eduardo; Kocian, Travis; Oliviero, Nicholas

2012-11-01

Our activities focus on the identification and understanding of the second-mode instability for representative configurations in hypersonic flight. These include the Langley 93-10 flared cone and the Purdue compression cone, both at 0 degrees angle of attack at Mach 6. Through application of nonlinear parabolized stability equations (NPSE) and linear parabolized stability equations (PSE) to both geometries, it is concluded that mean-flow distortion tends to amplify frequencies less than the peak frequency and stabilize those greater by modifying the boundary-layer thickness. As initial disturbance amplitude is increased and/or a broad spectrum disturbance is introduced, direct numerical simulations (DNS) or NPSE appear to be the proper choices to model the evolution, and relative evolution, because these computational tools include these nonlinear effects (mean-flow distortion). Support from AFOSR/NASA National Center for Hypersonic Research in Laminar-Turbulent Transition through Grant FA9550-09-1-0341 is gratefully acknowledged. The authors also thank Pointwise, AeroSoft, and Texas Advanced Computing Center (TACC).

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.

An inventory of aeronautical ground research facilities. Volume 1: Wind tunnels

NASA Technical Reports Server (NTRS)

Pirrello, C. J.; Hardin, R. D.; Heckart, M. V.; Brown, K. R.

1971-01-01

A survey of wind tunnel research facilities in the United States is presented. The inventory includes all subsonic, transonic, and hypersonic wind tunnels operated by governmental and private organizations. Each wind tunnel is described with respect to size, mechanical operation, construction, testing capabilities, and operating costs. Facility performance data are presented in charts and tables.

Fluorescence Imaging and Streamline Visualization of Hypersonic Flow over Rapid Prototype Wind-Tunnel Models

NASA Technical Reports Server (NTRS)

Danehy, Paul M.; Alderfer, David W.; Inman, Jennifer A.; Berger, Karen T.; Buck, Gregory M.; Schwartz, Richard J.

2008-01-01

Reentry models for use in hypersonic wind tunnel tests were fabricated using a stereolithography apparatus. These models were produced in one day or less, which is a significant time savings compared to the manufacture of ceramic or metal models. The models were tested in the NASA Langley Research Center 31-Inch Mach 10 Air Tunnel. Only a few of the models survived repeated tests in the tunnel, and several failure modes of the models were identified. Planar laser-induced fluorescence (PLIF) of nitric oxide (NO) was used to visualize the flowfields in the wakes of these models. Pure NO was either seeded through tubes plumbed into the model or via a tube attached to the strut holding the model, which provided localized addition of NO into the model s wake through a porous metal cylinder attached to the end of the tube. Models included several 2- inch diameter Inflatable Reentry Vehicle Experiment (IRVE) models and 5-inch diameter Crew Exploration Vehicle (CEV) models. Various model configurations and NO seeding methods were used, including a new streamwise visualization method based on PLIF. Virtual Diagnostics Interface (ViDI) technology, developed at NASA Langley Research Center, was used to visualize the data sets in post processing. The use of calibration "dotcards" was investigated to correct for camera perspective and lens distortions in the PLIF images.

National-Scale Wind Resource Assessment for Power Generation (Presentation)

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

Baring-Gould, E. I.

2013-08-01

This presentation describes the current standards for conducting a national-scale wind resource assessment for power generation, along with the risk/benefit considerations to be considered when beginning a wind resource assessment. The presentation describes changes in turbine technology and viable wind deployment due to more modern turbine technology and taller towers and shows how the Philippines national wind resource assessment evolved over time to reflect changes that arise from updated technologies and taller towers.

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.

NREL Researchers Play Integral Role in National Offshore Wind Strategy |

Science.gov Websites

News | NREL Researchers Play Integral Role in National Offshore Wind Strategy NREL Researchers Play Integral Role in National Offshore Wind Strategy December 16, 2016 A photo of three offshore wind turbines in turbulent water. Offshore wind energy in the United States is just getting started, with the

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.

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.

Heat transfer phase change paint tests of 0.0175-scale models (nos. 21-0 and 46-0) of the Rockwell International space shuttle orbiter in the AEDC tunnel B hypersonic wind tunnel (test OH25A)

NASA Technical Reports Server (NTRS)

Dye, W. H.

1975-01-01

Tests were conducted in a hypersonic wind tunnel using various truncated space shuttle orbiter configurations in an attempt to establish the optimum model size for other tests examining body shock-wing leading edge interference effects. The tests were conducted at Mach number 8 using the phase change paint technique. A test description, tabulated data, and tracings of isotherms made from photographs taken during the test are presented.

Aerothermodynamics of expert ballistic vehicle at hypersonic speeds

NASA Astrophysics Data System (ADS)

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

2012-01-01

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

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

Multiple Velocity Profile Measurements in Hypersonic Flows using Sequentially-Imaged Fluorescence Tagging

NASA Technical Reports Server (NTRS)

Bathel, Brett F.; Danehy, Paul M.; Inmian, Jennifer A.; Jones, Stephen B.; Ivey, Christopher B.; Goyne, Christopher P.

2010-01-01

Nitric-oxide planar laser-induced fluorescence (NO PLIF) was used to perform velocity measurements in hypersonic flows by generating multiple tagged lines which fluoresce as they convect downstream. For each laser pulse, a single interline, progressive scan intensified CCD camera was used to obtain separate images of the initial undelayed and delayed NO molecules that had been tagged by the laser. The CCD configuration allowed for sub-microsecond acquisition of both images, resulting in sub-microsecond temporal resolution as well as sub-mm spatial resolution (0.5-mm x 0.7-mm). Determination of axial velocity was made by application of a cross-correlation analysis of the horizontal shift of individual tagged lines. Quantification of systematic errors, the contribution of gating/exposure duration errors, and influence of collision rate on fluorescence to temporal uncertainty were made. Quantification of the spatial uncertainty depended upon the analysis technique and signal-to-noise of the acquired profiles. This investigation focused on two hypersonic flow experiments: (1) a reaction control system (RCS) jet on an Orion Crew Exploration Vehicle (CEV) wind tunnel model and (2) a 10-degree half-angle wedge containing a 2-mm tall, 4-mm wide cylindrical boundary layer trip. The experiments were performed at the NASA Langley Research Center's 31-inch Mach 10 wind tunnel.

75 FR 2844 - Green Mountam National Forest; Vermont; Deerfield Wind Project

Federal Register 2010, 2011, 2012, 2013, 2014

2010-01-19

... DEPARTMENT OF AGRICULTURE Forest Service Green Mountam National Forest; Vermont; Deerfield Wind... authorization from Deerfield Wind, LLC for the installation and operation of wind turbines on National Forest... original notice of intent to prepare the Deerfield Wind Project EIS was published in the Federal Register...

Hypersonic transport technology

NASA Technical Reports Server (NTRS)

Kirkham, F. S.; Hunt, J. L.

1977-01-01

The possibility of developing a hypersonic commercial transport (HST) for the early 21st century is explored in terms of potential performance characteristics and recent advances in propulsion and structures technology. Range-gross weight characteristics indicate that a 200 passenger, Mach 6 aircraft with a range of 9200 km (5000 n.mi.) would have a gross take-off weight not too different from that of current wide-body subsonic transports. The low cruise sonic boom overpressures generated by the HST opens the possibility of supersonic overland flight which, if permitted, would greatly increase the attractiveness of this type of aircraft. Recent advances in hypersonic propulsion systems and long-lived hypersonic aircraft structure are also discussed. The airframe-integrated scramjet and the actively-cooled airframe structure are identified as the most promising candidates for the HST and current approaches are described in some detail.

2015 Key Wind Program and National Laboratory Accomplishments

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

Office of Energy Efficiency and Renewable Energy

The U.S. Department of Energy (DOE) Wind Program is committed to helping the nation secure cost-competitive sources of renewable energy through the development and deployment of innovative wind power technologies. By investing in improvements to wind plant design, technology development, and operation as well as developing tools to identify the highest quality wind resources, the Wind Program serves as a leader in making wind energy technologies more competitive with traditional sources of energy and a larger part of our nation’s renewable energy portfolio.

Overview and Meteorological Validation of the Wind Integration National Dataset toolkit

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

Draxl, C.; Hodge, B. M.; Clifton, A.

2015-04-13

The Wind Integration National Dataset (WIND) Toolkit described in this report fulfills these requirements, and constitutes a state-of-the-art national wind resource data set covering the contiguous United States from 2007 to 2013 for use in a variety of next-generation wind integration analyses and wind power planning. The toolkit is a wind resource data set, wind forecast data set, and wind power production and forecast data set derived from the Weather Research and Forecasting (WRF) numerical weather prediction model. WIND Toolkit data are available online for over 116,000 land-based and 10,000 offshore sites representing existing and potential wind facilities.

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.

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.

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

National Wind Technology Center Video (Text Version) | NREL

Science.gov Websites

blades of wind turbines and the National Renewable Energy Laboratory logo. It then cuts to images of able to do the testing." (Voiceover) To create energy, the wind turns the blades, which rotate the testing." Inside the Industrial User Facility, wind turbine blades are shaken and flexed. (Voiceover

Space shuttle orbiter rear mounted reaction control system jet interaction study. [hypersonic wind tunnel tests

NASA Technical Reports Server (NTRS)

Rausch, J. R.

1977-01-01

The effect of interaction between the reaction control system (RCS) jets and the flow over the space shuttle orbiter in the atmosphere was investigated in the NASA Langley 31-inch continuous flow hypersonic tunnel at a nominal Mach number of 10.3 and in the AEDC continuous flow hypersonic tunnel B at a nominal Mach number of 6, using 0.01 and .0125 scale force models with aft RCS nozzles mounted both on the model and on the sting of the force model balance. The data show that RCS nozzle exit momentum ratio is the primary correlating parameter for effects where the plume impinges on an adjacent surface and mass flow ratio is the parameter when the plume interaction is primarily with the external stream. An analytic model of aft mounted RCS units was developed in which the total reaction control moments are the sum of thrust, impingement, interaction, and cross-coupling terms.

Computational and experimental investigation of two-dimensional scramjet inlets and hypersonic flow over a sharp flat plate

NASA Astrophysics Data System (ADS)

Messitt, Donald G.

1999-11-01

The WIND code was employed to compute the hypersonic flow in the shock wave boundary layer merged region near the leading edge of a sharp flat plate. Solutions were obtained at Mach numbers from 9.86 to 15.0 and free stream Reynolds numbers of 3,467 to 346,700 in-1 (1.365 · 105 to 1.365 · 107 m-1) for perfect gas conditions. The numerical results indicated a merged shock wave and viscous layer near the leading edge. The merged region grew in size with increasing free stream Mach number, proportional to Minfinity 2/Reinfinity. Profiles of the static pressure in the merged region indicated a strong normal pressure gradient (∂p/∂y). The normal pressure gradient has been neglected in previous analyses which used the boundary layer equations. The shock wave near the leading edge was thick, as has been experimentally observed. Computed shock wave locations and surface pressures agreed well within experimental error for values of the rarefaction parameter, chi/M infinity2 < 0.3. A preliminary analysis using kinetic theory indicated that rarefied flow effects became important above this value. In particular, the WIND solution agreed well in the transition region between the merged flow, which was predicted well by the theory of Li and Nagamatsu, and the downstream region where the strong interaction theory applied. Additional computations with the NPARC code, WIND's predecessor, demonstrated the ability of the code to compute hypersonic inlet flows at free stream Mach numbers up to 20. Good qualitative agreement with measured pressure data indicated that the code captured the important physical features of the shock wave - boundary layer interactions. The computed surface and pitot pressures fell within the combined experimental and numerical error bounds for most points. The calculations demonstrated the need for extremely fine grids when computing hypersonic interaction flows.

Parametric Analysis of a Hypersonic Inlet using Computational Fluid Dynamics

NASA Astrophysics Data System (ADS)

Oliden, Daniel

For CFD validation, hypersonic flow fields are simulated and compared with experimental data specifically designed to recreate conditions found by hypersonic vehicles. Simulated flow fields on a cone-ogive with flare at Mach 7.2 are compared with experimental data from NASA Ames Research Center 3.5" hypersonic wind tunnel. A parametric study of turbulence models is presented and concludes that the k-kl-omega transition and SST transition turbulence model have the best correlation. Downstream of the flare's shockwave, good correlation is found for all boundary layer profiles, with some slight discrepancies of the static temperature near the surface. Simulated flow fields on a blunt cone with flare above Mach 10 are compared with experimental data from CUBRC LENS hypervelocity shock tunnel. Lack of vibrational non-equilibrium calculations causes discrepancies in heat flux near the leading edge. Temperature profiles, where non-equilibrium effects are dominant, are compared with the dissociation of molecules to show the effects of dissociation on static temperature. Following the validation studies is a parametric analysis of a hypersonic inlet from Mach 6 to 20. Compressor performance is investigated for numerous cowl leading edge locations up to speeds of Mach 10. The variable cowl study showed positive trends in compressor performance parameters for a range of Mach numbers that arise from maximizing the intake of compressed flow. An interesting phenomenon due to the change in shock wave formation for different Mach numbers developed inside the cowl that had a negative influence on the total pressure recovery. Investigation of the hypersonic inlet at different altitudes is performed to study the effects of Reynolds number, and consequently, turbulent viscous effects on compressor performance. Turbulent boundary layer separation was noted as the cause for a change in compressor performance parameters due to a change in Reynolds number. This effect would not be

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.

National Wind Tunnel Complex (NWTC)

NASA Technical Reports Server (NTRS)

1996-01-01

The National Wind Tunnel Complex (NWTC) Final Report summarizes the work carried out by a unique Government/Industry partnership during the period of June 1994 through May 1996. The objective of this partnership was to plan, design, build and activate 'world class' wind tunnel facilities for the development of future-generation commercial and military aircraft. The basis of this effort was a set of performance goals defined by the National Facilities Study (NFS) Task Group on Aeronautical Research and Development Facilities which established two critical measures of improved wind tunnel performance; namely, higher Reynolds number capability and greater productivity. Initial activities focused upon two high-performance tunnels (low-speed and transonic). This effort was later descoped to a single multipurpose tunnel. Beginning in June 1994, the NWTC Project Office defined specific performance requirements, planned site evaluation activities, performed a series of technical/cost trade studies, and completed preliminary engineering to support a proposed conceptual design. Due to budget uncertainties within the Federal government, the NWTC project office was directed to conduct an orderly closure following the Systems Design Review in March 1996. This report provides a top-level status of the project at that time. Additional details of all work performed have been archived and are available for future reference.

Vortex lattice prediction of subsonic aerodynamics of hypersonic vehicle concepts

NASA Technical Reports Server (NTRS)

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

1977-01-01

The vortex lattice method introduced by Lamar and Gloss (1975) was applied to the prediction of subsonic aerodynamic characteristics of hypersonic body-wing configurations. The reliability of the method was assessed through comparison of the calculated and observed aerodynamic performances of two National Hypersonic Flight Research Facility craft at Mach 0.2. The investigation indicated that a vortex lattice model involving 120 or more panel elements can give good results for the lift and induced drag coefficients of the craft, as well as for the pitching moment at angles of attack below 10 to 15 deg. Automated processes for calculating the local slopes of mean-camber surfaces may also render the method suitable for use in preliminary design phases.

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

NASA Technical Reports Server (NTRS)

Bui, Trong T.; Lux, David P.; Stenger, Michael T.; Munson, Michael J.; Teate, George F.

2007-01-01

The Phoenix Air-Launched Small Missile (ALSM) flight testbed was conceived and is proposed to help address the lack of quick-turnaround and cost-effective hypersonic flight research capabilities. The Phoenix ALSM testbed results from utilization of the United States Navy Phoenix AIM-54 (Hughes Aircraft Company, now Raytheon Company, Waltham, Massachusetts) long-range, guided air-to-air missile and the National Aeronautics and Space Administration (NASA) Dryden Flight Research Center (Edwards, California) F-15B (McDonnell Douglas, now the Boeing Company, Chicago, Illinois) testbed airplane. The retirement of the Phoenix AIM-54 missiles from fleet operation has presented an opportunity for converting this flight asset into a new flight testbed. This cost-effective new platform will fill the gap in the test and evaluation of 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 assist the development and operation of future higher-performance ALSM flight testbeds as well as responsive microsatellite-small-payload air-launched space boosters.

A study of transonic aerodynamic analysis methods for use with a hypersonic aircraft synthesis code

NASA Technical Reports Server (NTRS)

Sandlin, Doral R.; Davis, Paul Christopher

1992-01-01

A means of performing routine transonic lift, drag, and moment analyses on hypersonic all-body and wing-body configurations were studied. The analysis method is to be used in conjunction with the Hypersonic Vehicle Optimization Code (HAVOC). A review of existing techniques is presented, after which three methods, chosen to represent a spectrum of capabilities, are tested and the results are compared with experimental data. The three methods consist of a wave drag code, a full potential code, and a Navier-Stokes code. The wave drag code, representing the empirical approach, has very fast CPU times, but very limited and sporadic results. The full potential code provides results which compare favorably to the wind tunnel data, but with a dramatic increase in computational time. Even more extreme is the Navier-Stokes code, which provides the most favorable and complete results, but with a very large turnaround time. The full potential code, TRANAIR, is used for additional analyses, because of the superior results it can provide over empirical and semi-empirical methods, and because of its automated grid generation. TRANAIR analyses include an all body hypersonic cruise configuration and an oblique flying wing supersonic transport.

Establishment of a National Wind Energy Center at University of Houston

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

Wang, Su Su

The DOE-supported project objectives are to: establish a national wind energy center (NWEC) at University of Houston and conduct research to address critical science and engineering issues for the development of future large MW-scale wind energy production systems, especially offshore wind turbines. The goals of the project are to: (1) establish a sound scientific/technical knowledge base of solutions to critical science and engineering issues for developing future MW-scale large wind energy production systems, (2) develop a state-of-the-art wind rotor blade research facility at the University of Houston, and (3) through multi-disciplinary research, introducing technology innovations on advanced wind-turbine materials, processing/manufacturingmore » technology, design and simulation, testing and reliability assessment methods related to future wind turbine systems for cost-effective production of offshore wind energy. To achieve the goals of the project, the following technical tasks were planned and executed during the period from April 15, 2010 to October 31, 2014 at the University of Houston: (1) Basic research on large offshore wind turbine systems (2) Applied research on innovative wind turbine rotors for large offshore wind energy systems (3) Integration of offshore wind-turbine design, advanced materials and manufacturing technologies (4) Integrity and reliability of large offshore wind turbine blades and scaled model testing (5) Education and training of graduate and undergraduate students and post- doctoral researchers (6) Development of a national offshore wind turbine blade research facility The research program addresses both basic science and engineering of current and future large wind turbine systems, especially offshore wind turbines, for MW-scale power generation. The results of the research advance current understanding of many important scientific issues and provide technical information for solving future large wind turbines with advanced

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.

NREL and Sandia National Laboratories to Sharpen Wind Farm Turbine Controls

Science.gov Websites

| News | NREL NREL and Sandia National Laboratories to Sharpen Wind Farm Turbine Controls NREL and Sandia National Laboratories to Sharpen Wind Farm Turbine Controls April 1, 2016 Researchers at wind turbine modeling. The NREL controls team have been evaluating their control theory in simulations

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.

Sounding Rockets as a Real Flight Platform for Aerothermodynamic Cfd Validation of Hypersonic Flight Experiments

NASA Astrophysics Data System (ADS)

Stamminger, A.; Turner, J.; Hörschgen, M.; Jung, W.

2005-02-01

This paper describes the possibilities of sounding rockets to provide a platform for flight experiments in hypersonic conditions as a supplement to wind tunnel tests. Real flight data from measurement durations longer than 30 seconds can be compared with predictions from CFD calculations. This paper will regard projects flown on sounding rockets, but mainly describe the current efforts at Mobile Rocket Base, DLR on the SHarp Edge Flight EXperiment SHEFEX.

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.

National Offshore Wind Energy Grid Interconnection Study Full Report

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

Daniel, John P.; Liu, Shu; Ibanez, Eduardo

2014-07-30

The National Offshore Wind Energy Grid Interconnection Study (NOWEGIS) considers the availability and potential impacts of interconnecting large amounts of offshore wind energy into the transmission system of the lower 48 contiguous United States.

National Wind Technology Center sitewide, Golden, CO: Environmental assessment

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

NONE

1996-11-01

The National Renewable Energy Laboratory (NREL), the nation`s primary solar and renewable energy research laboratory, proposes to expand its wind technology research and development program activities at its National Wind Technology Center (NWTC) near Golden, Colorado. NWTC is an existing wind energy research facility operated by NREL for the US Department of Energy (DOE). Proposed activities include the construction and reuse of buildings and facilities, installation of up to 20 wind turbine test sites, improvements in infrastructure, and subsequent research activities, technology testing, and site operations. In addition to wind turbine test activities, NWTC may be used to support othermore » NREL program activities and small-scale demonstration projects. This document assesses potential consequences to resources within the physical, biological, and human environment, including potential impacts to: air quality, geology and soils, water resources, biological resources, cultural and historic resources, socioeconomic resources, land use, visual resources, noise environment, hazardous materials and waste management, and health and safety conditions. Comment letters were received from several agencies in response to the scoping and predecisional draft reviews. The comments have been incorporated as appropriate into the document with full text of the letters contained in the Appendices. Additionally, information from the Rocky Flats Environmental Technology Site on going sitewide assessment of potential environmental impacts has been reviewed and discussed by representatives of both parties and incorporated into the document as appropriate.« less

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.

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.

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.

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.

Hypersonic flow analysis

NASA Technical Reports Server (NTRS)

Chow, Chuen-Yen; Ryan, James S.

1987-01-01

While the zonal grid system of Transonic Navier-Stokes (TNS) provides excellent modeling of complex geometries, improved shock capturing, and a higher Mach number range will be required if flows about hypersonic aircraft are to be modeled accurately. A computational fluid dynamics (CFD) code, the Compressible Navier-Stokes (CNS), is under development to combine the required high Mach number capability with the existing TNS geometry capability. One of several candidate flow solvers for inclusion in the CNS is that of F3D. This upwinding flow solver promises improved shock capturing, and more accurate hypersonic solutions overall, compared to the solver currently used in TNS.

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

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.

Final Report on the Creation of the Wind Integration National Dataset (WIND) Toolkit and API: October 1, 2013 - September 30, 2015

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

Hodge, Bri-Mathias

2016-04-08

The primary objective of this work was to create a state-of-the-art national wind resource data set and to provide detailed wind plant output data for specific sites based on that data set. Corresponding retrospective wind forecasts were also included at all selected locations. The combined information from these activities was used to create the Wind Integration National Dataset (WIND), and an extraction tool was developed to allow web-based data access.

Shuttle orbiter boundary layer transition at flight and wind tunnel conditions

NASA Technical Reports Server (NTRS)

Goodrich, W. D.; Derry, S. M.; Bertin, J. J.

1983-01-01

Hypersonic boundary layer transition data obtained on the windward centerline of the Shuttle orbiter during entry for the first five flights are presented and analyzed. Because the orbiter surface is composed of a large number of thermal protection tiles, the transition data include the effects of distributed roughness arising from tile misalignment and gaps. These data are used as a benchmark for assessing and improving the accuracy of boundary layer transition predictions based on correlations of wind tunnel data taken on both aerodynamically rough and smooth orbiter surfaces. By comparing these two data bases, the relative importance of tunnel free stream noise and surface roughness on orbiter boundary layer transition correlation parameters can be assessed. This assessment indicates that accurate predications of transition times can be made for the orbiter at hypersonic flight conditions by using roughness dominated wind tunnel data. Specifically, times of transition onset and completion is accurately predicted using a correlation based on critical and effective values of a roughness Reynolds number previously derived from wind tunnel data.

Instrumentation Development for Large Scale Hypersonic Inflatable Aerodynamic Decelerator Characterization

NASA Technical Reports Server (NTRS)

Swanson, Gregory T.; Cassell, Alan M.

2011-01-01

Hypersonic Inflatable Aerodynamic Decelerator (HIAD) technology is currently being considered for multiple atmospheric entry applications as the limitations of traditional entry vehicles have been reached. The Inflatable Re-entry Vehicle Experiment (IRVE) has successfully demonstrated this technology as a viable candidate with a 3.0 m diameter vehicle sub-orbital flight. To further this technology, large scale HIADs (6.0 8.5 m) must be developed and tested. To characterize the performance of large scale HIAD technology new instrumentation concepts must be developed to accommodate the flexible nature inflatable aeroshell. Many of the concepts that are under consideration for the HIAD FY12 subsonic wind tunnel test series are discussed below.

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.

A Feasibility Study to Evaluate Wind Energy Potential on the Navajo Nation

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

Terry Battiest

2012-11-30

The project, A Feasibility Study to Evaluate Wind Energy Potential on the Navajo Nation, is funded under a solicitation issued by the U.S. Department of Energy Tribal Energy Program. Funding provided by the grant allowed the Navajo Nation to measure wind potential at two sites, one located within the boundaries of the Navajo Nation and the other off-reservation during the project period (September 5, 2005 - September 30, 2009). The recipient for the grant award is the Navajo Tribal Utility Authority (NTUA). The grant allowed the Navajo Nation and NTUA manage the wind feasibility from initial site selection through themore » decision-making process to commit to a site for wind generation development. The grant activities help to develop human capacity at NTUA and help NTUA to engage in renewable energy generation activities, including not only wind but also solar and biomass. The final report also includes information about development activities regarding the sited included in the grant-funded feasibility study.« less

Rekindled vision of hypersonic travel

NASA Technical Reports Server (NTRS)

Colladay, Raymond S.

1987-01-01

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

An Experimental Investigation of Wall-Cooling Effects on Hypersonic Boundary-Layer Stability in a Quiet Wind Tunnel

NASA Technical Reports Server (NTRS)

Blanchard, Alan E.; Selby, Gregory V.

1996-01-01

One of the primary reasons for developing quiet tunnels is for the investigation of high-speed boundary-layer stability and transition phenomena without the transition-promoting effects of acoustic radiation from tunnel walls. In this experiment, a flared-cone model under adiabatic- and cooled-wall conditions was placed in a calibrated, 'quiet' Mach 6 flow and the stability of the boundary layer was investigated using a prototype constant-voltage anemometer. The results were compared with linear-stability theory predictions and good agreement was found in the prediction of second-mode frequencies and growth. In addition, the same 'N=10' criterion used to predict boundary-layer transition in subsonic, transonic, and supersonic flows was found to be applicable for the hypersonic flow regime as well. Under cooled-wall conditions, a unique set of continuous spectra data was acquired that documents the linear, nonlinear, and breakdown regions associated with the transition of hypersonic flow under low-noise conditions.

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.

Limitations of the method of characteristics when applied to axisymmetric hypersonic nozzle design

NASA Technical Reports Server (NTRS)

Edwards, Anne C.; Perkins, John N.; Benton, James R.

1990-01-01

A design study of axisymmetric hypersonic wind tunnel nozzles was initiated by NASA Langley Research Center with the objective of improving the flow quality of their ground test facilities. Nozzles for Mach 6 air, Mach 13.5 nitrogen, and Mach 17 nitrogen were designed using the Method of Characteristics/Boundary Layer (MOC/BL) approach and were analyzed with a Navier-Stokes solver. Results of the analysis agreed well with design for the Mach 6 case, but revealed oblique shock waves of increasing strength originating from near the inflection point of the Mach 13.5 and Mach 17 nozzles. The findings indicate that the MOC/BL design method has a fundamental limitation that occurs at some Mach number between 6 an 13.5. In order to define the limitation more exactly and attempt to discover the cause, a parametric study of hypersonic ideal air nozzles designed with the current MOC/BL method was done. Results of this study indicate that, while stagnations conditions have a moderate affect on the upper limit of the method, the method fails at Mach numbers above 8.0.

National Wind Technology Center Dynamic 5-Megawatt Dynamometer

ScienceCinema

Felker, Fort

2018-06-06

The National Wind Technology Center (NWTC) offers wind industry engineers a unique opportunity to conduct a wide range of tests. Its custom-designed dynamometers can test wind turbine systems from 1 kilowatt (kW) to 5 megawatts (MW). The NWTC's new dynamometer facility simulates operating field conditions to assess the reliability and performance of wind turbine prototypes and commercial machines, thereby reducing deployment time, failures, and maintenance or replacement costs. Funded by the U.S. Department of Energy with American Recovery and Reinvestment Act (ARRA) funds, the 5-MW dynamometer will provide the ability to test wind turbine drivetrains and connect those drivetrains directly to the electricity grid or through a controllable grid interface (CGI). The CGI tests the low-voltage ride-through capability of a drivetrain as well as its response to faults and other abnormal grid conditions.

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.

Research and educational initiatives at the Syracuse University Center for Hypersonics

NASA Technical Reports Server (NTRS)

Spina, E.; Lagraff, J.; Davidson, B.; Bogucz, E.; Dang, T.

1995-01-01

The Department of Mechanical, Aerospace, and Manufacturing Engineering and the Northeast Parallel Architectures Center of Syracuse University have been funded by NASA to establish a program to educate young engineers in the hypersonic disciplines. This goal is being achieved through a comprehensive five-year program that includes elements of undergraduate instruction, advanced graduate coursework, undergraduate research, and leading-edge hypersonics research. The research foci of the Syracuse Center for Hypersonics are three-fold; high-temperature composite materials, measurements in turbulent hypersonic flows, and the application of high-performance computing to hypersonic fluid dynamics.

Magnetic Control of Hypersonic Flow

NASA Astrophysics Data System (ADS)

Poggie, Jonathan; Gaitonde, Datta

2000-11-01

Electromagnetic control is an appealing possibility for mitigating the thermal loads that occur in hypersonic flight, in particular for the case of atmospheric entry. There was extensive research on this problem between about 1955 and 1970,(M. F. Romig, ``The Influence of Electric and Magnetic Fields on Heat Transfer to Electrically Conducting Fluids,'' \\underlineAdvances In Heat Transfer), Vol. 1, Academic Press, NY, 1964. and renewed interest has arisen due to developments in the technology of super-conducting magnets and the understanding of the physics of weakly-ionized, non-equilibrium plasmas. In order to examine the physics of this problem, and to evaluate the practicality of electromagnetic control in hypersonic flight, we have developed a computer code to solve the three-dimensional, non-ideal magnetogasdynamics equations. We have applied the code to the problem of magnetically-decelerated hypersonic flow over a sphere, and observed a reduction, with an applied dipole field, in heat flux and skin friction near the nose of the body, as well as an increase in shock standoff distance. The computational results compare favorably with the analytical predictions of Bush.(W. B. Bush, ``Magnetohydrodynamic-Hypersonic Flow Past a Blunt Body'', Journal of the Aero/Space Sciences, Vol. 25, No. 11, 1958; ``The Stagnation-Point Boundary Layer in the Presence of an Applied Magnetic Field'', Vol. 28, No. 8, 1961.)

Stereoscopic Imaging in Hypersonics Boundary Layers using Planar Laser-Induced Fluorescence

NASA Technical Reports Server (NTRS)

Danehy, Paul M.; Bathel, Brett; Inman, Jennifer A.; Alderfer, David W.; Jones, Stephen B.

2008-01-01

Stereoscopic time-resolved visualization of three-dimensional structures in a hypersonic flow has been performed for the first time. Nitric Oxide (NO) was seeded into hypersonic boundary layer flows that were designed to transition from laminar to turbulent. A thick laser sheet illuminated and excited the NO, causing spatially-varying fluorescence. Two cameras in a stereoscopic configuration were used to image the fluorescence. The images were processed in a computer visualization environment to provide stereoscopic image pairs. Two methods were used to display these image pairs: a cross-eyed viewing method which can be viewed by naked eyes, and red/blue anaglyphs, which require viewing through red/blue glasses. The images visualized three-dimensional information that would be lost if conventional planar laser-induced fluorescence imaging had been used. Two model configurations were studied in NASA Langley Research Center's 31-Inch Mach 10 Air Wind tunnel. One model was a 10 degree half-angle wedge containing a small protuberance to force the flow to transition. The other model was a 1/3-scale, truncated Hyper-X forebody model with blowing through a series of holes to force the boundary layer flow to transition to turbulence. In the former case, low flowrates of pure NO seeded and marked the boundary layer fluid. In the latter, a trace concentration of NO was seeded into the injected N2 gas. The three-dimensional visualizations have an effective time resolution of about 500 ns, which is fast enough to freeze this hypersonic flow. The 512x512 resolution of the resulting images is much higher than high-speed laser-sheet scanning systems with similar time response, which typically measure 10-20 planes.

National Offshore Wind Strategy: Facilitating the Development of the Offshore Wind Industry in the United States

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

Patrick Gilman; Maurer, Ben; Feinberg, Luke

2016-09-01

The U.S. Department of Energy, through its Wind Energy Technologies Office, and U.S. Department of the Interior, through its Bureau of Ocean Energy Management, have jointly produced this updated national strategy to facilitate the responsible development of offshore wind energy in the United States.

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.

Evaluation of candidate working fluid formulations for the electrothermal - chemical wind tunnel

NASA Technical Reports Server (NTRS)

Akyurtlu, Jale F.; Akyurtlu, Ates

1991-01-01

Various candidate chemical formulations are evaluated as a precursor for the working fluid to be used in the electrothermal hypersonic test facility which was under study at the NASA LaRC Hypersonic Propulsion Branch, and the formulations which would most closely satisfy the goals set for the test facility are identified. Out of the four tasks specified in the original proposal, the first two, literature survey and collection of kinetic data, are almost completed. The third task, work on a mathematical model of the ET wind tunnel operation, was started and concentrated on the expansion in the nozzle with finite rate kinetics.

National Wind Technology Center & Local Information | NREL

Science.gov Websites

NWTC is not located at the main NREL facility in Golden, Colorado; it is approximately 25 miles north lodging for hotels and bed and breakfasts within 15-20 miles driving distance of the National Wind

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.

Application of CFD to a generic hypersonic flight research study

NASA Technical Reports Server (NTRS)

Green, Michael J.; Lawrence, Scott L.; Dilley, Arthur D.; Hawkins, Richard W.; Walker, Mary M.; Oberkampf, William L.

1993-01-01

Computational analyses have been performed for the initial assessment of flight research vehicle concepts that satisfy requirements for potential hypersonic experiments. Results were obtained from independent analyses at NASA Ames, NASA Langley, and Sandia National Labs, using sophisticated time-dependent Navier-Stokes and parabolized Navier-Stokes methods. Careful study of a common problem consisting of hypersonic flow past a slightly blunted conical forebody was undertaken to estimate the level of uncertainty in the computed results, and to assess the capabilities of current computational methods for predicting boundary-layer transition onset. Results of this study in terms of surface pressure and heat transfer comparisons, as well as comparisons of boundary-layer edge quantities and flow-field profiles are presented here. Sensitivities to grid and gas model are discussed. Finally, representative results are presented relating to the use of Computational Fluid Dynamics in the vehicle design and the integration/support of potential experiments.

Unstart coupling mechanism analysis of multiple-modules hypersonic inlet.

PubMed

Hu, Jichao; Chang, Juntao; Wang, Lei; Cao, Shibin; Bao, Wen

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.

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

Design and Development of a Real-Time Model Attitude Measurement System for Hypersonic Facilities

NASA Technical Reports Server (NTRS)

Jones, Thomas W.; Lunsford, Charles B.

2005-01-01

A series of wind tunnel tests have been conducted to evaluate a multi-camera videogrammetric system designed to measure model attitude in hypersonic facilities. The technique utilizes processed video data and applies photogrammetric principles for point tracking to compute model position including pitch, roll and yaw variables. A discussion of the constraints encountered during the design, development, and testing process, including lighting, vibration, operational range and optical access is included. Initial measurement results from the NASA Langley Research Center (LaRC) 31-Inch Mach 10 tunnel are presented.

Design and Development of a Real-Time Model Attitude Measurement System for Hypersonic Facilities

NASA Technical Reports Server (NTRS)

Jones, Thomas W.; Lunsford, Charles B.

2004-01-01

A series of wind tunnel tests have been conducted to evaluate a multi-camera videogrammetric system designed to measure model attitude in hypersonic facilities. The technique utilizes processed video data and applies photogrammetric principles for point tracking to compute model position including pitch, roll and yaw variables. A discussion of the constraints encountered during the design, development, and testing process, including lighting, vibration, operational range and optical access is included. Initial measurement results from the NASA Langley Research Center (LaRC) 31-Inch Mach 10 tunnel are presented.

Retooling CFD for hypersonic aircraft

NASA Technical Reports Server (NTRS)

Dwoyer, Douglas L.; Kutler, Paul; Povinelli, Louis A.

1987-01-01

The CFD facility requirements of hypersonic aircraft configuration design development are different from those thus far employed for reentry vehicle design, because (1) the airframe and the propulsion system must be fully integrated to achieve the desired performance; (2) the vehicle must be reusable, with minimum refurbishment requirements between flights; and (3) vehicle performance must be optimized for a wide range of Mach numbers. An evaluation is presently made of flow resolution within shock waves, transition and turbulence phenomenon tractability, chemical reaction modeling, and hypersonic boundary layer transition, with state-of-the-art CFD.

NASA's Advanced Space Transportation Hypersonic Program

NASA Technical Reports Server (NTRS)

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

2002-01-01

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

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.

Configuration development study of the X-24C hypersonic research airplane, phase 3

NASA Technical Reports Server (NTRS)

Combs, H. G.

1977-01-01

The conclusion evolved from the three phased study on the configuration development of the X-24C Hypersonic Research Airplane makes it evident that it is practical to design and build the high performance National Hypersonic Flight Research Facility airplane with today's state of the art within the cost and operational constraints established by NASA. The vehicle launched at 31.75 Mg from the B-52 can cruise for 40 seconds at Mach 6.78 on scramjets. Without scramjets it can approach Mach 8 with a 453.6 Kg payload or do 70 seconds of cruise at Mach 6 with a 2.27 Mg payload. Reduction in cost is possible with a vehicle scaled to a lesser mass and capability.

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.

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.

75 FR 23666 - Huron-Manistee National Forests, White Pines Wind Farm Project, Mason County, MI

Federal Register 2010, 2011, 2012, 2013, 2014

2010-05-04

... DEPARTMENT OF AGRICULTURE Forest Service Huron-Manistee National Forests, White Pines Wind Farm... environmental impact statement for the White Pines Wind Farm Project on National Forest System (NFS) lands... terminates the environmental analysis process for the White Pines Wind Farm Project. DATES: The Notice of...

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.

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.

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.

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

Results of a M = 5.3 heat transfer test of the integrated vehicle using phase-change paint techniques on the 0.0175-scale model 56-OTS in the NASA/Ames Research Center 3.5-foot hypersonic wind tunnel

NASA Technical Reports Server (NTRS)

Marroquin, J.

1985-01-01

An experimental investigation was performed in the NASA/Ames Research Center 3.5-foot Hypersonic Wind Tunnel to obtain supersonic heat-distribution data in areas between the orbiter and external tank using phase-change paint techniques. The tests used Novamide SSV Model 56-OTS in the first and second-stage ascent configurations. Data were obtained at a nominal Mach number of 5.3 and a Reynolds number per foot of 5 x 10 to the 6th power with angles of attack of 0 deg, +/- 5 deg, and sideslip angles of 0 deg and +/- 5 deg.

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

NASA Technical Reports Server (NTRS)

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

1990-01-01

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

winderosionnetwork.org – Portal to the National Wind Erosion Research Network

USDA-ARS?s Scientific Manuscript database

The National Wind Erosion Research Network was established in 2014 as a collaborative effort led by the USDA Agricultural Research Service and Natural Resources Conservation Service, and USDI Bureau of Land Management, to address the need for standardized measurements of wind erosion and its control...

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.

Vibrational relaxation in hypersonic flow fields

NASA Technical Reports Server (NTRS)

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

1993-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1975-01-01

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

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.

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.

Nano-ADEPT Aeroloads Wind Tunnel Test

NASA Technical Reports Server (NTRS)

Smith, Brandon; Yount, Bryan; Kruger, Carl; Brivkalns, Chad; Makino, Alberto; Cassell, Alan; Zarchi, Kerry; McDaniel, Ryan; Ross, James; Wercinski, Paul;

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.

Multiscale Computational Analysis of Nitrogen and Oxygen Gas-Phase Thermochemistry in Hypersonic Flows

NASA Astrophysics Data System (ADS)

Bender, Jason D.

Understanding hypersonic aerodynamics is important for the design of next-generation aerospace vehicles for space exploration, national security, and other applications. Ground-level experimental studies of hypersonic flows are difficult and expensive; thus, computational science plays a crucial role in this field. Computational fluid dynamics (CFD) simulations of extremely high-speed flows require models of chemical and thermal nonequilibrium processes, such as dissociation of diatomic molecules and vibrational energy relaxation. Current models are outdated and inadequate for advanced applications. We describe a multiscale computational study of gas-phase thermochemical processes in hypersonic flows, starting at the atomic scale and building systematically up to the continuum scale. The project was part of a larger effort centered on collaborations between aerospace scientists and computational chemists. We discuss the construction of potential energy surfaces for the N4, N2O2, and O4 systems, focusing especially on the multi-dimensional fitting problem. A new local fitting method named L-IMLS-G2 is presented and compared with a global fitting method. Then, we describe the theory of the quasiclassical trajectory (QCT) approach for modeling molecular collisions. We explain how we implemented the approach in a new parallel code for high-performance computing platforms. Results from billions of QCT simulations of high-energy N2 + N2, N2 + N, and N2 + O2 collisions are reported and analyzed. Reaction rate constants are calculated and sets of reactive trajectories are characterized at both thermal equilibrium and nonequilibrium conditions. The data shed light on fundamental mechanisms of dissociation and exchange reactions -- and their coupling to internal energy transfer processes -- in thermal environments typical of hypersonic flows. We discuss how the outcomes of this investigation and other related studies lay a rigorous foundation for new macroscopic models for

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.

Issues Associated with a Hypersonic Maglev Sled

NASA Technical Reports Server (NTRS)

Haney, Joseph W.; Lenzo, J.

1996-01-01

Magnetic levitation has been explored for application from motors to transportation. All of these applications have been at velocities where the physics of the air or operating fluids are fairly well known. Application of Maglev to hypersonic velocities (Mach greater than 5) presents many opportunities, but also issues that require understanding and resolution. Use of Maglev to upgrade the High Speed Test Track at Holloman Air Force Base in Alamogordo New Mexico is an actual hypersonic application that provides the opportunity to improve test capabilities. However, there are several design issues that require investigation. This paper presents an overview of the application of Maglev to the test track and the issues associated with developing a hypersonic Maglev sled. The focus of this paper is to address the issues with the Maglev sled design, rather than the issues with the development of superconducting magnets of the sled system.

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

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.

Largest Ever Wind Power Commitments Sets Clean Energy Example for Nation

Science.gov Websites

'Largest Ever' Wind Power Commitments Sets Clean Energy Example for Nation NEWS MEDIA CONTACTS ;Federal agencies in Colorado are setting an example for the rest of the nation and the leaders of business

Progress in modeling hypersonic turbulent boundary layers

NASA Technical Reports Server (NTRS)

Zeman, Otto

1993-01-01

A good knowledge of the turbulence structure, wall heat transfer, and friction in turbulent boundary layers (TBL) at high speeds is required for the design of hypersonic air breathing airplanes and reentry space vehicles. This work reports on recent progress in the modeling of high speed TBL flows. The specific research goal described here is the development of a second order closure model for zero pressure gradient TBL's for the range of Mach numbers up to hypersonic speeds with arbitrary wall cooling requirements.

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.

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.

Hypersonic Viscous Flow Over Large Roughness Elements

NASA Technical Reports Server (NTRS)

Chang, Chau-Lyan; Choudhari, Meelan M.

2009-01-01

Viscous flow over discrete or distributed surface roughness has great implications for hypersonic flight due to aerothermodynamic considerations related to laminar-turbulent transition. Current prediction capability is greatly hampered by the limited knowledge base for such flows. To help fill that gap, numerical computations are used to investigate the intricate flow physics involved. An unstructured mesh, compressible Navier-Stokes code based on the space-time conservation element, solution element (CESE) method is used to perform time-accurate Navier-Stokes calculations for two roughness shapes investigated in wind tunnel experiments at NASA Langley Research Center. It was found through 2D parametric study that at subcritical Reynolds numbers of the boundary layers, absolute instability resulting in vortex shedding downstream, is likely to weaken at supersonic free-stream conditions. On the other hand, convective instability may be the dominant mechanism for supersonic boundary layers. Three-dimensional calculations for a rectangular or cylindrical roughness element at post-shock Mach numbers of 4.1 and 6.5 also confirm that no self-sustained vortex generation is present.

Turbulence measurements in hypersonic boundary layers using constant-temperature anemometry and Reynolds stress measurements in hypersonic boundary layers

NASA Technical Reports Server (NTRS)

Spina, Eric F.

1995-01-01

The primary objective in the two research investigations performed under NASA Langley sponsorship (Turbulence measurements in hypersonic boundary layers using constant temperature anemometry and Reynolds stress measurements in hypersonic boundary layers) has been to increase the understanding of the physics of hypersonic turbulent boundary layers. The study began with an extension of constant-temperature thermal anemometry techniques to a Mach 11 helium flow, including careful examinations of hot-wire construction techniques, system response, and system calibration. This was followed by the application of these techniques to the exploration of a Mach 11 helium turbulent boundary layer (To approximately 290 K). The data that was acquired over the course of more than two years consists of instantaneous streamwise mass flux measurements at a frequency response of about 500 kHz. The data are of exceptional quality in both the time and frequency domain and possess a high degree of repeatability. The data analysis that has been performed to date has added significantly to the body of knowledge on hypersonic turbulence, and the data reduction is continuing. An attempt was then made to extend these thermal anemometry techniques to higher enthalpy flows, starting with a Mach 6 air flow with a stagnation temperature just above that needed to prevent liquefaction (To approximately 475 F). Conventional hot-wire anemometry proved to be inadequate for the selected high-temperature, high dynamic pressure flow, with frequent wire breakage and poor system frequency response. The use of hot-film anemometry has since been investigated for these higher-enthalpy, severe environment flows. The difficulty with using hot-film probes for dynamic (turbulence) measurements is associated with construction limitations and conduction of heat into the film substrate. Work continues under a NASA GSRP grant on the development of a hot film probe that overcomes these shortcomings for hypersonic

National Wind Technology Center | NREL

Science.gov Websites

. Wind Energy Research Wind turbine blade Wind energy research at the NWTC allows for validation and verification of large and small components and wind turbine systems. Photo by Dennis Schroeder / NREL 40935 Wind energy research at the NWTC has pioneered wind turbine components, systems, and modeling methods

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

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.

Aerodynamic heating in transitional hypersonic boundary layers: Role of second-mode instability

NASA Astrophysics Data System (ADS)

Zhu, Yiding; Chen, Xi; Wu, Jiezhi; Chen, Shiyi; Lee, Cunbiao; Gad-el-Hak, Mohamed

2018-01-01

The evolution of second-mode instabilities in hypersonic boundary layers and its effects on aerodynamic heating are investigated. Experiments are conducted in a Mach 6 wind tunnel using fast-response pressure sensors, fluorescent temperature-sensitive paint, and particle image velocimetry. Calculations based on parabolic stability equations and direct numerical simulations are also performed. It is found that second-mode waves, accompanied by high-frequency alternating fluid compression and expansion, produce intense aerodynamic heating in a small region that rapidly heats the fluid passing through it. As the second-mode waves decay downstream, the dilatation-induced aerodynamic heating decreases while its shear-induced counterpart keeps growing. The latter brings about a second growth of the surface temperature when transition is completed.

RDHWT/MARIAH II Hypersonic Wind Tunnel Research Program

DTIC Science & Technology

2008-09-01

Diagnostics Dr. Gary Brown – Gas Dynamics Dr. Ihab Girgis – Modeling Dr. Dennis Mansfield – Experimental Ring Technical Services Dr. Leon Ring – Systems...wind tunnel (MSHWT) with Mach 8 to 15, true -temperature flight test capabilities. This research program was initiated in fiscal year (FY) 1998 and is...Force test capabilities that exist today. Performance goals of the MSHWT are true temperature, Mach 8 to 15, dynamic pressure of 500 to 2000 psf (24 to

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.

Videometric Applications in Wind Tunnels

NASA Technical Reports Server (NTRS)

Burner, A. W.; Radeztsky, R. H.; Liu, Tian-Shu

1997-01-01

Videometric measurements in wind tunnels can be very challenging due to the limited optical access, model dynamics, optical path variability during testing, large range of temperature and pressure, hostile environment, and the requirements for high productivity and large amounts of data on a daily basis. Other complications for wind tunnel testing include the model support mechanism and stringent surface finish requirements for the models in order to maintain aerodynamic fidelity. For these reasons nontraditional photogrammetric techniques and procedures sometimes must be employed. In this paper several such applications are discussed for wind tunnels which include test conditions with Mach number from low speed to hypersonic, pressures from less than an atmosphere to nearly seven atmospheres, and temperatures from cryogenic to above room temperature. Several of the wind tunnel facilities are continuous flow while one is a short duration blowdown facility. Videometric techniques and calibration procedures developed to measure angle of attack, the change in wing twist and bending induced by aerodynamic load, and the effects of varying model injection rates are described. Some advantages and disadvantages of these techniques are given and comparisons are made with non-optical and more traditional video photogrammetric techniques.

Hypersonic Shock/Boundary-Layer Interaction Database

NASA Technical Reports Server (NTRS)

Settles, G. S.; Dodson, L. J.

1991-01-01

Turbulence modeling is generally recognized as the major problem obstructing further advances in computational fluid dynamics (CFD). A closed solution of the governing Navier-Stokes equations for turbulent flows of practical consequence is still far beyond grasp. At the same time, the simplified models of turbulence which are used to achieve closure of the Navier-Stokes equations are known to be rigorously incorrect. While these models serve a definite purpose, they are inadequate for the general prediction of hypersonic viscous/inviscid interactions, mixing problems, chemical nonequilibria, and a range of other phenomena which must be predicted in order to design a hypersonic vehicle computationally. Due to the complexity of turbulence, useful new turbulence models are synthesized only when great expertise is brought to bear and considerable intellectual energy is expended. Although this process is fundamentally theoretical, crucial guidance may be gained from carefully-executed basic experiments. Following the birth of a new model, its testing and validation once again demand comparisons with data of unimpeachable quality. This report concerns these issues which arise from the experimental aspects of hypersonic modeling and represents the results of the first phase of an effort to develop compressible turbulence models.

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.

Optimal trajectories for hypersonic launch vehicles

NASA Technical Reports Server (NTRS)

Ardema, Mark D.; Bowles, Jeffrey V.; Whittaker, Thomas

1994-01-01

In this paper, we derive a near-optimal guidance law for the ascent trajectory from earth surface to earth orbit of a hypersonic, dual-mode propulsion, lifting vehicle. Of interest are both the optical flight path and the optimal operation of the propulsion system. The guidance law is developed from the energy-state approximation of the equations of motion. Because liquid hydrogen fueled hypersonic aircraft are volume sensitive, as well as weight sensitive, the cost functional is a weighted sum of fuel mass and volume; the weighting factor is chosen to minimize gross take-off weight for a given payload mass and volume in orbit.

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.

A concept of a hypersonic flight experiment of a winged vehicle

NASA Astrophysics Data System (ADS)

Shirouzu, Masao; Watanabe, Shigeya

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

National Wind Distance Learning Collaborative

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

Dr. James B. Beddow

2013-03-29

Executive Summary The energy development assumptions identified in the Department of Energy's position paper, 20% Wind Energy by 2030, projected an exploding demand for wind energy-related workforce development. These primary assumptions drove a secondary set of assumptions that early stage wind industry workforce development and training paradigms would need to undergo significant change if the workforce needs were to be met. The current training practice and culture within the wind industry is driven by a relatively small number of experts with deep field experience and knowledge. The current training methodology is dominated by face-to-face, classroom based, instructor present training. Givenmore » these assumptions and learning paradigms, the purpose of the National Wind Distance Learning Collaborative was to determine the feasibility of developing online learning strategies and products focused on training wind technicians. The initial project scope centered on (1) identifying resources that would be needed for development of subject matter and course design/delivery strategies for industry-based (non-academic) training, and (2) development of an appropriate Learning Management System (LMS). As the project unfolded, the initial scope was expanded to include development of learning products and the addition of an academic-based training partner. The core partners included two training entities, industry-based Airstreams Renewables and academic-based Lake Area Technical Institute. A third partner, Vision Video Interactive, Inc. provided technology-based learning platforms (hardware and software). The revised scope yielded an expanded set of results beyond the initial expectation. Eight learning modules were developed for the industry-based Electrical Safety course. These modules were subsequently redesigned and repurposed for test application in an academic setting. Software and hardware developments during the project's timeframe enabled redesign providing

CFD for hypersonic propulsion

NASA Technical Reports Server (NTRS)

Povinelli, Louis A.

1991-01-01

An overview is given of research activity on the application of computational fluid dynamics (CDF) for hypersonic propulsion systems. After the initial consideration of the highly integrated nature of air-breathing hypersonic engines and airframe, attention is directed toward computations carried out for the components of the engine. A generic inlet configuration is considered in order to demonstrate the highly three dimensional viscous flow behavior occurring within rectangular inlets. Reacting flow computations for simple jet injection as well as for more complex combustion chambers are then discussed in order to show the capability of viscous finite rate chemical reaction computer simulations. Finally, the nozzle flow fields are demonstrated, showing the existence of complex shear layers and shock structure in the exhaust plume. The general issues associated with code validation as well as the specific issue associated with the use of CFD for design are discussed. A prognosis for the success of CFD in the design of future propulsion systems is offered.

CFD for hypersonic propulsion

NASA Technical Reports Server (NTRS)

Povinelli, Louis A.

1990-01-01

An overview is given of research activity on the application of computational fluid dynamics (CDF) for hypersonic propulsion systems. After the initial consideration of the highly integrated nature of air-breathing hypersonic engines and airframe, attention is directed toward computations carried out for the components of the engine. A generic inlet configuration is considered in order to demonstrate the highly three dimensional viscous flow behavior occurring within rectangular inlets. Reacting flow computations for simple jet injection as well as for more complex combustion chambers are then discussed in order to show the capability of viscous finite rate chemical reaction computer simulations. Finally, the nozzle flow fields are demonstrated, showing the existence of complex shear layers and shock structure in the exhaust plume. The general issues associated with code validation as well as the specific issue associated with the use of CFD for design are discussed. A prognosis for the success of CFD in the design of future propulsion systems is offered.

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.

Uncertainty Propagation in Hypersonic Vehicle Aerothermoelastic Analysis

NASA Astrophysics Data System (ADS)

Lamorte, Nicolas Etienne

Hypersonic vehicles face a challenging flight environment. The aerothermoelastic analysis of its components requires numerous simplifying approximations. Identifying and quantifying the effect of uncertainties pushes the limits of the existing deterministic models, and is pursued in this work. An uncertainty quantification framework is used to propagate the effects of identified uncertainties on the stability margins and performance of the different systems considered. First, the aeroelastic stability of a typical section representative of a control surface on a hypersonic vehicle is examined. Variability in the uncoupled natural frequencies of the system is modeled to mimic the effect of aerodynamic heating. Next, the stability of an aerodynamically heated panel representing a component of the skin of a generic hypersonic vehicle is considered. Uncertainty in the location of transition from laminar to turbulent flow and the heat flux prediction is quantified using CFD. In both cases significant reductions of the stability margins are observed. A loosely coupled airframe--integrated scramjet engine is considered next. The elongated body and cowl of the engine flow path are subject to harsh aerothermodynamic loading which causes it to deform. Uncertainty associated with deformation prediction is propagated to the engine performance analysis. The cowl deformation is the main contributor to the sensitivity of the propulsion system performance. Finally, a framework for aerothermoelastic stability boundary calculation for hypersonic vehicles using CFD is developed. The usage of CFD enables one to consider different turbulence conditions, laminar or turbulent, and different models of the air mixture, in particular real gas model which accounts for dissociation of molecules at high temperature. The system is found to be sensitive to turbulence modeling as well as the location of the transition from laminar to turbulent flow. Real gas effects play a minor role in the

Correlation parameters for the study of leeside heating on a lifting body at hypersonic speeds

NASA Technical Reports Server (NTRS)

Vidal, R. J.

1974-01-01

Leeside heating was studied with the aim of gaining some insight into: (1) the magnitude of the leeside heating rates and (2) the methods to be used to extrapolate wind tunnel leeside heating rates to the full scale flight condition. This study was based on existing experimental data obtained in a hypersonic shock tunnel on lifting body configurations that are typical of shuttle orbiter vehicles. Heat transfer was first measured on the windward side to determine the boundary layer type. Then the leeside heating was investigated with the classified boundary layer. Correlation data are given on the windward turbulent boundary layer, the windward laminar boundary layer, and the leeside surfaces.

A reduced theoretical model for estimating condensation effects in combustion-heated hypersonic tunnel

NASA Astrophysics Data System (ADS)

Lin, L.; Luo, X.; Qin, F.; Yang, J.

2018-03-01

As one of the combustion products of hydrocarbon fuels in a combustion-heated wind tunnel, water vapor may condense during the rapid expansion process, which will lead to a complex two-phase flow inside the wind tunnel and even change the design flow conditions at the nozzle exit. The coupling of the phase transition and the compressible flow makes the estimation of the condensation effects in such wind tunnels very difficult and time-consuming. In this work, a reduced theoretical model is developed to approximately compute the nozzle-exit conditions of a flow including real-gas and homogeneous condensation effects. Specifically, the conservation equations of the axisymmetric flow are first approximated in the quasi-one-dimensional way. Then, the complex process is split into two steps, i.e., a real-gas nozzle flow but excluding condensation, resulting in supersaturated nozzle-exit conditions, and a discontinuous jump at the end of the nozzle from the supersaturated state to a saturated state. Compared with two-dimensional numerical simulations implemented with a detailed condensation model, the reduced model predicts the flow parameters with good accuracy except for some deviations caused by the two-dimensional effect. Therefore, this reduced theoretical model can provide a fast, simple but also accurate estimation of the condensation effect in combustion-heated hypersonic tunnels.

Analysis of a six-component, flow-through, strain-gage, force balance used for hypersonic wind tunnel models with scramjet exhaust flow simulation. M.S. Thesis Final Report

NASA Technical Reports Server (NTRS)

Kniskern, Marc W.

1990-01-01

The thermal effects of simulant gas injection and aerodynamic heating at the model's surface on the measurements of a non-watercooled, flow through balance were investigated. A stainless steel model of a hypersonic air breathing propulsion cruise missile concept (HAPCM-50) was used to evaluate this balance. The tests were conducted in the 20-inch Mach 6 wind tunnel at NASA-Langley. The balance thermal effects were evaluated at freestream Reynolds numbers ranging from .5 to 7 x 10(exp 6) ft and angles of attack between -3.5 to 5 deg at Mach 6. The injection gases considered included cold air, hot air, and a mixture of 50 percent Argon and 50 percent Freon-12. The stagnation temperatures of the cold air, hot air, and Ar-Fr(12) reached 111, 214, and 283 F, respectively within the balance. A bakelite sleeve was inserted into the inner tube of the balance to minimize the thermal effects of these injection gases. Throughout the tests, the normal force, side force, yaw moment, roll moment, and pitching moment balance measurements were unaffected by the balance thermal effects of the injection gases and the wind tunnel flow. However, the axial force (AF) measurement was significantly affected by balance heating. The average zero shifts in the AF measurements were 1.9, 3.8, and 5.9 percent for cold air, hot air, and Ar-Fr(12) injection, respectively. The AF measurements decreased throughout these tests which lasted from 70 to 110 seconds. During the cold air injection tests, the AF measurements were accurate up to at least ten seconds after the model was injected into the wind tunnel test section. For the hot air and Ar-Fr(12) tests, the AF measurements were accurate up to at least five seconds after model injection.

Wind Vision: A New Era for Wind Power in the United States (Highlights); U.S. Department of Energy (DOE), NREL (National Renewable Energy Laboratory)

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

None

2015-03-01

This is a four-part Wind Vision project, consisting of Wind Vision Highlights, Executive Summary, a Full Report, and Appendix. The U.S. Department of Energy (DOE) Wind Program, in close cooperation with the wind industry, led a comprehensive analysis to evaluate future pathways for the wind industry. The Wind Vision report updates and expands upon the DOE's 2008 report, 20% Wind Energy by 2030, and defines the societal, environmental, and economic benefits of wind power in a scenario with wind energy supplying 10% of national end-use electricity demand by 2020, 20% by 2030, and 35% by 2050.

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.

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

NASA Technical Reports Server (NTRS)

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

1984-01-01

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

Design, Validation, and Testing of a Hot-Film Anemometer for Hypersonic Flow

NASA Astrophysics Data System (ADS)

Sheplak, Mark

The application of constant-temperature hot-film anemometry to hypersonic flow has been reviewed and extended in this thesis. The objective of this investigation was to develop a measurement tool capable of yielding continuous, high-bandwidth, quantitative, normal mass-flux and total -temperature measurements in moderate-enthalpy environments. This research has produced a probe design that represents a significant advancement over existing designs, offering the following improvements: (1) a five-fold increase in bandwidth; (2) true stagnation-line sensor placement; (3) a two order-of-magnitude decrease in sensor volume; and (4) over a 70% increase in maximum film temperature. These improvements were achieved through substrate design, sensor placement, the use of high-temperature materials, and state -of-the-art microphotolithographic fabrication techniques. The experimental study to characterize the probe was performed in four different hypersonic wind tunnels at NASA-Langley Research Center. The initial test consisted of traversing the hot film through a Mach 6, flat-plate, turbulent boundary layer in air. The detailed static-calibration measurements that followed were performed in two different hypersonic flows: a Mach 11 helium flow and Mach 6 air flow. The final test of this thesis consisted of traversing the probe through the Mach 6 wake of a 70^ circ blunt body. The goal of this test was to determine the state (i.e., laminar or turbulent) of the wake. These studies indicate that substrate conduction effects result in instrumentation characteristics that prevent the hot-film anemometer from being used as a quantitative tool. The extension of this technique to providing quantitative information is dependent upon the development of lower thermal-conductivity substrate materials. However, the probe durability, absence of strain gauging, and high bandwidth represent significant improvements over the hot-wire technique for making qualitative measurements. Potential

Multigrid schemes for viscous hypersonic flows

NASA Technical Reports Server (NTRS)

Swanson, R. C.; Radespiel, R.

1993-01-01

Several multigrid schemes are considered for the numerical computation of viscous hypersonic flows. For each scheme, the basic solution algorithm employs upwind spatial discretization with explicit multistage time stepping. Two-level versions of the various multigrid algorithms are applied to the two-dimensional advection equation, and Fourier analysis is used to determine their damping properties. The capabilities of the multigrid methods are assessed by solving two different hypersonic flow problems. Some new multigrid schemes, based on semicoarsening strategies, are shown to be quite effective in relieving the stiffness caused by the high-aspect-ratio cells required to resolve high Reynolds number flows. These schemes exhibit good convergence rates for Reynolds numbers up to 200 x 10(exp 6).

Photogrammetry of a Hypersonic Inflatable Aerodynamic Decelerator

NASA Technical Reports Server (NTRS)

Kushner, Laura Kathryn; Littell, Justin D.; Cassell, Alan M.

2013-01-01

In 2012, two large-scale models of a Hypersonic Inflatable Aerodynamic decelerator were tested in the National Full-Scale Aerodynamic Complex at NASA Ames Research Center. One of the objectives of this test was to measure model deflections under aerodynamic loading that approximated expected flight conditions. The measurements were acquired using stereo photogrammetry. Four pairs of stereo cameras were mounted inside the NFAC test section, each imaging a particular section of the HIAD. The views were then stitched together post-test to create a surface deformation profile. The data from the photogram- metry system will largely be used for comparisons to and refinement of Fluid Structure Interaction models. This paper describes how a commercial photogrammetry system was adapted to make the measurements and presents some preliminary results.

Research in robust control for hypersonic aircraft

NASA Technical Reports Server (NTRS)

Calise, A. J.

1994-01-01

The research during the third reporting period focused on fixed order robust control design for hypersonic vehicles. A new technique was developed to synthesize fixed order H(sub infinity) controllers. A controller canonical form is imposed on the compensator structure and a homotopy algorithm is employed to perform the controller design. Various reduced order controllers are designed for a simplified version of the hypersonic vehicle model used in our previous studies to demonstrate the capabilities of the code. However, further work is needed to investigate the issue of numerical ill-conditioning for large order systems and to make the numerical approach more reliable.

Advances in Computational Capabilities for Hypersonic Flows

NASA Technical Reports Server (NTRS)

Kumar, Ajay; Gnoffo, Peter A.; Moss, James N.; Drummond, J. Philip

1997-01-01

The paper reviews the growth and advances in computational capabilities for hypersonic applications over the period from the mid-1980's to the present day. The current status of the code development issues such as surface and field grid generation, algorithms, physical and chemical modeling, and validation is provided. A brief description of some of the major codes being used at NASA Langley Research Center for hypersonic continuum and rarefied flows is provided, along with their capabilities and deficiencies. A number of application examples are presented, and future areas of research to enhance accuracy, reliability, efficiency, and robustness of computational codes are discussed.

Body weight of hypersonic aircraft, part 1

NASA Technical Reports Server (NTRS)

Ardema, Mark D.

1988-01-01

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

Heat transfer tests on a 0.01-scale Rockwell configuration 3 space shuttle orbiter and tank (37-OT) in the Calspan 48-inch hypersonic shock tunnel (OH12/IH21), volume 1

NASA Technical Reports Server (NTRS)

Kotch, M.

1975-01-01

Model information and data are presented from wind tunnel tests conducted on 0.01-scale models of the space shuttle orbiter and external tank. These tests were conducted in a hypersonic shock tunnel to determine heating rates on ascent and reentry configurations at various Reynolds numbers, Mach numbers, and angles of attack.

United States Supports Distributed Wind Technology Improvements; NREL (National Renewable Energy Laboratory)

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

Sinclair, Karin

2015-06-15

This presentation provides information on the activities conducted through the Competitiveness Improvement Project (CIP), initiated in 2012 by the U.S. Department of Energy (DOE) and executed through the National Renewable Energy Laboratory (NREL) to support the distributed wind industry. The CIP provides research and development funding and technical support to improve distributed wind turbine technology and increase the competitiveness of U.S. small and midsize wind turbine manufacturers. Through this project, DOE/NREL assists U.S. manufacturers to lower the levelized cost of energy of wind turbines through component improvements, manufacturing process upgrades, and turbine testing. Ultimately, this support is expected to leadmore » to turbine certification through testing to industry-recognized wind turbine performance and safety standards.« less

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.

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

NASA Technical Reports Server (NTRS)

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

1997-01-01

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

CFD code calibration and inlet-fairing effects on a 3D hypersonic powered-simulation model

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Tatum, Kenneth E.

1993-01-01

A three-dimensional (3D) computational study has been performed addressing issues related to the wind tunnel testing of a hypersonic powered-simulation model. The study consisted of three objectives. The first objective was to calibrate a state-of-the-art computational fluid dynamics (CFD) code in its ability to predict hypersonic powered-simulation flows by comparing CFD solutions with experimental surface pressure dam. Aftbody lower surface pressures were well predicted, but lower surface wing pressures were less accurately predicted. The second objective was to determine the 3D effects on the aftbody created by fairing over the inlet; this was accomplished by comparing the CFD solutions of two closed-inlet powered configurations with a flowing-inlet powered configuration. Although results at four freestream Mach numbers indicate that the exhaust plume tends to isolate the aftbody surface from most forebody flowfield differences, a smooth inlet fairing provides the least aftbody force and moment variation compared to a flowing inlet. The final objective was to predict and understand the 3D characteristics of exhaust plume development at selected points on a representative flight path. Results showed a dramatic effect of plume expansion onto the wings as the freestream Mach number and corresponding nozzle pressure ratio are increased.

CFD Code Calibration and Inlet-Fairing Effects On a 3D Hypersonic Powered-Simulation Model

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Tatum, Kenneth E.

1993-01-01

A three-dimensional (3D) computational study has been performed addressing issues related to the wind tunnel testing of a hypersonic powered-simulation model. The study consisted of three objectives. The first objective was to calibrate a state-of-the-art computational fluid dynamics (CFD) code in its ability to predict hypersonic powered-simulation flows by comparing CFD solutions with experimental surface pressure data. Aftbody lower surface pressures were well predicted, but lower surface wing pressures were less accurately predicted. The second objective was to determine the 3D effects on the aftbody created by fairing over the inlet; this was accomplished by comparing the CFD solutions of two closed-inlet powered configurations with a flowing- inlet powered configuration. Although results at four freestream Mach numbers indicate that the exhaust plume tends to isolate the aftbody surface from most forebody flow- field differences, a smooth inlet fairing provides the least aftbody force and moment variation compared to a flowing inlet. The final objective was to predict and understand the 3D characteristics of exhaust plume development at selected points on a representative flight path. Results showed a dramatic effect of plume expansion onto the wings as the freestream Mach number and corresponding nozzle pressure ratio are increased.

Multiple Velocity Profile Measurements in Hypersonic Flows Using Sequentially-Imaged Fluorescence Tagging

NASA Technical Reports Server (NTRS)

Bathel, Brett F.; Danehy, Paul M.; Inman, Jennifer A.; Jones, Stephen B.; Ivey,Christopher b.; Goyne, Christopher P.

2010-01-01

Nitric-oxide planar laser-induced fluorescence (NO PLIF) was used to perform velocity measurements in hypersonic flows by generating multiple tagged lines which fluoresce as they convect downstream. For each laser pulse, a single interline, progressive scan intensified CCD (charge-coupled device) camera was used to obtain two sequential images of the NO molecules that had been tagged by the laser. The CCD configuration allowed for sub-microsecond acquisition of both images, resulting in sub-microsecond temporal resolution as well as sub-mm spatial resolution (0.5-mm horizontal, 0.7-mm vertical). Determination of axial velocity was made by application of a cross-correlation analysis of the horizontal shift of individual tagged lines. A numerical study of measured velocity error due to a uniform and linearly-varying collisional rate distribution was performed. Quantification of systematic errors, the contribution of gating/exposure duration errors, and the influence of collision rate on temporal uncertainty were made. Quantification of the spatial uncertainty depended upon the signal-to-noise ratio of the acquired profiles. This velocity measurement technique has been demonstrated for two hypersonic flow experiments: (1) a reaction control system (RCS) jet on an Orion Crew Exploration Vehicle (CEV) wind tunnel model and (2) a 10-degree half-angle wedge containing a 2-mm tall, 4-mm wide cylindrical boundary layer trip. The experiments were performed at the NASA Langley Research Center's 31-Inch Mach 10 Air Tunnel.

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.

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.

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.

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.

Three dimensional viscous analysis of a hypersonic inlet

NASA Technical Reports Server (NTRS)

Reddy, D. R.; Smith, G. E.; Liou, M.-F.; Benson, Thomas J.

1989-01-01

The flow fields in supersonic/hypersonic inlets are currently being studied at NASA Lewis Research Center using 2- and 3-D full Navier-Stokes and Parabolized Navier-Stokes solvers. These tools have been used to analyze the flow through the McDonnell Douglas Option 2 inlet which has been tested at Calspan in support of the National Aerospace Plane Program. Comparisons between the computational and experimental results are presented. These comparisons lead to better overall understanding of the complex flows present in this class of inlets. The aspects of the flow field emphasized in this work are the 3-D effects, the transition from laminar to turbulent flow, and the strong nonuniformities generated within the inlet.

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

The Mothball, Sustainment, and Proposed Reactivation of the Hypersonic Tunnel Facility (HTF) at NASA Glenn Research Center Plum Brook Station

NASA Technical Reports Server (NTRS)

Thomas, Scott R.; Lee, Jinho; Stephens, John W.; Hostler, Robert W., Jr.; VonKamp, William D.

2010-01-01

The Hypersonic Tunnel Facility (HTF) located at the NASA Glenn Research Center s Plum Brook Station in Sandusky, Ohio, is the nation s only large-scale, non-vitiated, hypersonic propulsion test facility. The HTF, with its 4-story graphite induction heater, is capable of duplicating Mach 5, 6, and 7 flight conditions. This unique propulsion system test facility has experienced several standby and reactivation cycles. The intent of the paper is to overview the HTF capabilities to the propulsion community, present the current status of HTF, and share the lessons learned from putting a large-scale facility into mothball status for a later restart

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.

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.

Constrained Aerothermodynamic Design of Hypersonic Vehicles

NASA Technical Reports Server (NTRS)

Gally, Tom; Campbell, Dick

2002-01-01

An investigation was conducted into possible methods of incorporating a hypersonic design capability with aerothermodynamic constraints into the CDISC aerodynamic design tool. The work was divided into two distinct phases: develop relations between surface curvature and hypersonic pressure coefficient which are compatible with CDISC's direct-iterative design method; and explore and implement possible methods of constraining the heat transfer rate over all or portions of the design surface. The main problem in implementing this method has been the weak relationship between surface shape and pressure coefficient at the stagnation point and the need to design around the surface blunt leading edge where there is a slope singularity. The final results show that some success has been achieved, but further improvements are needed.

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.

Hypersonic Viscous Flow Over Large Roughness Elements

NASA Technical Reports Server (NTRS)

Chang, Chau-Lyan; Choudhari, Meelan M.

2009-01-01

Viscous flow over discrete or distributed surface roughness has great implications for hypersonic flight due to aerothermodynamic considerations related to laminar-turbulent transition. Current prediction capability is greatly hampered by the limited knowledge base for such flows. To help fill that gap, numerical computations are used to investigate the intricate flow physics involved. An unstructured mesh, compressible Navier-Stokes code based on the space-time conservation element, solution element (CESE) method is used to perform time-accurate Navier-Stokes calculations for two roughness shapes investigated in wind tunnel experiments at NASA Langley Research Center. It was found through 2D parametric study that at subcritical Reynolds numbers, spontaneous absolute instability accompanying by sustained vortex shedding downstream of the roughness is likely to take place at subsonic free-stream conditions. On the other hand, convective instability may be the dominant mechanism for supersonic boundary layers. Three-dimensional calculations for both a rectangular and a cylindrical roughness element at post-shock Mach numbers of 4.1 and 6.5 also confirm that no self-sustained vortex generation from the top face of the roughness is observed, despite the presence of flow unsteadiness for the smaller post-shock Mach number case.

Aerodynamic prediction techniques for hypersonic configuration design

NASA Technical Reports Server (NTRS)

1981-01-01

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

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

NASA Technical Reports Server (NTRS)

Harloff, Gary J.; Berkowitz, Brian M.

1988-01-01

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

Rudolf Hermann, wind tunnels and aerodynamics

NASA Astrophysics Data System (ADS)

Lundquist, Charles A.; Coleman, Anne M.

2008-04-01

Rudolf Hermann was born on December 15, 1904 in Leipzig, Germany. He studied at the University of Leipzig and at the Aachen Institute of Technology. His involvement with wind tunnels began in 1934 when Professor Carl Wieselsberger engaged him to work at Aachen on the development of a supersonic wind tunnel. On January 6, 1936, Dr. Wernher von Braun visited Dr. Hermann to arrange for use of the Aachen supersonic wind tunnel for Army problems. On April 1, 1937, Dr. Hermann became Director of the Supersonic Wind Tunnel at the Army installation at Peenemunde. Results from the Aachen and Peenemunde wind tunnels were crucial in achieving aerodynamic stability for the A-4 rocket, later designated as the V-2. Plans to build a Mach 10 'hypersonic' wind tunnel facility at Kochel were accelerated after the Allied air raid on Peenemunde on August 17, 1943. Dr. Hermann was director of the new facility. Ignoring destruction orders from Hitler as WWII approached an end in Europe, Dr. Hermann and his associates hid documents and preserved wind tunnel components that were acquired by the advancing American forces. Dr. Hermann became a consultant to the Air Force at its Wright Field in November 1945. In 1951, he was named professor of Aeronautical Engineering at the University of Minnesota. In 1962, Dr. Hermann became the first Director of the Research Institute at the University of Alabama in Huntsville (UAH), a position he held until he retired in 1970.

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.

The Variable Density Wind Tunnel of the National Advisory Committee for Aeronautics

NASA Technical Reports Server (NTRS)

Munk, Max M; Miller, Elton W

1926-01-01

This report contains an exact description of the new wind tunnel of the National Advisory Committee for Aeronautics. This is the first american type wind tunnel. It differs from ordinary wind tunnels by its being surrounded by a strong steel shell, 35 feet long and 15 feet in diameter. A compressor system is provided to fill this shell - and hence the entire wind tunnel - with air compressed to a density up to 25 times the ordinary atmospheric density. It is demonstrated in the report that the increase of the air density makes up for a corresponding decrease in the scale of the model. Hence such american type wind tunnel is free from scale effect. The report is illustrated by many drawings and photographs. All construction details are described, and many dimensions given. The method of conducting tests is also described and some preliminary results given in the report. So far, the tests have confirmed the chief feature of this wind tunnel - absence of scale effect.

Quantification of precipitation measurement discontinuity induced by wind shields on national gauges

USGS Publications Warehouse

Yang, Daqing; Goodison, Barry E.; Metcalfe, John R.; Louie, Paul; Leavesley, George H.; Emerson, Douglas G.; Hanson, Clayton L.; Golubev, Valentin S.; Elomaa, Esko; Gunther, Thilo; Pangburn, Timothy; Kang, Ersi; Milkovic, Janja

1999-01-01

Various combinations of wind shields and national precipitation gauges commonly used in countries of the northern hemisphere have been studied in this paper, using the combined intercomparison data collected at 14 sites during the World Meteorological Organization's (WMO) Solid Precipitation Measurement Intercomparison Project. The results show that wind shields improve gauge catch of precipitation, particularly for snow. Shielded gauges, on average, measure 20–70% more snow than unshielded gauges. Without a doubt, the use of wind shields on precipitation gauges has introduced a significant discontinuity into precipitation records, particularly in cold and windy regions. This discontinuity is not constant and it varies with wind speed, temperature, and precipitation type. Adjustment for this discontinuity is necessary to obtain homogenous precipitation data for climate change and hydrological studies. The relation of the relative catch ratio (RCR, ratio of measurements of shielded gauge to unshielded gauge) versus wind speed and temperature has been developed for Alter and Tretyakov wind shields. Strong linear relations between measurements of shielded gauge and unshielded gauge have also been found for different precipitation types. The linear relation does not fully take into account the varying effect of wind and temperature on gauge catch. Overadjustment by the linear relation may occur at those sites with lower wind speeds, and underadjustment may occur at those stations with higher wind speeds. The RCR technique is anticipated to be more applicable in a wide range of climate conditions. The RCR technique and the linear relation have been tested at selected WMO intercomparison stations, and reasonable agreement between the adjusted amounts and the shielded gauge measurements was obtained at most of the sites. Test application of the developed methodologies to a regional or national network is therefore recommended to further evaluate their applicability in

Heat transfer test of an 0.006-scale thin-skin thermocouple space shuttle model (50-0, 41-T) in the NASA-Ames Research Center 3.5-foot hypersonic wind tunnel at Mach 5.3 (IH28), volume 1

NASA Technical Reports Server (NTRS)

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

1976-01-01

Data obtained from a heat transfer test conducted on an 0.006-scale space shuttle orbiter and external tank in the NASA-Ames Research Center 3.5-foot Hypersonic Wind Tunnel are presented. The purpose of this test was to obtain data under simulated return-to-launch-site abort conditions. Configurations tested were integrated orbiter and external tank, orbiter alone, and external tank alone at angles of attack of 0, + or - 30, + or - 60, + or - 90, and + or - 120 degrees. Runs were conducted at Mach numbers of 5.2 and 5.3 for Reynolds numbers of 1.0 and 4.0 million per foot, respectively. Heat transfer data were obtained from 75 orbiter and 75 external tank iron-constantan thermocouples.

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.

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.

A study of hypersonic small-disturbance theory

NASA Technical Reports Server (NTRS)

Van Dyke, Milton D

1954-01-01

A systematic study is made of the approximate inviscid theory of thin bodies moving at such high supersonic speeds that nonlinearity is an essential feature of the equations of flow. The first-order small-disturbance equations are derived for three-dimensional motions involving shock waves, and estimates are obtained for the order of error involved in the approximation. The hypersonic similarity rule of Tsien and Hayes, and Hayes' unsteady analogy appear in the course of the development. It is shown that the hypersonic theory can be interpreted so that it applies also in the range of linearized supersonic flow theory. Several examples are solved according to the small-disturbance theory, and compared with the full solutions when available.

Multi-Exciter Vibroacoustic Simulation of Hypersonic Flight Vibration

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

GREGORY,DANNY LYNN; CAP,JEROME S.; TOGAMI,THOMAS C.

1999-11-11

Many aerospace structures must survive severe high frequency, hypersonic, random vibration during their flights. The random vibrations are generated by the turbulent boundary layer developed along the exterior of the structures during flight. These environments have not been simulated very well in the past using a fixed-based, single exciter input with an upper frequency range of 2 kHz. This study investigates the possibility of using acoustic ardor independently controlled multiple exciters to more accurately simulate hypersonic flight vibration. The test configuration, equipment, and methodology are described. Comparisons with actual flight measurements and previous single exciter simulations are also presented.

A hypersonic vehicle approach to planetary exploration

NASA Technical Reports Server (NTRS)

Murbach, Marcus S.

1993-01-01

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

High speed digital holographic interferometry for hypersonic flow visualization

NASA Astrophysics Data System (ADS)

Hegde, G. M.; Jagdeesh, G.; Reddy, K. P. J.

2013-06-01

Optical imaging techniques have played a major role in understanding the flow dynamics of varieties of fluid flows, particularly in the study of hypersonic flows. Schlieren and shadowgraph techniques have been the flow diagnostic tools for the investigation of compressible flows since more than a century. However these techniques provide only the qualitative information about the flow field. Other optical techniques such as holographic interferometry and laser induced fluorescence (LIF) have been used extensively for extracting quantitative information about the high speed flows. In this paper we present the application of digital holographic interferometry (DHI) technique integrated with short duration hypersonic shock tunnel facility having 1 ms test time, for quantitative flow visualization. Dynamics of the flow fields in hypersonic/supersonic speeds around different test models is visualized with DHI using a high-speed digital camera (0.2 million fps). These visualization results are compared with schlieren visualization and CFD simulation results. Fringe analysis is carried out to estimate the density of the flow field.

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.

Experimental Study of Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Aeroshell with Axisymmetric Surface Deflection Patterns

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Hollingsworth, Kevin E.

2017-01-01

A wind tunnel test program was conducted to obtain aeroheating environment data on Hypersonic Inflatable Aerodynamic Decelerator aeroshells with flexible thermal protection systems. Data were obtained on a set of rigid wind tunnel models with surface deflection patterns of various heights that simulated a range of potential in-flight aeroshell deformations. Wind tunnel testing was conducted at Mach 6 at unit Reynolds numbers from 2.1 × 10(exp 6)/ft to 8.3 × 10(exp 6)/ft and angles of attack from 0 deg to 18 deg. Boundary-layer transition onset and global surface heating distribution measurements were performed using phosphor thermography and flow field images were obtained through schlieren photography. Surface deflections were found to both promote early transition of the boundary layer and to augment heating levels for both laminar and turbulent flows. A complimentary computational flow field study was also performed to provide heating predictions for comparison with the measurements as well as boundary layer flow field properties for use in correlating the data. Correlations of the wind tunnel data were developed to predict deflection effects on boundary layer transition and surface heating and were applied to both the wind tunnel test conditions and to the trajectory of NASA's successful IRVE-3 flight test. In general, the correlations produced at least qualitative agreement with the wind tunnel data, although the heating levels were underpredicted for some of the larger surface deflections. For the flight conditions, the correlations suggested that peak heating levels on the leeward side conical flank of the IRVE-3 vehicle may have exceeded those at nose for times late in the trajectory after the peak heating time point. However, the flight estimates were based on a conservative assumption of surface deflection magnitude (i.e., larger) than likely was produced in flight.

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.

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.

Results of investigation on an 0.004-scale 140c modified configuration space shuttle vehicle orbiter model (74-0) in the NASA/Langley Research Center hypersonic nitrogen tunnel (OA89)

NASA Technical Reports Server (NTRS)

Hawthorne, P. J.

1975-01-01

Wind tunnel test data for the modified space shuttle vehicle orbiter is documented. Tests were made at various elevon settings and additionally in wing off/bodyflap off configuration at angles of attack from -5 to 42.5 degrees at zero yaw. Data obtained on high hypersonic longitudinal and lateral directional stability and control characteristics of the updated SSV configuration in an initially diatomic medium are included.

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.

SENECA NATION FINAL REPORT Installation of a 1.5MW WIND TURBINE

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

Giacobbe, Anthony

In February 2014, Seneca Energy and the Seneca Nation officially kicked off the Cattaraugus wind turbine project. The project began with a nine month NEPA study, which resulted in a Finding of No Significant Impact on 10/5/2015. The Seneca Construction Management Company was responsible for the oversight of the civil and foundation construction which commenced shortly after the FONSI was finalized. In April 2017, the Seneca Nation paid its first $25 credit to residents enrolled in the “Seneca Electricity Credit Program”. The $25 credit is applied directly to resident’s utility invoices and will total $300 of savings every year andmore » ~$8,000 of savings over the project lifecycle. The Seneca Nation has already paid $40,000 towards resident’s electricity bills since the commissioning of the wind turbine in April 2017. The Seneca Electricity Credit Program was one of the primary driving factors for the project and we continue to have additional residents enroll for the program each month. Our goal is to continue to add residents to our credit program and assist them in keeping electricity rates down, so that they can use their money for other necessary living expenses. The Seneca Nation and Seneca Energy are extremely proud of the Cattaraugus wind turbine as well as all of the other exciting energy projects taking place on territory. We are continuing to follow our Strategic Energy Plan as we move forward towards achieving energy self-sufficiency.« less

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.

NREL National Wind Technology Center (NWTC): M2 Tower; Boulder, Colorado (Data)

DOE Data Explorer

Jager, D.; Andreas, A.

1996-09-24

The National Wind Technology Center (NWTC), located at the foot of the Rocky Mountains near Boulder, Colorado, is a world-class research facility managed by NREL for the U.S. Department of Energy. NWTC researchers work with members of the wind energy industry to advance wind power technologies that lower the cost of wind energy through research and development of state-of-the-art wind turbine designs. NREL's Measurement and Instrument Data Center provides data from NWTC's M2 tower which are derived from instruments mounted on or near an 82 meter (270 foot) meteorological tower located at the western edge of the NWTC site and about 11 km (7 miles) west of Broomfield, and approximately 8 km (5 miles) south of Boulder, Colorado. The data represent the mean value of readings taken every two seconds and averaged over one minute. The wind speed and direction are measured at six heights on the tower and air temperature is measured at three heights. The dew point temperature, relative humidity, barometric pressure, totalized liquid precipitation, and global solar radiation are also available.

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.

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

Hypersonic Vehicle Propulsion System Simplified Model Development

NASA Technical Reports Server (NTRS)

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

2007-01-01

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

Hypersonic Boundary Layer Instability Over a Corner

NASA Technical Reports Server (NTRS)

Balakumar, Ponnampalam; Zhao, Hong-Wu; McClinton, Charles (Technical Monitor)

2001-01-01

A boundary-layer transition study over a compression corner was conducted under a hypersonic flow condition. Due to the discontinuities in boundary layer flow, the full Navier-Stokes equations were solved to simulate the development of disturbance in the boundary layer. A linear stability analysis and PSE method were used to get the initial disturbance for parallel and non-parallel flow respectively. A 2-D code was developed to solve the full Navier-stokes by using WENO(weighted essentially non-oscillating) scheme. The given numerical results show the evolution of the linear disturbance for the most amplified disturbance in supersonic and hypersonic flow over a compression ramp. The nonlinear computations also determined the minimal amplitudes necessary to cause transition at a designed location.

A Free-flight Wind Tunnel for Aerodynamic Testing at Hypersonic Speeds

NASA Technical Reports Server (NTRS)

Seiff, Alvin

1954-01-01

The supersonic free-flight wind tunnel is a facility at the Ames Laboratory of the NACA in which aerodynamic test models are gun-launched at high speed and directed upstream through the test section of a supersonic wind tunnel. In this way, test Mach numbers up to 10 have been attained and indications are that still higher speeds will be realized. An advantage of this technique is that the air and model temperatures simulate those of flight through the atmosphere. Also the Reynolds numbers are high. Aerodynamic measurements are made from photographic observation of the model flight. Instruments and techniques have been developed for measuring the following aerodynamic properties: drag, initial lift-curve slope, initial pitching-moment-curve slope, center of pressure, skin friction, boundary-layer transition, damping in roll, and aileron effectiveness. (author)

Control integration concept for hypersonic cruise-turn maneuvers

NASA Technical Reports Server (NTRS)

Raney, David L.; Lallman, Frederick J.

1992-01-01

Piloting difficulties associated with conducting aircraft maneuvers in hypersonic flight are caused in part by the nonintuitive nature of the aircraft response and the stringent constraints anticipated on allowable angle of attack and dynamic pressure variations. An approach is documented that provides precise, coordinated maneuver control during excursions from a hypersonic cruise flight path and the necessary flight condition constraints. The approach is to achieve specified guidance commands by resolving altitude and cross range errors into a load factor and bank angle command by using a coordinate transformation that acts as an interface between outer and inner loop flight controls. This interface, referred to as a 'resolver', applies constraints on angle of attack and dynamic pressure perturbations while prioritizing altitude regulation over cross range. An unpiloted test simulation, in which the resolver was used to drive inner loop flight controls, produced time histories of responses to guidance commands and atmospheric disturbances at Mach numbers of 6, 10, 15, and 20. Angle of attack and throttle perturbation constraints, combined with high speed flight effects and the desire to maintain constant dynamic pressure, significantly impact the maneuver envelope for a hypersonic vehicle.

Supersonic and hypersonic shock/boundary-layer interaction database

NASA Technical Reports Server (NTRS)

Settles, Gary S.; Dodson, Lori J.

1994-01-01

An assessment is given of existing shock wave/tubulent boundary-layer interaction experiments having sufficient quality to guide turbulence modeling and code validation efforts. Although the focus of this work is hypersonic, experiments at Mach numbers as low as 3 were considered. The principal means of identifying candidate studies was a computerized search of the AIAA Aerospace Database. Several hundred candidate studies were examined and over 100 of these were subjected to a rigorous set of acceptance criteria for inclusion in the data-base. Nineteen experiments were found to meet these criteria, of which only seven were in the hypersonic regime (M is greater than 5).

A technique for measuring hypersonic flow velocity profiles

NASA Technical Reports Server (NTRS)

Gartrell, L. R.

1973-01-01

A technique for measuring hypersonic flow velocity profiles is described. This technique utilizes an arc-discharge-electron-beam system to produce a luminous disturbance in the flow. The time of flight of this disturbance was measured. Experimental tests were conducted in the Langley pilot model expansion tube. The measured velocities were of the order of 6000 m/sec over a free-stream density range from 0.000196 to 0.00186 kg/cu m. The fractional error in the velocity measurements was less than 5 percent. Long arc discharge columns (0.356 m) were generated under hypersonic flow conditions in the expansion-tube modified to operate as an expansion tunnel.

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.;

Results of wind tunnel RCS interaction tests on a 0.010-scale space shuttle orbiter model (51-0) in the Calspan Corporation 48-inch hypersonic shock tunnel (test 0A93)

NASA Technical Reports Server (NTRS)

Daileda, J. J.; Marroquin, J.; Rogers, C. E.

1976-01-01

A hypersonic shock tunnel test on a 0.010 scale SSV orbital configuration was performed to determine the effects of RCS jet/flow field interactions on SSV aerodynamic stability and control characteristics at various hypersonic Mach and Reynolds numbers. Flow field interaction data were obtained using pitch and roll jets. In addition, direct impingement data were obtained at a Mach number of zero with the test section pumped down to below 10 microns of mercury pressure.

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.

A review of design issues specific to hypersonic flight vehicles

NASA Astrophysics Data System (ADS)

Sziroczak, D.; Smith, H.

2016-07-01

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

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.

Direct Numerical Simulation of Acoustic Noise Generation from the Nozzle Wall of a Hypersonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Huang, Junji; Duan, Lian; Choudhari, Meelan M.

2017-01-01

The acoustic radiation from the turbulent boundary layer on the nozzle wall of a Mach 6 Ludwieg Tube is simulated using Direct Numerical Simulations (DNS), with the flow conditions falling within the operational range of the Mach 6 Hypersonic Ludwieg Tube, Braunschweig (HLB). The mean and turbulence statistics of the nozzle-wall boundary layer show good agreement with those predicted by Pate's correlation and Reynolds Averaged Navier-Stokes (RANS) computations. The rms pressure fluctuation P'(rms)/T(w) plateaus in the freestream core of the nozzle. The intensity of the freestream noise within the nozzle is approximately 20% higher than that radiated from a single at pate with a similar freestream Mach number, potentially because of the contributions to the acoustic radiation from multiple azimuthal segments of the nozzle wall.

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

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.

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.

A Numerical Study of 2-D Surface Roughness Effects on the Growth of Wave Modes in Hypersonic Boundary Layers

NASA Astrophysics Data System (ADS)

Fong, Kahei Danny

to find the most unstable second-mode frequency that can lead to transition. And by doing a theoretical analysis such as LST, the synchronization location for the most unstable frequency can be found. Roughness elements are then strategically placed downstream of the synchronization point to damp out this dangerous second-mode wave, thus stabilizing the boundary layer and suppressing the transition process. This method is later experimentally validated in Purdue's Mach 6 quiet wind tunnel. Overall, this research has not only provided details of when and how 2-D roughness stabilizes a hypersonic boundary layer, it also has led to a successful application of numerical simulation data to the development of a new roughness-based transition delay method, which could potentially have significant contributions to the design of future generation hypersonic vehicles.

The National Wind Energy Skills Assessment and Preparing for the Future Wind Workforce; NREL (National Renewable Energy Laboratory)

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

Tegen, Suzanne

2015-07-10

A robust workforce is essential to growing domestic wind manufacturing capabilities. This presentation provides an overview of an NREL analysis of wind-focused education at American colleges and universities. The second part of the presentation discusses DOE/NREL workforce-related projects, such as the Wind Career Map, the Collegiate Wind Competition, and the Wind for Schools project.

Perspectives on hypersonic viscous and nonequilibrium flow research

NASA Technical Reports Server (NTRS)

Cheng, H. K.

1992-01-01

An attempt is made to reflect on current focuses in certain areas of hypersonic flow research by examining recent works and their issues. Aspects of viscous interaction, flow instability, and nonequilibrium aerothermodynamics pertaining to theoretical interest are focused upon. The field is a diverse one, and many exciting works may have either escaped the writer's notice or been abandoned for the sake of space. Students of hypersonic viscous flow must face the transition problems towards the two opposite ends of the Reynolds or Knudsen number range, which represents two regimes where unresolved fluid/gas dynamic problems abound. Central to the hypersonic flow studies is high-temperature physical gas dynamics; here, a number of issues on modelling the intermolecular potentials and inelastic collisions remain the obstacles to quantitative predictions. Research in combustion and scramjet propulsion will certainly be benefitted by advances in turbulent mixing and new computational fluid dynamics (CFD) strategies on multi-scaled complex reactions. Even for the sake of theoretical development, the lack of pertinent experimental data in the right energy and density ranges is believed to be among the major obstacles to progress in aerothermodynamic research for hypersonic flight. To enable laboratory simulation of nonequilibrium effects anticipated for transatmospheric flight, facilities capable of generating high enthalpy flow at density levels higher than in existing laboratories are needed (Hornung 1988). A new free-piston shock tunnel capable of realizing a test-section stagnation temperature of 10(exp 5) at Reynolds number 50 x 10(exp 6)/cm is being completed and preliminary tests has begun (H. Hornung et al. 1992). Another laboratory study worthy of note as well as theoretical support is the nonequilibrium flow experiment of iodine vapor which has low activation energies for vibrational excitation and dissociation, and can be studied in a laboratory with modest

75 FR 52374 - National Environmental Policy Act; NASA Glenn Research Center Plum Brook Station Wind Farm Project

Federal Register 2010, 2011, 2012, 2013, 2014

2010-08-25

...; NASA Glenn Research Center Plum Brook Station Wind Farm Project AGENCY: National Aeronautics and Space... Environmental Impact Statement (EIS) for the NASA GRC Plum Brook Station Wind Farm Project located near Sandusky... obtain public comments on construction and operation of the wind farm. The purpose of constructing and...

A feasibility study of a hypersonic real-gas facility

NASA Technical Reports Server (NTRS)

Gully, J. H.; Driga, M. D.; Weldon, W. F.

1987-01-01

A four month feasibility study of a hypersonic real-gas free flight test facility for NASA Langley Research Center (LARC) was performed. The feasibility of using a high-energy electromagnetic launcher (EML) to accelerate complex models (lifting and nonlifting) in the hypersonic, real-gas facility was examined. Issues addressed include: design and performance of the accelerator; design and performance of the power supply; design and operation of the sabot and payload during acceleration and separation; effects of high current, magnetic fields, temperature, and stress on the sabot and payload; and survivability of payload instrumentation during acceleration, flight, and soft catch.

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.

eLaunch Hypersonics: An Advanced Launch System

NASA Technical Reports Server (NTRS)

Starr, Stanley

2010-01-01

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

Models of Interacting Stellar Winds

NASA Astrophysics Data System (ADS)

Wilkin, Francis Patrick

Stars drive supersonic winds which interact violently with their surroundings. Analytic and numerical models of hypersonic, interacting circumstellar flows are presented for several important astrophysical problems. A new solution method for steady-state, axisymmetric, wind collision problems is applied to radiative bow shocks from moving stars and to the collision of two spherical winds in a binary star system. The solutions obtained describe the shape of the geometrically thin, shocked shell of matter, as well as its mass surface density and the tangential velocity within it. Analytic solutions are also obtained for non-axisymmetric bow shocks, where the asymmetry arises due to either a transverse gradient in the ambient medium, or a misaligned, axisymmetric stellar wind. While the solutions are all easily scaled in terms of their relevant dimensional parameters, the important assumption of radiative shocks implies that the models are most applicable towards systems with dense environments and low preshock velocities. The bow shock model has previously been applied to cometary, ultracompact HII regions by Van Buren et al. (1990), who discussed extensively the applicability of the thin shell approximation. I next model the collision between a protostellar wind and supersonic infall from a rotating cloud, employing a quasi-steady, thin-shell formulation. The spherical wind is initially crushed to the protostellar surface by nearly spherical infall. The centrifugal distortion of infalling matter eventually permits a wind-supported, trapped bubble to slowly expand on an evolutionary (~ 105 yr) time. The shell becomes progressively more extended along the rotational axis, due to the asymmetry of the infall. When the quasi-steady assumption breaks down, the shell has become a needle-like, bipolar configuration that may represent a precursor to protostellar jets. I stress, however, the likelihood of instability for the shell, and the possibility of oscillatory behavior

Combined LAURA-UPS hypersonic solution procedure

NASA Technical Reports Server (NTRS)

Wood, William A.; Thompson, Richard A.

1993-01-01

A combined solution procedure for hypersonic flowfields around blunted slender bodies was implemented using a thin-layer Navier-Stokes code (LAURA) in the nose region and a parabolized Navier-Stokes code (UPS) on the after body region. Perfect gas, equilibrium air, and non-equilibrium air solutions to sharp cones and a sharp wedge were obtained using UPS alone as a preliminary step. Surface heating rates are presented for two slender bodies with blunted noses, having used LAURA to provide a starting solution to UPS downstream of the sonic line. These are an 8 deg sphere-cone in Mach 5, perfect gas, laminar flow at 0 and 4 deg angles of attack and the Reentry F body at Mach 20, 80,000 ft equilibrium gas conditions for 0 and 0.14 deg angles of attack. The results indicate that this procedure is a timely and accurate method for obtaining aerothermodynamic predictions on slender hypersonic vehicles.

The Blunt Plate In Hypersonic Flow

NASA Technical Reports Server (NTRS)

Baradell, Donald L.; Bertram, Mitchel H.

1960-01-01

The sonic-wedge characteristics method has been used to obtain the shock shapes and surface pressure distributions on several blunt two-dimensional shapes in a hypersonic stream for several values of the ratio of specific heats. These shapes include the blunt slab at angle of attack and power profiles of the form yb = a)P, where 0 les than m less than 1, Yb and x are coordinates of the body surface, and a is a constant. These numerical results have been compared with the results of blast-wave theory, and methods of predicting the pressure distributions and shock shapes are proposed in each case. The effects of a free-stream conical-flow gradient on the pressure distribution on a blunt slab in hypersonic flow were investigated by the sonic-wedge characteristics method and were found to be sizable in many cases. Procedures which are satisfactory for reducing pressure data obtained in conical flows with small gradients are presented.

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.

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.

Assessment of CFD capability for prediction of hypersonic shock interactions

NASA Astrophysics Data System (ADS)

Knight, Doyle; Longo, José; Drikakis, Dimitris; Gaitonde, Datta; Lani, Andrea; Nompelis, Ioannis; Reimann, Bodo; Walpot, Louis

2012-01-01

The aerothermodynamic loadings associated with shock wave boundary layer interactions (shock interactions) must be carefully considered in the design of hypersonic air vehicles. The capability of Computational Fluid Dynamics (CFD) software to accurately predict hypersonic shock wave laminar boundary layer interactions is examined. A series of independent computations performed by researchers in the US and Europe are presented for two generic configurations (double cone and cylinder) and compared with experimental data. The results illustrate the current capabilities and limitations of modern CFD methods for these flows.

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.

Remote Infrared Imaging of the Space Shuttle During Hypersonic Flight: HYTHIRM Mission Operations and Coordination

NASA Technical Reports Server (NTRS)

Schwartz, Richard J.; McCrea, Andrew C.; Gruber, Jennifer R.; Hensley, Doyle W.; Verstynen, Harry A.; Oram, Timothy D.; Berger, Karen T.; Splinter, Scott C.; Horvath, Thomas J.; Kerns, Robert V.

2011-01-01

The Hypersonic Thermodynamic Infrared Measurements (HYTHIRM) project has been responsible for obtaining spatially resolved, scientifically calibrated in-flight thermal imagery of the Space Shuttle Orbiter during reentry. Starting with STS-119 in March of 2009 and continuing through to the majority of final flights of the Space Shuttle, the HYTHIRM team has to date deployed during seven Shuttle missions with a mix of airborne and ground based imaging platforms. Each deployment of the HYTHIRM team has resulted in obtaining imagery suitable for processing and comparison with computational models and wind tunnel data at Mach numbers ranging from over 18 to under Mach 5. This paper will discuss the detailed mission planning and coordination with the NASA Johnson Space Center Mission Control Center that the HYTHIRM team undergoes to prepare for and execute each mission.

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

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.

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

Detached Eddy Simulations of Hypersonic Transition

NASA Technical Reports Server (NTRS)

Yoon, S.; Barnhardt, M.; Candler, G.

2010-01-01

This slide presentation reviews the use of Detached Eddy Simulation (DES) of hypersonic transistion. The objective of the study was to investigate the feasibility of using CFD in general, DES in particular, for prediction of roughness-induced boundary layer transition to turbulence and the resulting increase in heat transfer.

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

NASA Technical Reports Server (NTRS)

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

2010-01-01

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

Tropospheric Wind Monitoring During Day-of-Launch Operations for National Aeronautics and Space Administration's Space Shuttle Program

NASA Technical Reports Server (NTRS)

Decker, Ryan K.; Leach, Richard

2004-01-01

The Environments Group at the National Aeronautics and Space Administration's Marshall Space Flight Center (NASA/MSFC) monitors the winds aloft at Kennedy Space Center (KSC) during the countdown for all Space Shuttle launches. Assessment of tropospheric winds is used to support the ascent phase of launch. Three systems at KSC are used to generate independent tropospheric wind profiles prior to launch; 1) high resolution Jimsphere balloon system, 2) 50-MHz Doppler Radar Wind Profiler (DRWP) and 3) low resolution radiosonde system. Data generated by the systems are used to assess spatial and temporal wind variability during launch countdown to ensure wind change observed does not violate wind change criteria constraints.

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.

Computational analysis of hypersonic flows past elliptic-cone waveriders

NASA Technical Reports Server (NTRS)

Yoon, Bok-Hyun; Rasmussen, Maurice L.

1991-01-01

A comprehensive study for the inviscid numerical calculation of the hypersonic flow past a class of elliptic-cone derived waveriders is presented. The theoretical background associated with hypersonic small-disturbance theory (HSDT) is reviewed. Several approximation formulas for the waverider compression surface are established. A CFD algorithm is used to calculate flow fields for the on-design case and a variety of off-design cases. The results are compared with HSDT, experiment, and other available CFD results. For the waverider shape used in previous investigations, the bow shock for the on-design condition stands off from the leading-edge tip of the waverider. It was found that this occurs because the tip was too thick according to the approximating shape formula that was used to describe the compression surface. When this was corrected, the bow shock became closer to attached as it should be. At Mach numbers greater than the design condition, a lambda-shock configuration develops near the tip of the compression surface. At negative angles of attack, other complicated shock patterns occur near the leading-edge tip. These heretofore unknown flow patterns show the power and utility of CFD for investigating novel hypersonic configurations such as waveriders.

National Wind Tecnology Center Provides Dual Axis Resonant Blade Testing

ScienceCinema

Felker, Fort

2018-01-16

NREL's Structural Testing Laboratory at the National Wind Technology Center (NWTC) provides experimental laboratories, computer facilities for analytical work, space for assembling components and turbines for atmospheric testing as well as office space for industry researchers. Fort Felker, center director at the NWTC, discusses NREL's state-of-the-art structural testing capabilities and shows a flapwise and edgewise blade test in progress.

National Wind Tecnology Center Provides Dual Axis Resonant Blade Testing

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

Felker, Fort

2013-11-13

NREL's Structural Testing Laboratory at the National Wind Technology Center (NWTC) provides experimental laboratories, computer facilities for analytical work, space for assembling components and turbines for atmospheric testing as well as office space for industry researchers. Fort Felker, center director at the NWTC, discusses NREL's state-of-the-art structural testing capabilities and shows a flapwise and edgewise blade test in progress.

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.

Santa Ana Winds and Fire Regimes of Southern California National Forests

NASA Astrophysics Data System (ADS)

Bendix, J.

2015-12-01

In Southern California, it has long been understood that foehn-type Santa Ana winds are an important factor in the occurrence of large wildfires. Although a variety of anecdotal observations and statistical analyses have confirmed the importance of these winds to wildfire, particularly in the Fall months when Santa Ana winds overlap with dry fuels from summer drought, many of the details of those winds' impacts on fire remain obscure. This paper uses data regarding individual fires from California's Fire and Resource Assessment Program database and a compilation of Santa Ana Wind days (SAW days) published by Abatzoglou et al. in 2013 to assess the relationship of Santa Ana winds to fire occurrence and size in Southern California. The analysis included 474 fires larger than 20 ha (~50 acres).that burned on the four Southern California national forests (Angeles, Cleveland, Los Padres and San Bernardino) between 1948 and 2010. Overall, just 10.3% of the fires started on SAW days, and 14.4% experienced at least one SAW day between start and containment dates. The impact of Santa Ana winds is greater, however, with increasing fire size. For fires > 4000 ha, 18.4% began on SAW days, with 30.4% experiencing at least one SAW day before containment. And 20% of fires > 20000 ha started on SAW days, with 50% including one or more SAW days. Fires beginning on SAW days were larger, with a mean of 6239 ha compared to 2150 ha for fires that began on non-SAW days. Only 2% of the fires that began on SAW days were started by lightning, suggesting that the impact of Santa Ana winds on Southern California fire regimes may be enhanced by humans' role in ignitions.

Hypersonic Wake Diagnostics Using Laser Induced Fluorescence Techniques

NASA Technical Reports Server (NTRS)

Mills, Jack L.; Sukenik, Charles I.; Balla, Robert J.

2011-01-01

A review of recent research performed in iodine that involves a two photon absorption of light at 193 nm will be discussed, and it's potential application to velocimetry measurements in a hypersonic flow field will be described. An alternative seed atom, Krypton, will be presented as a good candidate for performing nonintrusive hypersonic flow diagnostics. Krypton has a metastable state with a lifetime of approximately 43 s which would prove useful for time of flight measurement (TOF) and a sensitivity to collisions that can be utilized for density measurements. Calculations using modest laser energies and experimental values show an efficiency of excited state production to be on the order of 10(exp -6) for a two photon absorption at 193 nm.

Progress with multigrid schemes for hypersonic flow problems

NASA Technical Reports Server (NTRS)

Radespiel, R.; Swanson, R. C.

1991-01-01

Several multigrid schemes are considered for the numerical computation of viscous hypersonic flows. For each scheme, the basic solution algorithm uses upwind spatial discretization with explicit multistage time stepping. Two level versions of the various multigrid algorithms are applied to the two dimensional advection equation, and Fourier analysis is used to determine their damping properties. The capabilities of the multigrid methods are assessed by solving three different hypersonic flow problems. Some new multigrid schemes based on semicoarsening strategies are shown to be quite effective in relieving the stiffness caused by the high aspect ratio cells required to resolve high Reynolds number flows. These schemes exhibit good convergence rates for Reynolds numbers up to 200 x 10(exp 6) and Mach numbers up to 25.

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.

Probabilistic Assessment of National Wind Tunnel

NASA Technical Reports Server (NTRS)

Shah, A. R.; Shiao, M.; Chamis, C. C.

1996-01-01

A preliminary probabilistic structural assessment of the critical section of National Wind Tunnel (NWT) is performed using NESSUS (Numerical Evaluation of Stochastic Structures Under Stress) computer code. Thereby, the capabilities of NESSUS code have been demonstrated to address reliability issues of the NWT. Uncertainties in the geometry, material properties, loads and stiffener location on the NWT are considered to perform the reliability assessment. Probabilistic stress, frequency, buckling, fatigue and proof load analyses are performed. These analyses cover the major global and some local design requirements. Based on the assumed uncertainties, the results reveal the assurance of minimum 0.999 reliability for the NWT. Preliminary life prediction analysis results show that the life of the NWT is governed by the fatigue of welds. Also, reliability based proof test assessment is performed.

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.

Wind Vision: A New Era for Wind Power in the United States

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

U.S. Department of Energy

With more than 4.5% of the nation's electricity supplied by wind energy today, the Department of Energy has collaborated with industry, environmental organizations, academic institutions, and national laboratories to develop a renewed Wind Vision, documenting the contributions of wind to date and envisioning a future where wind continues to provide key contributions to the nation’s energy portfolio. Building on and updating the 2008 20% Wind Energy by 2030 report, the new Wind Vision Report quantifies the economic, environmental, and social benefits of a robust wind energy future and the actions that wind stakeholders can take to make it a reality.

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.

Computational study of generic hypersonic vehicle flow fields

NASA Technical Reports Server (NTRS)

Narayan, Johnny R.

1994-01-01

The geometric data of the generic hypersonic vehicle configuration included body definitions and preliminary grids for the forebody (nose cone excluded), midsection (propulsion system excluded), and afterbody sections. This data was to be augmented by the nose section geometry (blunt conical section mated with the noncircular cross section of the forebody initial plane) along with a grid and a detailed supersonic combustion ramjet (scramjet) geometry (inlet and combustor) which should be merged with the nozzle portion of the afterbody geometry. The solutions were to be obtained by using a Navier-Stokes (NS) code such as TUFF for the nose portion, a parabolized Navier-Stokes (PNS) solver such as the UPS and STUFF codes for the forebody, a NS solver with finite rate hydrogen-air chemistry capability such as TUFF and SPARK for the scramjet and a suitable solver (NS or PNS) for the afterbody and external nozzle flows. The numerical simulation of the hypersonic propulsion system for the generic hypersonic vehicle is the major focus of this entire work. Supersonic combustion ramjet is such a propulsion system, hence the main thrust of the present task has been to establish a solution procedure for the scramjet flow. The scramjet flow is compressible, turbulent, and reacting. The fuel used is hydrogen and the combustion process proceeds at a finite rate. As a result, the solution procedure must be capable of addressing such flows.

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

NASA Technical Reports Server (NTRS)

Chamitoff, Gregory Errol

1992-01-01

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

Wind for Schools: A Wind Powering America Project

ERIC Educational Resources Information Center

US Department of Energy, 2007

2007-01-01

The U.S. Department of Energy's (DOE's) Wind Powering America program (based at the National Renewable Energy Laboratory) sponsors the Wind for Schools Project to raise awareness in rural America about the benefits of wind energy while simultaneously educating college seniors regarding wind energy applications. The three primary project goals of…

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.

1974-01-01

The configuration and performance of the propulsion system for the hypersonic research vehicle are discussed. A study of the interactions between propulsion and aerodynamics of the highly integrated vehicle was conducted. The hypersonic research vehicle is configured to test the technology of structural and thermal protection systems concepts and the operation of the propulsion system under true flight conditions for most of the hypersonic flight regime. The subjects considered are: (1) research vehicle and scramjet engine configurations to determine fundamental engine sizing constraints, (2) analytical methods for computing airframe and propulsion system components, and (3) characteristics of a candidate nozzle to investigate vehicle stability and acceleration performance.

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.

Multigrid for hypersonic viscous two- and three-dimensional flows

NASA Technical Reports Server (NTRS)

Turkel, E.; Swanson, R. C.; Vatsa, V. N.; White, J. A.

1991-01-01

The use of a multigrid method with central differencing to solve the Navier-Stokes equations for hypersonic flows is considered. The time dependent form of the equations is integrated with an explicit Runge-Kutta scheme accelerated by local time stepping and implicit residual smoothing. Variable coefficients are developed for the implicit process that removes the diffusion limit on the time step, producing significant improvement in convergence. A numerical dissipation formulation that provides good shock capturing capability for hypersonic flows is presented. This formulation is shown to be a crucial aspect of the multigrid method. Solutions are given for two-dimensional viscous flow over a NACA 0012 airfoil and three-dimensional flow over a blunt biconic.

Protection of surface assets on Mars from wind blown jettisoned spacecraft components

NASA Astrophysics Data System (ADS)

Paton, Mark

2017-07-01

Jettisoned Entry, Descent and Landing System (EDLS) hardware from landing spacecraft have been observed by orbiting spacecraft, strewn over the Martian surface. Future Mars missions that land spacecraft close to prelanded assets will have to use a landing architecture that somehow minimises the possibility of impacts from these jettisoned EDLS components. Computer modelling is used here to investigate the influence of wind speed and direction on the distribution of EDLS components on the surface. Typical wind speeds encountered in the Martian Planetary Boundary Layer (PBL) were found to be of sufficient strength to blow items having a low ballistic coefficient, i.e. Hypersonic Inflatable Aerodynamic Decelerators (HIADs) or parachutes, onto prelanded assets even when the lander itself touches down several kilometres away. Employing meteorological measurements and careful characterisation of the Martian PBL, e.g. appropriate wind speed probability density functions, may then benefit future spacecraft landings, increase safety and possibly help reduce the delta v budget for Mars landers that rely on aerodynamic decelerators.

@NWTC Newsletter | Wind | NREL

Science.gov Websites

Mutually Beneficial NREL Researchers Play Integral Role in National Offshore Wind Strategy NWTC's Grid . More NREL Researchers Play Integral Role in National Offshore Wind Strategy The national energy

National facilities study. Volume 2: Task group on aeronautical research and development facilities report

NASA Technical Reports Server (NTRS)

1994-01-01

The Task Group on Aeronautics R&D Facilities examined the status and requirements for aeronautics facilities against the competitive need. Emphasis was placed on ground-based facilities for subsonic, supersonic and hypersonic aerodynamics, and propulsion. Subsonic and transonic wind tunnels were judged to be most critical and of highest priority. Results of the study are presented.

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.

Flow field measurements for cylindrical configurations in a hypersonic wind tunnel: Windward and leeward flow fields

NASA Technical Reports Server (NTRS)

Bertin, J. J.; Lamb, J. P.; Center, K. R.; Graumann, B. W.

1971-01-01

Windward and leeward measurements were made for a variety of simulated infinite cylinders exposed to hypersonic streams over an angle of attack from 30 deg to 90 deg. For the range of conditions included in the study, the following conclusions are made: (1) Swept cylinder theory provides a reasonable correlation of the measured laminar heat transfer rates from the plane of symmetry. (2) The boundary layer transition criteria in the plane of symmetry are a function of the transverse curvature. (3) Relaminarization of the circumferential boundary layer for a right circular cylinder was observed at the highest Reynolds number tested. (4) The effect of leeside geometry on the average heat transfer rate can be correlated with a single geometric parameter which is dependent on the location of separation. (5) The relationship of leeward heating to angle of attack is virtually linear for each cross section. (6) No systematic effect of free stream Reynolds number was observed.

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.

On the instability of hypersonic flow past a flat plate

NASA Technical Reports Server (NTRS)

Blackaby, Nicholas; Cowley, Stephen; Hall, Philip

1990-01-01

The instability of hypersonic boundary-layer flows over flat plates is considered. The viscosity of the fluid is taken to be governed by Sutherland's law, which gives a much more accurate representation of the temperature dependence of fluid viscosity at hypersonic speeds than Chapman's approximate linear law; although at lower speeds the temperature variation of the mean state is less pronounced so that the Chapman law can be used with some confidence. Attention is focussed on the so-called (vorticity) mode of instability of the viscous hypersonic boundary layer. This is thought to be the fastest growing inviscid disturbance at hypersonic speeds; it is also believed to have an asymptotically larger growth rate than any viscous or centrifugal instability. As a starting point the instability of the hypersonic boundary layer which exists far downstream from the leading edge of the plate is investigated. In this regime the shock that is attached to the leading edge of the plate plays no role, so that the basic boundary layer is non-interactive. It is shown that the vorticity mode of instability of this flow operates on a significantly different lengthscale than that obtained if a Chapman viscosity law is assumed. In particular, it is found that the growth rate predicted by a linear viscosity law overestimates the size of the growth rate by O(M(exp 2). Next, the development of the vorticity mode as the wavenumber decreases is described, and it is shown that acoustic modes emerge when the wavenumber has decreased from it's O(1) initial value to O(M (exp -3/2). Finally, the inviscid instability of the boundary layer near the leading edge in the interaction zone is discussed and particular attention is focussed on the strong interaction region which occurs sufficiently close to the leading edge. It is found that the vorticity mode in this regime is again unstable, and that it is concentrated in the transition layer at the edge of the boundary layer where the temperature

Systems Challenges for Hypersonic Vehicles

NASA Technical Reports Server (NTRS)

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

1997-01-01

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

Wind Maps | Geospatial Data Science | NREL

Science.gov Websites

Wind Maps Wind Maps Wind Prospector This GIS application supports resource assessment and data exploration for wind development. This collection of wind maps and assessments details the wind resource in Geospatial Data Science Team. National Wind Resource Assessment The national wind resource assessment was

Performance analysis of a new hypersonic vitrector system.

PubMed

Stanga, Paulo Eduardo; Pastor-Idoate, Salvador; Zambrano, Isaac; Carlin, Paul; McLeod, David

2017-01-01

To evaluate porcine vitreous flow and water flow rates in a new prototype hypersonic vitrectomy system compared to currently available pneumatic guillotine vitrectors (GVs) systems. Two vitrectors were tested, a prototype, ultrasound-powered, hypersonic vitrector (HV) and a GV. Porcine vitreous was obtained within 12 to 24 h of sacrifice and kept at 4°C. A vial of vitreous or water was placed on a precision balance and its weight measured before and after the use of each vitrector. Test parameters included changes in aspiration levels, vitrector gauge, cut rates for GVs, % ultrasound (US) power for HVs, and port size for HVs. Data was analysed using linear regression and t-tests. There was no difference in the total average mean water flow between the 25-gauge GV and the 25-gauge HV (t-test: P = 0.363); however, 25-gauge GV was superior (t-test: P < 0.001) in vitreous flow. The 23-gauge GV was only more efficient in water and vitreous removal than 23-gauge HV needle-1 (Port 0.0055) (t-test: P < 0.001). For HV, wall thickness and gauge had no effect on flow rates. Water and vitreous flows showed a direct correlation with increasing aspiration levels and % US power (p<0.05). The HV produced consistent water and vitreous flow rates across the range of US power and aspiration levels tested. Hypersonic vitrectomy may be a promising new alternative to the currently available guillotine-based technologies.

Hypersonic blunt body computations including real gas effects

NASA Technical Reports Server (NTRS)

Montagne, J.-L.; Yee, H. C.; Klopfer, G. H.; Vinokur, M.

1989-01-01

Various second-order explicit and implicit TVD shock-capturing methods, a generalization of Roe's approximate Riemann solver, and a generalized flux-vector splitting scheme are used to study two-dimensional hypersonic real-gas flows. Special attention is given to the identification of some of the elements and parameters which can affect the convergence rate for high Mach numbers or real gases, but have negligible effect for low Mach numbers, for cases involving steady-state inviscid blunt flows. Blunt body calculations at Mach numbers of greater than 15 are performed to treat real-gas effects, and impinging shock results are obtained to test the treatment of slip surfaces and complex structures. Even with the addition of improvements, the convergence rate of algorithms in the hypersonic flow regime is found to be generally slower for a real gas than for a perfect gas.

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

Research on Aeroheating of Hypersonic Reentry Vehicle Base Flow Fields

NASA Astrophysics Data System (ADS)

Xuguo, Qin; Yongtao, Shui; Yonghai, Wang; Gang, Chen; Qiang, Li

2017-09-01

The structure of the base flow of a hypersonic reentry vehicle and the resulting base pressure and heat transfer have been studied by numerical study. The compressible Navier-Stokes equations are solved by the finite-volume method. SST k-ω turbulence model is used, and comparisons are made with flight test. Attention was focused on assessing the effects of angle of attack and Mach number. It was found that angle of attack can significantly alter the wake flow structure and reentry vehicle base pressure and heating distributions. The results of the simulation may provide a theoretical basis for the design of the thermal protection system of hypersonic reentry vehicles.

The HIAD stands for Hypersonic Inflatable Aerodynamic Decelerato

NASA Image and Video Library

2013-05-31

The HIAD stands for Hypersonic Inflatable Aerodynamic Decelerator, an inflatable spacecraft technology that allows payloads to survive the harsh conditions of atmospheric re-entry. This photo was taken at NASA Langley in Building 1250 when sensors were being applied.

The HIAD stands for Hypersonic Inflatable Aerodynamic Decelerato

NASA Image and Video Library

2013-04-25

The HIAD stands for Hypersonic Inflatable Aerodynamic Decelerator, an inflatable spacecraft technology that allows payloads to survive the harsh conditions of atmospheric re-entry. This photo was taken at NASA Langley in Building 1250 when sensors were being applied.

Results of tests of a Rockwell International space shuttle orbiter (-139 configuration) 0.0175-scale model (no. 29-0) in AEDC tunnel F to determine hypersonic heating effects (OH11)

NASA Technical Reports Server (NTRS)

Quan, M.

1975-01-01

Results from wind tunnel tests to determine hypersonic aerodynamic heating rates on a NASA/Rockwell Space Shuttle Orbiter are reported. The tests were to determine Mach number effects, if any, and to obtain overall heating rate data at high Mach numbers from 10.5 to 16. The model used was a 0.0175-scale model built to Rockwell Orbiter lines VL70-000139. The model identity number is 29-0. These tests, designated OH11, were conducted in the AEDC Tunnel F.

Formulation of aerodynamic prediction techniques for hypersonic configuration design

NASA Technical Reports Server (NTRS)

1979-01-01

An investigation of approximate theoretical techniques for predicting aerodynamic characteristics and surface pressures for relatively slender vehicles at moderate hypersonic speeds was performed. Emphasis was placed on approaches that would be responsive to preliminary configuration design level of effort. Supersonic second order potential theory was examined in detail to meet this objective. Shock layer integral techniques were considered as an alternative means of predicting gross aerodynamic characteristics. Several numerical pilot codes were developed for simple three dimensional geometries to evaluate the capability of the approximate equations of motion considered. Results from the second order computations indicated good agreement with higher order solutions and experimental results for a variety of wing like shapes and values of the hypersonic similarity parameter M delta approaching one.

On the instability of hypersonic flow past a wedge

NASA Technical Reports Server (NTRS)

Cowley, Stephen; Hall, Philip

1988-01-01

The instability of a compressible flow past a wedge is investigated in the hypersonic limit. Particular attention is given to the Tollmien-Schlichting waves governed by triple-deck theory though some discussion of inviscid modes is given. It is shown that the attached shock has a significant effect on the growth rates of Tollmien-Schlichting waves. Moreover, the presence of the shock allows for more than one unstable Tollmien-Schlichting wave. Indeed, an infinite discrete spectrum of unstable waves is induced by the shock, but these modes are unstable over relatively small but high frequency ranges. The shock is shown to have little effect on the inviscid modes considered by previous authors and an asymptotic description of inviscid modes in the hypersonic limit is given.

Results of aerodynamic heat transfer tests of a 0.0175-scale model of the Rockwell International Space Shuttle Orbiter 139 (model number 22-0) in the NASA/Ames 3.5-foot hypersonic wind tunnel (test OH6)

NASA Technical Reports Server (NTRS)

Dye, W. H.; Lockman, W. K.

1975-01-01

The results of a hypersonic wind tunnel test program conducted using a 0.0175 scale thin-skin thermocouple model of the Space Shuttle Orbiter to obtain aerodynamic heat transfer data on the Orbiter under simulated reentry conditions were presented. The test program was conducted at a Mach number of 7.3 and a freestream Reynolds number ranging between 1.0 and 6.0 million/foot. The model was tested for angles of attack ranging between 10 deg and 30 deg and a sideslip angle of 0 deg. The model was constructed of 15-5 PH stainless steel with the instrumented areas machined to a nominal skin thickness of 0.030 in. The model instrumentation consisted of 288 iron-constantan thermocouples spot welded to the skin inner surface, but only 75 of these were used in this test program. A high-speed, analog-to-digital data acquisition system was used to record data on magnetic tape.

Experimental Study Of SHEFEX II Hypersonic Aerodynamics And Canard Efficiency In H2K

NASA Astrophysics Data System (ADS)

Neeb, D.; Gulhan, A.

2011-05-01

One main objective of the DLR SHEFEX programme is to prove that sharp edged vehicles are capable of performing a re-entry into earth atmosphere by using a simple thermal protection system consisting of flat ceramic tiles. In comparison to blunt nose configurations like the Space shuttle, which are normally used for re-entry configurations, the SHEFEX TPS design is able to significantly reduce the costs and complexity of TPS structures and simultaneously increase the aerodynamic performance of the flight vehicle [1], [2]. To study its characteristics and perform several defined in-flight experiments during re-entry, the vehicle’s attitude will be controlled actively by canards [3]. In the framework of the SHEFEX II project an experimental investigation has been conducted in the hypersonic wind tunnel H2K to characterize the aerodynamic performance of the vehicle in hypersonic flow regime. The model has a modular design to enable the study of a variety of different influencing parameters. Its 4 circumferential canards have been made independently adjustable to account for the simulation of different manoeuvre conditions. To study the control behaviour of the vehicle and validate CFD data, a variation of canard deflections, angle of attack and angle of sideslip have been applied. Tests have been carried out at Mach 7 and 8.7 with a Reynolds number sensitivity study at the lower Mach number. The model was equipped with a six component internal balance to realize accurate coefficient measurements. The flow topology has been analyzed using Schlieren images. Beside general aerodynamic performance and canard efficiencies, flow phenomena like shock impingement on the canards could be determined by Schlieren images as well as by the derived coefficients.

Simulation of Stagnation Region Heating in Hypersonic Flow on Tetrahedral Grids

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.

2007-01-01

Hypersonic flow simulations using the node based, unstructured grid code FUN3D are presented. Applications include simple (cylinder) and complex (towed ballute) configurations. Emphasis throughout is on computation of stagnation region heating in hypersonic flow on tetrahedral grids. Hypersonic flow over a cylinder provides a simple test problem for exposing any flaws in a simulation algorithm with regard to its ability to compute accurate heating on such grids. Such flaws predominantly derive from the quality of the captured shock. The importance of pure tetrahedral formulations are discussed. Algorithm adjustments for the baseline Roe / Symmetric, Total-Variation-Diminishing (STVD) formulation to deal with simulation accuracy are presented. Formulations of surface normal gradients to compute heating and diffusion to the surface as needed for a radiative equilibrium wall boundary condition and finite catalytic wall boundary in the node-based unstructured environment are developed. A satisfactory resolution of the heating problem on tetrahedral grids is not realized here; however, a definition of a test problem, and discussion of observed algorithm behaviors to date are presented in order to promote further research on this important problem.

Experiments on Hypersonic Roughness Induced Transition by Means of Infrared Thermography

NASA Astrophysics Data System (ADS)

Schrijer, F. F. J.; Scarano, F.; van Oudheusden, B. W.; Bannink, W. J.

2005-02-01

Roughness induced boundary layer transition is experimentally investigated in the hypersonic flow regime at M = 9. The primary interest is the possible effect of stepwise geometry imperfections (2D isolated roughness) on (boundary layer) transition which may be caused on the EXPERT vehicle by the difference in thermal expansion due to the different materials used in the vehicle-nose construction. Also 3D isolated and 3D distributed roughness configurations were studied. Quantitative Infra-Red Thermography (QIRT) is used as primary diagnostic technique to measure the surface convective heat transfer and to detect boundary layer laminar-to-turbulent transition. The investigation shows that for a given height of the roughness element, the boundary layer is least sensitive to a step-like disturbance, whereas distributed 3D roughness was found to be effective in triggering transition. The experimental results have been compared to existing hypersonic transition correlations (PANT and Shuttle). Finally a transition criterion is evaluated which is based on the critical roughness height Reynolds number. Usage of this criterion enables a straightforward extrapolation to flight. Key words: hypersonic flow, boundary layer transition.

Prediction of forces and moments for hypersonic flight vehicle control effectors

NASA Technical Reports Server (NTRS)

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

1991-01-01

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

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

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

Gao, Xiaotian; Jiang, Binhao, E-mail: jiangbh@hit.edu.cn

2015-06-21

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

Experimental Hypersonic Aerodynamic Characteristics of the 2001 Mars Surveyor Precision Lander with Flap

NASA Technical Reports Server (NTRS)

Horvath, Thomas J.; OConnell, Tod F.; Cheatwood, F. McNeil; Prabhu, Ramadas K.; Alter, Stephen J.

2002-01-01

Aerodynamic wind-tunnel screening tests were conducted on a 0.029 scale model of a proposed Mars Surveyor 2001 Precision Lander (70 deg half angle spherically blunted cone with a conical afterbody). The primary experimental objective was to determine the effectiveness of a single flap to trim the vehicle at incidence during a lifting hypersonic planetary entry. The laminar force and moment data, presented in the form of coefficients, and shock patterns from schlieren photography were obtained in the NASA Langley Aerothermodynamic Laboratory for post-normal shock Reynolds numbers (based on forebody diameter) ranging from 2,637 to 92,350, angles of attack ranging from 0 tip to 23 degrees at 0 and 2 degree sideslip, and normal-shock density ratios of 5 and 12. Based upon the proposed entry trajectory of the 2001 Lander, the blunt body heavy gas tests in CF, simulate a Mach number of approximately 12 based upon a normal shock density ratio of 12 in flight at Mars. The results from this experimental study suggest that when traditional means of providing aerodynamic trim for this class of planetary entry vehicle are not possible (e.g. offset c.g.), a single flap can provide similar aerodynamic performance. An assessment of blunt body aerodynamic effects attributed to a real gas were obtained by synergistic testing in Mach 6 ideal-air at a comparable Reynolds number. From an aerodynamic perspective, an appropriately sized flap was found to provide sufficient trim capability at the desired L/D for precision landing. Inviscid hypersonic flow computations using an unstructured grid were made to provide a quick assessment of the Lander aerodynamics. Navier-Stokes computational predictions were found to be in very good agreement with experimental measurement.

Calculation of hypersonic shock structure using flux-split algorithms

NASA Technical Reports Server (NTRS)

Eppard, W. M.; Grossman, B.

1991-01-01

There exists an altitude regime in the atmosphere that is within the continuum domain, but wherein the conventional Navier-Stokes equations cease to be accurate. The altitude limits for this so called continuum transition regime depend on vehicle size and speed. Within this regime the thickness of the bow shock wave is no longer negligible when compared to the shock stand-off distance and the peak radiation intensity occurs within the shock wave structure itself. For this reason it is no longer valid to treat the shock wave as a discontinuous jump and it becomes necessary to compute through the shock wave itself. To accurately calculate hypersonic flowfields, the governing equations must be capable of yielding realistic profiles of flow variables throughout the structure of a hypersonic shock wave. The conventional form of the Navier-Stokes equations is restricted to flows with only small departures from translational equilibrium; it is for this reason they do not provide the capability to accurately predict hypersonic shock structure. Calculations in the continuum transition regime, therefore, require the use of governing equations other than Navier-Stokes. Several alternatives to Navier-Stokes are discussed; first for the case of a monatomic gas and then for the case of a diatomic gas where rotational energy must be included. Results are presented for normal shock calculations with argon and nitrogen.

Reduced-Order Aerothermoelastic Analysis of Hypersonic Vehicle Structures

NASA Astrophysics Data System (ADS)

Falkiewicz, Nathan J.

Design and simulation of hypersonic vehicles require consideration of a variety of disciplines due to the highly coupled nature of the flight regime. In order to capture all of the potential effects on vehicle dynamics, one must consider the aerodynamics, aerodynamic heating, heat transfer, and structural dynamics as well as the interactions between these disciplines. The problem is further complicated by the large computational expense involved in capturing all of these effects and their interactions in a full-order sense. While high-fidelity modeling techniques exist for each of these disciplines, the use of such techniques is computationally infeasible in a vehicle design and control system simulation setting for such a highly coupled problem. Early in the design stage, many iterations of analyses may need to be carried out as the vehicle design matures, thus requiring quick analysis turnaround time. Additionally, the number of states used in the analyses must be small enough to allow for efficient control simulation and design. As a result, alternatives to full-order models must be considered. This dissertation presents a fully coupled, reduced-order aerothermoelastic framework for the modeling and analysis of hypersonic vehicle structures. The reduced-order transient thermal solution is a modal solution based on the proper orthogonal decomposition. The reduced-order structural dynamic model is based on projection of the equations of motion onto a Ritz modal subspace that is identified a priori. The reduced-order models are assembled into a time-domain aerothermoelastic simulation framework which uses a partitioned time-marching scheme to account for the disparate time scales of the associated physics. The aerothermoelastic modeling framework is outlined and the formulations associated with the unsteady aerodynamics, aerodynamic heating, transient thermal, and structural dynamics are outlined. Results demonstrate the accuracy of the reduced-order transient

Mach 4 Performance of a Fixed-Geometry Hypersonic Inlet with Rectangular-to-Elliptical Shape Transition

NASA Technical Reports Server (NTRS)

Smart, Michael K.; Trexler, Carl A.

2003-01-01

Wind-tunnel testing of a hypersonic inlet with rectangular-to-elliptical shape transition has been conducted at Mach 4.0. These tests were performed to investigate the starting and back-pressure limits of this fixed-geometry inlet at conditions well below the Mach 5.7 design point. Results showed that the inlet required side spillage holes in order to self-start at Mach 4.0. Once started, the inlet generated a compression ratio of 12.6, captured almost 80% of available air and withstood a back-pressure ratio of 30.3 relative to tunnel static pressure. The spillage penalty for self-starting was estimated to be 4% of available air. These experimental results, along with previous experimental results at Mach 6.2 indicate that fixed-geometry inlets with rectangular-to-elliptical shape transition are a viable configuration for airframe- integrated scramjets that operate over a significant Mach number range.

Thermal-Mechanical Testing of Hypersonic Vehicle Structures

NASA Technical Reports Server (NTRS)

Hudson, Larry; Stephens, Craig

2007-01-01

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

Numerical analysis of exhaust jet secondary combustion in hypersonic flow field

NASA Astrophysics Data System (ADS)

Yang, Tian-Peng; Wang, Jiang-Feng; Zhao, Fa-Ming; Fan, Xiao-Feng; Wang, Yu-Han

2018-05-01

The interaction effect between jet and control surface in supersonic and hypersonic flow is one of the key problems for advanced flight control system. The flow properties of exhaust jet secondary combustion in a hypersonic compression ramp flow field were studied numerically by solving the Navier-Stokes equations with multi-species and combustion reaction effects. The analysis was focused on the flow field structure and the force amplification factor under different jet conditions. Numerical results show that a series of different secondary combustion makes the flow field structure change regularly, and the temperature increases rapidly near the jet exit.

Thermal Nonequilibrium in Hypersonic Separated Flow

DTIC Science & Technology

2014-12-22

flow duration and steadiness. 15. SUBJECT TERMS Hypersonic Flowfield Measurements, Laser Diagnostics of Gas Flow, Laser Induced...extent than the NS computation. While it would be convenient to believe that the more physically realistic flow modeling of the DSMC gas - surface...index and absorption coefficient. Each of the curves was produced assuming a 0.5 % concentration of lithium at the Condition A nozzle exit conditions

Geometry Modeling and Adaptive Control of Air-Breathing Hypersonic Vehicles

NASA Astrophysics Data System (ADS)

Vick, Tyler Joseph

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

Numerical Computations of Hypersonic Boundary-Layer over Surface Irregularities

NASA Technical Reports Server (NTRS)

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

2010-01-01

Surface irregularities such as protuberances inside a hypersonic boundary layer may lead to premature transition on the vehicle surface. Early transition in turn causes large localized surface heating that could damage the thermal protection system. Experimental measurements as well as numerical computations aimed at building a knowledge base for transition Reynolds numbers with respect to different protuberance sizes and locations have been actively pursued in recent years. This paper computationally investigates the unsteady wake development behind large isolated cylindrical roughness elements and the scaled wind-tunnel model of the trip used in a recent flight measurement during the reentry of space shuttle Discovery. An unstructured mesh, compressible flow solver based on the space-time conservation element, solution element (CESE) method is used to perform time-accurate Navier-Stokes calculations for the flow past a roughness element under several wind-tunnel conditions. For a cylindrical roughness element with a height to the boundary-layer thickness ratio from 0.8 to 2.5, the wake flow is characterized by a mushroom-shaped centerline streak and horse-shoe vortices. While time-accurate solutions converged to a steady-state for a ratio of 0.8, strong flow unsteadiness is present for a ratio of 1.3 and 2.5. Instability waves marked by distinct disturbance frequencies were found in the latter two cases. Both the centerline streak and the horse-shoe vortices become unstable downstream. The oscillatory vortices eventually reach an early breakdown stage for the largest roughness element. Spectral analyses in conjunction with the computed root mean square variations suggest that the source of the unsteadiness and instability waves in the wake region may be traced back to possible absolute instability in the front-side separation region.

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.

Unstructured Mesh Methods for the Simulation of Hypersonic Flows

NASA Technical Reports Server (NTRS)

Peraire, Jaime; Bibb, K. L. (Technical Monitor)

2001-01-01

This report describes the research work undertaken at the Massachusetts Institute of Technology. The aim of this research is to identify effective algorithms and methodologies for the efficient and routine solution of hypersonic viscous flows about re-entry vehicles. For over ten years we have received support from NASA to develop unstructured mesh methods for Computational Fluid Dynamics. As a result of this effort a methodology based on the use, of unstructured adapted meshes of tetrahedra and finite volume flow solvers has been developed. A number of gridding algorithms flow solvers, and adaptive strategies have been proposed. The most successful algorithms developed from the basis of the unstructured mesh system FELISA. The FELISA system has been extensively for the analysis of transonic and hypersonic flows about complete vehicle configurations. The system is highly automatic and allows for the routine aerodynamic analysis of complex configurations starting from CAD data. The code has been parallelized and utilizes efficient solution algorithms. For hypersonic flows, a version of the, code which incorporates real gas effects, has been produced. One of the latest developments before the start of this grant was to extend the system to include viscous effects. This required the development of viscous generators, capable of generating the anisotropic grids required to represent boundary layers, and viscous flow solvers. In figures I and 2, we show some sample hypersonic viscous computations using the developed viscous generators and solvers. Although these initial results were encouraging, it became apparent that in order to develop a fully functional capability for viscous flows, several advances in gridding, solution accuracy, robustness and efficiency were required. As part of this research we have developed: 1) automatic meshing techniques and the corresponding computer codes have been delivered to NASA and implemented into the GridEx system, 2) a finite

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

NASA Technical Reports Server (NTRS)

Berry, Scott A.; Nowak, Robert J.

2003-01-01

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

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.

Vorticity interaction effects on blunt bodies. [hypersonic viscous shock layers

NASA Technical Reports Server (NTRS)

Anderson, E. C.; Wilcox, D. C.

1977-01-01

Numerical solutions of the viscous shock layer equations governing laminar and turbulent flows of a perfect gas and radiating and nonradiating mixtures of perfect gases in chemical equilibrium are presented for hypersonic flow over spherically blunted cones and hyperboloids. Turbulent properties are described in terms of the classical mixing length. Results are compared with boundary layer and inviscid flowfield solutions; agreement with inviscid flowfield data is satisfactory. Agreement with boundary layer solutions is good except in regions of strong vorticity interaction; in these flow regions, the viscous shock layer solutions appear to be more satisfactory than the boundary layer solutions. Boundary conditions suitable for hypersonic viscous shock layers are devised for an advanced turbulence theory.

Assessment of predictive capabilities for aerodynamic heating in hypersonic flow

NASA Astrophysics Data System (ADS)

Knight, Doyle; Chazot, Olivier; Austin, Joanna; Badr, Mohammad Ali; Candler, Graham; Celik, Bayram; Rosa, Donato de; Donelli, Raffaele; Komives, Jeffrey; Lani, Andrea; Levin, Deborah; Nompelis, Ioannis; Panesi, Marco; Pezzella, Giuseppe; Reimann, Bodo; Tumuklu, Ozgur; Yuceil, Kemal

2017-04-01

The capability for CFD prediction of hypersonic shock wave laminar boundary layer interaction was assessed for a double wedge model at Mach 7.1 in air and nitrogen at 2.1 MJ/kg and 8 MJ/kg. Simulations were performed by seven research organizations encompassing both Navier-Stokes and Direct Simulation Monte Carlo (DSMC) methods as part of the NATO STO AVT Task Group 205 activity. Comparison of the CFD simulations with experimental heat transfer and schlieren visualization suggest the need for accurate modeling of the tunnel startup process in short-duration hypersonic test facilities, and the importance of fully 3-D simulations of nominally 2-D (i.e., non-axisymmmetric) experimental geometries.

Hypersonic airframe structures: Technology needs and flight test requirements

NASA Technical Reports Server (NTRS)

Stone, J. E.; Koch, L. C.

1979-01-01

Hypersonic vehicles, that may be produced by the year 2000, were identified. Candidate thermal/structural concepts that merit consideration for these vehicles were described. The current status of analytical methods, materials, manufacturing techniques, and conceptual developments pertaining to these concepts were reviewed. Guidelines establishing meaningful technology goals were defined and twenty-eight specific technology needs were identified. The extent to which these technology needs can be satisfied, using existing capabilities and facilities without the benefit of a hypersonic research aircraft, was assessed. The role that a research aircraft can fill in advancing this technology was discussed and a flight test program was outlined. Research aircraft thermal/structural design philosophy was also discussed. Programs, integrating technology advancements with the projected vehicle needs, were presented. Program options were provided to reflect various scheduling and cost possibilities.

Finite Element Analysis of a NASA National Transonic Facility Wind Tunnel Balance

NASA Technical Reports Server (NTRS)

Lindell, Michael C.

1996-01-01

This paper presents the results of finite element analyses and correlation studies performed on a NASA National Transonic Facility (NTF) Wind Tunnel balance. In the past NASA has relied primarily on classical hand analyses, coupled with relatively large safety factors, for predicting maximum stresses in wind tunnel balances. Now, with the significant advancements in computer technology and sophistication of general purpose analysis codes, it is more reasonable to pursue finite element analyses of these balances. The correlation studies of the present analyses show very good agreement between the analyses and data measured with strain gages and therefore the studies give higher confidence for using finite element analyses to analyze and optimize balance designs in the future.

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.

Wind and Water Power Fact Sheets | Wind | NREL

Science.gov Websites

Advanced Control Turbine Systems to Increase Performance, Decrease Structural Loading of Wind Turbines and and Water Power Fact Sheets Wind and Water Power Fact Sheets The capabilities for research at the National Wind Technology Center (NWTC) are numerous. Below you will find fact sheets about the many

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.

Growing a Wind Workforce: The National Wind Energy Skills Assessment Report (Poster)

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

Tegen, S.

This poster summarizes results from the first published investigation into the detailed makeup of the wind energy workforce as well as a glance at the educational infrastructure and training needs of the wind industry. Insights from this research into the domestic wind workforce allow the private sector, educational institutions, and federal and state governments to make better informed workforce-related decisions based on the current data and future projections.

Modeling of Electron Transpiration Cooling for Leading Edges of Hypersonic Vehicles

NASA Astrophysics Data System (ADS)

Hanquist, Kyle Matthew

The development of aeronautics has been largely driven by the passion to fly faster. From the flight of the Wright Flyer that flew 48 km/hr to the recent advances in hypersonic flight, most notably NASA's X-43A that flew at over 3 km/s, the velocity of flight has steadily increased. However, as these hypersonic speeds are reached and increased, contradicting aerothermodynamic design requirements present themselves. For example, a hypersonic cruise vehicle requires sharp leading edges to decrease the drag in order to maximize the range. However, the aerodynamic performance gains obtained by having a sharp leading edge come at the cost of very high, localized heating rates. There is currently no ideal way to manage these heating loads for sustained hypersonic flight, especially as flight velocities continue to increase. An approach that has been recently proposed involves using thermo-electric materials on these sharp leading edges to manage the heating loads. When exposed to high convective heating rates, these materials emit a current of electrons that leads to a cooling effect of the surface of the vehicle called electron transpiration cooling (ETC). This dissertation focuses on developing a modeling approach to investigate this phenomenon. The research includes developing and implementing an approach for ETC into a computational fluid dynamics code for simulation of hypersonic flow that accounts for electron emission from the surface. Models for space-charge-limited emission are also developed and implemented in order to accurately determine the level of emission from the surface. This work involves developing analytic models and assessing them using a direct-kinetic plasma sheath solver. Electric field effects are also implemented in the modeling approach, which accounts for forced diffusion and Joule heating. Finally, the modeling approach is coupled to a material response code in order to model the heat transfer into the material surface. Using this modeling

Hot-wire anemometry in hypersonic helium flow

NASA Technical Reports Server (NTRS)

Wagner, R. D.; Weinstein, L. M.

1974-01-01

Hot-wire anemometry techniques are described that have been developed and used for hypersonic-helium-flow studies. The short run time available dictated certain innovations in applying conventional hot-wire techniques. Some examples are given to show the application of the techniques used. Modifications to conventional equipment are described, including probe modifications and probe heating controls.

Multi-Dimensional, Inviscid Flux Reconstruction for Simulation of Hypersonic Heating on Tetrahedral Grids

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.

2009-01-01

The quality of simulated hypersonic stagnation region heating on tetrahedral meshes is investigated by using a three-dimensional, upwind reconstruction algorithm for the inviscid flux vector. Two test problems are investigated: hypersonic flow over a three-dimensional cylinder with special attention to the uniformity of the solution in the spanwise direction and hypersonic flow over a three-dimensional sphere. The tetrahedral cells used in the simulation are derived from a structured grid where cell faces are bisected across the diagonal resulting in a consistent pattern of diagonals running in a biased direction across the otherwise symmetric domain. This grid is known to accentuate problems in both shock capturing and stagnation region heating encountered with conventional, quasi-one-dimensional inviscid flux reconstruction algorithms. Therefore the test problem provides a sensitive test for algorithmic effects on heating. This investigation is believed to be unique in its focus on three-dimensional, rotated upwind schemes for the simulation of hypersonic heating on tetrahedral grids. This study attempts to fill the void left by the inability of conventional (quasi-one-dimensional) approaches to accurately simulate heating in a tetrahedral grid system. Results show significant improvement in spanwise uniformity of heating with some penalty of ringing at the captured shock. Issues with accuracy near the peak shear location are identified and require further study.

Prediction of forces and moments for hypersonic flight vehicle control effectors

NASA Technical Reports Server (NTRS)

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

1993-01-01

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

A fine-wire thermocouple probe for measurement of stagnation temperatures in real gas hypersonic flows of nitrogen

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Griffith, Wayland C.; Yanta, William J.

1991-01-01

A fine-wire thermocouple probe was used to determine freestream stagnation temperatures in hypersonic flows. Data were gathered in a N2 blowdown wind tunnel with runtimes of 1-5 s. Tests were made at supply pressures between 30 and 1400 atm and supply temperatures between 700 and 1900 K, with Mach numbers of 14 to 16. An iterative procedure requiring thermocouple data, pilot pressure measurements, and supply conditions was used to determine test cell stagnation temperatures. Probe conduction and radiation losses, as well as real gas behavior of N2, were accounted for during analysis. Temperature measurement error was found to be 5 to 10 percent. A correlation was drawn between thermocouple diameter Reynolds number and temperature recovery ratio. Transient probe behavior was studied and was found to be adequate in temperature gradients up to 1000 K/s.

A new computational method for reacting hypersonic flows

NASA Astrophysics Data System (ADS)

Niculescu, M. L.; Cojocaru, M. G.; Pricop, M. V.; Fadgyas, M. C.; Pepelea, D.; Stoican, M. G.

2017-07-01

Hypersonic gas dynamics computations are challenging due to the difficulties to have reliable and robust chemistry models that are usually added to Navier-Stokes equations. From the numerical point of view, it is very difficult to integrate together Navier-Stokes equations and chemistry model equations because these partial differential equations have different specific time scales. For these reasons, almost all known finite volume methods fail shortly to solve this second order partial differential system. Unfortunately, the heating of Earth reentry vehicles such as space shuttles and capsules is very close linked to endothermic chemical reactions. A better prediction of wall heat flux leads to smaller safety coefficient for thermal shield of space reentry vehicle; therefore, the size of thermal shield decreases and the payload increases. For these reasons, the present paper proposes a new computational method based on chemical equilibrium, which gives accurate prediction of hypersonic heating in order to support the Earth reentry capsule design.

Impacts | Wind | NREL

Science.gov Websites

in hard hats standing on top of a large wind turbine overlooking several other wind turbines in the Framework Transforms FAST Wind Turbine Modeling Tool NREL Assesses National Design Standards for Offshore Wind Resource NREL Identifies Investments for Wind Turbine Drivetrain Technologies Awards R&D 100

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.

A unique high heat flux facility for testing hypersonic engine components

NASA Technical Reports Server (NTRS)

Melis, Matthew E.; Gladden, Herbert J.

1990-01-01

This paper describes the Hot Gas Facility, a unique, reliable, and cost-effective high-heat-flux facility for testing hypersonic engine components developed at the NASA Lewis Research Center. The Hot Gas Facility is capable of providing heat fluxes ranging from 200 Btu/sq ft per sec on flat surfaces up to 8000 Btu/sq ft per sec at a leading edge stagnation point. The usefulness of the Hot Gas Facility for the NASP community was demonstrated by testing hydrogen-cooled structures over a range of temperatures and pressures. Ranges of the Reynolds numbers, Prandtl numbers, enthalpy, and heat fluxes similar to those expected during hypersonic flights were achieved.

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.

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.

Wind Tunnel and Propulsion Test Facilities: An Assessment of NASA's Capabilities to Serve National Needs

NASA Technical Reports Server (NTRS)

Anton, Philip S.; Gritton, Eugene C.; Mesic, Richard; Steinberg, Paul; Johnson, Dana J.

2004-01-01

This monograph reveals and discusses the National Aeronautics and Space Administration's (NASA's) wind tunnel and propulsion test facility management issues that are creating real risks to the United States' competitive aeronautics advantage.

Downstream Effects on Orbiter Leeside Flow Separation for Hypersonic Flows

NASA Technical Reports Server (NTRS)

Buck, Gregory M.; Pulsonetti, Maria V.; Weilmuenster, K. James

2005-01-01

Discrepancies between experiment and computation for shuttle leeside flow separation, which came to light in the Columbia accident investigation, are resolved. Tests were run in the Langley Research Center 20-Inch Hypersonic CF4 Tunnel with a baseline orbiter model and two extended trailing edge models. The extended trailing edges altered the wing leeside separation lines, moving the lines toward the fuselage, proving that wing trailing edge modeling does affect the orbiter leeside flow. Computations were then made with a wake grid. These calculations more closely matched baseline experiments. Thus, the present findings demonstrate that it is imperative to include the wake flow domain in CFD calculations in order to accurately predict leeside flow separation for hypersonic vehicles at high angles of attack.

Flow quality experiment in a tandem nozzle wind tunnel at Mach 3

NASA Astrophysics Data System (ADS)

Wu, Jie; Zamre, Pradip; Radespiel, Rolf

2015-01-01

In this study, the disturbance characterization and flow quality improvement of a newly designed Tandem Nozzle Mach 3 Wind Tunnel are presented. Firstly, a combined modal analysis is conducted to characterize the freestream disturbances with initial set-up of the settling chamber by using a Pitot probe and a hot-wire anemometry. Then, disturbance reduction in the supersonic wind tunnel is investigated by inserting various damping materials into the settling chamber, while a Pitot probe instrumented with Kulite sensor is employed to monitor the variation of the Pitot pressure fluctuation in the test section. Eventually, an optimized configuration of the settling chamber is determined by a combination of certain damping materials. Afterward, the freestream disturbances are re-characterized with the optimized set-up of the settling chamber, and the disturbance level is found to be significantly reduced. Through this study, valuable experience has been acquired for the disturbance reduction in tandem nozzle type supersonic wind tunnel for the first time, which enhances the feasibility of extending the operation range of conventional hypersonic Ludwieg tubes.

An Approach to Establishing System Benefits for Technology in NASA's Hypersonics Investment Area

NASA Technical Reports Server (NTRS)

Hueter, Uwe; Pannell, Bill; Cook, Stephen (Technical Monitor)

2001-01-01

NASA's has established long term goals for access-to-space. The third generation launch systems are to be fully reusable and operational around 2025. The goals for the third generation launch system are to significantly reduce cost and improve safety over current systems. The Advanced Space Transportation Program (ASTP) Office at the NASA's Marshall Space Flight Center in Huntsville, AL has the agency lead to develop space transportation technologies. Within ASTP, under the Hypersonics Investment Area, third generation technologies are being pursued. The Hypersonics Investment Area's primary objective is to mature vehicle technologies to enable substantial increases in the design and operating margins of third generation RLVs (current Space Shuttle is considered the first generation RLV) by incorporating advanced propulsion systems, materials, structures, thermal protection systems, power, and avionics technologies. The paper describes the system process, tools and concepts used to determine the technology benefits. Preliminary results will be presented along with the current technology investments that are being made by ASTP's Hypersonics Investment Area.

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-43A Project Overview: Adventures in Hypersonics

NASA Technical Reports Server (NTRS)

Davis, Mark; Grindle, Laurie

2007-01-01

A viewgraph presentation describing the hypersonics program at NASA Dryden Flight Research Center is shown. The topics include: 1) X-43A Program Overview; 2) Vehicle Description; 3) Flight 1, MIB & Return to Flight; 4) Flight 2 and Results; 5) Flight 3 and Results; and 6) Concluding Remarks

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.

AFRL Ludwieg Tube Initial Performance

DTIC Science & Technology

2017-11-01

Nov 2016 14. ABSTRACT The Air Force Research Laboratory has developed and constructed a Ludwieg tube wind tunnel for hypersonic experimental ...Ludwieg tube wind tunnel for hypersonic experimental research. This wind tunnel is now operational and its initial performance has been...opening and the first expansion wave reflection was 100 ms, as expected. About 80 ms of quasi -steady pressure was obtained after the valve-opening

Heat-transfer test results for a .0275-scale space shuttle external tank with a 10 deg/40 deg double cone-ogive nose in the NASA/AMES 3.5-foot hypersonic wind tunnel (FH14), volume 2

NASA Technical Reports Server (NTRS)

Carroll, H. R.

1977-01-01

A .0275 scale forebody model of the new baseline configuration of the space shuttle external tank vent cap configuration was tested to determine the flow field due to the double cone configuration. The tests were conducted in a 3.5 foot hypersonic wind tunnel at alpha = -5 deg, -4.59 deg, 0 deg, 5 deg, and 10 deg; beta = 0 deg, -3 deg, -5.51 deg, -6 deg, -9 deg, and +6 deg; nominal freestream Reynolds numbers per foot of 1.5 x 1 million, 3.0 x 1 million, and 5.0 x 1 million; and a nominal Mach number of 5. Separation and reattached flow from thermocouple data, shadowgraphs, and oil flows indicate that separation begins about 80% from the tip of the 10 deg cone, then reattaches on the vent cap and produces fully turbulent flow over most of the model forebody. The hardware disturbs the flow over a much larger area than present TPS application has assumed. A correction to the flow disturbance was experimentally suggested from the results of an additional test run.

Application of Pressure Sensitive Paint in Hypersonic Flows

NASA Technical Reports Server (NTRS)

Jules, Kenol; Carbonaro, Mario; Zemsch, Stephan

1995-01-01

It is well known in the aerodynamic field that pressure distribution measurement over the surface of an aircraft model is a problem in experimental aerodynamics. For one thing, a continuous pressure map can not be obtained with the current experimental methods since they are discrete. Therefore, interpolation or CFD methods must be used for a more complete picture of the phenomenon under study. For this study, a new technique was investigated which would provide a continuous pressure distribution over the surface under consideration. The new method is pressure sensitive paint. When pressure sensitive paint is applied to an aerodynamic surface and placed in an operating wind-tunnel under appropriate lighting, the molecules luminesce as a function of the local pressure of oxygen over the surface of interest during aerodynamic flow. The resulting image will be brightest in the areas of low pressure (low oxygen concentration), and less intense in the areas of high pressure (where oxygen is most abundant on the surface). The objective of this investigation was to use pressure sensitive paint samples from McDonnell Douglas (MDD) for calibration purpose in order to assess the response of the paint under appropriate lighting and to use the samples over a flat plate/conical fin mounted at 75 degrees from the center of the plate in order to study the shock/boundary layer interaction at Mach 6 in the Von Karman wind-tunnel. From the result obtained it was concluded that temperature significantly affects the response of the paint and should be given the uppermost attention in the case of hypersonic flows. Also, it was found that past a certain temperature threshold, the paint intensity degradation became irreversible. The comparison between the pressure tap measurement and the pressure sensitive paint showed the right trend. However, there exists a shift when it comes to the actual value. Therefore, further investigation is under way to find the cause of the shift.

Wind Powering America Podcasts, Wind Powering America (WPA)

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

Not Available

2012-04-01

Wind Powering America and the National Association of Farm Broadcasters produce a series of radio interviews featuring experts discussing wind energy topics. The interviews are aimed at a rural stakeholder audience and are available as podcasts. On the Wind Powering America website, you can access past interviews on topics such as: Keys to Local Wind Energy Development Success, What to Know about Installing a Wind Energy System on Your Farm, and Wind Energy Development Can Revitalize Rural America. This postcard is a marketing piece that stakeholders can provide to interested parties; it will guide them to this online resource formore » podcast episodes.« less

Hypersonic Global Strike Feasibility and Options

DTIC Science & Technology

2012-02-15

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

High heat flux actively cooled honeycomb sandwich structural panel for a hypersonic aircraft

NASA Technical Reports Server (NTRS)

Koch, L. C.; Pagel, L. L.

1978-01-01

The results of a program to design and fabricate an unshielded actively cooled structural panel for a hypersonic aircraft are presented. The design is an all-aluminum honeycomb sandwich with embedded cooling passages soldered to the inside of the outer moldline skin. The overall finding is that an actively cooled structure appears feasible for application on a hypersonic aircraft, but the fabrication process is complex and some material and manufacturing technology developments are required. Results from the program are summarized and supporting details are presented.

Results of investigations (OA77 and OA78) on an 0.015-scale 140A/B configuration space shuttle vehicle orbiter model 49-0 in the AEDC VKF B and C wind tunnels, revision A

NASA Technical Reports Server (NTRS)

Gillins, R. L.

1975-01-01

Aerodynamic data obtained from wind tunnel tests of an 0.015-scale 140A/B configuration SSV Orbiter model in the AEDC VKF B and C wind tunnels are presented. Tests were conducted at Mach numbers of 6 and 8 in the B tunnel and at a Mach number of 10 to in the C tunnel to verify hypersonic stability and control characteristics, determine control surface effectiveness, and investigate Reynolds number effects of the 140A/B configuration. Force data were obtained for various control surface settings and Reynolds numbers in the angle-of-attack range of 15 deg to 45 deg and at angles of sideslip of -5 deg to +10 deg. Data were obtained for a few configurations at angles of attack from -27 deg to 45 deg. Control surface variables included elevon, rudder, speedbrake and bodyflap deflections. The effects of an alternate wing leading edge shape were investigated to determine its hypersonic stability and control characteristics.

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

NASA Technical Reports Server (NTRS)

Morelli, Eugene A.

2008-01-01

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

An Analysis of Wintertime Winds in Washington, D.C.

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

Berg, Larry K.; Allwine, K Jerry

This report consists of a description of the wintertime climatology of wind speed and wind direction around the National Mall in Washington, D.C. Meteorological data for this study were collected at Ronald Reagan Washington National Airport (Reagan National), Dulles International Airport (Dulles), and a set of surface meteorological stations that are located on a number of building tops around the National Mall. A five-year wintertime climatology of wind speed and wind direction measured at Reagan National and Dulles are presented. A more detailed analysis was completed for the period December 2003 through February 2004 using data gathered from stations locatedmore » around the National Mall, Reagan National, and Dulles. Key findings of our study include the following: * There are systematic differences between the wind speed and wind direction observed at Reagan National and the wind speed and wind direction measured by building top weather stations located in the National Mall. Although Dulles is located much further from the National Mall than Reagan National, there is better agreement between the wind speed and wind direction measured at Dulles and the weather stations in the National Mall. * When the winds are light (less than 3 ms-1 or 7 mph), there are significant differences in the wind directions reported at the various weather stations within the Mall. * Although the mean characteristics of the wind are similar at the various locations, significant, short-term differences are found when the time series are compared. These differences have important implications for the dispersion of airborne contaminants. In support of wintertime special events in the area of the National Mall, we recommend placing four additional meteorological instruments: three additional surface stations, one on the east bank of the Potomac River, one south of the Reflecting Pool (to better define the flow within the Mall), and a surface station near the Herbert C. Hoover Building

Hypersonic research engine/aerothermodynamic integration model: Experimental results. Volume 3: Mach 7 component integration and performance

NASA Technical Reports Server (NTRS)

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

1976-01-01

The NASA Hypersonic Research Engine Project was undertaken to design, develop, and construct a hypersonic research ramjet engine for high performance and to flight test the developed concept on the X-15-2A airplane over the speed range from Mach 3 to 8. Computer program results are presented here for the Mach 7 component integration and performance tests.

Evolution of supersonic corner vortex in a hypersonic inlet/isolator model

NASA Astrophysics Data System (ADS)

Huang, He-Xia; Tan, Hui-Jun; Sun, Shu; Ling, Yu

2016-12-01

There are complex corner vortex flows in a rectangular hypersonic inlet/isolator. The corner vortex propagates downstream and interacts with the shocks and expansion waves in the isolator repeatedly. The supersonic corner vortex in a generic hypersonic inlet/isolator model is theoretically and numerically analyzed at a freestream Mach number of 4.92. The cross-flow topology of the corner vortex flow is found to obey Zhang's theory ["Analytical analysis of subsonic and supersonic vortex formation," Acta Aerodyn. Sin. 13, 259-264 (1995)] strictly, except for the short process with the vortex core situated in a subsonic flow which is surrounded by a supersonic flow. In general, the evolution history of the corner vortex under the influence of the background waves in the hypersonic inlet/isolator model can be classified into two types, namely, from the adverse pressure gradient region to the favorable pressure gradient region and the reversed one. For type 1, the corner vortex is a one-celled vortex with the cross-sectional streamlines spiraling inwards at first. Then the Hopf bifurcation occurs and the streamlines in the outer part of the limit cycle switch to spiraling outwards, yielding a two-celled vortex. The limit cycle shrinks gradually and finally vanishes with the streamlines of the entire corner vortex spiraling outwards. For type 2, the cross-sectional streamlines of the corner vortex spiral outwards first. Then a stable limit cycle is formed, yielding a two-celled vortex. The short-lived limit cycle forces the streamlines in the corner vortex to change the spiraling trends rapidly. Although it is found in this paper that there are some defects on the theoretical proof of the limit cycle, Zhang's theory is proven useful for the prediction and qualitative analysis of the complex corner vortex in a hypersonic inlet/isolator. In addition, three conservation laws inside the limit cycle are obtained.

Shuttle Damage/Repair from the Perspective of Hypersonic Boundary Layer Transition - Experimental Results

NASA Technical Reports Server (NTRS)

Horvath, Thomas J.; Berry, Scott A.; Merski, N. Ronald; Berger, Karen T.; Buck, Gregory M.; Liechty, Derek S.; Schneider, Steven P.

2006-01-01

An overview is provided of the experimental wind tunnel program conducted at the NASA Langley Research Center Aerothermodynamics Laboratory in support of an agency-wide effort to prepare the Shuttle Orbiter for Return-to-Flight. The effect of an isolated protuberance and an isolated rectangular cavity on hypersonic boundary layer transition onset on the windward surface of the Shuttle Orbiter has been experimentally characterized. These experimental studies were initiated to provide a protuberance and cavity effects database for developing hypersonic transition criteria to support on-orbit disposition of thermal protection system damage or repair. In addition, a synergistic experimental investigation was undertaken to assess the impact of an isolated mass-flow entrainment source (simulating pyrolysis/outgassing from a proposed tile repair material) on boundary layer transition. A brief review of the relevant literature regarding hypersonic boundary layer transition induced from cavities and localized mass addition from ablation is presented. Boundary layer transition results were obtained using 0.0075-scale Orbiter models with simulated tile damage (rectangular cavities) of varying length, width, and depth and simulated tile damage or repair (protuberances) of varying height. Cavity and mass addition effects were assessed at a fixed location (x/L = 0.3) along the model centerline in a region of near zero pressure gradient. Cavity length-to-depth ratio was systematically varied from 2.5 to 17.7 and length-to-width ratio of 1 to 8.5. Cavity depth-to-local boundary layer thickness ranged from 0.5 to 4.8. Protuberances were located at several sites along the centerline and port/starboard attachment lines along the chine and wing leading edge. Protuberance height-to-boundary layer thickness was varied from approximately 0.2 to 1.1. Global heat transfer images and heating distributions of the Orbiter windward surface using phosphor thermography were used to infer the

Publications | Wind | NREL

Science.gov Websites

-specific analysis can be used to assess the risk induced by loss of a wind turbine blade. The study used for different wind turbine configurations. The authors used assumptions specific to the National Wind ., failure rate for wind turbine rotors) are based on a 13-year-old report on wind turbines installed in

Preliminary results of the large experimental wind turbine phase of the national wind energy program

NASA Technical Reports Server (NTRS)

Thomas, R. L.; Sholes, J. E.

1975-01-01

A major phase of the wind energy program is the development of reliable wind turbines for supplying cost-competitive electrical energy. This paper discusses the preliminary results of two projects in this phase of the program. First an experimental 100 kW wind turbine design and its status are reviewed. Also discussed are the results of two parallel design studies for determining the configurations and power levels for wind turbines with minimum energy costs. These studies show wind energy costs of 7 to 1.5 c/kWH for wind turbines produced in quantities of 100 to 1000 a year and located at sites having average winds of 12 to 18 mph.

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

NASA Technical Reports Server (NTRS)

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

1976-01-01

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

Effect of surface roughness on the heating rates of large-angled hypersonic blunt cones

NASA Astrophysics Data System (ADS)

Irimpan, Kiran Joy; Menezes, Viren

2018-03-01

Surface-roughness caused by the residue of an ablative Thermal Protection System (TPS) can alter the turbulence level and surface heating rates on a hypersonic re-entry capsule. Large-scale surface-roughness that could represent an ablated TPS, was introduced over the forebody of a 120° apex angle blunt cone, in order to test for its influence on surface heating rates in a hypersonic freestream of Mach 8.8. The surface heat transfer rates measured on smooth and roughened models under the same freestream conditions were compared. The hypersonic flow-fields of the smooth and rough-surfaced models were visualized to analyse the flow physics. Qualitative numerical simulations and pressure measurements were carried out to have an insight into the high-speed flow physics. Experimental observations under moderate Reynolds numbers indicated a delayed transition and an overall reduction of 17-46% in surface heating rates on the roughened model.

Boundary-layer instability & transition on a flared cone in a Mach 6 quiet wind tunnel

NASA Astrophysics Data System (ADS)

Hofferth, Jerrod; Saric, William

2011-11-01

Measurements of boundary-layer transition location and instability growth on a sharp-tipped 5°-half-angle flared cone were conducted in a low-disturbance Mach 6 wind tunnel at a freestream unit Reynolds number of 10 × 106/m. Under quiet flow at these conditions, the boundary layer becomes transitional near the base of the cone, where significant second-mode instability growth is evident. Transition location is determined using an array of embedded thermocouples, and instability development is observed in mean and fluctuating mass flux data using hotwire anemometry. The present work seeks to reproduce and build upon previous experiments which used the same test article and similar diagnostics in the facility's former installation at NASA Langley. Together with comprehensive measurements of the freestream disturbance environment, these baseline cone data characterize the facility's performance relative to that in its previous installation. In addition, the current campaign establishes experimental readiness for future research, which will study the effects of periodic surface roughness and controlled-input disturbances. AFOSR/NASA National Center for Hypersonic Research in Laminar-Turbulent Transition; Grant FA9550-09-1-0341.

Design of a pulse-type strain gauge balance for a long-test-duration hypersonic shock tunnel

NASA Astrophysics Data System (ADS)

Wang, Y.; Liu, Y.; Jiang, Z.

2016-11-01

When the measurement of aerodynamic forces is conducted in a hypersonic shock tunnel, the inertial forces lead to low-frequency vibrations of the model, and its motion cannot be addressed through digital filtering because a sufficient number of cycles cannot be obtained during a tunnel run. This finding implies restrictions on the model size and mass as the natural frequencies are inversely proportional to the length scale of the model. Therefore, the force measurement still has many problems, particularly for large and heavy models. Different structures of a strain gauge balance (SGB) are proposed and designed, and the measurement element is further optimized to overcome the difficulties encountered during the measurement of aerodynamic forces in a shock tunnel. The motivation for this study is to assess the structural performance of the SGB used in a long-test-duration JF12 hypersonic shock tunnel, which has more than 100 ms of test time. Force tests were conducted for a large-scale cone with a 10^° semivertex angle and a length of 0.75 m in the JF12 long-test-duration shock tunnel. The finite element method was used for the analysis of the vibrational characteristics of the Model-Balance-Sting System (MBSS) to ensure a sufficient number of cycles, particularly for the axial force signal during a shock tunnel run. The higher-stiffness SGB used in the test shows good performance, wherein the frequency of the MBSS increases because of the stiff construction of the balance. The experimental results are compared with the data obtained in another wind tunnel and exhibit good agreement at M = 7 and α =5°.

Heat-Pipe-Cooled Leading Edges for Hypersonic Vehicles

NASA Technical Reports Server (NTRS)

Glass, David E.

2006-01-01

Heat pipes can be used to effectively cool wing leading edges of hypersonic vehicles. . Heat-pipe leading edge development. Design validation heat pipe testing confirmed design. Three heat pipes embedded and tested in C/C. Single J-tube heat pipe fabricated and testing initiated. HPCLE work is currently underway at several locations.

Qualitative Assessment of the Acoustic Disturbance Environment in the NASA LaRC 20-Inch MACH 6 Wind Tunnel

NASA Technical Reports Server (NTRS)

Horvath, Thomas J.; Berry, Scott A.; Hamilton, H. Harris

2001-01-01

An experimental investigation was conducted on a 5-degree-half-angle cone with a flare in a conventional Mach 6 wind tunnel to examine the effect of facility noise on boundary layer transition. The effect of tunnel noise was inferred by comparing transition onset locations determined from the present test to that previously obtained in a Mach 6 quiet tunnel. Together, the two sets of experiments are believed to represent the first direct comparison of transition onset between a conventional and a quiet hypersonic wind tunnel using a common test model. In the present conventional hypersonic tunnel experiment, adiabatic wall temperatures were measured and heat transfer distributions were inferred on the cone flare model at zero degree angle of attack over a range of length Reynolds numbers (2 x 10(exp 6) to 10 x 10(exp 6)) which resulted in laminar and turbulent flow. Wall-to-total temperature ratio for the transient heating measurements and the adiabatic wall temperature measurements were 0.69 and 0.86, respectively. The cone flare nosetip radius was varied from 0.0001 to 0.125-inch to examine the effects of bluntness on transition onset. At comparable freestream conditions the transition onset Reynolds number obtained on the cone flare model in the conventional "noisy" tunnel was approximately 25% lower than that measured in the low disturbance tunnel.

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.

Preliminary results of the large experimental wind turbine phase of the national wind energy program

NASA Technical Reports Server (NTRS)

Thomas, R. L.; Sholes, T.; Sholes, J. E.

1975-01-01

The preliminary results of two projects in the development phase of reliable wind turbines designed to supply cost-competitive electrical energy were discussed. An experimental 100 kW wind turbine design and its status are first reviewed. The results of two parallel design studies for determining the configurations and power levels for wind turbines with minimum energy costs are also discussed. These studies predict wind energy costs of 1.5 to 7 cents per kW-h for wind turbines produced in quantities of 100 to 1000 per year and located at sites having average winds of 12 to 18 mph.

Control of supersonic wind-tunnel noise by laminarization of nozzle-wall boundary layer

NASA Technical Reports Server (NTRS)

Beckwith, I. E.; Harvey, W. D.; Harris, J. E.; Holley, B. B.

1973-01-01

One of the principal design requirements for a quiet supersonic or hypersonic wind tunnel is to maintain laminar boundary layers on the nozzle walls and thereby reduce disturbance levels in the test flow. The conditions and apparent reasons for laminar boundary layers which have been observed during previous investigations on the walls of several nozzles for exit Mach numbers from 2 to 20 are reviewed. Based on these results, an analysis and an assessment of nozzle design requirements for laminar boundary layers including low Reynolds numbers, high acceleration, suction slots, wall temperature control, wall roughness, and area suction are presented.

Force Measurements in Short-Duration Hypersonic Facilities

DTIC Science & Technology

1975-11-01

method which depends upon a double differentiation of experimental data, albeit smoothed. When the motion is non -plarnar, the tricyclic method of...J.0. & Hamaker, F.M. An experimental investigation of the base pressure characteristics of (1955) non -lifting bodies of revolution at Mach numbers...AD-AO18 842 FORCE MEASUREMENTS IN SHORT-DURATION HYPERSONIC FACILITIES Leonard Bernstein, et al Advisory Group for Aerospace Research and Development

Airbreathing hypersonic vehicle design and analysis methods

NASA Technical Reports Server (NTRS)

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

1996-01-01

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

CFD Validation Studies for Hypersonic Flow Prediction

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.

2001-01-01

A series of experiments to measure pressure and heating for code validation involving hypersonic, laminar, separated flows was conducted at the Calspan-University at Buffalo Research Center (CUBRC) in the Large Energy National Shock (LENS) tunnel. The experimental data serves as a focus for a code validation session but are not available to the authors until the conclusion of this session. The first set of experiments considered here involve Mach 9.5 and Mach 11.3 N2 flow over a hollow cylinder-flare with 30 degree flare angle at several Reynolds numbers sustaining laminar, separated flow. Truncated and extended flare configurations are considered. The second set of experiments, at similar conditions, involves flow over a sharp, double cone with fore-cone angle of 25 degrees and aft-cone angle of 55 degrees. Both sets of experiments involve 30 degree compressions. Location of the separation point in the numerical simulation is extremely sensitive to the level of grid refinement in the numerical predictions. The numerical simulations also show a significant influence of Reynolds number on extent of separation. Flow unsteadiness was easily introduced into the double cone simulations using aggressive relaxation parameters that normally promote convergence.

CFD Validation Studies for Hypersonic Flow Prediction

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.

2001-01-01

A series of experiments to measure pressure and heating for code validation involving hypersonic, laminar, separated flows was conducted at the Calspan-University at Buffalo Research Center (CUBRC) in the Large Energy National Shock (LENS) tunnel. The experimental data serves as a focus for a code validation session but are not available to the authors until the conclusion of this session. The first set of experiments considered here involve Mach 9.5 and Mach 11.3 N, flow over a hollow cylinder-flare with 30 deg flare angle at several Reynolds numbers sustaining laminar, separated flow. Truncated and extended flare configurations are considered. The second set of experiments, at similar conditions, involves flow over a sharp, double cone with fore-cone angle of 25 deg and aft-cone angle of 55 deg. Both sets of experiments involve 30 deg compressions. Location of the separation point in the numerical simulation is extremely sensitive to the level of grid refinement in the numerical predictions. The numerical simulations also show a significant influence of Reynolds number on extent of separation. Flow unsteadiness was easily introduced into the double cone simulations using aggressive relaxation parameters that normally promote convergence.

Compressibility Considerations for kappa-omega Turbulence Models in Hypersonic Boundary Layer Applications

NASA Technical Reports Server (NTRS)

Rumsey, C. L.

2009-01-01

The ability of kappa-omega models to predict compressible turbulent skin friction in hypersonic boundary layers is investigated. Although uncorrected two-equation models can agree well with correlations for hot-wall cases, they tend to perform progressively worse - particularly for cold walls - as the Mach number is increased in the hypersonic regime. Simple algebraic models such as Baldwin-Lomax perform better compared to experiments and correlations in these circumstances. Many of the compressibility corrections described in the literature are summarized here. These include corrections that have only a small influence for kappa-omega models, or that apply only in specific circumstances. The most widely-used general corrections were designed for use with jet or mixing-layer free shear flows. A less well-known dilatation-dissipation correction intended for boundary layer flows is also tested, and is shown to agree reasonably well with the Baldwin-Lomax model at cold-wall conditions. It exhibits a less dramatic influence than the free shear type of correction. There is clearly a need for improved understanding and better overall physical modeling for turbulence models applied to hypersonic boundary layer flows.

Solution-Space Screening of a Hypersonic Endurance Demonstrator

NASA Technical Reports Server (NTRS)

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

2012-01-01

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

Hypersonic Boundary Layer Stability over a Flared Cone in a Quiet Tunnel

NASA Technical Reports Server (NTRS)

Lachowicz, Jason T.; Chokani, Ndaona; Wilkinson, Stephen P.

1996-01-01

Hypersonic boundary layer measurements were conducted over a flared cone in a quiet wind tunnel. The flared cone was tested at a freestream unit Reynolds number of 2.82x106/ft in a Mach 6 flow. This Reynolds number provided laminar-to-transitional flow over the model in a low-disturbance environment. Point measurements with a single hot wire using a novel constant voltage anemometry system were used to measure the boundary layer disturbances. Surface temperature and schlieren measurements were also conducted to characterize the laminar-to-transitional state of the boundary layer and to identify instability modes. Results suggest that the second mode disturbances were the most unstable and scaled with the boundary layer thickness. The integrated growth rates of the second mode compared well with linear stability theory in the linear stability regime. The second mode is responsible for transition onset despite the existence of a second mode sub-harmonic. The sub-harmonic wavelength also scales with the boundary layer thickness. Furthermore, the existence of higher harmonics of the fundamental suggests that non-linear disturbances are not associated with high free stream disturbance levels.

Hypersonic Magneto-Fluid-Dynamic Compression in Cylindrical Inlet

NASA Technical Reports Server (NTRS)

Shang, Joseph S.; Chang, Chau-Lyan

2007-01-01

Hypersonic magneto-fluid-dynamic interaction has been successfully performed as a virtual leading-edge strake and a virtual cowl of a cylindrical inlet. In a side-by-side experimental and computational study, the magnitude of the induced compression was found to be depended on configuration and electrode placement. To better understand the interacting phenomenon the present investigation is focused on a direct current discharge at the leading edge of a cylindrical inlet for which validating experimental data is available. The present computational result is obtained by solving the magneto-fluid-dynamics equations at the low magnetic Reynolds number limit and using a nonequilibrium weakly ionized gas model based on the drift-diffusion theory. The numerical simulation provides a detailed description of the intriguing physics. After validation with experimental measurements, the computed results further quantify the effectiveness of a magnet-fluid-dynamic compression for a hypersonic cylindrical inlet. At a minuscule power input to a direct current surface discharge of 8.14 watts per square centimeter of electrode area produces an additional compression of 6.7 percent for a constant cross-section cylindrical inlet.

Numerical analysis of hypersonic turbulent film cooling flows

NASA Technical Reports Server (NTRS)

Chen, Y. S.; Chen, C. P.; Wei, H.

1992-01-01

As a building block, numerical capabilities for predicting heat flux and turbulent flowfields of hypersonic vehicles require extensive model validations. Computational procedures for calculating turbulent flows and heat fluxes for supersonic film cooling with parallel slot injections are described in this study. Two injectant mass flow rates with matched and unmatched pressure conditions using the database of Holden et al. (1990) are considered. To avoid uncertainties associated with the boundary conditions in testing turbulence models, detailed three-dimensional flowfields of the injection nozzle were calculated. Two computational fluid dynamics codes, GASP and FDNS, with the algebraic Baldwin-Lomax and k-epsilon models with compressibility corrections were used. It was found that the B-L model which resolves near-wall viscous sublayer is very sensitive to the inlet boundary conditions at the nozzle exit face. The k-epsilon models with improved wall functions are less sensitive to the inlet boundary conditions. The testings show that compressibility corrections are necessary for the k-epsilon model to realistically predict the heat fluxes of the hypersonic film cooling problems.

Airframe Research and Technology for Hypersonic Airbreathing Vehicles

NASA Technical Reports Server (NTRS)

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

2002-01-01

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

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

PubMed

Liu, Yanbin; Liu, Mengying; Sun, Peihua

2014-01-01

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

Preliminary Studies on Aerodynamic Control with Direct Current Discharge at Hypersonic Speed

NASA Astrophysics Data System (ADS)

Watanabe, Yasumasa; Takama, Yoshiki; Imamura, Osamu; Watanuki, Tadaharu; Suzuki, Kojiro

A new idea of an aerodynamic control device for hypersonic vehicles using plasma discharges is presented. The effect of DC plasma discharge on a hypersonic flow is examined with both experiments and CFD analyses. It is revealed that the surface pressure upstream of plasma area significantly increases, which would be preferable in realizing a new aerodynamic control devices. Such pressure rise is also observed in the result of analyses of the Navier-Stokes equations with energy addition that simulates the Joule heating of a plasma discharge. It is revealed that the pressure rise due to the existence of the plasma discharge can be qualitatively explained as an effect of Joule heating.

Development of a non-linear simulation for generic hypersonic vehicles - ASUHS1

NASA Technical Reports Server (NTRS)

Salas, Juan; Lovell, T. Alan; Schmidt, David K.

1993-01-01

A nonlinear simulation is developed to model the longitudinal motion of a vehicle in hypersonic flight. The equations of motion pertinent to this study are presented. Analytic expressions for the aerodynamic forces acting on a hypersonic vehicle which were obtained from Newtonian Impact Theory are further developed. The control surface forces are further examined to incorporate vehicle elastic motion. The purpose is to establish feasible equations of motion which combine rigid body, elastic, and aeropropulsive dynamics for use in nonlinear simulations. The software package SIMULINK is used to implement the simulation. Also discussed are issues needing additional attention and potential problems associated with the implementation (with proposed solutions).

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.

Study of the coupling between real gas effects and rarefied effects on hypersonic aerodynamics

NASA Astrophysics Data System (ADS)

Chen, Song; Hu, Yuan; Sun, Quanhua

2012-11-01

Hypersonic vehicles travel across the atmosphere at very high speed, and the surrounding gas experiences complicated physical and chemical processes. These processes produce real gas effects at high temperature and rarefied gas effects at high altitude where the two effects are coupled through molecular collisions. In this study, we aim to identify the individual real gas and rarefied gas effects by simulating hypersonic flow over a 2D cylinder, a sphere and a blunted cone using a continuum-based CFD approach and the direct simulation Monte Carlo method. It is found that physical processes such as vibrational excitation and chemical reaction will reduce significantly the shock stand-off distance and flow temperature for flows having small Knudsen number. The calculated skin friction and surface heat flux will decrease when the real gas effects are considered in simulations. The trend, however, gets weakened as the Knudsen number increases. It is concluded that the rarefied gas effects weaken the real gas effects on hypersonic flows.

Investigation of piloting aids for manual control of hypersonic maneuvers

NASA Technical Reports Server (NTRS)

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

1995-01-01

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

High-Order Finite-Difference Schemes for Numerical Simulation of Hypersonic Boundary-Layer Transition

NASA Astrophysics Data System (ADS)

Zhong, Xiaolin

1998-08-01

Direct numerical simulation (DNS) has become a powerful tool in studying fundamental phenomena of laminar-turbulent transition of high-speed boundary layers. Previous DNS studies of supersonic and hypersonic boundary layer transition have been limited to perfect-gas flow over flat-plate boundary layers without shock waves. For hypersonic boundary layers over realistic blunt bodies, DNS studies of transition need to consider the effects of bow shocks, entropy layers, surface curvature, and finite-rate chemistry. It is necessary that numerical methods for such studies are robust and high-order accurate both in resolving wide ranges of flow time and length scales and in resolving the interaction between the bow shocks and flow disturbance waves. This paper presents a new high-order shock-fitting finite-difference method for the DNS of the stability and transition of hypersonic boundary layers over blunt bodies with strong bow shocks and with (or without) thermo-chemical nonequilibrium. The proposed method includes a set of new upwind high-order finite-difference schemes which are stable and are less dissipative than a straightforward upwind scheme using an upwind-bias grid stencil, a high-order shock-fitting formulation, and third-order semi-implicit Runge-Kutta schemes for temporal discretization of stiff reacting flow equations. The accuracy and stability of the new schemes are validated by numerical experiments of the linear wave equation and nonlinear Navier-Stokes equations. The algorithm is then applied to the DNS of the receptivity of hypersonic boundary layers over a parabolic leading edge to freestream acoustic disturbances.

Wind Vision. A New Era for Wind Power in the United States (Executive Summary, Full Report, and Appendices); U.S. Department of Energy (DOE), NREL (National Renewable Energy Laboratory)

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

None, None

The Wind Vision analysis demonstrates the economic value that wind power can bring to the nation, a value exceeding the costs of deployment. Wind’s environmental benefits can address key societal challenges such as climate change, air quality and public health, and water scarcity. Wind deployment can provide U.S. jobs, U.S. manufacturing, and lease and tax revenues in local communities to strengthen and support a transition of the nation’s electricity sector towards a low-carbon U.S. economy. The path needed to achieve 10% wind by 2020, 20% by 2030, and 35% by 2050 requires new tools, priorities, and emphases beyond those forgedmore » by the wind industry in growing to 4.5% of current U.S. electricity demand. Consideration of new strategies and updated priorities as identified in the Wind Vision could provide substantial positive outcomes for future generations.« less

Recent advances in hypersonic technology

NASA Technical Reports Server (NTRS)

Dwoyer, Douglas L.

1990-01-01

This paper will focus on recent advances in hypersonic aerodynamic prediction techniques. Current capabilities of existing numerical methods for predicting high Mach number flows will be discussed and shortcomings will be identified. Physical models available for inclusion into modern codes for predicting the effects of transition and turbulence will also be outlined and their limitations identified. Chemical reaction models appropriate to high-speed flows will be addressed, and the impact of their inclusion in computational fluid dynamics codes will be discussed. Finally, the problem of validating predictive techniques for high Mach number flows will be addressed.

Experimental investigation of hypersonic aerodynamics

NASA Technical Reports Server (NTRS)

Heinemann, K.; Intrieri, Peter F.

1987-01-01

An extensive series of ballistic range tests are currently being conducted at the Ames Research Center. These tests are intended to investigate the hypersonic aerodynamic characteristics of two basic configurations, which are: the blunt-cone Galileo probe which is scheduled to be launched in late 1989 and will enter the atmosphere of Jupiter in 1994, and a generic slender cone configuration to provide experimental aerodynamic data including good flow-field definition which computational aerodynamicists could use to validate their computer codes. Some of the results obtained thus far are presented and work for the near future is discussed.

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.

Internal hypersonic flow. [in thin shock layer

NASA Technical Reports Server (NTRS)

Lin, T. C.; Rubin, S. G.

1974-01-01

An approach for studying hypersonic internal flow with the aid of a thin-shock-layer approximation is discussed, giving attention to a comparison of thin-shock-layer results with the data obtained on the basis of the imposition theory or a finite-difference integration of the Euler equations. Relations in the case of strong interaction are considered together with questions of pressure distribution and aspects of the boundary-layer solution.

Grid Integration Research | Wind | NREL

Science.gov Websites

-generated simulation of a wind turbine. Wind Power Plant Modeling and Simulation Engineers at the National computer-aided engineering tool, FAST, as well as their wind power plant simulation tool, Wind-Plant

Distributed Wind Competitiveness Improvement Project

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

The Competitiveness Improvement Project (CIP) is a periodic solicitation through the U.S. Department of Energy and its National Renewable Energy Laboratory. The Competitiveness Improvement Project (CIP) is a periodic solicitation through the U.S. Department of Energy and its National Renewable Energy Laboratory. Manufacturers of small and medium wind turbines are awarded cost-shared grants via a competitive process to optimize their designs, develop advanced manufacturing processes, and perform turbine testing. The goals of the CIP are to make wind energy cost competitive with other distributed generation technology and increase the number of wind turbine designs certified to national testing standards. Thismore » fact sheet describes the CIP and funding awarded as part of the project.ufacturers of small and medium wind turbines are awarded cost-shared grants via a competitive process to optimize their designs, develop advanced manufacturing processes, and perform turbine testing. The goals of the CIP are to make wind energy cost competitive with other distributed generation technology and increase the number of wind turbine designs certified to national testing standards. This fact sheet describes the CIP and funding awarded as part of the project.« less

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.

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

PubMed Central

Liu, Mengying; Sun, Peihua

2014-01-01

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

Blowing in the wind: evaluating wind energy projects on the national forests

Treesearch

Kerry Schlichting; Evan Mercer

2011-01-01

The 650 million ac of federal lands are facing increased scrutiny for wind energy development. As a result, the US Forest Service has been directed to develop policies and procedures for siting wind energy projects. We incorporate geospatial site suitability analysis with applicable policy and management principles to illustrate the use of a Spatial Decision Support...

Updates to Multi-Dimensional Flux Reconstruction for Hypersonic Simulations on Tetrahedral Grids

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.

2010-01-01

The quality of simulated hypersonic stagnation region heating with tetrahedral meshes is investigated by using an updated three-dimensional, upwind reconstruction algorithm for the inviscid flux vector. An earlier implementation of this algorithm provided improved symmetry characteristics on tetrahedral grids compared to conventional reconstruction methods. The original formulation however displayed quantitative differences in heating and shear that were as large as 25% compared to a benchmark, structured-grid solution. The primary cause of this discrepancy is found to be an inherent inconsistency in the formulation of the flux limiter. The inconsistency is removed by employing a Green-Gauss formulation of primitive gradients at nodes to replace the previous Gram-Schmidt algorithm. Current results are now in good agreement with benchmark solutions for two challenge problems: (1) hypersonic flow over a three-dimensional cylindrical section with special attention to the uniformity of the solution in the spanwise direction and (2) hypersonic flow over a three-dimensional sphere. The tetrahedral cells used in the simulation are derived from a structured grid where cell faces are bisected across the diagonal resulting in a consistent pattern of diagonals running in a biased direction across the otherwise symmetric domain. This grid is known to accentuate problems in both shock capturing and stagnation region heating encountered with conventional, quasi-one-dimensional inviscid flux reconstruction algorithms. Therefore the test problems provide a sensitive indicator for algorithmic effects on heating. Additional simulations on a sharp, double cone and the shuttle orbiter are then presented to demonstrate the capabilities of the new algorithm on more geometrically complex flows with tetrahedral grids. These results provide the first indication that pure tetrahedral elements utilizing the updated, three-dimensional, upwind reconstruction algorithm may be used for the

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.

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.

The National Wind Erosion Research Network: Building a standardized long-term data resource for aeolian research, modeling and land management

USDA-ARS?s Scientific Manuscript database

The National Wind Erosion Research Network was established in 2014 as a collaborative effort led by the USDA Agricultural Research Service and Natural Resources Conservation Service, and USDI Bureau of Land Management, to address the need for a broad and coordinated research program to develop wind ...

Uncertainty Assessment of Hypersonic Aerothermodynamics Prediction Capability

NASA Technical Reports Server (NTRS)

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

2011-01-01

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

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

NWTC Helps Chart the World's Wind Resource Potential

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

2015-09-01

Researchers at the National Renewable Energy Laboratory's (NREL's) National Wind Technology Center (NWTC) provide the wind industry, policymakers, and other stakeholders with applied wind resource data, information, maps, and technical assistance. These tools, which emphasize wind resources at ever-increasing heights, help stakeholders evaluate the wind resource and development potential for a specific area.

X-RAY EMISSION LINE PROFILES FROM WIND CLUMP BOW SHOCKS IN MASSIVE STARS

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

Ignace, R.; Waldron, W. L.; Cassinelli, J. P.

2012-05-01

The consequences of structured flows continue to be a pressing topic in relating spectral data to physical processes occurring in massive star winds. In a preceding paper, our group reported on hydrodynamic simulations of hypersonic flow past a rigid spherical clump to explore the structure of bow shocks that can form around wind clumps. Here we report on profiles of emission lines that arise from such bow shock morphologies. To compute emission line profiles, we adopt a two-component flow structure of wind and clumps using two 'beta' velocity laws. While individual bow shocks tend to generate double-horned emission line profiles,more » a group of bow shocks can lead to line profiles with a range of shapes with blueshifted peak emission that depends on the degree of X-ray photoabsorption by the interclump wind medium, the number of clump structures in the flow, and the radial distribution of the clumps. Using the two beta law prescription, the theoretical emission measure and temperature distribution throughout the wind can be derived. The emission measure tends to be power law, and the temperature distribution is broad in terms of wind velocity. Although restricted to the case of adiabatic cooling, our models highlight the influence of bow shock effects for hot plasma temperature and emission measure distributions in stellar winds and their impact on X-ray line profile shapes. Previous models have focused on geometrical considerations of the clumps and their distribution in the wind. Our results represent the first time that the temperature distribution of wind clump structures are explicitly and self-consistently accounted for in modeling X-ray line profile shapes for massive stars.« less

A Numerical Study of Hypersonic Forebody/Inlet Integration Problem

NASA Technical Reports Server (NTRS)

Kumar, Ajay

1991-01-01

A numerical study of hypersonic forebody/inlet integration problem is presented in the form of the view-graphs. The following topics are covered: physical/chemical modeling; solution procedure; flow conditions; mass flow rate at inlet face; heating and skin friction loads; 3-D forebogy/inlet integration model; and sensitivity studies.

Permanent-Change Thermal Paints for Hypersonic Flight-Test

DTIC Science & Technology

2010-09-24

thermochromic liquid crystals (Ireland et al. 1999, Ireland & Jones 2000), and temperature sensitive paints (Liu & Sullivan 2005), thermal paints are...surfaces and fin-fuselage junctions, shock boundary layer interactions, scramjet combustion chambers and around control thrusters. Thermal paints can...use of discrete wired sensors may also not practical on some locations on a hypersonic vehicle. Thermochromic liquid crystals Coatings of

Connecting Communities to Wind Resources

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

Baring-Gould, Edward I

2017-08-18

WINDExchange is the platform for the U.S. Department of Energy's (DOE's) Wind Energy Technologies Office to disseminate credible wind energy information on a national level. Stakeholder engagement and outreach activities are designed to enable well-informed decisions about appropriate wind energy deployment. WINDExchange focuses on wind energy outreach at the national level while managing and supporting similar regional efforts through the implementation of DOE Regional Resource Centers (RRCs). This fact sheet provides an overview of DOE's WINDExchange initiative and the RRCs. Examples of RRC activities are provided.

NASA's Hypersonic Investment Area

NASA Technical Reports Server (NTRS)

Hueter, Uwe; Hutt, John; McClinton, Charles

2002-01-01

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

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.

Actively cooled plate fin sandwich structural panels for hypersonic aircraft

NASA Technical Reports Server (NTRS)

Smith, L. M.; Beuyukian, C. S.

1979-01-01

An unshielded actively cooled structural panel was designed for application to a hypersonic aircraft. The design was an all aluminum stringer-stiffened platefin sandwich structure which used a 60/40 mixture of ethylene glycol/water as the coolant. Eight small test specimens of the basic platefin sandwich concept and three fatigue specimens from critical areas of the panel design was fabricated and tested (at room temperature). A test panel representative of all features of the panel design was fabricated and tested to determine the combined thermal/mechanical performance and structural integrity of the system. The overall findings are that; (1) the stringer-stiffened platefin sandwich actively cooling concept results in a low mass design that is an excellent contender for application to a hypersonic vehicle, and (2) the fabrication processes are state of the art but new or modified facilities are required to support full scale panel fabrication.

Real-Gas Correction Factors for Hypersonic Flow Parameters in Helium

NASA Technical Reports Server (NTRS)

Erickson, Wayne D.

1960-01-01

The real-gas hypersonic flow parameters for helium have been calculated for stagnation temperatures from 0 F to 600 F and stagnation pressures up to 6,000 pounds per square inch absolute. The results of these calculations are presented in the form of simple correction factors which must be applied to the tabulated ideal-gas parameters. It has been shown that the deviations from the ideal-gas law which exist at high pressures may cause a corresponding significant error in the hypersonic flow parameters when calculated as an ideal gas. For example the ratio of the free-stream static to stagnation pressure as calculated from the thermodynamic properties of helium for a stagnation temperature of 80 F and pressure of 4,000 pounds per square inch absolute was found to be approximately 13 percent greater than that determined from the ideal-gas tabulation with a specific heat ratio of 5/3.