SEARCH · Search NASA

Results for “SUPERSONIC”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 361 records · Page 20

All sources NASA NTRS DOE OSTI

NASA NTRS1987

Integration of a supersonic unsteady aerodynamic code into the NASA FASTEX system

A supersonic unsteady aerodynamic loads prediction method based on the constant pressure method was integrated into the NASA FASTEX system. The updated FASTEX code can be employed for aeroelastic analyses in subsonic and supersonic flow regimes. A brief description of the supersonic constant pressure panel method, as applied to lifting surfaces and body configurations, is followed by a documentation of updates required to incorporate this method in the FASTEX code. Test cases showing correlations of predicted pressure distributions, flutter solutions, and stability derivatives with available data are reported.

Appa, Kari↗

NASA NTRS1987

Dual Stokes cars measurements in supersonic combusting flow

Coherent antiStokes Raman spectroscopy (CARS) was used to make simultaneous measurements of temperature, nitrogen and oxygen density in a reacting supersonic flow. A supersonic burner (SSB) was designed to provide supersonic flow in which combustion can be studied in the laboratory. Measurements made with the CARS system will be used for validation of computational fluid dynamic (CFD) codes. Preliminary results of the CFD calculations are presented.

Jarrett, Olin, Jr.↗

NASA NTRS1988

Supersonic jet plume interaction with a flat plate

Supersonic jet plume interaction with a flat plate was studied using a model scaled test apparatus designed to simulate plume/aircraft structure interaction for the cruise configuration. The generic configuration consisted of a rectangular supersonic nozzle of aspect ratio 7, and a large flat plate located beneath the nozzle at various nozzle plate distances; the plate was instrumented to measure surface dynamic pressure and mean wall temperature, with provisions for measurements of acceleration and strain on coupon size panels that could be inserted in the plate. Phase-averaged schlieren measurements revealed the presence of high-intensity acoustic emission from the supersonic plume above the plate, directed upstream; this radiation could be associated with the shock noise generation. Narrow band spectra of surface dynamic pressure show spectral peaks with amplitude levels reaching 1 psi, related to the screech tones. Temperature measurements indicated elevated surface temperatures in regions of high turbulence intensity.

Seiner, John M.↗

NASA NTRS1988

A preliminary design study of supersonic through-flow fan inlets

From Mach 3.20 cruise propulsion systems, preliminary design studies for two supersonic through-flow fan primary inlets and a single core inlet were undertaken. Method of characteristics and one-dimensional performance techniques were applied to assess the potential improvements supersonic through-flow fan technology has over more conventional systems. A fixed geometry supersonic through-flow fan primary inlet was found to have better performance than a conventional inlet design on the basis of total pressure recovery, air flow, aerodynamic drag and size and weight.

Barnhart, Paul J.↗

NASA NTRS1988

Experimental investigation of the performance of a supersonic compressor cascade

Supersonic cascade wind tunnel results are presented for a linear, supersonic compressor cascade derived from the near-tip section of a high-throughflow axial flow compressor rotor over the inlet Mach number range of 1.30-1.71. Laser anemometry was used to obtain flow-velocity measurements showing the wave pattern in the entrance region. Attention is given to the unique-incidence relationship for this cascade, which relates the supersonic inlet Mach number to the inlet flow direction. An empirical correlation is obtained for the influence of the independent parameters of back pressure, axial velocity density ratio, and blade element performance.

Tweedt, T. L.↗

NASA NTRS1988

Experimental investigation of the performance of a supersonic compressor cascade

Results are presented from an experimental investigation of a linear, supersonic, compressor cascade tested in the supersonic cascade wind tunnel facility at the DFVLR in Cologne, Federal Republic of Germany. The cascade design was derived from the near-tip section of a high-through-flow axial flow compressor rotor with a design relative inlet Mach number of 1.61. Test data were obtained over a range of inlet Mach numbers from 1.23 to 1.71, and a range of static pressure ratios and axial-velocity-density ratios (AVDR) at the design inlet condition. Flow velocity measurements showing the wave pattern in the cascade entrance region were obtained using a laser transit anemometer. From these measurements, some unique-incidence conditions were determined, thus relating the supersonic inlet Mach number to the inlet flow direction. The influence of static pressure ratio and AVDR on the blade passage flow and the blade-element performance is described, and an empirical correlation is used to show the influence of these two (independent) parameters on the exit flow angle and total-pressure loss for the design inlet condition.

Tweedt, D. L.↗

NASA NTRS1988

A preliminary design study of supersonic through-flow fan inlets

From Mach 3.20 cruise propulsion systems, preliminary design studies for two supersonic through-flow fan primary inlets and a single core inlet were undertaken. Method of characteristics and one dimensional performance techniques were applied to assess the potential improvements supersonic through-flow fan technology has over more conventional systems. A fixed geometry supersonic through-flow fan primary inlet was found to have better performance than a conventional inlet design on the basis of total pressure recovery, air flow, aerodynamic drag and size and weight.

Barnhart, Paul J.↗

NASA NTRS1988

A two-dimensional numerical simulation of a supersonic, chemically reacting mixing layer

Research has been undertaken to achieve an improved understanding of physical phenomena present when a supersonic flow undergoes chemical reaction. A detailed understanding of supersonic reacting flows is necessary to successfully develop advanced propulsion systems now planned for use late in this century and beyond. In order to explore such flows, a study was begun to create appropriate physical models for describing supersonic combustion, and to develop accurate and efficient numerical techniques for solving the governing equations that result from these models. From this work, two computer programs were written to study reacting flows. Both programs were constructed to consider the multicomponent diffusion and convection of important chemical species, the finite rate reaction of these species, and the resulting interaction of the fluid mechanics and the chemistry. The first program employed a finite difference scheme for integrating the governing equations, whereas the second used a hybrid Chebyshev pseudospectral technique for improved accuracy.

Drummond, J. Philip↗

NASA NTRS1984

Experimental results for a two-dimensional supersonic inlet used as a thrust deflecting nozzle

Nearly all supersonic V/STOL aircraft concepts are dependent on the thrust deflecting capability of a nozzle. In one unique concept, referred to as the reverse flow dual fan, not only is there a thrust deflecting nozzle for the fan and core engine exit flow, but because of the way the propulsion system operates during vertical takeoff and landing, the supersonic inlet is also used as a thrust deflecting nozzle. This paper presents results of an experimental study to evaluate the performance of a supersonic inlet used as a thrust deflecting nozzle for this reverse flow dual fan concept. Results are presented in terms of nozzle thrust coefficient and thrust vector angle for a number of inlet/nozzle configurations. Flow visualization and nozzle exit flow survey results are also shown.

Johns, Albert L.↗

NASA NTRS1989

Performance potential of air turbo-ramjet employing supersonic through-flow fan

A study was conducted to assess the performance potential of a supersonic through-flow fan in an advanced engine designed to power a Mach-5 cruise vehicle. It included a preliminary evaluation of fan performance requirements and the desirability of supersonic versus subsonic combustion, the design and performance of supersonic fans, and the conceptual design of a single-pass air-turbo-rocket/ramjet engine for a Mach 5 cruise vehicle. The study results showed that such an engine could provide high thrust over the entire speed range from sea-level takeoff to Mach 5 cruise, especially over the transonic speed range, and high fuel specific impulse at the Mach 5 cruise condition, with the fan windmilling.

Kepler, C. E.↗

NASA NTRS1989

Supersonic far-field boundary conditions for transonic small-disturbance theory

Characteristic far-field boundary conditions for supersonic freestream flow have been developed and implemented within a transonic small-disturbance code. The boundary conditions have been implemented within the CAP-TSD code which has been developed recently for aeroelastic analysis of complete aircraft configurations. These boundary conditions improve the accuracy of the solutions for supersonic freestream applications. They also allow the extent of the grid to be much smaller, thus providing savings in the computational time required to obtain solutions. Comparisons are shown between surface pressures computed using large and small grid extents for the NACA 0012 airfoil and the F-5 wing at various Mach numbers and angles of attack. Both steady and unsteady results are presented and comparisons are made with Euler results and with experimental data to assess the accuracy of the new far-field boundary conditions. Comparisons of these results show that the supersonic boundary conditions allow a much smaller grid to be used without losing accuracy.

Gibbons, Michael D.↗

NASA NTRS1988

Measurements of a supersonic turbulent vortex

Mean-flow measurements of a supersonic turbulent streamwise vortex are presented. This vortex was produced by the injection of a swirling flow along the centerline of a supersonic airstream at Mach 3. Directional Mach number distributions, obtained via a five-hole flow-angularity probe, reveal vortex characteristics similar to those of the incompressible case, even though rotational Mach numbers up to 0.8 were obtained. This work is the first step of a study of the supersonic vortex breakdown phenomenon.

Metwally, O. M.↗

NASA NTRS1989

Quantitative characterization of a nonreacting, supersonic combustor flowfield using unified, laser-induced iodine fluorescence

A calibrated, nonintrusive optical technique, laser-induced iodine fluorescence (LIIF) was used to quantify the steady, compressible flowfield of a nonreacting, supersonic combustor. The combustor was configured with single and staged, transverse-air injection into a supersonic-air freestream behind a rearward-facing step. Pressure, temperature, two-velocity components, and injectant mole fraction were measured with high spatial resolution in the three-dimensional flowfields. These experimental results provide a benchmark set of data for validation of computational fluid dynamic (CFD) codes being developed to model supersonic combustor flowfields.

Fletcher, D. G.↗

NASA NTRS1990

Preliminary design of a supersonic Short Takeoff and Vertical Landing (STOVL) fighter aircraft

The preliminary design study of a supersonic Short Takeoff and Vertical Landing (STOVL) fighter is presented. A brief historical survey of powered lift vehicles was presented, followed by a technology assessment of the latest supersonic STOVL engine cycles under consideration by industry and government in the U.S. and UK. A survey of operational fighter/attack aircraft and the modern battlefield scenario were completed to develop, respectively, the performance requirements and mission profiles for the study. Three configurations were initially investigated with the following engine cycles: a hybrid fan vectored thrust cycle, a lift+lift/cruise cycle, and a mixed flow vectored thrust cycle. The lift+lift/cruise aircraft configuration was selected for detailed design work which consisted of: (1) a material selection and structural layout, including engine removal considerations, (2) an aircraft systems layout, (3) a weapons integration model showing the internal weapons bay mechanism, (4) inlet and nozzle integration, (5) an aircraft suckdown prediction, (6) an aircraft stability and control analysis, including a takeoff, hover, and transition control analysis, (7) a performance and mission capability study, and (8) a life cycle cost analysis. A supersonic fighter aircraft with STOVL capability with the lift+lift/cruise engine cycle seems a viable option for the next generation fighter.

Cox, Brian↗

NASA NTRS1988

Modeling of supersonic reacting flow fields

A detailed understanding of the scramjet combustor flow field is critical to the achievement of a successful design. Even though the combustor flow field is quite complex, it can be realistically viewed as a collection of spatially developing and reacting supersonic mixing layers that are initially discrete, but that ultimately merge into larger more complex zones. These mixing layers begin downstream of a set of fuel injectors that introduce gaseous hydrogen in both a parallel and transverse direction into a supersonic air stream entering from the engine inlet. The behavior of the initial portion of the combustor flow, in the mixing layers near the fuel injectors, appears to be most critical, since this is where the mechanism for efficient high speed mixing must be established to achieve the required degree of combustion downstream. Because of the structure of the flow field in this initial portion of the combustor, a single supersonic, spatially developing and reacting mixing layer serves as an excellent physical model for the overall flow field. Even though this reacting mixing layer flow is geometrically simple, it can still be made to retain all of the fluid mechanical and chemical complexities present in the actual combustor flow field.

Drummond, J. P.↗

NASA NTRS1990

The aerodynamic design of the oblique flying wing supersonic transport

The aerodynamic design of a supersonic oblique flying wing is strongly influenced by the requirement that passengers must be accommodated inside the wing. It was revealed that thick oblique wings of very high sweep angle can be efficient at supersonic speeds when transonic normal Mach numbers are allowed on the upper surface of the wing. The goals were motivated by the ability to design a maximum thickness, minimum size oblique flying wing. A 2-D Navier-Stokes solver was used to design airfoils up to 16 percent thickness with specified lift, drag and pitching moment. A new method was developed to calculate the required pressure distribution on the wing based on the airfoil loading, normal Mach number distribution and theoretical knowledge of the minimum drag of oblique configurations at supersonic speeds. The wing mean surface for this pressure distribution was calculated using an inverse potential flow solver. The lift to drag ratio of this wing was significantly higher than that of a comparable delta wing for cruise speeds up to Mach 2.

Vandervelden, Alexander J. M.↗

NASA NTRS1990

Chemically reacting supersonic flow calculation using an assumed PDF model

This work is motivated by the need to develop accurate models for chemically reacting compressible turbulent flow fields that are present in a typical supersonic combustion ramjet (SCRAMJET) engine. In this paper the development of a new assumed probability density function (PDF) reaction model for supersonic turbulent diffusion flames and its implementation into an efficient Navier-Stokes solver are discussed. The application of this model to a supersonic hydrogen-air flame will be considered.

Farshchi, M.↗

NASA NTRS1990

Supersonic combustion engine testbed, heat lightning

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

Hoying, D.↗