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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.

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At least 703 records · Page 39
All sourcesNASA NTRSDOE OSTI
NASA NTRS1997

A Dual-Line Detection Rayleigh Scattering Diagnostic Technique for the Combustion of Hydrocarbon Fuels and Filtered UV Rayleigh Scattering for Gas Velocity Measurements

Non-intrusive techniques for the dynamic measurement of gas flow properties such as density, temperature and velocity, are needed in the research leading to the development of new generation high-speed aircraft. Accurate velocity, temperature and density data obtained in ground testing and in-flight measurements can help understand the flow physics leading to transition and turbulence in supersonic, high-altitude flight. Such non-intrusive measurement techniques can also be used to study combustion processes of hydrocarbon fuels in aircraft engines. Reliable, time and space resolved temperature measurements in various combustor configurations can lead to a better understanding of high temperature chemical reaction dynamics thus leading to improved modeling and better prediction of such flows. In view of this, a research program was initiated at Polytechnic University's Aerodynamics Laboratory with support from NASA Lewis Research Center through grants NAG3-1301 and NAG3-1690. The overall objective of this program has been to develop laser-based, non-contact, space- and time-resolved temperature and velocity measurement techniques. In the initial phase of the program a ND:YAG laser-based dual-line Rayleigh scattering technique was developed and tested for the accurate measurement of gas temperature in the presence of background laser glare. Effort was next directed towards the development of a filtered, spectrally-resolved Rayleigh/Mie scattering technique with the objective of developing an interferometric method for time-frozen velocity measurements in high-speed flows utilizing the uv line of an ND:YAG laser and an appropriate molecular absorption filter. This effort included both a search for an appropriate filter material for the 266 nm laser line and the development and testing of several image processing techniques for the fast processing of Fabry-Perot images for velocity and temperature information. Finally, work was also carried out for the development of a new laser-based strain-rate and vorticity technique for the time-resolved measurement of vorticity and strain-rates in turbulent flows.

Otugen, M. Volkan↗
NASA NTRS1978

Flight-measured pressure characteristics of aft-facing steps in thick boundary layer flow for transonic and supersonic Mach numbers

Aft-facing step base pressure flight data were obtained for three step heights for nominal transonic Mach numbers of 0.80, 0.90, and 0.95, and for supersonic Mach numbers of 2.2, 2.5, and 2.8 with a Reynolds number, based on the fuselage length ahead of the step, of about 10 to the 8th power. Surface static pressures were measured ahead of the step, behind the step, and on the step face (base), and a boundary layer rake was used to obtain boundary layer reference conditions. A comparison of the data from the present and previous experiments shows the same trend of increasing base pressure ratio (decreasing drag) with increasing values of momentum thickness to step height ratios. However, the absolute level of these data does not always agree at the supersonic Mach numbers. For momentum thickness to height ratios near 1.0, the differences in the base pressure ratios appear to be primarily a function of Reynolds number based on the momentum thickness. Thus, for Mach numbers above 2, the data analyzed show that the base pressure ratio decreases (drag increases) as Reynolds number based on momentum thickness increases for a given momentum thickness and step height.

Powers, S. G.↗
NASA NTRS1974

High altitude turbulence encountered by the supersonic YF-12A airplane

The present work describes the turbulence experienced by the YF-12A airplane on the basis of airplane acceleration data obtained at altitudes above 12.2 km. Data presented include the subjective intensities reported by the air crew, the portion of flight distance in turbulence, the variation of turbulence with season, and the thickness and length of turbulence patches as determined along the flight path. Compared with that experienced by subsonic jets below 12.2 km, turbulence above 12.2 km was mild, but the crew was more sensitive to gust accelerations during supersonic flight at altitudes above 12.2 km than during subsonic flight at lower altitudes. About 6-8% of the distance traveled was in turbulence between 12.2 and 16.8 km, as compared to less than 1% above 18.3 km. High-altitude turbulence increased by a factor of three from summer to winter. Turbulence patches were 0.4 km thick and 10 km long on the average.

Ehernberger, L. J.↗
NASA NTRS1983

The minimization of pylon-mounted store effects on air combat capability

Some effects of pylon-mounted missiles on aft-tail delta wing supersonic fighter concepts have been investigated. Whereas minimum drag penalties do occur with the addition of missiles, the effects at higher lifts, corresponding to maneuvering flight, are less severe and often favorable. Lower speeds and altitudes enhance the maneuvering capability and one-on-one air combat would probably tend to degenerate to subsonic speeds even though the combatants may be flying supersonic fighters. Higher speed (supersonic) flight might best be reserved for interceptors with long-range missiles where the weapon carriage effects at low angles of attack are of prime importance.

Spearman, M. L.↗
NASA NTRS1949

Flight Investigations at High-Subsonic, Transonic, and Supersonic Speeds to Determine Zero-Lift Drag of Fin-Stabilized Bodies of Revolution having Fineness Ratios of 12.5, 8.91, and 6.04 and Varying Positions of Maximum Diameter

Rocket-powered models were flown at high-subsonic, transonic, and supersonic speeds to determine the zero-lift drag of fin-stabilized parabolic bodies of revolution differing in fineness ratio and in position of maximum diameter. The present paper presents the results for fineness ratio 12.5, 8.91 and 6.04 bodies having maximum diameters located at stations of 20, 40, 60, and 80 percent of body length. All configurations had cut-off sterns and all had equal base, frontal, and exposed fin areas. For most of the supersonic-speed range models having their maximum diameters at the 60-percent station gave the lowest values of drag coefficient. At supersonic speeds, increasing the fineness ratio generally reduced the drag coefficient for a given position of maximum diameter.

Hart, Roger G.↗
NASA NTRS2005

Transonic-Small-Disturbance and Linear Analyses for the Active Aeroelastic Wing Program

Analysis serves many roles in the Active Aeroelastic Wing (AAW) program. It has been employed to ensure safe testing of both a flight vehicle and wind tunnel model, has formulated models for control law design, has provided comparison data for validation of experimental methods and has addressed several analytical research topics. Aeroelastic analyses using mathematical models of both the flight vehicle and the wind tunnel model configurations have been conducted. Static aeroelastic characterizations of the flight vehicle and wind tunnel model have been produced in the transonic regime and at low supersonic Mach numbers. The flight vehicle has been analyzed using linear aerodynamic theory and transonic small disturbance theory. Analyses of the wind-tunnel model were performed using only linear methods. Research efforts conducted through these analyses include defining regions of the test space where transonic effects play an important role and investigating transonic similarity. A comparison of these aeroelastic analyses for the AAW flight vehicle is presented in this paper. Results from a study of transonic similarity are also presented. Data sets from these analyses include pressure distributions, stability and control derivatives, control surface effectiveness, and vehicle deflections.

Wiesman, Carol D.↗
NASA NTRS2018

ASPIRE Aerodynamic Models and Flight Performance

NASA's Advanced Supersonic Parachute Inflation Research Experiments (ASPIRE) project was established to test full-scale supersonic parachutes at Mars-relevant conditions, as a risk-reduction activity for NASA's upcomingMars2020 mission. Deployment and inflation of Disk-Gap-Band (DGB) parachutes were examined at Mach number and dynamic pressure conditions relevant to Mars2020, using a sounding rocket platform. The flight tests examined two parachutes: a build-to-print version of the parachute used by the Mars Science Laboratory and a strengthened version of this parachute that has the same geometry but differs in materials and construction. The first flight test (SR01) of the built-to-print parachute took place on October 4, 2017, followed by the first test of the strengthened parachute during flight SR02 on March 31, 2018. A second test of the strengthened parachute with a higher target load, SR03, took place on September 7, 2018. Over the sequence of the three tests, the parachute was exposed to increasing aerodynamic loads with the peak load during SR03 being significantly larger than is expected during a Martian descent. All three tests were successful: the parachute deployment and inflation occurred at the intended conditions, the measurement systems performed as designed and provided the data, and the parachutes survived the aerodynamic loads they were exposed to. The flight tests yielded valuable data on parachute forces and high-speed imagery of the deployment and inflation process.

Parachute↗
NASA NTRS1993

Aeroelastic character of a National Aerospace Plane demonstrator concept

The paper provides an analytical assessment of the flutter character of an unclassified National Aerospace Plane configuration known as the demonstrator. Linear subsonic, supersonic, and hypersonic analysis indicate that the vehicle is prone to body-freedom flutter resulting from the decrease in vibration frequency of the all-moveable wing at high flight dynamic pressures. As the wing-pivot frequency decreases, it couples with the vehicle short-period mode resulting in dynamic instability. A similar instability sometimes occurs when the pivot mode couples with the fuselage-bending mode. Also assessed, for supersonic flight conditions, are configuration variations that include relocation of the wing further aft on the lifting-body fuselage, and the addition of body flaps to the rear of the vehicle. These changes are destabilizing because they result in severe wing-pivot/fuselage-bending instabilities at dynamic pressures lower than the instabilities indicated for the original demonstrator. Finally, a two-point wing support and actuation system concept is proposed for the National Aerospace Plane, which if developed may (according to cursory analysis) enhance overall stability.

Spain, Charles V.↗
NASA NTRS2012

Wind-Tunnel and Flight Test Results for the Measurements of Flow Variables at Supersonic Speeds Using Improved Wedge and Conical Probes

The results of supersonic wind-tunnel tests on three probes at nominal Mach numbers of 1.6, 1.8 and 2.0 and flight tests on two of these probes up to a Mach number of 1.9 are described. One probe is an 8 deg. half-angle wedge with two total-pressure measurements and one static. The second, a conical probe, is a cylinder that has a 15 deg., semi-angle cone tip with one total-pressure orifice at the apex and four static-pressure orifices on the surface of the cone, 90 deg. apart, and about two-thirds of the distance from the cone apex to the base of the cone. The third is a 2 deg. semi-angle cone that has two static ports located 180 deg. apart about 1.5 inches behind the apex of the cone. The latter probe was included since it has been the "probe of choice" for wind-tunnel flow-field pressure measurements (or one similar to it) for the past half-century. The wedge and 15 deg. conical probes used in these tests were designed for flight diagnostic measurements for flight Mach numbers down to 1.35 and 1.15 respectively, and have improved capabilities over earlier probes of similar shape. The 15. conical probe also has a temperature sensor that is located inside the cylindrical part of the probe that is exposed to free-stream flow through an annulus at the apex of the cone. It enables the determination of free-stream temperature, density, speed of sound, and velocity, in addition to free-stream pressure, Mach number, angle of attack and angle of sideslip. With the time-varying velocity, acceleration can be calculated. Wind-tunnel tests of the two probes were made in NASA Langley Research Center fs Unitary Plan Wind Tunnel (UPWT) at Mach numbers of 1.6, 1.8, and 2.0. Flight tests were carried out at the NASA Dryden Flight Research Center (DFRC) on its F-15B aircraft up to Mach numbers of 1.9. The probes were attached to a fixture, referred to as the Centerline Instrumented Pylon (CLIP), under the fuselage of the aircraft. Problems controlling the velocity of the flow through the conical probe required for accurate temperature measurements are noted, as well as some calibration problems of the miniature pressure sensors that required a re-calculation of the flow variables. Data are presented for angle of attack, pressure and Mach number obtained in the wind tunnel and in flight. In the wind tunnel some transient data were obtained by translating the probes through the shock flow field created by a bump on the wind-tunnel wall.

Bobbitt, Percy J.↗
NASA NTRS2010

Analysis of Nozzle Jet Plume Effects on Sonic Boom Signature

An axisymmetric full Navier-Stokes computational fluid dynamics (CFD) study was conducted to examine nozzle exhaust jet plume effects on the sonic boom signature of a supersonic aircraft. A simplified axisymmetric nozzle geometry, representative of the nozzle on the NASA Dryden NF-15B Lift and Nozzle Change Effects on Tail Shock (LaNCETS) research airplane, was considered. The highly underexpanded nozzle flow is found to provide significantly more reduction in the tail shock strength in the sonic boom N-wave pressure signature than perfectly expanded and overexpanded nozzle flows. A tail shock train in the sonic boom signature, similar to what was observed in the LaNCETS flight data, is observed for the highly underexpanded nozzle flow. The CFD results provide a detailed description of the nozzle flow physics involved in the LaNCETS nozzle at different nozzle expansion conditions and help in interpreting LaNCETS flight data as well as in the eventual CFD analysis of a full LaNCETS aircraft. The current study also provided important information on proper modeling of the LaNCETS aircraft nozzle. The primary objective of the current CFD research effort was to support the LaNCETS flight research data analysis effort by studying the detailed nozzle exhaust jet plume s imperfect expansion effects on the sonic boom signature of a supersonic aircraft. Figure 1 illustrates the primary flow physics present in the interaction between the exhaust jet plume shock and the sonic boom coming off of an axisymmetric body in supersonic flight. The steeper tail shock from highly expanded jet plume reduces the dip of the sonic boom N-wave signature. A structured finite-volume compressible full Navier-Stokes CFD code was used in the current study. This approach is not limited by the simplifying assumptions inherent in previous sonic boom analysis efforts. Also, this study was the first known jet plume sonic boom CFD study in which the full viscous nozzle flow field was modeled, without coupling to a sonic boom propagation analysis code, from the stagnation chamber of the nozzle to the far field external flow, taking into account all nonisentropic effects in the shocks, boundary layers, and free shear layers, and their interactions at distances up to 30 times the nozzle exit diameter from the jet centerline. A CFD solution is shown in Figure 2. The flow field is very complicated and multi-dimensional, with shock-shock and shockplume interactions. At the time of this reporting, a full three-dimensional CFD study was being conducted to evaluate the effects of nozzle vectoring on the aircraft tail shock strength.

Bui, Trong↗
NASA NTRS1972

Review of sonic-boom generation theory and prediction methods.

The prediction techniques reviedi he present paper permit the calculation of sonic booms produced by rather complex conventional supersonic aircraft designs performing level nonaccelerated flight in a quiet atmosphere. Basic concepts of supersonic flow analysis, for representation of an airplane as a linear distribution of disturbances and for determination of the resultant pressure field complete with shocks, are outlined. Numerical techniques for implementation of the theory are discussed briefly, and examples of the correlation of theory with experimental data from wind tunnel and flight tests are presented. Special attention is given to presentation of a simplified method for rapid 'first-cut' estimation of farfield bow-shock overpressure. Finally, some problems encountered in attempts at applying the prediction techniques for the nearfield at high supersonic Mach numbers are recognized, and the need for further refinement of present techniques or the development of new systems is discussed.

Carlson, H. W.↗
NASA NTRS1973

Navigation, guidance, and control systems for V/STOL aircraft.

The development of digital autopilots and integrated avionics systems, applicable to many classes of vehicles and missions, was undertaken by Sperry Flight Systems in the mid-sixties. The first application of the system was planned for automatic flight control in the U.S. supersonic transport; the termination of that program, however, thwarted any flight experience. The second application, which has additional navigation and energy management functions, is an airborne simulator of the space shuttle vehicle. The latter system underwent a series of successful flight tests in a CV-990 aircraft under contract with NASA. The third application, which has new electronic displays, air data computation, and time-constrained guidance (i.e., specified position and altitude at a specified time), is in the DOT/NASA STOLAND test program. The STOLAND system is described specifically in this paper.

Osder, S. S.↗