Test Capability Enhancements to the NASA Langley 8-Foot High Temperature Tunnel

NASA Technical Reports Server (NTRS)

Harvin, S. F.; Cabell, K. F.; Gallimore, S. D.; Mekkes, G. L.

2006-01-01

The NASA Langley 8-Foot High Temperature Tunnel produces true enthalpy environments simulating flight from Mach 4 to Mach 7, primarily for airbreathing propulsion and aerothermal/thermo-structural testing. Flow conditions are achieved through a methane-air heater and nozzles producing aerodynamic Mach numbers of 4, 5 or 7 and have exit diameters of 8 feet or 4.5 feet. The 12-ft long free-jet test section, housed inside a 26-ft vacuum sphere, accommodates large test articles. Recently, the facility underwent significant upgrades to support hydrocarbon fueled scramjet engine testing and to expand flight simulation capability. The upgrades were required to meet engine system development and flight clearance verification requirements originally defined by the joint NASA-Air Force X-43C Hypersonic Flight Demonstrator Project and now the Air Force X-51A Program. Enhancements to the 8-Ft. HTT were made in four areas: 1) hydrocarbon fuel delivery; 2) flight simulation capability; 3) controls and communication; and 4) data acquisition/processing. The upgrades include the addition of systems to supply ethylene and liquid JP-7 to test articles; a Mach 5 nozzle with dynamic pressure simulation capability up to 3200 psf, the addition of a real-time model angle-of-attack system; a new programmable logic controller sub-system to improve process controls and communication with model controls; the addition of MIL-STD-1553B and high speed data acquisition systems and a classified data processing environment. These additions represent a significant increase to the already unique test capability and flexibility of the facility, and complement the existing array of test support hardware such as a model injection system, radiant heaters, six-component force measurement system, and optical flow field visualization hardware. The new systems support complex test programs that require sophisticated test sequences and precise management of process fluids. Furthermore, the new systems, such as the real-time angle of attack system and the new programmable logic controller enhance the test efficiency of the facility. The motivation for the upgrades and the expanded capabilities is described here.

Atomization, drop size, and penetration for cross-stream water injection at high-altitude reentry conditions with application to the RAM C-1 and C-3 flights

NASA Technical Reports Server (NTRS)

Gooderum, P. B.; Bushnell, D. M.

1972-01-01

Atomization, drop size, and penetration data are presented for cross stream water injection at conditions simulating high altitude reentry (low Weber number, high static temperature, high Knudsen number, and low static pressure). These results are applied to the RAM C-1 and C-3 flights. Two primary breakup modes are considered, vapor pressure or flashing and aerodynamic atomization. Results are given for breakup boundaries and mean drop size for each of these atomization mechanisms. Both standard and flight orifice geometries are investigated. The data were obtained in both a static environment and in conventional aerodynamic facilities at Mach numbers of 4.5 and 8. The high temperature aspects of reentry were simulated in a Mach 5.5 cyanogen-oxygen tunnel with total temperature of 4500 K.

Forward flight effects on broadband shock associated noise of supersonic jets

NASA Technical Reports Server (NTRS)

Tam, Christopher K. W.

1989-01-01

The stochastic model theory of TAM (1987, 1989) for broadband shock associated noise was extended to include the effects of forward flight. The theory was applied to the forward flight simulation experiments of Norum and Shearin (1984, 1986, and 1988). Good agreement is found between calculated and measured far-field noise spectra over the flight Mach number range of 0.0 to 0.4.

In-pipe aerodynamic characteristics of a projectile in comparison with free flight for transonic Mach numbers

NASA Astrophysics Data System (ADS)

Hruschka, R.; Klatt, D.

2018-03-01

The transient shock dynamics and drag characteristics of a projectile flying through a pipe 3.55 times larger than its diameter at transonic speed are analyzed by means of time-of-flight and pipe wall pressure measurements as well as computational fluid dynamics (CFD). In addition, free-flight drag of the 4.5-mm-pellet-type projectile was also measured in a Mach number range between 0.5 and 1.5, providing a means for comparison against in-pipe data and CFD. The flow is categorized into five typical regimes the in-pipe projectile experiences. When projectile speed and hence compressibility effects are low, the presence of the pipe has little influence on the drag. Between Mach 0.5 and 0.8, there is a strong drag increase due to the presence of the pipe, however, up to a value of about two times the free-flight drag. This is exactly where the nose-to-base pressure ratio of the projectile becomes critical for locally sonic speed, allowing the drag to be estimated by equations describing choked flow through a converging-diverging nozzle. For even higher projectile Mach numbers, the drag coefficient decreases again, to a value slightly below the free-flight drag at Mach 1.5. This behavior is explained by a velocity-independent base pressure coefficient in the pipe, as opposed to base pressure decreasing with velocity in free flight. The drag calculated by CFD simulations agreed largely with the measurements within their experimental uncertainty, with some discrepancies remaining for free-flying projectiles at supersonic speed. Wall pressure measurements as well as measured speeds of both leading and trailing shocks caused by the projectile in the pipe also agreed well with CFD.

X-2 in flight after drop from B-50 mothership

NASA Technical Reports Server (NTRS)

1956-01-01

The Bell Aircraft Company X-2 (46-674) drops away from its Boeing B-50 mothership in this photo. Lt. Col. Frank 'Pete' Everest piloted 674 on its first unpowered flight on 5 August 1954. He made the first rocket-powered flight on 18 November 1955. Everest made the first supersonic X-2 flight in 674 on 25 April 1956, achieving a speed of Mach 1.40. In July, he reached Mach 2.87, just short of the Mach 3 goal. The other X-2, 675, was written off prior to making any powered flights. An explosion during a captive flight resulted in the death of Bell test pilot Jean 'Skip' Ziegler. The X-2 was jettisoned over Lake Ontario, and the launch aircraft was damaged beyond repair. The first X-2, 674, continued flying, making a total of 17 launches. On 7 September 1956, Capt. Iven Kincheloe became the first man to exceed 100,000 feet when he reached an altitude of 126,200 feet in 674. The X-2, initially an Air Force program, was scheduled to be transferred to the civilian National Advisory Committee for Aeronautics (NACA) for scientific research. The Air Force delayed turning the aircraft over to the NACA in the hope of attaining Mach 3 in the airplane. The service requested and received a two-month extension to qualify another Air Force test pilot, Capt. Miburn 'Mel' Apt, in the X-2 and attempt to exceed Mach 3. After several ground briefings in the simulator, Apt (with no previous rocket plane experience) made his flight on 27 September 1956. Apt raced away from the B-50 under full power, quickly outdistancing the F-100 chase planes. At high altitude, he nosed over, accelerating rapidly. The X-2 reached Mach 3.2 (2,094 mph) at 65,000 feet. Apt became the first man to fly more than three times the speed of sound. Still above Mach 3, he began an abrupt turn back to Edwards. This maneuver proved fatal as the X-2 began a series of diverging rolls and tumbled out of control. Apt tried to regain control of the aircraft. Unable to do so, Apt separated the escape capsule. Too late, he attempted to bail out and was killed when the capsule impacted on the Edwards bombing range. The rest of the X-2 crashed five miles away. The X-2 was a swept-wing, rocket-powered aircraft designed to fly faster than Mach 3 (three times the speed of sound). It was built for the U.S. Air Force by the Bell Aircraft Company, Buffalo, New York. The X-2 was flown to investigate the problems of aerodynamic heating as well as stability and control effectiveness at high altitudes and high speeds (in excess of Mach 3). Bell aircraft built two X-2 aircraft. These were constructed of K-monel (a copper and nickel alloy) for the fuselage and stainless steel for the swept wings and control surfaces. The aircraft had ejectable nose capsules instead of ejection seats because the development of ejection seats had not reached maturity at the time the X-2 was conceived. The X-2 ejection canopy was successfully tested using a German V-2 rocket. The X-2 used a skid-type landing gear to make room for more fuel. The airplane was air launched from a modified Boeing B-50 Superfortress Bomber. X-2 Number 1 made its first unpowered glide flight on Aug. 5, 1954, and made a total of 17 (4 glide and 13 powered) flights before it was lost Sept. 27, 1956. The pilot on Flight 17, Capt. Milburn Apt, had flown the aircraft to a record speed of Mach 3.2 (2,094 mph), thus becoming the first person to exceed Mach 3. During that last flight, inertial coupling occurred and the pilot was killed. The aircraft suffered little damage in the crash, resulting in proposals (never implemented) from the Langley Memorial Aeronautical Laboratory, Hampton, Virginia, to rebuild it for use in a hypersonic (Mach 5+) test program. In 1953, X-2 Number 2 was lost in an in-flight explosion while at the Bell Aircraft Company during captive flight trials and was jettisoned into Lake Ontario. The Air Force had previously flown the aircraft on three glide flights at Edwards Air Force Base, California, in 1952. Although the NACA's High-Speed Flight Station, Edwards, California, (predecessor of NASA's Dryden Flight Research Center) never actually flew the X-2 aircraft, the NACA did support the program primarily through Langley Memorial Aeronautical Laboratory wind-tunnel tests and Wallops Island, Virginia, rocket-model tests. The NACA High-Speed Flight Station also provided stability and control recording instrumentation and simulator support for the Air Force flights. In the latter regard, the NACA worked with the Air Force in using a special computer to extrapolate and predict aircraft behavior from flight data.

Exhaust emissions survey of a turbofan engine for flame holder swirl type augmentors at simulated altitude flight conditions

NASA Technical Reports Server (NTRS)

Moss, J. E., Jr.

1981-01-01

Emissions of carbon dioxide, total oxides of nitrogen, unburned hydrocarbons, and carbon monoxide from an F100 afterburning two spool turbofan engine at simulated flight conditions are reported. Tests were run at Mach 0.8 at altitudes of 10.97 and 13.71 km (36,000 and 45,000 ft), and at Mach 1.2 at 13.71 km (45,000 ft). Emission measurements were made from intermediate power (nonafterburning) through maximum afterburning, using a single point gas sample probe traversed across the horizontal diameter of the exhaust nozzle. The data show that emissions vary with flight speed, altitude, power level, and radial position across the nozzle. Carbon monoxide emissions were low for intermediate and partial afterburning power. Unburned hydrocarbons were near zero for most of the simulated flight conditions. At maximum afterburning, there were regions of NOx deficiency in regions of high CO. The results suggest that the low NOx levels observed in the tests are a result of interaction with high CO in the thermal converter. CO2 emissions were proportional to local fuel air ratio for all test conditions.

Validation of scramjet exhaust simulation technique at Mach 6

NASA Technical Reports Server (NTRS)

Hopkins, H. B.; Konopka, W.; Leng, J.

1979-01-01

Current design philosophy for hydrogen-fueled, scramjet-powered hypersonic aircraft results in configurations with strong couplings between the engine plume and vehicle aerodynamics. The experimental verification of the scramjet exhaust simulation is described. The scramjet exhaust was reproduced for the Mach 6 flight condition by the detonation tube simulator. The exhaust flow pressure profiles, and to a large extent the heat transfer rate profiles, were then duplicated by cool gas mixtures of Argon and Freon 13B1 or Freon 12. The results of these experiments indicate that a cool gas simulation of the hot scramjet exhaust is a viable simulation technique except for phenomena which are dependent on the wall temperature relative to flow temperature.

Langley Mach 4 scramjet test facility

NASA Technical Reports Server (NTRS)

Andrews, E. H., Jr.; Torrence, M. G.; Anderson, G. Y.; Northam, G. B.; Mackley, E. A.

1985-01-01

An engine test facility was constructed at the NASA Langley Research Center in support of a supersonic combustion ramjet (scramjet) technology development program. Hydrogen combustion in air with oxygen replenishment provides simulated air at Mach 4 flight velocity, pressure, and true total temperature for an altitude range from 57,000 to 86,000 feet. A facility nozzle with a 13 in square exit produces a Mach 3.5 free jet flow for engine propulsion tests. The facility is described and calibration results are presented which demonstrate the suitability of the test flow for conducting scramjet engine research.

Altitude Performance Characteristics of Tail-pipe Burner with Convergingconical Burner Section on J47 Turbojet Engine

NASA Technical Reports Server (NTRS)

Prince, William R; Mcaulay, John E

1950-01-01

An investigation of turbojet-engine thrust augmentation by means of tail-pipe burning was conducted in the NACA Lewis altitude wind tunnel. Performance data were obtained with a tail-pipe burner having a converging conical burner section installed on an axial-flow-compressor type turbojet engine over a range of simulated flight conditions and tail-pipe fuel-air ratios with a fixed-area exhaust nozzle. A maximum tail-pipe combustion efficiency of 0.86 was obtained at an altitude of 15,000 feet and a flight Mach number of 0.23. Tail-pipe burner operation was possible up to an altitude of 45,000 feet at a flight Mach number of 0.23.

Emission calibration of a J-58 afterburning turbojet engine at simulated supersonic stratospheric flight conditions

NASA Technical Reports Server (NTRS)

Holdeman, J. D.

1974-01-01

Emissions of total oxides of nitrogen, unburned hydrocarbons, and carbon monoxide from a J-58 engine at simulated flight conditions of Mach 2.0, 2.4, and 2.8 at 19.8 km altitude are reported. For each flight condition, measurements were made for four engine power levels from maximum power without afterburning through maximum afterburning. These measurements were made 7 cm downstream of the engine primary nozzle using a single point traversing gas sample probe. Results show that emissions vary with flight speed, engine power level, and with radial position across the exhaust.

Emission calibration of a J-58 afterburning turbojet engine at simulated supersonic, stratospheric flight conditions

NASA Technical Reports Server (NTRS)

Holdeman, J. D.

1974-01-01

Emissions of total oxides of nitrogen, unburned hydrocarbons, and carbon monoxide from a J-58 engine at simulated flight conditions of Mach 2.0, 2.4, and 2.8 at 19.8 km altitude are reported. For each flight condition, measurements were made for four engine power levels from maximum power without afterburning through maximum afterburning. These measurements were made 7 cm downstream of the engine primary nozzle using a single point traversing gas sample probe. Results show that emissions vary with flight speed, engine power level, and with radial position across the exhaust.

Parameter identification studies on the NASA/Ames Research Center Advanced Concepts Flight Simulator. M.S. Thesis

NASA Technical Reports Server (NTRS)

Mckavitt, Thomas P., Jr.

1990-01-01

The results of an aircraft parameters identification study conducted on the National Aeronautics and Space Administration/Ames Research Center Advanced Concepts Flight Simulator (ACFS) in conjunction with the Navy-NASA Joint Institute of Aeronautics are given. The ACFS is a commercial airline simulator with a design based on future technology. The simulator is used as a laboratory for human factors research and engineering as applied to the commercial airline industry. Parametric areas examined were engine pressure ratio (EPR), optimum long range cruise Mach number, flap reference speed, and critical take-off speeds. Results were compared with corresponding parameters of the Boeing 757 and 767 aircraft. This comparison identified two areas where improvements can be made: (1) low maximum lift coefficients (on the order of 20-25 percent less than those of a 757); and (2) low optimum cruise Mach numbers. Recommendations were made to those anticipated with the application of future technologies.

Experimental Investigation of an Integrated Strut-Rocket/Scramjet Operating at Mach 4.0 and 6.5 Conditions

NASA Technical Reports Server (NTRS)

Hawk, Clark; Nelson, Karl

1998-01-01

A series of tests were conducted to investigate RBCC performance at ramjet and scramjet conditions. The hardware consisted of a linear strut-rocket manufactured by Aerojet and a dual-mods scramjet combustor. The hardware was tested at NASA Langley Research Center in the Direct Connect Supersonic Combustion Test Facility at Mach 4.0 and 6.5 simulated flight conditions.

The X-43A (Hyper-X) Flies Into the Record Books

NASA Technical Reports Server (NTRS)

Grindle, Laurie; Bahm, Catherine

2006-01-01

The goal of the Hyper-X research program, conducted jointly by the NASA Dryden Flight Research Center and the NASA Langley Research Center, was to demonstrate and validate the technology, experimental techniques, and computation methods and tools for design and performance predictions of a hypersonic aircraft with an airframe-integrated, scramjet propulsion system. Three X-43A airframe-integrated, scramjet research vehicles were designed and fabricated to achieve that goal by flight test: two test flights at Mach 7 and one test flight at Mach 10. The first flight, conducted on June 2, 2001, experienced a launch vehicle failure and resulted in a 9-month mishap investigation. A two-year return-to-flight effort ensued and concluded when the second Mach 7 flight was successful on March 27, 2004. Just eight months later, on November 16, the X-43A successfully completed the third and final flight. These two flights were the first flight demonstrations, at Mach 7 and Mach 10 respectively, of an airframe-integrated, scramjet-powered, hypersonic vehicle.

Effect of afterbody geometry on aerodynamic characteristics of isolated nonaxisymmetric afterbodies at transonic Mach numbers

NASA Technical Reports Server (NTRS)

Bangert, Linda S.; Carson, George T., Jr.

1992-01-01

A parametric study was conducted in the Langley 16-Foot Transonic Tunnel on an isolated nonaxisymmetic fuselage model that simulates a twin-engine fighter. The effects of aft-end closure distribution (top/bottom) nozzle-flap boattail angle versus nozzle-sidewall boattail angle) and afterbody and nozzle corner treatment (sharp or radius) were investigated. Four different closure distributions with three different corner radii were tested. Tests were conducted over a range of Mach numbers from 0.40 to 1.25 and over a range of angles of attack from -3 to 9 degrees. Solid plume simulators were used to simulate the jet exhaust. For a given closure distribution in the range of Mach numbers tested, the sharp-corner nozzles generally had the highest drag, and the 2-in. corner-radius nozzles generally had the lowest drag. The effect of closure distribution on afterbody drag was highly dependent on configuration and flight condition.

Impact of Flight Enthalpy, Fuel Simulant, and Chemical Reactions on the Mixing Characteristics of Several Injectors at Hypervelocity Flow Conditions

NASA Technical Reports Server (NTRS)

Drozda, Tomasz G.; Baurle, Robert A.; Drummond, J. Philip

2016-01-01

The high total temperatures or total enthalpies required to duplicate the high-speed flight conditions in ground experiments often place stringent requirements on the material selection and cooling needs for the test articles and intrusive flow diagnostic equipment. Furthermore, for internal flows, these conditions often complicate the use of nonintrusive diagnostics that need optical access to the test section and interior portions of the flowpath. Because of the technical challenges and increased costs associated with experimentation at high values of total enthalpy, an attempt is often made to reduce it. This is the case for the Enhanced Injection and Mixing Project (EIMP) currently underway in the Arc-Heated Scramjet Test Facility at the NASA Langley Research Center. The EIMP aims to investigate supersonic combustion ramjet (scramjet) fuel injection and mixing physics, improve the understanding of underlying physical processes, and develop enhancement strategies and functional relationships between mixing performance and losses relevant to flight Mach numbers greater than 8. The experiments will consider a "direct-connect" approach and utilize a Mach 6 nozzle to simulate the combustor entrance flow of a scramjet engine. However, while the value of the Mach number is matched to that expected at the combustor entrance in flight, the maximum value of the total enthalpy for these experiments is limited by the thermal-structural limits of the uncooled experimental hardware. Furthermore, the fuel simulant is helium, not hydrogen. The use of "cold" flows and non-reacting mixtures of fuel simulants for mixing experiments is not new and has been extensively utilized as a screening technique for scramjet fuel injectors. In this study, Reynolds-averaged simulations are utilized (RAS) to systematically verify the implicit assumptions used by the EIMP. This is accomplished by first performing RAS of mixing for two injector configurations at planned nominal experimental conditions. The mixing parameters of interest, such as mixing efficiency and total pressure recovery, are then computed and compared to the values obtained from RAS under the true enthalpy conditions and using helium and hydrogen. Finally, the impact of combustion on mixing, often deemed small enough to neglect at hypervelocity conditions, is assessed by comparing the results obtained from the hydrogen-fueled reacting and non-reacting RAS. For reacting flows, in addition to mixing efficiency and total pressure recovery, the combustion efficiency and thrust potential are also considered. In all of the simulations, the incoming air Mach number and the fuel-to-air ratio are the same, while the total pressure, total enthalpy, and the fuel simulant vary depending on the case considered. It is found that under some conditions the "cold" flow experiments are a good approximation of the flight.

Flight investigation of an underwing nacelle installation of an auxiliary-inlet ejector nozzle with a clamshell flow diverter from Mach 0.6 to 1.3

NASA Technical Reports Server (NTRS)

Head, V. L.

1972-01-01

A nozzle installation of general interest is a podded engine mounted near the aft lower surface of the wing. The effect of this installation on the performance of an auxiliary-inlet ejector nozzle with a clamshell flow diverter was investigated over a Mach number range of 0.6 to 1.3 by using a modified F-106B aircraft. The clamshell flow diverter was tested in a 17 deg position with double-hinged synchronized floating doors. The ejector nozzle trailing-edge flaps were simulated in the closed position with a rigid structure which provided a boattail angle of 10 deg. Primary nozzle area was varied as exhaust gas temperature was varied between 975 and 1561 K. With the nozzle in a subsonic cruise position, the nozzle gross thrust coefficient was 0.918 at a flight Mach number of 0.9.

Development of X-43A Mach 10 Leading Edges

NASA Technical Reports Server (NTRS)

Ohlhorst, Craig W.; Glass, David E.; Bruce, Walter E., III; Lindell, Michael C.; Vaughn, Wallace L.; Dirling, R. B., Jr.; Hogenson, P. A.; Nichols, J. M.; Risner, N. W.; Thompson, D. R.

2005-01-01

The nose leading edge of the Hyper-X Mach 10 vehicle was orginally anticipated to reach temperatures near 4000 F at the leading-edge stagnation line. A SiC coated carbon/carbon (C/C) leading-edge material will not survive that extreme temperature for even a short duration single flight. To identify a suitable leading edge for the Mach 10 vehicle, arc-jet testing was performed on thirteen leading-edge segments fabricated from different material systems to evaluate their performance in a simulated flight environment. Hf, Zr, Si, and Ir based materials, in most cases as a coating on C/C, were included in the evaluation. Afterwards, MER, Tucson, AZ was selected as the supplier of the flight vehicle leading edges. The nose and the vertical and horizontal tail leading edges were fabricated out of a 3:1 biased high thermal conductivity C/C. The leading edges were coated with a three layer coating comprised of a SiC conversion of the top surface of the C/C, followed by a chemical vapor deposited layer of SiC, followed by a thin chemical vapor deposited layer of HfC. This paper will describe the fabrication of the Mach 10 C/C leading edges and the testing performed to validate performance.

Computational and Experimental Characterization of the Mach 6 Facility Nozzle Flow for the Enhanced Injection and Mixing Project at NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Drozda, Tomasz G.; Cabell, Karen F.; Passe, Bradley J.; Baurle, Robert A.

2017-01-01

Computational fluid dynamics analyses and experimental data are presented for the Mach 6 facility nozzle used in the Arc-Heated Scramjet Test Facility for the Enhanced Injection and Mixing Project (EIMP). This project, conducted at the NASA Langley Research Center, aims to investigate supersonic combustion ramjet (scramjet) fuel injection and mixing physics relevant to flight Mach numbers greater than 8. The EIMP experiments use a two-dimensional Mach 6 facility nozzle to provide the high-speed air simulating the combustor entrance flow of a scramjet engine. Of interest are the physical extent and the thermodynamic properties of the core flow at the nozzle exit plane. The detailed characterization of this flow is obtained from three-dimensional, viscous, Reynolds-averaged simulations. Thermodynamic nonequilibrium effects are also investigated. The simulations are compared with the available experimental data, which includes wall static pressures as well as in-stream static pressure, pitot pressure and total temperature obtained via in-stream probes positioned just downstream of the nozzle exit plane.

Simulation and Analyses of Stage Separation Two-Stage Reusable Launch Vehicles

NASA Technical Reports Server (NTRS)

Pamadi, Bandu N.; Neirynck, Thomas A.; Hotchko, Nathaniel J.; Tartabini, Paul V.; Scallion, William I.; Murphy, Kelly J.; Covell, Peter F.

2005-01-01

NASA has initiated the development of methodologies, techniques and tools needed for analysis and simulation of stage separation of next generation reusable launch vehicles. As a part of this activity, ConSep simulation tool is being developed which is a MATLAB-based front-and-back-end to the commercially available ADAMS(registered Trademark) solver, an industry standard package for solving multi-body dynamic problems. This paper discusses the application of ConSep to the simulation and analysis of staging maneuvers of two-stage-to-orbit (TSTO) Bimese reusable launch vehicles, one staging at Mach 3 and the other at Mach 6. The proximity and isolated aerodynamic database were assembled using the data from wind tunnel tests conducted at NASA Langley Research Center. The effects of parametric variations in mass, inertia, flight path angle, altitude from their nominal values at staging were evaluated. Monte Carlo runs were performed for Mach 3 staging to evaluate the sensitivity to uncertainties in aerodynamic coefficients.

Simulation and Analyses of Stage Separation of Two-Stage Reusable Launch Vehicles

NASA Technical Reports Server (NTRS)

Pamadi, Bandu N.; Neirynck, Thomas A.; Hotchko, Nathaniel J.; Tartabini, Paul V.; Scallion, William I.; Murphy, K. J.; Covell, Peter F.

2007-01-01

NASA has initiated the development of methodologies, techniques and tools needed for analysis and simulation of stage separation of next generation reusable launch vehicles. As a part of this activity, ConSep simulation tool is being developed which is a MATLAB-based front-and-back-end to the commercially available ADAMS(Registerd TradeMark) solver, an industry standard package for solving multi-body dynamic problems. This paper discusses the application of ConSep to the simulation and analysis of staging maneuvers of two-stage-to-orbit (TSTO) Bimese reusable launch vehicles, one staging at Mach 3 and the other at Mach 6. The proximity and isolated aerodynamic database were assembled using the data from wind tunnel tests conducted at NASA Langley Research Center. The effects of parametric variations in mass, inertia, flight path angle, altitude from their nominal values at staging were evaluated. Monte Carlo runs were performed for Mach 3 staging to evaluate the sensitivity to uncertainties in aerodynamic coefficients.

Development of the X-33 Aerodynamic Uncertainty Model

NASA Technical Reports Server (NTRS)

Cobleigh, Brent R.

1998-01-01

An aerodynamic uncertainty model for the X-33 single-stage-to-orbit demonstrator aircraft has been developed at NASA Dryden Flight Research Center. The model is based on comparisons of historical flight test estimates to preflight wind-tunnel and analysis code predictions of vehicle aerodynamics documented during six lifting-body aircraft and the Space Shuttle Orbiter flight programs. The lifting-body and Orbiter data were used to define an appropriate uncertainty magnitude in the subsonic and supersonic flight regions, and the Orbiter data were used to extend the database to hypersonic Mach numbers. The uncertainty data consist of increments or percentage variations in the important aerodynamic coefficients and derivatives as a function of Mach number along a nominal trajectory. The uncertainty models will be used to perform linear analysis of the X-33 flight control system and Monte Carlo mission simulation studies. Because the X-33 aerodynamic uncertainty model was developed exclusively using historical data rather than X-33 specific characteristics, the model may be useful for other lifting-body studies.

Analysis of sonic boom measurements near shock wave extremities for flight near Mach 1.0 and for airplane accelerations

NASA Technical Reports Server (NTRS)

Haglund, G. T.; Kane, E. J.

1974-01-01

The analysis of the 14 low-altitude transonic flights showed that the prevailing meteorological consideration of the acoustic disturbances below the cutoff altitude during threshold Mach number flight has shown that a theoretical safe altitude appears to be valid over a wide range of meteorological conditions and provides a reasonable estimate of the airplane ground speed reduction to avoid sonic boom noise during threshold Mach number flight. Recent theoretical results for the acoustic pressure waves below the threshold Mach number caustic showed excellent agreement with observations near the caustic, but the predicted overpressure levels were significantly lower than those observed far from the caustic. The analysis of caustics produced by inadvertent low-magnitude accelerations during flight at Mach numbers slightly greater than the threshold Mach number showed that folds and associated caustics were produced by slight changes in the airplane ground speed. These caustic intensities ranged from 1 to 3 time the nominal steady, level flight intensity.

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

DTIC Science & Technology

2013-03-01

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

Twin Jet

NASA Technical Reports Server (NTRS)

Henderson, Brenda; Bozak, Rick

2010-01-01

Many subsonic and supersonic vehicles in the current fleet have multiple engines mounted near one another. Some future vehicle concepts may use innovative propulsion systems such as distributed propulsion which will result in multiple jets mounted in close proximity. Engine configurations with multiple jets have the ability to exploit jet-by-jet shielding which may significantly reduce noise. Jet-by-jet shielding is the ability of one jet to shield noise that is emitted by another jet. The sensitivity of jet-by-jet shielding to jet spacing and simulated flight stream Mach number are not well understood. The current experiment investigates the impact of jet spacing, jet operating condition, and flight stream Mach number on the noise radiated from subsonic and supersonic twin jets.

Mitigation of Autoignition Due to Premixing in a Hypervelocity Flow Using Active Wall Cooling

NASA Technical Reports Server (NTRS)

Axdahl, Erik; Kumar, Ajay; Wilhite, Alan

2013-01-01

Preinjection of fuel on the forebody of an airbreathing vehicle is a proposed method to gain access to hypervelocity flight Mach numbers. However, this creates the possibility of autoignition either near the wall or in the core of the flow, thereby consuming fuel prematurely as well as increasing the amount of pressure drag on the vehicle. The computational fluid dynamics code VULCAN was used to conduct three dimensional simulations of the reacting flow in the vicinity of hydrogen injectors on a flat plate at conditions relevant to a Mach 12 notional flight vehicle forebody to determine the location where autoignition occurs. Active wall cooling strategies were formulated and simulated in response to regions of autoignition. It was found that tangential film cooling using hydrogen or helium were both able to nearly or completely eliminate wall autoignition in the flow domain of interest.

Mach 10 Stage Separation Analysis for the X43-A

NASA Technical Reports Server (NTRS)

Tartabini, Paul V.; Bose, David M.; Thornblom, Mark N.; Lien, J. P.; Martin, John G.

2007-01-01

This paper describes the pre-flight stage separation analysis that was conducted in support of the final flight of the X-43A. In that flight, which occurred less than eight months after the successful Mach 7 flight, the X-43A Research Vehicle attained a peak speed of Mach 9.6. Details are provided on how the lessons learned from the Mach 7 flight affected separation modeling and how adjustments were made to account for the increased flight Mach number. Also, the procedure for defining the feedback loop closure and feed-forward parameters employed in the separation control logic are described, and their effect on separation performance is explained. In addition, the range and nominal values of these parameters, which were included in the Mission Data Load, are presented. Once updates were made, the nominal pre-flight trajectory and Monte Carlo statistical results were determined and stress tests were performed to ensure system robustness. During flight the vehicle performed within the uncertainty bounds predicted in the pre-flight analysis and ultimately set the world record for airbreathing powered flight.

Model Selection for the Multiple Model Adaptive Algorithm for In-Flight Simulation.

DTIC Science & Technology

1987-12-01

of the two models, while the other model was given a probability of approximately zero. If the probabilties were exactly one and zero for the...Figures 6-103 through 6-107. From Figure 6-103, it can be seen that the probabilty of the model associated with the 10,000 ft, 0.35 Mach flight con

In-flight acoustic results from an advanced-design propeller at Mach numbers to 0.8

NASA Technical Reports Server (NTRS)

Mackall, K. G.; Lasagna, P. L.; Walsh, K.; Dittmar, J. H.

1982-01-01

Acoustic data for the advanced-design SR-3 propeller at Mach numbers to 0.8 and helical tip Mach numbers to 1.14 are presented. Several advanced-design propellers, previously tested in wind tunnels at the Lewis Research Center, are being tested in flight at the Dryden Flight Research Facility. The flight-test propellers are mounted on a pylon on the top of the fuselage of a JetStar airplane. Instrumentation provides near-field acoustic data for the SR-3. Acoustic data for the SR-3 propeller at Mach numbers up to 0.8, for propeller helical tip Mach numbers up to 1.14, and comparison of wind tunnel to flight data are included. Flowfield profiles measured in the area adjacent to the propeller are also included.

Overview With Results and Lessons Learned of the X-43A Mach 10 Flight

NASA Technical Reports Server (NTRS)

Marshall, Laurie A.; Bahm, Catherine; Corpening, Griffin P.; Sherrill, Robert

2005-01-01

This paper provides an overview of the final flight of the NASA X-43A project. The project consisted of three flights, two planned for Mach 7 and one for Mach 10. The third and final flight, November 16, 2004, was the first Mach 10 flight demonstration of an airframe-integrated, scramjet-powered, hypersonic vehicle. The goals and objectives for the project as well as those for the third flight are presented. The configuration of the Hyper-X stack including the X-43A, Hyper-X launch vehicle, and Hyper-X research vehicle adapter is discussed. The second flight of the X-43A was successfully conducted on March 27, 2004. Mission differences, vehicle modifications and lessons learned from the second flight as they applied to the third flight are also discussed. An overview of flight 3 results is presented.

Experimental Supersonic Combustion Research at NASA Langley

NASA Technical Reports Server (NTRS)

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

1998-01-01

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

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

NASA Technical Reports Server (NTRS)

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

2000-01-01

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

Combined Experimental and Numerical Investigation of Lightcraft no. 200 Aerodynamics at Mach 3

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

Droz, I. M.; Myrabo, L. N.; McInerney, J. P.

2008-04-28

The combined experimental and numerical research study investigated the supersonic aerodynamics of a Type 200 laser lightcraft at Mach 3 and {approx}18 km altitude. Several 1 inch (2.54 cm) and 1.25 inch (3.175 cm) diameter lightcraft models with 'closed' axisymmetric inlets were machined from 6061-T6 aluminum and tested in RPI's vacuum-driven Mach 3 wind tunnel. Schlieren photographs were taken of the unpowered models in both axial- and lateral-flight (i.e., 'Frisbee' mode) directions, then compared and contrasted with CFD predictions using Fluent registered . One 1.25 inch axial flight model was fitted with a piezoelectric load cell to measure axial dragmore » forces. Preliminary measurements of aerodynamic lift forces in the lateral flight mode were recorded as a function of angle of attack, using a special strain guage sting balance with an adjustable elbow. The bow shock structure captured in Schlieren photographs correlated well with CFD simulations, as well as with shockwave theory for common conical noses. In these axial flight model tests, slight differences were noted between the Schlieren photos and CFD density contour plots, especially with regard to the secondary shock structure; CFD results predicted these shocks closer to the shroud than nature would have it.« less

Preliminary Design of the Low Speed Propulsion Air Intake of the LAPCAT-MR2 Aircraft

NASA Astrophysics Data System (ADS)

Meerts, C.; Steelant, J.; Hendrick, P.

2011-08-01

A supersonic air intake has been designed for the low speed propulsion system of the LAPCAT-MR2 aircraft. Development has been based on the XB-70 aircraft air intake which achieves extremely high performances over a wide operation range through the combined use of variable geometry and porous wall suction for boundary layer control. Design of the LAPCAT-MR2 intake has been operated through CFD simulations using DLR TAU-Code (perfect gas model - Menter SST turbulence model). First, a new boundary condition has been validated into the DLR TAU-Code (perfect gas model) for porous wall suction modelling. Standard test cases have shown surprisingly good agreement with both theoretical predictions and experimental results. Based upon this validation, XB-70 air intake performances have been assessed through CFD simulations over the subsonic, transonic and supersonic operation regions and compared to available flight data. A new simulation strategy was deployed avoiding numerical instabilities when initiating the flow in both transonic and supersonic operation modes. First, the flow must be initiated with a far field Mach number higher than the target flight Mach number. Additionally, the inlet backpressure may only be increased to its target value once the oblique shock pattern downstream the intake compression ramps is converged. Simulations using that strategy have shown excellent agreement with in-flight measurements for both total pressure recovery ratio and variable geometry schedule prediction. The demarcation between stable and unstable operation could be well reproduced. Finally, a modified version of the XB-70 air intake has been integrated in the elliptical intake on the LAPCAT vehicle. Operation of this intake in the LAPCAT-MR2 environment is under evaluation using the same simulation strategy as the one developed for the XB-70. Performances are assessed at several key operation points to assess viability of this design. This information will allow in a next phase to better quantify the operation of the aerojet engines from take-off till the switch-over flight Mach number for the dual mode ramjet.

Summary of Rocketdyne Engine A5 Rocket Based Combined Cycle Testing

NASA Technical Reports Server (NTRS)

Ketchum. A.; Emanuel, Mark; Cramer, John

1998-01-01

Rocketdyne Propulsion and Power (RPP) has completed a highly successful experimental test program of an advanced rocket based combined cycle (RBCC) propulsion system. The test program was conducted as part of the Advanced Reusable Technology program directed by NASA-MSFC to demonstrate technologies for low-cost access to space. Testing was conducted in the new GASL Flight Acceleration Simulation Test (FAST) facility at sea level (Mach 0), Mach 3.0 - 4.0, and vacuum flight conditions. Significant achievements obtained during the test program include 1) demonstration of engine operation in air-augmented rocket mode (AAR), ramjet mode and rocket mode and 2) smooth transition from AAR to ramjet mode operation. Testing in the fourth mode (scramjet) is scheduled for November 1998.

Real-time aerodynamic heating and surface temperature calculations for hypersonic flight simulation

NASA Technical Reports Server (NTRS)

Quinn, Robert D.; Gong, Leslie

1990-01-01

A real-time heating algorithm was derived and installed on the Ames Research Center Dryden Flight Research Facility real-time flight simulator. This program can calculate two- and three-dimensional stagnation point surface heating rates and surface temperatures. The two-dimensional calculations can be made with or without leading-edge sweep. In addition, upper and lower surface heating rates and surface temperatures for flat plates, wedges, and cones can be calculated. Laminar or turbulent heating can be calculated, with boundary-layer transition made a function of free-stream Reynolds number and free-stream Mach number. Real-time heating rates and surface temperatures calculated for a generic hypersonic vehicle are presented and compared with more exact values computed by a batch aeroheating program. As these comparisons show, the heating algorithm used on the flight simulator calculates surface heating rates and temperatures well within the accuracy required to evaluate flight profiles for acceptable heating trajectories.

Comparisons Between NO PLIF Imaging and CFD Simulations of Mixing Flowfields for High-Speed Fuel Injectors

NASA Technical Reports Server (NTRS)

Drozda, Tomasz, G.; Cabell, Karen F.; Ziltz, Austin R.; Hass, Neil E.; Inman, Jennifer A.; Burns, Ross A.; Bathel, Brett F.; Danehy, Paul M.; Abul-Huda, Yasin M.; Gamba, Mirko

2017-01-01

The current work compares experimentally and computationally obtained nitric oxide (NO) planar laser-induced fluorescence (PLIF) images of the mixing flowfields for three types of high-speed fuel injectors: a strut, a ramp, and a rectangular flush-wall. These injection devices, which exhibited promising mixing performance at lower flight Mach numbers, are currently being studied as a part of the Enhanced Injection and Mixing Project (EIMP) at the NASA Langley Research Center. The EIMP aims to investigate scramjet fuel injection and mixing physics, and improve the understanding of underlying physical processes relevant to flight Mach numbers greater than eight. In the experiments, conducted in the NASA Langley Arc-Heated Scramjet Test Facility (AHSTF), the injectors are placed downstream of a Mach 6 facility nozzle, which simulates the high Mach number air flow at the entrance of a scramjet combustor. Helium is used as an inert substitute for hydrogen fuel. The PLIF is obtained by using a tunable laser to excite the NO, which is present in the AHSTF air as a direct result of arc-heating. Consequently, the absence of signal is an indication of pure helium (fuel). The PLIF images computed from the computational fluid dynamics (CFD) simulations are obtained by combining a fluorescence model for NO with the Reynolds-Averaged Simulation results carried out using the VULCAN-CFD solver to obtain a computational equivalent of the experimentally measured PLIF signal. The measured NO PLIF signal is mainly a function of NO concentration allowing for semi-quantitative comparisons between the CFD and the experiments. The PLIF signal intensity is also sensitive to pressure and temperature variations in the flow, allowing additional flow features to be identified and compared with the CFD. Good agreement between the PLIF and the CFD results provides increased confidence in the CFD simulations for investigations of injector performance.

An Analytical Explanation for the X-43A Flush Air Data Sensing System Pressure Mismatch Between Flight and Theory

NASA Technical Reports Server (NTRS)

Ellsworth, Joel C.

2010-01-01

Following the successful Mach 7 flight test of the X-43A, unexpectedly low pressures were measured by the aft set of the onboard Flush Air Data Sensing System s pressure ports. These in-flight aft port readings were significantly lower below Mach 3.5 than was predicted by theory. The same lower readings were also seen in the Mach 10 flight of the X-43A and in wind-tunnel data. The pre-flight predictions were developed based on 2-dimensional wedge flow, which fails to predict some of the significant 3-dimensional flow features in this geometry at lower Mach numbers. Using Volterra s solution to the wave equation as a starting point, a three-dimensional finite wedge approximation to flow over the X-43A forebody is presented. The surface pressures from this approximation compare favorably with the measured wind tunnel and flight data at speeds of Mach 2.5 and 3.

Local Flow Conditions for Propulsion Experiments on the NASA F-15B Propulsion Flight Test Fixture

NASA Technical Reports Server (NTRS)

Vachon, Michael J.; Moes, Timothy R.; Corda, Stephen

2005-01-01

Local flow conditions were measured underneath the National Aeronautics and Space Administration F-15B airplane to support development of future experiments on the Propulsion Flight Test Fixture (PFTF). The local Mach number and flow angles were measured using a conventional air data boom on a cone-cylinder mounted under the PFTF and compared with the airplane air data nose boom measurements. At subsonic flight speeds, the airplane and PFTF Mach numbers were approximately equal. Transonic Mach number values were up to 0.1 greater at the PFTF than the airplane, which is a counterintuitive result. The PFTF local supersonic Mach numbers were as much as 0.46 less than the airplane values. The maximum local Mach number at the PFTF was approximately 1.6 at an airplane Mach number near 2.0. The PFTF local angle of attack was negative at all Mach numbers, ranging from -3 to -8 degrees. When the airplane angle of sideslip was zero, the PFTF local value was zero between Mach 0.8 and Mach 1.1, -2 degrees between Mach 1.1 and Mach 1.5, and increased from zero to 1 degree from Mach 1.5 to Mach 2.0. Airplane inlet shock waves crossed the aerodynamic interface plane between Mach 1.85 and Mach 1.90.

Uncertainty Due to Unsteady Fluid/Structure Interaction for the Ares I Vehicle Traversing the Transonic Regime

NASA Technical Reports Server (NTRS)

Bartels, Robert E.

2012-01-01

Rapid reduced-order numerical models are being investigated as candidates to simulate the dynamics of a flexible launch vehicle during atmospheric ascent. There has also been the extension of these new approaches to include gust response. These methods are used to perform aeroelastic and gust response analyses at isolated Mach numbers. Such models require a method to time march through a succession of ascent Mach numbers. An approach is presented for interpolating reduced-order models of the unsteady aerodynamics at successive Mach numbers. The transonic Mach number range is considered here since launch vehicles can suffer the highest dynamic loads through this range. Realistic simulations of the flexible vehicle behavior as it traverses this Mach number range are presented. The response of the vehicle due to gusts is computed. Uncertainties in root mean square and maximum bending moment and crew module accelerations are presented due to assumed probability distributions in design parameters, ascent flight conditions, gusts. The primary focus is on the uncertainty introduced by modeling fidelity. It is found that an unsteady reduced order model produces larger excursions in the root mean square loading and accelerations than does a quasi-steady reduced order model.

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

NASA Astrophysics Data System (ADS)

Waltrup, Paul J.

1992-09-01

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

Testing of DLR C/C-SiC and C/C for HIFiRE 8 Scramjet Combustor

NASA Technical Reports Server (NTRS)

Glass, David E.; Capriotti, Diego P.; Reimer, Thomas; Kutemeyer, Marius; Smart, Michael K.

2014-01-01

Ceramic Matrix Composites (CMCs) have been proposed for use as lightweight hot structures in scramjet combustors. Previous studies have calculated significant weight savings by utilizing CMCs (active and passive) versus actively cooled metallic scramjet structures. Both a carbon/carbon (C/C) and a carbon/carbon-silicon carbide (C/C-SiC) material fabricated by DLR (Stuttgart, Germany) are being considered for use in a passively cooled combustor design for Hypersonic International Flight Research Experimentation (HIFiRE) 8, a joint Australia / Air Force Research Laboratory hypersonic flight program, expected to fly at Mach 7 for approximately 30 sec, at a dynamic pressure of 55 kilopascals. Flat panels of the DLR C/C and C/C-SiC materials were installed downstream of a hydrogen-fueled, dual-mode scramjet combustor and tested for several minutes at conditions simulating flight at Mach 5 and Mach 6. Gaseous hydrogen fuel was used to fuel the scramjet combustor. The test panels were instrumented with embedded Type K and Type S thermocouples. Zirconia felt insulation was used during some of the tests to reduce heat loss from the back surface and thus increase the heated surface temperature of the C/C-SiC panel approximately 177 C (350 F). The final C/C-SiC panel was tested for three cycles totaling over 135 sec at Mach 6 enthalpy. Slightly more erosion was observed on the C/C panel than the C/C-SiC panels, but both material systems demonstrated acceptable recession performance for the HIFiRE 8 flight.

Dual-Mode Combustion of Hydrogen in a Mach 5, Continuous-Flow Facility

NASA Technical Reports Server (NTRS)

Goyne, C. P.; McDaniel, J. C.; Quagliaroli, T. M.; Krauss, R. H.; Day, S. W.; Reubush, D. E. (Technical Monitor); McClinton, C. R. (Technical Monitor); Reubush, D. E.

2001-01-01

Results of an experimental and numerical study of a dual-mode scramjet combustor are reported. The experiment consisted of a direct-connect test of a Mach 2 hydrogen-air combustor with a single unswept-ramp fuel injector. The flow stagnation enthalpy simulated a flight Mach number of 5. Measurements were obtained using conventional wall instrumentation and a particle-imaging laser diagnostic technique. The particle imaging was enabled through the development of a new apparatus for seeding fine silicon dioxide particles into the combustor fuel stream. Numerical simulations of the combustor were performed using the GASP code. The modeling, and much of the experimental work, focused on the supersonic combustion mode. Reasonable agreement was observed between experimental and numerical wall pressure distributions. However, the numerical model was unable to predict accurately the effects of combustion on the fuel plume size, penetration, shape, and axial growth.

Hyper-X Flight Engine Ground Testing for X-43 Flight Risk Reduction

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Rock, Kenneth E.; Ruf, Edward G.; Witte, David W.; Andrews, Earl H., Jr.

2001-01-01

Airframe-integrated scramjet engine testing has been completed at Mach 7 flight conditions in the NASA Langley 8-Foot High Temperature Tunnel as part of the NASA Hyper-X program. This test provided engine performance and operability data, as well as design and database verification, for the Mach 7 flight tests of the Hyper-X research vehicle (X-43), which will provide the first-ever airframe-integrated scramjet data in flight. The Hyper-X Flight Engine, a duplicate Mach 7 X-43 scramjet engine, was mounted on an airframe structure that duplicated the entire three-dimensional propulsion flowpath from the vehicle leading edge to the vehicle trailing edge. This model was also tested to verify and validate the complete flight-like engine system. This paper describes the subsystems that were subjected to flight-like conditions and presents supporting data. The results from this test help to reduce risk for the Mach 7 flights of the X-43.

Flight Test Techniques Used to Evaluate Performance Benefits During Formation Flight

NASA Technical Reports Server (NTRS)

Ray, Ronald J.; Cobleigh, Brent R.; Vachon, M. Jake; SaintJohn, Clinton

2002-01-01

The Autonomous Formation Flight research project has been implemented at the NASA Dryden Flight Research Center to demonstrate the benefits of formation flight and develop advanced technologies to facilitate exploiting these benefits. Two F/A-18 aircraft have been modified to precisely control and monitor relative position, and to determine performance of the trailing airplane. Flight test maneuvers and analysis techniques have been developed to determine the performance advantages, including drag and fuel flow reductions and improvements in range factor. By flying the trailing airplane through a matrix of lateral, longitudinal, and vertical offset positions, a detailed map of the performance benefits has been obtained at two flight conditions. Significant performance benefits have been obtained during this flight test phase. Drag reductions of more than 20 percent and fuel flow reductions of more than 18 percent have been measured at flight conditions of Mach 0.56 and an altitude of 25,000 ft. The results show favorable agreement with published theory and generic predictions. An F/A-18 long-range cruise mission at Mach 0.8 and an altitude of 40,000 ft has been simulated in the optimum formation position and has demonstrated a 14-percent fuel reduction when compared with a controlled chase airplane of similar configuration.

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

NASA Technical Reports Server (NTRS)

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

2005-01-01

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

X-38 Application of Dynamic Inversion Flight Control

NASA Technical Reports Server (NTRS)

Wacker, Roger; Munday, Steve; Merkle, Scott

2001-01-01

This paper summarizes the application of a nonlinear dynamic inversion (DI) flight control system (FCS) to an autonomous flight test vehicle in NASA's X-38 Project, a predecessor to the International Space Station (ISS) Crew Return Vehicle (CRV). Honeywell's Multi-Application Control-H (MACH) is a parameterized FCS design architecture including both model-based DI rate-compensation and classical P+I command-tracking. MACH was adopted by X-38 in order to shorten the design cycle time for different vehicle shapes and flight envelopes and evolving aerodynamic databases. Specific design issues and analysis results are presented for the application of MACH to the 3rd free flight (FF3) of X-38 Vehicle 132 (V132). This B-52 drop test, occurring on March 30, 2000, represents the first flight test of MACH and one of the first few known applications of DI in the primary FCS of an autonomous flight test vehicle.

Exhaust emission calibration of two J-58 afterburning turbojet engines at simulated high-altitude, supersonic flight conditions

NASA Technical Reports Server (NTRS)

Holdeman, J. D.

1976-01-01

Emissions of total oxides of nitrogen, nitric oxide, unburned hydrocarbons, carbon monoxide, and carbon dioxide from two J-58 afterburning turbojet engines at simulated high-altitude flight conditions are reported. Test conditions included flight speeds from Mach 2 to 3 at altitudes from 16.0 to 23.5 km. For each flight condition exhaust measurements were made for four or five power levels, from maximum power without afterburning through maximum afterburning. The data show that exhaust emissions vary with flight speed, altitude, power level, and radial position across the exhaust. Oxides of nitrogen emissions decreased with increasing altitude and increased with increasing flight speed. Oxides of nitrogen emission indices with afterburning were less than half the value without afterburning. Carbon monoxide and hydrocarbon emissions increased with increasing altitude and decreased with increasing flight speed. Emissions of these species were substantially higher with afterburning than without.

CFD Simulations of the Supersonic Inflatable Aerodynamic Decelerator (SIAD) Ballistic Range Tests

NASA Technical Reports Server (NTRS)

Brock, Joseph; Stern, Eric; Wilder, Michael

2017-01-01

A series of ballistic range tests were performed on a scaled model of the Supersonic Flight Demonstration Test (SFDT) intended to test the Supersonic Inflatable Aerodynamic Decelerator (SIAD) geometry. The purpose of these experiments were to provide aerodynamic coefficients of the vehicle to aid in mission and vehicle design. The experimental data spans the moderate Mach number range, $3.8-2.0$, with a total angle of attack ($alpha_T$) range, $10o-20o$. These conditions are intended to span the Mach-$alpha$ space for the majority of the SFDT experiment. In an effort to validate the predictive capabilities of Computational Fluid Dynamics (CFD) for free-flight aerodynamic behavior, numerical simulations of the ballistic range experiment are performed using the unstructured finite volume Navier-Stokes solver, US3D. Comparisons to raw vehicle attitude, and post-processed aerodynamic coefficients are made between simulated results and experimental data. The resulting comparisons for both raw model attitude and derived aerodynamic coefficients show good agreement with experimental results. Additionally, near body pressure field values for each trajectory simulated are investigated. Extracted surface and wake pressure data gives further insights into dynamic flow coupling leading to a potential mechanism for dynamic instability.

X-43C Flight Demonstrator Project Overview

NASA Technical Reports Server (NTRS)

Moses, Paul L.

2003-01-01

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

Extrapolation From Wind Tunnel to Flight: Shuttle Orbiter Aerodynamics

NASA Technical Reports Server (NTRS)

Muylaert, J.; Walpot, L.; Rostand, P.; Rapuc, M.; Brauckmann, G.; Paulson, J.; Trockmorton, D.; Weilmuenster, K.

1998-01-01

The paper reviews a combined numerical and experimental activity on the Shuttle Orbiter, first performed at NASA Langley within the Orbiter Experiment (OEX) and subsequently at ESA, as part of the AGARD FDP WG 18 activities. The study at Langley was undertaken to resolve the pitch up anomaly observed during the entry of the first flight of the Shuttle Orbiter. The present paper will focus on real gas effects on aerodynamics and not on heating. The facilities used at NASA Langley were the 15-in. Mach 6, the 20-in, Mach 6, the 31-in. Mach 10 and the 20-in. Mach 6 CF4 facility. The paper focuses on the high Mach, high altitude portion of the first entry of the Shuttle where the vehicle exhibited a nose-up pitching moment relative to pre-flight prediction of (Delta C(sub m)) = 0.03. In order to study the relative contribution of compressibility, viscous interaction and real gas effects on basic body pitching moment and flap efficiency, an experimental study was undertaken to examine the effects of Mach, Reynolds and ratio of specific heats at NASA. At high Mach, a decrease of gamma occurs in the shock layer due to high temperature effects. The primary effect of this lower specific heat ratio is a decrease of the pressure on the aft windward expansion surface of the Orbiter causing the nose-up pitching moment. Testing in the heavy gas, Mach 6 CF4 tunnel, gave a good simulation of high temperature effects. The facilities used at ESA were the lm Mach 10 at ONERA Modane, the 0.7 m hot shot F4 at ONERA Le Fauga and the 0.88 m piston driven shock tube HEG at DLR Goettingen. Encouraging good force measurements were obtained in the F4 facility on the Orbiter configuration. Testing of the same model in the perfect gas Mach 10 S4 Modane facility was performed so as to have "reference" conditions. When one compares the F4 and S4 test results, the data suggests that the Orbiter "pitch up" is due to real gas effects. In addition, pressure measurements, performed on the aft portion of the windward side of the Halis configuration in HEG and F4, confirm that the pitch up is mainly attributed to a reduction of pressure due to a local decrease in gamma.

Space shuttle plume simulation application. Results and math model. [Ames unitary plan wind tunnel test

NASA Technical Reports Server (NTRS)

Boyle, W.; Conine, B.

1978-01-01

Pressure and gauge wind tunnel data from a transonic test of a 0.02 scale model of the space shuttle launch vehicle was analyzed to define the aerodynamic influence of the main propulsion system and solid rocket booster plumes during the transonic portion of ascent flight. Air was used as a simulant gas to develop the model exhaust plumes. A math model of the plume induced aerodynamic characteristics was developed for a range of Mach numbers to match the forebody aerodynamic math model. The base aerodynamic characteristics are presented in terms of forces and moments versus attitude. Total vehicle base and forebody aerodynamic characteristics are presented in terms of aerodynamic coefficients for Mach number from 0.6 to 1.4 Element and component base and forebody aerodynamic characteristics are presented for Mach numbers of 0.6, 1.05, 1.1, 1.25 and 1.4. The forebody data is available at Mach 1.55. Tolerances for all plume induced aerodynamic characteristics are developed in terms of a math model.

A Chief Engineer's View of the NASA X-43A Scramjet Flight Test

NASA Technical Reports Server (NTRS)

Marshall, Laurie A.; Corpening, Griffin P.; Sherrill, Robert

2005-01-01

This paper presents an overview of the preparation and execution of the first two flights of the NASA X-43A scramjet flight test project. The project consisted of three flights, two planned for Mach 7 and one for Mach 10. The first flight, conducted on June 2, 2001, was unsuccessful and resulted in a nine-month mishap investigation. A two-year return to flight effort ensued and concluded when the second Mach 7 flight was successfully conducted on March 27, 2004. The challenges faced by the project team as they prepared the first ever scramjet-powered airplane for flight are presented. Modifications made to the second flight vehicle as a result of the first flight failure and the return to flight activities are discussed. Flight results and lessons learned are also presented.

Ground Measurements of Airplane Shock-Wave Noise at Mach Numbers to 2.0 and at Altitudes to 60,000 Feet

NASA Technical Reports Server (NTRS)

Lina, Lindsay J.; Maglieri, Domenic J.

1960-01-01

The intensity of shock-wave noise at the ground resulting from flights at Mach numbers to 2.0 and altitudes to 60,000 feet was measured. Meagurements near the ground track for flights of a supersonic fighter and one flight of a supersonic bomber are presented. Level cruising flight at an altitude of 60,000 feet and a Mach number of 2.0 produced sonic booms which were considered to be tolerable, and it is reasonable t o expect that cruising flight at higher altitudes will produce booms of tolerable intensity for airplanes of the size and weight of the test airplanes. The measured variation of sonic-boom intensity with altitude was in good agreement with the variation calculated by an equation given in NASA Technical Note D-48. The effect of Mach number on the ground overpressure is small between Mach numbers of 1.4 and 2.0, a result in agreement with the theory. No amplification of the shock-wave overpressures due to refraction effects was apparent near the cutoff Mach number. A method for estimating the effect of fligh-path angle on cutoff Mach number is shown. Experimental results indicate agreement with the method, since a climb maneuver produced booms of a much decreased intensity as compared with the intensity of those measured in level flight at about the same altitude and Mach number. Comparison of sound pressure levels for the fighter and bomber airp lanes indicated little effect of either airplane size or weight at an altitude of 40,000 feet.

Cryogenic wind tunnels: Unique capabilities for the aerodynamicist

NASA Technical Reports Server (NTRS)

Hall, R. M.

1976-01-01

The cryogenic wind-tunnel concept as a practical means for improving ground simulation of transonic flight conditions. The Langley 1/3-meter transonic cryogenic tunnel is operational, and the design of a cryogenic National Transonic Facility is undertaken. A review of some of the unique capabilities of cryogenic wind tunnels is presented. In particular, the advantages of having independent control of tunnel Mach number, total pressure, and total temperature are highlighted. This separate control over the three tunnel parameters will open new frontiers in Mach number, Reynolds number, aeroelastic, and model-tunnel interaction studies.

Hyper-X Program Status

NASA Technical Reports Server (NTRS)

McClinton, Charles R.; Rausch, Vincent L.; Sitz, Joel; Reukauf, Paul

2001-01-01

This paper provides an overview of the objectives and status of the Hyper-X program, which is tailored to move hypersonic, airbreathing vehicle technology from the laboratory environment to the flight environment. The first Hyper-X research vehicle (HXRV), designated X-43, is being prepared at the Dryden Flight Research Center for flight at Mach 7. Extensive risk reduction activities for the first flight are completed, and non-recurring design activities for the Mach 10 X-43 (3rd flight) are nearing completion. The Mach 7 flight of the X-43, in the spring of 2001, will be the first flight of an airframe-integrated scramjet-powered vehicle. The Hyper-X program is continuing to plan follow-on activities to focus an orderly continuation of hypersonic technology development through flight research.

Hyper-X Program Status

NASA Technical Reports Server (NTRS)

McClinton, Charles R.; Reubush, David E.; Sitz, Joel; Reukauf, Paul

2001-01-01

This paper provides an overview of the objectives and status of the Hyper-X program, which is tailored to move hypersonic, airbreathing vehicle technology from the laboratory environment to the flight environment. The first Hyper-X research vehicle (HXRV), designated X-43, is being prepared at the Dryden Flight Research Center for flight at Mach 7. Extensive risk reduction activities for the first flight are completed, and non-recurring design activities for the Mach 10 X-43 (third flight) are nearing completion. The Mach 7 flight of the X-43, in the spring of 2001, will be the first flight of an airframe-integrated scramjet-powered vehicle. The Hyper-X program is continuing to plan follow-on activities to focus an orderly continuation of hypersonic technology development through flight research.

X-31 quasi-tailless flight demonstration

NASA Technical Reports Server (NTRS)

Huber, Peter; Schellenger, Harvey G.

1994-01-01

The primary objective of the quasi-tailless flight demonstration is to demonstrate the feasibility of using thrust vectoring for directional control of an unstable aircraft. By using this low-cost, low-risk approach it is possible to get information about required thrust vector control power and deflection rates from an inflight experiment as well as insight in low-power thrust vectoring issues. The quasi-tailless flight demonstration series with the X-31 began in March 1994. The demonstration flight condition was Mach 1.2 at 37,500 feet. A series of basic flying quality maneuvers, doublets, bank to bank rolls, and wind-up-turns have been performed with a simulated 100% vertical tail reduction. Flight test and supporting simulation demonstrated that the quasi-tailless approach is effective in representing the reduced stability of tailless configurations. The flights also demonstrated that thrust vectoring could be effectively used to stabilize a directionally unstable configuration and provide control power for maneuver coordination.

Computations of Viking Lander Capsule Hypersonic Aerodynamics with Comparisons to Ground and Flight Data

NASA Technical Reports Server (NTRS)

Edquist, Karl T.

2006-01-01

Comparisons are made between the LAURA Navier-Stokes code and Viking Lander Capsule hypersonic aerodynamics data from ground and flight measurements. Wind tunnel data are available for a 3.48 percent scale model at Mach 6 and a 2.75 percent scale model at Mach 10.35, both under perfect gas air conditions. Viking Lander 1 aerodynamics flight data also exist from on-board instrumentation for velocities between 2900 and 4400 m/sec (Mach 14 to 23.3). LAURA flowfield solutions are obtained for the geometry as tested or flown, including sting effects at tunnel conditions and finite-rate chemistry effects in flight. Using the flight vehicle center-of-gravity location (trim angle approx. equals -11.1 deg), the computed trim angle at tunnel conditions is within 0.31 degrees of the angle derived from Mach 6 data and 0.13 degrees from the Mach 10.35 trim angle. LAURA Mach 6 trim lift and drag force coefficients are within 2 percent of measured data, and computed trim lift-to-drag ratio is within 4 percent of the data. Computed trim lift and drag force coefficients at Mach 10.35 are within 5 percent and 3 percent, respectively, of wind tunnel data. Computed trim lift-to-drag ratio is within 2 percent of the Mach 10.35 data. Using the nominal density profile and center-of-gravity location, LAURA trim angle at flight conditions is within 0.5 degrees of the total angle measured from on-board instrumentation. LAURA trim lift and drag force coefficients at flight conditions are within 7 and 5 percent, respectively, of the flight data. Computed trim lift-to-drag ratio is within 4 percent of the data. Computed aerodynamics sensitivities to center-of-gravity location, atmospheric density, and grid refinement are generally small. The results will enable a better estimate of aerodynamics uncertainties for future Mars entry vehicles where non-zero angle-of-attack is required.

NASA's Hyper-X Program

NASA Technical Reports Server (NTRS)

Rausch, Vincent L.; McClinton, Charles R.; Sitz, Joel; Reukauf, Paul

2000-01-01

This paper provides an overview of the objectives and status of the Hyper-X program which is tailored to move hypersonic, airbreathing vehicle technology from the laboratory environment to the flight environment, the last stage preceding prototype development. The first Hyper-X research vehicle (HXRV), designated X-43, is being prepared at the Dryden Flight Research Center for flight at Mach 7 in the near future. In addition, the associated booster and vehicle-to-booster adapter are being prepared for flight and flight test preparations are well underway. Extensive risk reduction activities for the first flight and non-recurring design for the Mach 10 X-43 (3rd flight) are nearing completion. The Mach 7 flight of the X-43 will be the first flight of an airframe-integrated scramjet-powered vehicle.

CFD Analysis of Mixing Characteristics of Several Fuel Injectors at Hypervelocity Flow Conditions

NASA Technical Reports Server (NTRS)

Drozda, Tomasz G.; Drummond, J. Philip; Baurle, Robert A.

2016-01-01

CFD analysis is presented of the mixing characteristics and performance of three fuel injectors at hypervelocity flow conditions. The calculations were carried out using the VULCAN-CFD solver and Reynolds-Averaged Simulations (RAS). The high Mach number flow conditions match those proposed for the planned experiments conducted as a part of the Enhanced Injection and Mixing Project (EIMP) at the NASA Langley Research Center. The EIMP aims to investigate scramjet fuel injection and mixing physics, improve the understanding of underlying physical processes, and develop enhancement strategies and functional relationships relevant to flight Mach numbers greater than eight. Because of the high Mach number flow considered, the injectors consist of a fuel placement device, a strut; and a fluidic vortical mixer, a ramp. These devices accomplish the necessary task of distributing and mixing fuel into the supersonic cross-flow albeit via different strategies. Both of these devices were previously studied at lower flight Mach numbers where they exhibited promising performance in terms of mixing efficiency and total pressure recovery. For comparison, a flush-wall injector is also included. This type of injector generally represents the simplest method of introducing fuel into a scramjet combustor, however, at high flight Mach number conditions, the dynamic pressure needed to induce sufficient fuel penetration may be difficult to achieve along with other requirements such as achieving desired levels of fuel-to-air mixing at the required equivalence ratio. The three injectors represent the baseline configurations planned for the experiments. The current work discusses the mixing flow field behavior and differences among the three fuel injectors, mixing performance as described by the mixing efficiency and the total pressure recovery, and performance considerations based on the thrust potential.

Development and Implementation of Software for Visualizing and Editing Multidimensional Flight Simulation Input Data

NASA Technical Reports Server (NTRS)

Whelan, Todd Michael

1996-01-01

In a real-time or batch mode simulation that is designed to model aircraft dynamics over a wide range of flight conditions, a table look- up scheme is implemented to determine the forces and moments on the vehicle based upon the values of parameters such as angle of attack, altitude, Mach number, and control surface deflections. Simulation Aerodynamic Variable Interface (SAVI) is a graphical user interface to the flight simulation input data, designed to operate on workstations that support X Windows. The purpose of the application is to provide two and three dimensional visualization of the data, to allow an intuitive sense of the data set. SAVI also allows the user to manipulate the data, either to conduct an interactive study of the influence of changes on the vehicle dynamics, or to make revisions to data set based on new information such as flight test. This paper discusses the reasons for developing the application, provides an overview of its capabilities, and outlines the software architecture and operating environment.

Development and flight test results of an autothrottle control system at Mach 3 cruise

NASA Technical Reports Server (NTRS)

Gilyard, G. B.; Burken, J. J.

1980-01-01

Flight test results obtained with the original Mach hold autopilot designed the YF-12C airplane which uses elevator control and a newly developed Mach hold system having an autothrottle integrated with an altitude hold autopilot system are presented. The autothrottle tests demonstrate good speed control at high Mach numbers and high altitudes while simultaneously maintaining control over altitude and good ride qualities. The autothrottle system was designed to control either Mach number or knots equivalent airspeed (KEAS). Excellent control of Mach number or KEAS was obtained with the autothrottle system when combined with altitude hold. Ride qualities were significantly better than with the conventional Mach hold system.

Full-envelope aerodynamic modeling of the Harrier aircraft

NASA Technical Reports Server (NTRS)

Mcnally, B. David

1986-01-01

A project to identify a full-envelope model of the YAV-8B Harrier using flight-test and parameter identification techniques is described. As part of the research in advanced control and display concepts for V/STOL aircraft, a full-envelope aerodynamic model of the Harrier is identified, using mathematical model structures and parameter identification methods. A global-polynomial model structure is also used as a basis for the identification of the YAV-8B aerodynamic model. State estimation methods are used to ensure flight data consistency prior to parameter identification.Equation-error methods are used to identify model parameters. A fixed-base simulator is used extensively to develop flight test procedures and to validate parameter identification software. Using simple flight maneuvers, a simulated data set was created covering the YAV-8B flight envelope from about 0.3 to 0.7 Mach and about -5 to 15 deg angle of attack. A singular value decomposition implementation of the equation-error approach produced good parameter estimates based on this simulated data set.

Flight Demonstration Of Low Overpressure N-Wave Sonic Booms And Evanescent Waves

NASA Astrophysics Data System (ADS)

Haering, Edward A.; Smolka, James W.; Murray, James E.; Plotkin, Kenneth J.

2006-05-01

The recent flight demonstration of shaped sonic booms shows the potential for quiet overland supersonic flight, which could revolutionize air transport. To successfully design quiet supersonic aircraft, the upper limit of an acceptable noise level must be determined through quantitative recording and subjective human response measurements. Past efforts have concentrated on the use of sonic boom simulators to assess human response, but simulators often cannot reproduce a realistic sonic boom sound. Until now, molecular relaxation effects on low overpressure rise time had never been compared with flight data. Supersonic flight slower than the cutoff Mach number, which generates evanescent waves, also prevents loud sonic booms from impacting the ground. The loudness of these evanescent waves can be computed, but flight measurement validation is needed. A novel flight demonstration technique that generates low overpressure N-waves using conventional military aircraft is outlined, in addition to initial quantitative flight data. As part of this demonstration, evanescent waves also will be recorded.

Flight Demonstration Of Low Overpressure N-Wave Sonic Booms And Evanescent Waves

NASA Technical Reports Server (NTRS)

Haering, Edward A., Jr.; Smolka, James W.; Murray, James E.; Plotkin, Kenneth J.

2005-01-01

The recent flight demonstration of shaped sonic booms shows the potential for quiet overland supersonic flight, which could revolutionize air transport. To successfully design quiet supersonic aircraft, the upper limit of an acceptable noise level must be determined through quantitative recording and subjective human response measurements. Past efforts have concentrated on the use of sonic boom simulators to assess human response, but simulators often cannot reproduce a realistic sonic boom sound. Until now, molecular relaxation effects on low overpressure rise time had never been compared with flight data. Supersonic flight slower than the cutoff Mach number, which generates evanescent waves, also prevents loud sonic booms from impacting the ground. The loudness of these evanescent waves can be computed, but flight measurement validation is needed. A novel flight demonstration technique that generates low overpressure N-waves using conventional military aircraft is outlined, in addition to initial quantitative flight data. As part of this demonstration, evanescent waves also will be recorded.

The X-43A Flush Airdata Sensing System Flight Test Results

NASA Technical Reports Server (NTRS)

Baumann, Ethan; Pahle, Joseph W.; Davis, Mark; White, John Terry

2008-01-01

The National Aeronautics and Space Administration (NASA) has flight-tested a flush airdata sensing (FADS) system on the Hyper-X Research Vehicle (X-43A) at hypersonic speeds during the course of two successful flights. For this series of tests, the FADS system was calibrated to operate between Mach 3 and Mach 8, and flight test data was collected between Mach 1 and Mach 10. The FADS system acquired pressure data from surface-mounted ports and generated a real-time angle-of-attack (alpha) estimate on board the X-43A. The collected data were primarily intended to evaluate the FADS system performance, and the estimated alpha was used by the flight control algorithms on the X-43A for only a portion of the first successful flight. This paper provides an overview of the FADS system and alpha estimation algorithms, presents the in-flight alpha estimation algorithm performance, and provides comparisons to wind tunnel results and theory. Results indicate that the FADS system adequately estimated the alpha of the vehicle during the hypersonic portions of the two flights.

An experimental investigation of the NASA space shuttle external tank at hypersonic Mach numbers

NASA Technical Reports Server (NTRS)

Wittliff, C. E.

1975-01-01

Pressure and heat transfer tests were conducted simulating flight conditions which the space shuttle external tank will experience prior to break-up. The tests were conducted in the Calspan 48-inch Hypersonic Shock Tunnel and simulated entry conditions for nominal, abort-once-around (AOA), and return to launch site (RTLS) launch occurrences. Surface pressure and heat-transfer-rate distributions were obtained with and without various protuberences (or exterior hardware) on the model at Mach numbers from 15.2 to 17.7 at angles of attack from -15 deg to -180 deg and at several roll angles. The tests were conducted over a Reynolds number range from 1300 to 58,000, based on model length.

Piloted simulation study of a balloon-assisted deployment of an aircraft at high altitude

NASA Technical Reports Server (NTRS)

Murray, James; Moes, Timothy; Norlin, Ken; Bauer, Jeffrey; Geenen, Robert; Moulton, Bryan; Hoang, Stephen

1992-01-01

A piloted simulation was used to study the feasibility of a balloon assisted deployment of a research aircraft at high altitude. In the simulation study, an unmanned, modified sailplane was carried to 110,000 ft with a high altitude balloon and released in a nose down attitude. A remote pilot controlled the aircraft through a pullout and then executed a zoom climb to a trimmed, 1 g flight condition. A small parachute was used to limit the Mach number during the pullout to avoid adverse transonic effects. The use of small rocket motor was studied for increasing the maximum attainable altitude. Aerodynamic modifications to the basic sailplane included applying supercritical airfoil gloves over the existing wing and tail surfaces. The aerodynamic model of the simulated aircraft was based on low Reynolds number wind tunnel tests and computational techniques, and included large Mach number and Reynolds number effects at high altitude. Parametric variations were performed to study the effects of launch altitude, gross weight, Mach number limit, and parachute size on the maximum attainable stabilized altitude. A test altitude of approx. 95,000 ft was attained, and altitudes in excess of 100,000 ft was attained.

Numerical Analysis of the Trailblazer Inlet Flowfield for Hypersonic Mach Numbers

NASA Technical Reports Server (NTRS)

Steffen, C. J., Jr.; DeBonis, J. R.

1999-01-01

A study of the Trailblazer vehicle inlet was conducted using the Global Air Sampling Program (GASP) code for flight Mach numbers ranging from 4-12. Both perfect gas and finite rate chemical analysis were performed with the intention of making detailed comparisons between the two results. Inlet performance was assessed using total pressure recovery and kinetic energy efficiency. These assessments were based upon a one-dimensional stream-thrust-average of the axisymmetric flowfield. Flow visualization utilized to examine the detailed shock structures internal to this mixed-compression inlet. Kinetic energy efficiency appeared to be the least sensitive to differences between the perfect gas and finite rate chemistry results. Total pressure recovery appeared to be the most sensitive discriminator between the perfect gas and finite rate chemistry results for flight Mach numbers above Mach 6. Adiabatic wall temperature was consistently overpredicted by the perfect gas model for flight Mach numbers above Mach 4. The predicted shock structures were noticeably different for Mach numbers from 6-12. At Mach 4, the perfect gas and finite rate chemistry models collapse to the same result.

Wind-tunnel development of an SR-71 aerospike rocket flight test configuration

NASA Technical Reports Server (NTRS)

Smith, Stephen C.; Shirakata, Norm; Moes, Timothy R.; Cobleigh, Brent R.; Conners, Timothy H.

1996-01-01

A flight experiment has been proposed to investigate the performance of an aerospike rocket motor installed in a lifting body configuration. An SR-71 airplane would be used to carry the aerospike configuration to the desired flight test conditions. Wind-tunnel tests were completed on a 4-percent scale SR-71 airplane with the aerospike pod mounted in various locations on the upper fuselage. Testing was accomplished using sting and blade mounts from Mach 0.6 to Mach 3.2. Initial test objectives included assessing transonic drag and supersonic lateral-directional stability and control. During these tests, flight simulations were run with wind-tunnel data to assess the acceptability of the configurations. Early testing demonstrated that the initial configuration with the aerospike pod near the SR-71 center of gravity was unsuitable because of large nosedown pitching moments at transonic speeds. The excessive trim drag resulting from accommodating this pitching moment far exceeded the excess thrust capability of the airplane. Wind-tunnel testing continued in an attempt to find a configuration suitable for flight test. Multiple configurations were tested. Results indicate that an aft-mounted model configuration possessed acceptable performance, stability, and control characteristics.

Engine performance analysis and optimization of a dual-mode scramjet with varied inlet conditions

NASA Astrophysics Data System (ADS)

Tian, Lu; Chen, Li-Hong; Chen, Qiang; Zhong, Feng-Quan; Chang, Xin-Yu

2016-02-01

A dual-mode scramjet can operate in a wide range of flight conditions. Higher thrust can be generated by adopting suitable combustion modes. Based on the net thrust, an analysis and preliminary optimal design of a kerosene-fueled parameterized dual-mode scramjet at a crucial flight Mach number of 6 were investigated by using a modified quasi-one-dimensional method and simulated annealing strategy. Engine structure and heat release distributions, affecting the engine thrust, were chosen as analytical parameters for varied inlet conditions (isolator entrance Mach number: 1.5-3.5). Results show that different optimal heat release distributions and structural conditions can be obtained at five different inlet conditions. The highest net thrust of the parameterized dual-mode engine can be achieved by a subsonic combustion mode at an isolator entrance Mach number of 2.5. Additionally, the effects of heat release and scramjet structure on net thrust have been discussed. The present results and the developed analytical method can provide guidance for the design and optimization of high-performance dual-mode scramjets.

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

NASA Technical Reports Server (NTRS)

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

1992-01-01

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

PAB3D Simulations for the CAWAPI F-16XL

NASA Technical Reports Server (NTRS)

Elmiligui, Alaa; Abdol-Hamid, K. S.; Massey, Steven J.

2007-01-01

Numerical simulations of the flow around F-16XL are performed as a contribution to the Cranked Arrow Wing Aerodynamic Project International (CAWAPI) using the PAB3D CFD code. Two turbulence models are used in the calculations: a standard k-! model, and the Shih-Zhu-Lumley (SZL) algebraic stress model. Seven flight conditions are simulated for the flow around the F-16XL where the free stream Mach number varies from 0.242 to 0.97. The range of angles of attack varies from 0deg to 20deg. Computational results, surface static pressure, boundary layer velocity profiles, and skin friction are presented and compared with flight data. Numerical results are generally in good agreement with flight data, considering that only one grid resolution is utilized for the different flight conditions simulated in this study. The ASM results are closer to the flight data than the k-! model results. The ASM predicted a stronger primary vortex, however, the origin of the vortex and footprint is approximately the same as in the k-! predictions.

Mach Cutoff Analysis and Results from NASA's Farfield Investigation of No-Boom Thresholds

NASA Technical Reports Server (NTRS)

Cliatt, Larry J., II; Hill, Michael A.; Haering, Edward A., Jr.

2016-01-01

In support of the ongoing effort by the National Aeronautics and Space Administration (NASA) to bring supersonic commercial travel to the public, the NASA Armstrong Flight Research Center and the NASA Langley Research Center, in partnership with other industry organizations and academia, conducted a flight research experiment to analyze acoustic propagation in the Mach cutoff shadow zone. The effort was conducted in the fall of 2012 and named the Farfield Investigation of No-boom Thresholds (FaINT). The test helped to build a dataset that will go toward further understanding of the unique acoustic propagation characteristics below Mach cutoff altitude. FaINT was able to correlate sonic boom noise levels measured below cutoff altitude with precise airplane flight conditions, potentially increasing the accuracy over previous studies. A NASA F-18B airplane made supersonic passes such that its Mach cutoff caustic would be at varying distances above a linear 60-microphone, 7375-ft (2247.9 m) long array. A TG-14 motor glider equipped with a microphone on its wing-tip also attempted to capture the same sonic boom waves above ground, but below the Mach cutoff altitude. This paper identified an appropriate metric for sonic boom waveforms in the Mach cutoff shadow zone called Perceived Sound Exposure Level; derived an empirical relationship between Mach cutoff flight conditions and noise levels in the shadow zone; validated a safe cutoff altitude theory presented by previous studies; analyzed the sensitivity of flight below Mach cutoff to unsteady atmospheric conditions and realistic aircraft perturbations; and demonstrated the ability to record sonic boom measurements over 5000 ft (1524.0 m) above ground level, but below Mach cutoff altitude.

Farfield inflight measurements of high-speed turboprop noise

NASA Technical Reports Server (NTRS)

Balombin, J. R.; Loeffler, I. J.

1983-01-01

A flight program was carried out to determine the variation of noise level with distance from a model high-speed propeller. Noise measurements were obtained at different distances from a SR-3 propeller mounted on a JetStar aircraft, with the test instrumentation mounted on a Learjet flown in formation. The propeller was operated at 0.8 m flight Mach number, 1.12 helical tip Mach number and at 0.7 flight Mach number, 1.0 helical tip Mach number. The instantaneous pressure from individual blades was observed to rise faster at the 0.8 flight speed, than at the 0.7 M flight speed. The measured levels appeared to decrease in good agreement with a 6 dB/doubling of distance decay, over the measurement range of approximately 16 m to 100 m distance. Further extrapolation, to the distances represented by a community, would suggest that the propagated levels during cruise would not cause a serious community annoyance.

Measurement uncertainty and feasibility study of a flush airdata system for a hypersonic flight experiment

NASA Technical Reports Server (NTRS)

Whitmore, Stephen A.; Moes, Timothy R.

1994-01-01

Presented is a feasibility and error analysis for a hypersonic flush airdata system on a hypersonic flight experiment (HYFLITE). HYFLITE heating loads make intrusive airdata measurement impractical. Although this analysis is specifically for the HYFLITE vehicle and trajectory, the problems analyzed are generally applicable to hypersonic vehicles. A layout of the flush-port matrix is shown. Surface pressures are related airdata parameters using a simple aerodynamic model. The model is linearized using small perturbations and inverted using nonlinear least-squares. Effects of various error sources on the overall uncertainty are evaluated using an error simulation. Error sources modeled include boundarylayer/viscous interactions, pneumatic lag, thermal transpiration in the sensor pressure tubing, misalignment in the matrix layout, thermal warping of the vehicle nose, sampling resolution, and transducer error. Using simulated pressure data for input to the estimation algorithm, effects caused by various error sources are analyzed by comparing estimator outputs with the original trajectory. To obtain ensemble averages the simulation is run repeatedly and output statistics are compiled. Output errors resulting from the various error sources are presented as a function of Mach number. Final uncertainties with all modeled error sources included are presented as a function of Mach number.

Inlet Aerodynamics and Ram Drag of Laser-Propelled Lightcraft Vehicles

NASA Astrophysics Data System (ADS)

Langener, Tobias; Myrabo, Leik; Rusak, Zvi

2010-05-01

Numerical simulations are used to study the aerodynamic inlet properties of three axisymmetric configurations of laser-propelled Lightcraft vehicles operating at subsonic, transonic and supersonic speeds up to Mach 5. The 60 cm vehicles were sized for launching 0.1-1.0 kg nanosatellites with combined-cycle airbreathing/rocket engines, transitioning between propulsion modes at roughly Mach 5-6. Results provide the pressure, temperature, density, and velocity flowfields around and through the three representative vehicle/engine configurations, as well as giving the resulting ram drag and total drag coefficients—all as a function of flight Mach number. Simulations with rotating boundaries were also carried out, since for stability reasons, Lightcraft are normally spun up before lift-off. Given the three alternatives, it is demonstrated that the optimal geometry for minimum drag is the configuration with a parabola nose; hence, these inlet flow conditions are being applied in subsequent "direct connect" 2D laser propulsion experiments in a small transonic flow facility.

Comparisons Between NO PLIF Imaging and CFD Simulations of Mixing Flowfields for High-Speed Fuel Injectors

NASA Technical Reports Server (NTRS)

Drozda, Tomasz G.; Cabell, Karen F.; Ziltz, Austin R.; Hass, Neal E.; Inman, Jennifer A.; Burns, Ross A.; Bathel, Brett F.; Danehy, Paul M.

2017-01-01

The current work compares experimentally and computationally obtained nitric oxide (NO) planar laser induced fluorescence (PLIF) images of the mixing flowfields for three types of high-speed fuel injectors: a strut, a ramp, and a rectangular flushwall. These injection devices, which exhibited promising mixing performance at lower flight Mach numbers, are currently being studied as a part of the Enhanced Injection and Mixing Project (EIMP) at the NASA Langley Research Center. The EIMP aims to investigate scramjet fuel injection and mixing physics, and improve the understanding of underlying physical processes relevant to flight Mach numbers greater than eight. In the experiments, conducted in the NASA Langley Arc-Heated Scramjet Test Facility (AHSTF), the injectors are placed downstream of a Mach 6 facility nozzle, which simulates the high Mach number air flow at the entrance of a scramjet combustor. Helium is used as an inert substitute for hydrogen fuel. Both schlieren and PLIF techniques are applied to obtain mixing flowfield flow visualizations. The experimental PLIF is obtained by using a UV laser sheet to interrogate a plane of the flow by exciting fluorescence from the NO molecules, which are present in the AHSTF air. Consequently, the absence of signal in the resulting PLIF images is an indication of pure helium (fuel). The computational PLIF is obtained by applying a fluorescence model for NO to the results of the Reynolds-averaged simulations (RAS) of the mixing flow field carried out using the VULCAN-CFD solver. This approach is required because the PLIF signal is a nonlinear function of not only NO concentration, but also pressure, temperature, and the flow velocity. This complexity allows additional flow features to be identified and compared with those obtained from the computational fluid dynamics (CFD) simulations, however, such comparisons are only semiquantitative. Three-dimensional image reconstruction, similar to that used in magnetic resonance imaging, is also used to obtain images in the streamwise and spanwise planes from select cross-stream PLIF plane data. Synthetic schlieren is also computed from the RAS data. Good agreement between the experimental and computational results provides increased confidence in the CFD simulations for investigations of injector performance.

X-43A Final Flight Observations

NASA Technical Reports Server (NTRS)

Grindle, Laurie

2011-01-01

The presentation will provide an overview of the final flight of the NASA X-43A project. The project consisted of three flights, two planned for Mach 7 and one for Mach 10. The first flight, conducted on June 2, 2001, was unsuccessful and resulted in a nine-month mishap investigation. A two-year return to flight effort ensued and concluded when the second Mach 7 flight was successfully conducted on March 27, 2004. The third and final flight, which occurred on November 16, 2004, was the first Mach 10 flight demonstration of an airframe-integrated, scramjet-powered, hypersonic vehicle. As such, the final flight presented first time technical challenges in addition to final flight project closeout concerns. The goals and objectives for the third flight as well as those for the project will be presented. The configuration of the Hyper-X stack including the X-43A, Hyper-X launch vehicle, and Hyper-X research vehicle adapter wil also be presented. Mission differences, vehicle modifications and lessons learned from the first and second flights as they applied to the third flight will also be discussed. Although X-43A flight 3 was always planned to be the final flight of the X-43A project, the X-43 program had two other vehicles and corresponding flight phases in X-43C and X-43B. Those other projects never manifested under the X-43 banner and X-43A flight 3 also became the final flight of X-43 program.

Combustor Operability and Performance Verification for HIFiRE Flight 2

NASA Technical Reports Server (NTRS)

Storch, Andrea M.; Bynum, Michael; Liu, Jiwen; Gruber, Mark

2011-01-01

As part of the Hypersonic International Flight Research Experimentation (HIFiRE) Direct-Connect Rig (HDCR) test and analysis activity, three-dimensional computational fluid dynamics (CFD) simulations were performed using two Reynolds-Averaged Navier Stokes solvers. Measurements obtained from ground testing in the NASA Langley Arc-Heated Scramjet Test Facility (AHSTF) were used to specify inflow conditions for the simulations and combustor data from four representative tests were used as benchmarks. Test cases at simulated flight enthalpies of Mach 5.84, 6.5, 7.5, and 8.0 were analyzed. Modeling parameters (e.g., turbulent Schmidt number and compressibility treatment) were tuned such that the CFD results closely matched the experimental results. The tuned modeling parameters were used to establish a standard practice in HIFiRE combustor analysis. Combustor performance and operating mode were examined and were found to meet or exceed the objectives of the HIFiRE Flight 2 experiment. In addition, the calibrated CFD tools were then applied to make predictions of combustor operation and performance for the flight configuration and to aid in understanding the impacts of ground and flight uncertainties on combustor operation.

Transition Induced by Fence Geometrics on Shuttle Orbiter at Mach 10

NASA Technical Reports Server (NTRS)

Everhart, Joel L.

2010-01-01

Fence-induced transition data simulating a raised gap filler have been acquired on the wing lower surface of a Shuttle Orbiter model in the Langley 31-Inch Mach 10 Tunnel to compare with the Shuttle Boundary Layer Transition Flight and HYTHIRM Experiments, and to provide additional correlation data for the Boundary Layer Transition Tool. In a qualitative assessment, the data exhibit the expected response to all parameter variations; however, it is unclear whether fully effective tripping at the fence was ever realized at any test condition with the present model hardware. A preliminary, qualitative comparison of the ground-based transition measurements with those obtained from the STS-128 HYTHIRM imagery at Mach 15 reveal similar transition-wake response characteristics in terms of the spreading and the path along the vehicle surface.

Wreckage of the X-2 rocket plane was taken to NACA's High Speed Flight Station for analysis following the 1956 crash that killed Air Force pilot Capt. Mel Apt

NASA Image and Video Library

1956-11-21

The X-2, initially an Air Force program, was scheduled to be transferred to the civilian National Advisory Committee for Aeronautics (NACA) for scientific research. The Air Force delayed turning the aircraft over to the NACA in the hope of attaining Mach 3 in the airplane. The service requested and received a two-month extension to qualify another Air Force test pilot, Capt. Miburn "Mel" Apt, in the X-2 and attempt to exceed Mach 3. After several ground briefings in the simulator, Apt (with no previous rocket plane experience) made his flight on 27 September 1956. Apt raced away from the B-50 under full power, quickly outdistancing the F-100 chase planes. At high altitude, he nosed over, accelerating rapidly. The X-2 reached Mach 3.2 (2,094 mph) at 65,000 feet. Apt became the first man to fly more than three times the speed of sound. Still above Mach 3, he began an abrupt turn back to Edwards. This maneuver proved fatal as the X-2 began a series of diverging rolls and tumbled out of control. Apt tried to regain control of the aircraft. Unable to do so, Apt separated the escape capsule. Too late, he attempted to bail out and was killed when the capsule impacted on the Edwards bombing range. The rest of the X-2 crashed five miles away. The wreckage of the X-2 rocket plane was later taken to NACA's High Speed Flight Station for analysis following the crash.

A piloted evaluation of an oblique-wing research aircraft motion simulation with decoupling control laws

NASA Technical Reports Server (NTRS)

Kempel, Robert W.; Mcneill, Walter E.; Gilyard, Glenn B.; Maine, Trindel A.

1988-01-01

The NASA Ames Research Center developed an oblique-wing research plane from NASA's digital fly-by-wire airplane. Oblique-wing airplanes show large cross-coupling in control and dynamic behavior which is not present on conventional symmetric airplanes and must be compensated for to obtain acceptable handling qualities. The large vertical motion simulator at NASA Ames-Moffett was used in the piloted evaluation of a proposed flight control system designed to provide decoupled handling qualities. Five discrete flight conditions were evaluated ranging from low altitude subsonic Mach numbers to moderate altitude supersonic Mach numbers. The flight control system was effective in generally decoupling the airplane. However, all participating pilots objected to the high levels of lateral acceleration encountered in pitch maneuvers. In addition, the pilots were more critical of left turns (in the direction of the trailing wingtip when skewed) than they were of right turns due to the tendency to be rolled into the left turns and out of the right turns. Asymmetric side force as a function of angle of attack was the primary cause of lateral acceleration in pitch. Along with the lateral acceleration in pitch, variation of rolling and yawing moments as functions of angle of attack caused the tendency to roll into left turns and out of right turns.

Flight investigation of XB-70 structural response to oscillatory aerodynamic shaker excitation and correlation with analytical results

NASA Technical Reports Server (NTRS)

Mckay, J. M.; Kordes, E. E.; Wykes, J. H.

1973-01-01

The low frequency symmetric structural response and damping characteristics of the XB-70 airplane were measured at four flight conditions: heavyweight at a Mach number of 0.87 at an altitude of 7620 meters (25,000 feet); lightweight at a Mach number of 0.86 at an altitude of 7620 meters (25,000 feet); a Mach number of 1.59 at an altitude of 11,918 meters (39.100 feet); and a Mach number of 2.38 and an altitude of 18,898 meters (62,000 feet). The flight data are compared with the response calculated by using early XB-70 design data and with the response calculated with mass, structural, and aerodynamic data updated to reflect as closely as possible the airplane characteristics at three of the flight conditions actually flown.

Base pressure and heat transfer tests of the 0.0225-scale space shuttle plume simulation model (19-OTS) in yawed flight conditions in the NASA-Lewis 10x10-foot supersonic wind tunnel (test IH83)

NASA Technical Reports Server (NTRS)

Foust, J. W.

1979-01-01

Wind tunnel tests were performed to determine pressures, heat transfer rates, and gas recovery temperatures in the base region of a rocket firing model of the space shuttle integrated vehicle during simulated yawed flight conditions. First and second stage flight of the space shuttle were simulated by firing the main engines in conjunction with the SRB rocket motors or only the SSME's into the continuous tunnel airstream. For the correct rocket plume environment, the simulated altitude pressures were halved to maintain the rocket chamber/altitude pressure ratio. Tunnel freestream Mach numbers from 2.2 to 3.5 were simulated over an altitude range of 60 to 130 thousand feet with varying angle of attack, yaw angle, nozzle gimbal angle and SRB chamber pressure. Gas recovery temperature data derived from nine gas temperature probe runs are presented. The model configuration, instrumentation, test procedures, and data reduction are described.

X-43C Plans and Status

NASA Technical Reports Server (NTRS)

Moses, Paul L.

2003-01-01

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

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.

Effects of reaction control system jet simulation on the stability and control characteristics of a 0.015 scale space shuttle orbiter model tested in the Langley Research Center unitary plan wind tunnel

NASA Technical Reports Server (NTRS)

Daileda, J. J.; Marroquin, J.

1974-01-01

An experimental investigation was performed in the Langley Research Center Unitary Plan Wind Tunnel (Test 0A70) to obtain the detailed effects that RCS jet flow interactions with local orbiter flow field have on supersonic stability and control characteristics of the space shuttle orbiter. Six-component force data were obtained through an angle-of-attack range from 15 to 35 degrees at angles of sideslip of 0, +5, and -5 degrees. The test was conducted with yaw jet simulation at free-stream Mach numbers of 2.5 and 4.6, simulating SSV re-entry flight conditions at these Mach numbers. In addition to the basic force measurements, fuselage base pressures and pressures on the non-metric RCS pods were obtained.

CFD Simulation of the Space Shuttle Launch Vehicle with Booster Separation Motor and Reaction Control System Plumes

NASA Technical Reports Server (NTRS)

Gea, L. M.; Vicker, D.

2006-01-01

The primary objective of this paper is to demonstrate the capability of computational fluid dynamics (CFD) to simulate a very complicated flow field encountered during the space shuttle ascent. The flow field features nozzle plumes from booster separation motor (BSM) and reaction control system (RCS) jets with a supersonic incoming cross flow at speed of Mach 4. The overset Navier-Stokes code OVERFLOW, was used to simulate the flow field surrounding the entire space shuttle launch vehicle (SSLV) with high geometric fidelity. The variable gamma option was chosen due to the high temperature nature of nozzle flows and different plume species. CFD predicted Mach contours are in good agreement with the schlieren photos from wind tunnel test. Flow fields are discussed in detail and the results are used to support the debris analysis for the space shuttle Return To Flight (RTF) task.

CFD Simulation of the Space Shuttle Launch Vehicle with Booster Separation Motor and Reaction Control Plumes

NASA Technical Reports Server (NTRS)

Gea, L. M.; Vicker, D.

2006-01-01

The primary objective of this paper is to demonstrate the capability of computational fluid dynamics (CFD) to simulate a very complicated flow field encountered during the space shuttle ascent. The flow field features nozzle plumes from booster separation motor (BSM) and reaction control system (RCS) jets with a supersonic incoming cross flow at speed of Mach 4. The overset Navier-Stokes code OVERFLOW, was used to simulate the flow field surrounding the entire space shuttle launch vehicle (SSLV) with high geometric fidelity. The variable gamma option was chosen due to the high temperature nature of nozzle flows and different plume species. CFD predicted Mach contours are in good agreement with the schlieren photos from wind tunnel test. Flow fields are discussed in detail and the results are used to support the debris analysis for the space shuttle Return To Flight (RTF) task.

High speed commercial transport fuels considerations and research needs

NASA Technical Reports Server (NTRS)

Lee, C. M.; Niedzwiecki, R. W.

1989-01-01

NASA is currently evaluating the potential of incorporating High Speed Civil Transport (HSCT) aircraft in the commercial fleet in the beginning of the 21st century. NASA sponsored HSCT enabling studies currently underway with airframers and engine manufacturers, are addressing a broad range of technical, environmental, economic, and related issues. Supersonic cruise speeds for these aircraft were originally focused in the Mach 2 to 5 range. At these flight speeds, both jet fuels and liquid methane were considered potential fuel candidates. For the year 2000 to 2010, cruise Mach numbers of 2 to 3+ are projected for aircraft fuel with thermally stable liquid jet fuels. For 2015 and beyond, liquid methane fueled aircraft cruising at Mach numbers of 4+ may be viable candidates. Operation at supersonic speeds will be much more severe than those encountered at subsonic flight. One of the most critical problems is the potential deterioration of the fuel due to the high temperature environment. HSCT fuels will not only be required to provide the energy necessary for flight, but will also be subject to aerodynamic heating and, will be required to serve as the primary heat sink for cooling the engine and airframe. To define fuel problems for high speed flight, a fuels workshop was conducted at NASA Lewis Research Center. The purpose of the workshop was to gather experts on aviation fuels, airframe fuel systems, airport infrastructure, and combustion systems to discuss high speed fuel alternatives, fuel supply scenarios, increased thermal stability approaches and measurements, safety considerations, and to provide directional guidance for future R and D efforts. Subsequent follow-up studies defined airport infrastructure impacts of high speed fuel candidates. The results of these activities are summarized. In addition, an initial case study using modified in-house refinery simulation model Gordian code (1) is briefly discussed. This code can be used to simulate different types of refineries, emphasizing jet fuel production and relative cost factors.

The NASA Hyper-X Program

NASA Technical Reports Server (NTRS)

Freeman, Delman C., Jr.; Reubush, Daivd E.; McClinton, Charles R.; Rausch, Vincent L.; Crawford, J. Larry

1997-01-01

This paper provides an overview of NASA's Hyper-X Program; a focused hypersonic technology effort designed to move hypersonic, airbreathing vehicle technology from the laboratory environment to the flight environment. This paper presents an overview of the flight test program, research objectives, approach, schedule and status. Substantial experimental database and concept validation have been completed. The program is currently concentrating on the first, Mach 7, vehicle development, verification and validation in preparation for wind-tunnel testing in 1998 and flight testing in 1999. Parallel to this effort the Mach 5 and 10 vehicle designs are being finalized. Detailed analytical and experimental evaluation of the Mach 7 vehicle at the flight conditions is nearing completion, and will provide a database for validation of design methods once flight test data are available.

Selected results of the F-15 propulsion interactions program

NASA Technical Reports Server (NTRS)

Webb, L. D.; Nugent, J.

1982-01-01

A better understanding of propulsion system/airframe flow interactions could aid in the reduction of aircraft drag. For this purpose, NASA and the United States Air Force have conducted a series of wind-tunnel and flight tests on the F-15 airplane. This paper presents a correlation of flight test data from tests conducted at the NASA Dryden Flight Research Facility of the Ames Research Center, with data obtained from wind-tunnel tests. Flights were made at stabilized Mach numbers around 0.6, 0.9, 1.2, and 1.5 with accelerations up to near Mach number 2. Wind-tunnel tests used a 7.5 percent-scale F-15 inlet/airframe model. Flight and wind-tunnel pressure coefficients showed good agreement in most cases. Correlation of interaction effects caused by changes in cowl angle, angle-of-attack, and Mach number are presented. For the afterbody region, the pressure coefficients on the nozzle surfaces were influenced by boattail angles and Mach number. Boundary-layer thickness decreased as angle of attack increased above 4 deg.

Farfield inflight measurement of high-speed turboprop noise

NASA Technical Reports Server (NTRS)

Balombin, J. R.; Loeffler, I. J.

1982-01-01

A flight program was carried out to determine the variation of noise level with distance from a model high speed propeller. Noise measurements were obtained at different distances from a SR-3 propeller mounted on a JetStar aircraft, with the test instrumentation mounted on a Lear jet flown in formation. The propeller was operated at 0.8 flight Mach number, 1.12 helical tip Mach number and at 0.7 flight Mach number, 1.0 helical tip Mach number. The instantaneous pressure from individual blades was observed to rise faster at the 0.8 M flight speed, than at the 0.7 M flight speed. The measured levels appeared to decrease in good agreement with a 6 dB/doubling of distance decay, over the measurement range of approximately 16 m to 100 m distance. Further extrapolation, to the distances represented by a community, would suggest that the propagated levels during cruise would not cause a serious community annoyance.

Aerodynamic Loading Characteristics at Mach Numbers from 0.80 to 1.20 of a 1/10-Scale Three-Stage Scout Model

NASA Technical Reports Server (NTRS)

Kelly, Thomas C.

1961-01-01

Aerodynamic loads results have been obtained in the Langley 8-foot transonic pressure tunnel at Mach numbers from 0.80 to 1.20 for a 1/10-scale model of the upper three stages of the Scout vehicle. Tests were conducted through an angle-of-attack range from -8 deg to 8 deg at an average test Reynolds number per foot of about 4.0 x 10(exp 6). Results indicated that the peak negative pressures associated with expansion corners at the nose and transition flare exhibit sizeable variations which occur over a relatively small Mach number range. The magnitude of the variations may cause the critical local loading condition for the full-scale vehicle to occur at a Mach number considerably lower than that at which the maximum dynamic pressure occurs in flight. The addition of protuberances simulating antennas and wiring conduits had slight, localized effects. The lift carryover from the nose and transition flare on the cylindrical portions of the model generally increased with an increase in Mach number.

Recent Flight Test Results of the Joint CIAM-NASA Mach 6.5 Scramjet Flight Program

NASA Technical Reports Server (NTRS)

Roudakov, Alexander S.; Semenov, Vyacheslav L.; Hicks, John W.

1998-01-01

Under a contract with NASA, a joint Central Institute of Aviation Motors (CIAM) and NASA team recently conducted the fourth flight test of a dual-mode scramjet aboard the CIAM Hypersonic Flying Laboratory, 'Kholod'. With an aim test Mach 6.5 objective, the successful launch was conducted at the Sary Shagan test range in central Kazakstan on February 12, 1998. Ground-launch, rocket boosted by a modified Russian SA5 missile, the redesigned scramjet was accelerated to a new maximum velocity greater than Mach 6.4. This launch allowed for the measurement of the fully supersonic combustion mode under actual flight conditions. The primary program objective was the flight-to-ground correlation of measured data with preflight analysis and wind-tunnel tests in Russia and potentially in the United States. This paper describes the development and objectives of the program as well as the technical details of the scramjet and SA5 redesign to achieve the Mach 6.5 aim test condition. An overview of the launch operation is also given. Finally, preliminary flight test results are presented and discussed.

Flight Test Results from the Rake Airflow Gage Experiment on the F-15B

NASA Technical Reports Server (NTRS)

Frederick, Michael; Ratnayake, Nalin

2011-01-01

The results are described of the Rake Airflow Gage Experiment (RAGE), which was designed and fabricated to support the flight test of a new supersonic inlet design using Dryden's Propulsion Flight Test Fixture (PFTF) and F-15B testbed airplane (see figure). The PFTF is a unique pylon that was developed for flight-testing propulsion-related experiments such as inlets, nozzles, and combustors over a range of subsonic and supersonic flight conditions. The objective of the RAGE program was to quantify the local flowfield at the aerodynamic interface plane of the Channeled Centerbody Inlet Experiment (CCIE). The CCIE is a fixed representation of a conceptual mixed-compression supersonic inlet with a translating biconic centerbody. The primary goal of RAGE was to identify the relationship between free-stream and local Mach number in the low supersonic regime, with emphasis on the identification of the particular free-stream Mach number that produced a local Mach number of 1.5. Measurements of the local flow angularity, total pressure distortion, and dynamic pressure over the interface plane were also desired. The experimental data for the RAGE program were obtained during two separate research flights. During both flights, local flowfield data were obtained during straight and level acceleration segments out to steady-state test points. The data obtained from the two flights showed small variations in Mach number, flow angularity, and dynamic pressure across the interface plane at all flight conditions. The data show that a free-stream Mach number of 1.65 will produce the desired local Mach number of 1.5 for CCIE. The local total pressure distortion over the interface plane at this condition was approximately 1.5%. At this condition, there was an average of nearly 2 of downwash over the interface plane. This small amount of downwash is not expected to adversely affect the performance of the CCIE inlet.

Simulation of Flow Through Breach in Leading Edge at Mach 24

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.; Alter, Stephen J.

2004-01-01

A baseline solution for CFD Point 1 (Mach 24) in the STS-107 accident investigation was modified to include effects of holes through the leading edge into a vented cavity. The simulations were generated relatively quickly and early in the investigation by making simplifications to the leading edge cavity geometry. These simplifications in the breach simulations enabled: 1) A very quick grid generation procedure; 2) High fidelity corroboration of jet physics with internal surface impingements ensuing from a breach through the leading edge, fully coupled to the external shock layer flow at flight conditions. These simulations provided early evidence that the flow through a 2 inch diameter (or larger) breach enters the cavity with significant retention of external flow directionality. A normal jet directed into the cavity was not an appropriate model for these conditions at CFD Point 1 (Mach 24). The breach diameters were of the same order or larger than the local, external boundary-layer thickness. High impingement heating and pressures on the downstream lip of the breach were computed. It is likely that hole shape would evolve as a slot cut in the direction of the external streamlines. In the case of the 6 inch diameter breach the boundary layer is fully ingested.

Supersonic Jet Exhaust Noise at High Subsonic Flight Speed

NASA Technical Reports Server (NTRS)

Norum, Thomas D.; Garber, Donald P.; Golub, Robert A.; Santa Maria, Odilyn L.; Orme, John S.

2004-01-01

An empirical model to predict the effects of flight on the noise from a supersonic transport is developed. This model is based on an analysis of the exhaust jet noise from high subsonic flights of the F-15 ACTIVE Aircraft. Acoustic comparisons previously attainable only in a wind tunnel were accomplished through the control of both flight operations and exhaust nozzle exit diameter. Independent parametric variations of both flight and exhaust jet Mach numbers at given supersonic nozzle pressure ratios enabled excellent correlations to be made for both jet broadband shock noise and jet mixing noise at flight speeds up to Mach 0.8. Shock noise correlated with flight speed and emission angle through a Doppler factor exponent of about 2.6. Mixing noise at all downstream angles was found to correlate well with a jet relative velocity exponent of about 7.3, with deviations from this behavior only at supersonic eddy convection speeds and at very high flight Mach numbers. The acoustic database from the flight test is also provided.

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.

Hyper-X Engine Testing in the NASA Langley 8-Foot High Temperature Tunnel

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Rock, Kenneth E.; Witte, David W.; Ruf, Edward G.; Andrews, Earl H., Jr.

2000-01-01

Airframe-integrated scramjet engine tests have 8 completed at Mach 7 in the NASA Langley 8-Foot High Temperature Tunnel under the Hyper-X program. These tests provided critical engine data as well as design and database verification for the Mach 7 flight tests of the Hyper-X research vehicle (X-43), which will provide the first-ever airframe- integrated scramjet flight data. The first model tested was the Hyper-X Engine Model (HXEM), and the second was the Hyper-X Flight Engine (HXFE). The HXEM, a partial-width, full-height engine that is mounted on an airframe structure to simulate the forebody features of the X-43, was tested to provide data linking flowpath development databases to the complete airframe-integrated three-dimensional flight configuration and to isolate effects of ground testing conditions and techniques. The HXFE, an exact geometric representation of the X-43 scramjet engine mounted on an airframe structure that duplicates the entire three-dimensional propulsion flowpath from the vehicle leading edge to the vehicle base, was tested to verify the complete design as it will be flight tested. This paper presents an overview of these two tests, their importance to the Hyper-X program, and the significance of their contribution to scramjet database development.

F-16XL ship #1 (#849) during first flight of the Digital Flight Control System (DFCS)

NASA Technical Reports Server (NTRS)

1997-01-01

After completing its first flight with the Digital Flight Control System on December 16, 1997, the F-16XL #1 aircraft began a series of envelope expansion flights. On January 27 and 29, 1998, it successfully completed structural clearance tests, as well as most of the load testing Only flights at Mach 1.05 at 10,000 feet, Mach 1.1 at 15,000 feet, and Mach 1.2 at 20,000 feet remained. During the next flight, on February 4, an instrumentation problem cut short the planned envelope expansion tests. After the problem was corrected, the F-16XL returned to flight status, and on February 18 and 20, flight control and evaluation flights were made. Two more research flights were planned for the following week, but another problem appeared. During the ground start up, project personnel noticed that the leading edge flap moved without being commanded. The Digital Flight Control Computer was sent to the Lockheed-Martin facility at Fort Worth, where the problem was traced to a defective chip in the computer. After it was replaced, the F-16XL #1 flew a highly successful flight controls and handling qualities evaluation flight on March 26, clearing the way for the final tests. The final limited loads expansion flight occurred on March 31, and was fully successful. As a result, the on-site Lockheed-Martin loads engineer cleared the aircraft to Mach 1.8. The remaining two handling qualities and flight control evaluation flights were both made on April 3, 1998. These three flights concluded the flight test portion of the DFCS upgrade.

A Concept for the HIFiRE 8 Flight Test

NASA Astrophysics Data System (ADS)

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

2015-09-01

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

Altitude Performance Characteristics of Tail-pipe Burner with Variable-area Exhaust Nozzle

NASA Technical Reports Server (NTRS)

Jansen, Emmert T; Thorman, H Carl

1950-01-01

An investigation was conducted in the NACA Lewis altitude wind tunnel to determine effect of altitude and flight Mach number on performance of tail-pipe burner equipped with variable-area exhaust nozzle and installed on full-scale turbojet engine. At a given flight Mach number, with constant exhaust-gas and turbine-outlet temperatures, increasing altitude lowered the tail-pipe combustion efficiency and raised the specific fuel consumption while the augmented thrust ratio remained approximately constant. At a given altitude, increasing flight Mach number raised the combustion efficiency and augmented thrust ratio and lowered the specific fuel consumption.

Flight measured and calculated exhaust jet conditions for an F100 engine in an F-15 airplane

NASA Technical Reports Server (NTRS)

Hernandez, Francisco J.; Burcham, Frank W., Jr.

1988-01-01

The exhaust jet conditions, in terms of temperature and Mach number, were determined for a nozzle-aft end acoustic study flown on an F-15 aircraft. Jet properties for the F100 EMD engines were calculated using the engine manufacturer's specification deck. The effects of atmospheric temperature on jet Mach number, M10, were calculated. Values of turbine discharge pressure, PT6M, jet Mach number, and jet temperature were calculated as a function of aircraft Mach number, altitude, and power lever angle for the test day conditions. At a typical test point with a Mach number of 0.9, intermediate power setting, and an altitude of 20,000 ft, M10 was equal to 1.63. Flight measured and calculated values of PT6M were compared for intermediate power at altitudes of 15500, 20500, and 31000 ft. It was found that at 31000 ft, there was excellent agreement between both, but for lower altitudes the specification deck overpredicted the flight data. The calculated jet Mach numbers were believed to be accurate to within 2 percent.

Hyper-X Hot Structures Comparison of Thermal Analysis and Flight Data

NASA Technical Reports Server (NTRS)

Amundsen, Ruth M.; Leonard, Charles P.; Bruce, Walter E., III

2004-01-01

The Hyper-X (X-43A) program is a flight experiment to demonstrate scramjet performance and operability under controlled powered free-flight conditions at Mach 7 and 10. The Mach 7 flight was successfully completed on March 27, 2004. Thermocouple instrumentation in the hot structures (nose, horizontal tail, and vertical tail) recorded the flight thermal response of these components. Preflight thermal analysis was performed for design and risk assessment purposes. This paper will present a comparison of the preflight thermal analysis and the recorded flight data.

Flight-determined lag of angle-of-attack and angle-of-sideslip sensors in the YF-12A airplane from analysis of dynamic maneuvers

NASA Technical Reports Server (NTRS)

Gilyard, G. B.; Belte, D.

1974-01-01

Magnitudes of lags in the pneumatic angle-of-attack and angle-of-sideslip sensor systems of the YF-12A airplane were determined for a variety of flight conditions by analyzing stability and control data. The three analysis techniques used are described. An apparent trend with Mach number for measurements from both of the differential-pressure sensors showed that the lag ranged from approximately 0.15 second at subsonic speed to 0.4 second at Mach 3. Because Mach number was closely related to altitude for the available flight data, the individual effects of Mach number and altitude on the lag could not be separated clearly. However, the results indicated the influence of factors other than simple pneumatic lag.

Free-jet testing at Mach 3.44 in GASL's aero/thermo test facility

NASA Technical Reports Server (NTRS)

Cresci, D.; Koontz, S.; Tsai, C. Y.

1993-01-01

A supersonic blow-down tunnel has been used to conduct tests of a hydrogen burning ramjet engine at simulated Mach 3.44 conditions. A pebble-bed type storage heater, a free standing test cabin, and a 48 foot diameter vacuum sphere are used to simulate the flight conditions at nearly matched enthalpy and dynamic pressure. A two dimensional nozzle with a nominal 13.26 inch square exit provides a free-jet test environment. The facility used for these tests is described as are the results of a flow calibration performed on the M = 3.44 nozzle. Some facility/model interactions are discussed as are the eventual hardware modifications and operational procedures required to alleviate the interactions. Some engine test results are discussed briefly to document the success of the test program.

Simulations of 6-DOF Motion with a Cartesian Method

NASA Technical Reports Server (NTRS)

Murman, Scott M.; Aftosmis, Michael J.; Berger, Marsha J.; Kwak, Dochan (Technical Monitor)

2003-01-01

Coupled 6-DOF/CFD trajectory predictions using an automated Cartesian method are demonstrated by simulating a GBU-32/JDAM store separating from an F-18C aircraft. Numerical simulations are performed at two Mach numbers near the sonic speed, and compared with flight-test telemetry and photographic-derived data. Simulation results obtained with a sequential-static series of flow solutions are contrasted with results using a time-dependent flow solver. Both numerical methods show good agreement with the flight-test data through the first half of the simulations. The sequential-static and time-dependent methods diverge over the last half of the trajectory prediction. after the store produces peak angular rates. A cost comparison for the Cartesian method is included, in terms of absolute cost and relative to computing uncoupled 6-DOF trajectories. A detailed description of the 6-DOF method, as well as a verification of its accuracy, is provided in an appendix.

Predicted and flight test results of the performance, stability and control of the space shuttle from reentry to landing

NASA Technical Reports Server (NTRS)

Kirsten, P. W.; Richardson, D. F.; Wilson, C. M.

1983-01-01

Aerodynaic performance, stability and control data obtained from the first five reentries of the Space Shuttle orbiter are given. Flight results are compared to pedicted data from Mach 26.4 to Mach 0.4. Differences between flight and predicted data as well as probable causes for the discrepancies are given.

Flight and wind-tunnel measurements showing base drag reduction provided by a trailing disk for high Reynolds number turbulent flow for subsonic and transonic Mach numbers

NASA Technical Reports Server (NTRS)

Powers, Sheryll Goecke; Huffman, Jarrett K.; Fox, Charles H., Jr.

1986-01-01

The effectiveness of a trailing disk, or trapped vortex concept, in reducing the base drag of a large body of revolution was studied from measurements made both in flight and in a wind tunnel. Pressure data obtained for the flight experiment, and both pressure and force balance data were obtained for the wind tunnel experiment. The flight test also included data obtained from a hemispherical base. The experiment demonstrated the significant base drag reduction capability of the trailing disk to Mach 0.93 and to Reynolds numbers up to 80 times greater than for earlier studies. For the trailing disk data from the flight experiment, the maximum decrease in base drag ranged form 0.08 to 0.07 as Mach number increased from 0.70 to 0.93. Aircraft angles of attack ranged from 3.9 to 6.6 deg for the flight data. For the trailing disk data from the wind tunnel experiment, the maximum decrease in base and total drag ranged from 0.08 to 0.05 for the approximately 0 deg angle of attack data as Mach number increased from 0.30 to 0.82.

Supercomputer modeling of flow past hypersonic flight vehicles

NASA Astrophysics Data System (ADS)

Ermakov, M. K.; Kryukov, I. A.

2017-02-01

A software platform for MPI-based parallel solution of the Navier-Stokes (Euler) equations for viscous heat-conductive compressible perfect gas on 3-D unstructured meshes is developed. The discretization and solution of the Navier-Stokes equations are constructed on generalized S.K. Godunov’s method and the second order approximation in space and time. Developed software platform allows to carry out effectively flow past hypersonic flight vehicles simulations for the Mach numbers 6 and higher, and numerical meshes with up to 1 billion numerical cells and with up to 128 processors.

Numerical Simulations of the Boundary Layer Transition Flight Experiment

NASA Technical Reports Server (NTRS)

Tang, Chun Y.; Trumble, Kerry A.; Campbell, Charles H.; Lessard, Victor R.; Wood, William A.

2010-01-01

Computational Fluid Dynamics (CFD) simulations were used to study the possible effects that the Boundary Layer Transition (BLT) Flight Experiments may have on the heating environment of the Space Shuttle during its entry to Earth. To investigate this issue, hypersonic calculations using the Data-Parallel Line Relaxation (DPLR) and Langley Aerothermodynamic Upwind Relaxation (LAURA) CFD codes were computed for a 0.75 tall protuberance at flight conditions of Mach 15 and 18. These initial results showed high surface heating on the BLT trip and the areas surrounding the protuberance. Since the predicted peak heating rates would exceed the thermal limits of the materials selected to construct the BLT trip, many changes to the geometry were attempted in order to reduce the surface heat flux. The following paper describes the various geometry revisions and the resulting heating environments predicted by the CFD codes.

The Viking parachute qualification test technique.

NASA Technical Reports Server (NTRS)

Raper, J. L.; Lundstrom, R. R.; Michel, F. C.

1973-01-01

The parachute system for NASA's Viking '75 Mars lander was flight qualified in four high-altitude flight tests at the White Sands Missile range (WSMR). A balloon system lifted a full-scale simulated Viking spacecraft to an altitude where a varying number of rocket motors were used to propel the high drag, lifting test vehicle to test conditions which would simulate the range of entry conditions expected at Mars. A ground-commanded cold gas pointing system located on the balloon system provided powered vehicle azimuth control to insure that the flight trajectory remained within the WSMR boundaries. A unique ground-based computer-radar system was employed to monitor inflight performance of the powered vehicle and insure that command ignition of the parachute mortar occurred at the required test conditions of Mach number and dynamic pressure. Performance data were obtained from cameras, telemetry, and radar.

X-2 on ramp with B-50 mothership and support crew

NASA Technical Reports Server (NTRS)

1956-01-01

Air Force test pilot Capt. Iven Kincheloe stands in front of the Bell X-2 (46-674) on the ramp at Edwards Air Force Base, California. Behind the X-2 are ground support personnel, the B-50 launch aircraft and crew, chase planes, and support vehicles. Kincheloe had flown nearly 100 combat missions in Korea in an F-86 and was credited with shooting down 10 enemy aircraft. He then graduated from the Empire Test Pilot's School in Great Britain in December 1954, whereupon he was assigned to Edwards Air Force Base. He made four powered flights in the X-2. On September 7, 1956, he reached an altitude of 126,200 feet. After the death of Capt. Mel Apt and the loss of the X-2 #1 on September 27, 1956, in the first Mach 3 flight, Kincheloe was assigned as the Air Force project pilot for the X-15. Before he had a chance to fly that rocket-powered aircraft, Kincheloe himself lost his life on July 26, 1958, in an F-104 accident. The X-2 was a swept-wing, rocket-powered aircraft designed to fly faster than Mach 3 (three times the speed of sound). It was built for the U.S. Air Force by the Bell Aircraft Company, Buffalo, New York. The X-2 was flown to investigate the problems of aerodynamic heating as well as stability and control effectiveness at high altitudes and high speeds (in excess of Mach 3). Bell aircraft built two X-2 aircraft. These were constructed of K-monel (a copper and nickel alloy) for the fuselage and stainless steel for the swept wings and control surfaces. The aircraft had ejectable nose capsules instead of ejection seats because the development of ejection seats had not reached maturity at the time the X-2 was conceived. The X-2 ejection canopy was successfully tested using a German V-2 rocket. The X-2 used a skid-type landing gear to make room for more fuel. The airplane was air launched from a modified Boeing B-50 Superfortress Bomber. X-2 Number 1 made its first unpowered glide flight on Aug. 5, 1954, and made a total of 17 (4 glide and 13 powered) flights before it was lost Sept. 27, 1956. The pilot on Flight 17, Capt. Milburn Apt, had flown the aircraft to a record speed of Mach 3.2 (2,094 mph), thus becoming the first person to exceed Mach 3. During that last flight, inertial coupling occurred and the pilot was killed. The aircraft suffered little damage in the crash, resulting in proposals (never implemented) from the Langley Memorial Aeronautical Laboratory, Hampton, Virginia, to rebuild it for use in a hypersonic (Mach 5+) test program. In 1953, X-2 Number 2 was lost in an in-flight explosion while at the Bell Aircraft Company during captive flight trials and was jettisoned into Lake Ontario. The Air Force had previously flown the aircraft on three glide flights at Edwards Air Force Base, California, in 1952. Although the NACA's High-Speed Flight Station, Edwards, California, (predecessor of NASA's Dryden Flight Research Center) never actually flew the X-2 aircraft, the NACA did support the program primarily through Langley Memorial Aeronautical Laboratory wind-tunnel tests and Wallops Island, Virginia, rocket-model tests. The NACA High-Speed Flight Station also provided stability and control recording instrumentation and simulator support for the Air Force flights. In the latter regard, the NACA worked with the Air Force in using a special computer to extrapolate and predict aircraft behavior from flight data.

A simple dynamic engine model for use in a real-time aircraft simulation with thrust vectoring

NASA Technical Reports Server (NTRS)

Johnson, Steven A.

1990-01-01

A simple dynamic engine model was developed at the NASA Ames Research Center, Dryden Flight Research Facility, for use in thrust vectoring control law development and real-time aircraft simulation. The simple dynamic engine model of the F404-GE-400 engine (General Electric, Lynn, Massachusetts) operates within the aircraft simulator. It was developed using tabular data generated from a complete nonlinear dynamic engine model supplied by the manufacturer. Engine dynamics were simulated using a throttle rate limiter and low-pass filter. Included is a description of a method to account for axial thrust loss resulting from thrust vectoring. In addition, the development of the simple dynamic engine model and its incorporation into the F-18 high alpha research vehicle (HARV) thrust vectoring simulation. The simple dynamic engine model was evaluated at Mach 0.2, 35,000 ft altitude and at Mach 0.7, 35,000 ft altitude. The simple dynamic engine model is within 3 percent of the steady state response, and within 25 percent of the transient response of the complete nonlinear dynamic engine model.

Comparative Flow Path Analysis and Design Assessment of an Axisymmetric Hydrogen Fueled Scramjet Flight Test Engine at a Mach Number of 6.5

NASA Technical Reports Server (NTRS)

McClinton, C.; Rondakov, A.; Semenov, V.; Kopehenov, V.

1991-01-01

NASA has contracted with the Central Institute of Aviation Motors CIAM to perform a flight test and ground test and provide a scramjet engine for ground test in the United States. The objective of this contract is to obtain ground to flight correlation for a supersonic combustion ramjet (scramjet) engine operating point at a Mach number of 6.5. This paper presents results from a flow path performance and thermal evaluation performed on the design proposed by the CIAM. This study shows that the engine will perform in the scramjet mode for stoichiometric operation at a flight Mach number of 6.5. Thermal assessment of the structure indicates that the combustor cooling liner will provide adequate cooling for a Mach number of 6.5 test condition and that optional material proposed by CIAM for the cowl leading-edge design are required to allow operation with or without a type IV shock-shock interaction.

Aerodynamic Parameter Estimation for the X-43A (Hyper-X) from Flight Data

NASA Technical Reports Server (NTRS)

Morelli, Eugene A.; Derry, Stephen D.; Smith, Mark S.

2005-01-01

Aerodynamic parameters were estimated based on flight data from the third flight of the X-43A hypersonic research vehicle, also called Hyper-X. Maneuvers were flown using multiple orthogonal phase-optimized sweep inputs applied as simultaneous control surface perturbations at Mach 8, 7, 6, 5, 4, and 3 during the vehicle descent. Aerodynamic parameters, consisting of non-dimensional longitudinal and lateral stability and control derivatives, were estimated from flight data at each Mach number. Multi-step inputs at nearly the same flight conditions were also flown to assess the prediction capability of the identified models. Prediction errors were found to be comparable in magnitude to the modeling errors, which indicates accurate modeling. Aerodynamic parameter estimates were plotted as a function of Mach number, and compared with estimates from the pre-flight aerodynamic database, which was based on wind-tunnel tests and computational fluid dynamics. Agreement between flight estimates and values computed from the aerodynamic database was excellent overall.

Copy Right for Flight: Patterns of Technological Adaptation in Military Aviation

DTIC Science & Technology

2012-06-01

Heppenheimer , A Brief History of Flight : From Balloons to Mach 3 and Beyond (New York: Wiley, 2001), 238. 4 Ronald E. Miller and David Sawers...Technical Development of Modern Aviation (London,: Routledge & K. Paul, 1968), 159. 7 T. A . Heppenheimer , A Brief History of Flight : From Balloons to...Modern Aviation (London,: Routledge & K. Paul, 1968), 165. 10 T. A . Heppenheimer , A Brief History of Flight : From Balloons to Mach 3 and Beyond (New

CIAM/NASA Mach 6.5 Scramjet Flight and Ground Test

NASA Technical Reports Server (NTRS)

Voland, R. T.; Auslender, A. H.; Smart, M. K.; Roudakov, A. S.; Semenov, V. L.; Kopchenov, V.

1999-01-01

The Russian Central Institute of Aviation Motors (CIAM) performed a flight test of a CIAM-designed, hydrogen-cooled/fueled dual-mode scramjet engine over a Mach number range of approximately 3.5 to 6.4 on February 12, 1998, at the Sary Shagan test range in Kazakhstan. This rocket-boosted, captive-carry test of the axisymmetric engine reached the highest Mach number of any scramjet engine flight test to date. The flight test and the accompanying ground test program, conducted in a CIAM test facility near Moscow, were performed under a NASA contract administered by the Dryden Flight Research Center with technical assistance from the Langley Research Center. Analysis of the flight and ground data by both CIAM and NASA resulted in the following preliminary conclusions. An unexpected control sensor reading caused non-optimal fueling of the engine, and flowpath modifications added to the engine inlet during manufacture caused markedly reduced inlet performance. Both of these factors appear to have contributed to the dual-mode scramjet engine operating primarily in a subsonic combustion mode. At the maximum Mach number test point, combustion caused transition from supersonic flow at the fuel injector station to primarily subsonic flow in the combustor. Ground test data were obtained at similar conditions to the flight test, allowing for a meaningful comparison between the ground and flight data. The results of this comparison indicate that the differences in engine performance are small.

Hypersonic lateral and directional stability characteristics of aeroassist flight experiment configuration in air and CF4

NASA Technical Reports Server (NTRS)

Micol, John R.; Wells, William L.

1993-01-01

Hypersonic lateral and directional stability characteristics measured on a 60 deg half-angle elliptical cone, which was raked at an angle of 73 deg from the cone centerline and with an ellipsoid nose (ellipticity equal to 2.0 in the symmetry plane), are presented for angles of attack from -10 to 10 deg. The high normal-shock density ratio of a real gas was simulated by tests at a Mach number of 6 in air and CF4 (density ratio equal to 5.25 and 12.0, respectively). Tests were conducted in air at Mach 6 and 10 and in CF4 at Mach 6 to examine the effects of Mach number, Reynolds number, and normal-shock density ratio. Changes in Mach number from 6 to 10 in air or in Reynolds number by a factor of 4 at Mach 6 had a negligible effect on lateral and directional stability characteristics. Variations in normal-shock density ratio had a measurable effect on lateral and directional aerodynamic coefficients, but no significant effect on lateral and directional stability characteristics. Tests in air and CF4 indicated that the configuration was laterally and directionally stable through the test range of angle of attack.

Lessons Learned from Numerical Simulations of the F-16XL Aircraft at Flight Conditions

NASA Technical Reports Server (NTRS)

Rizzi, Arthur; Jirasek, Adam; Lamar, John; Crippa, Simone; Badcock, Kenneth; Boelens, Oklo

2009-01-01

Nine groups participating in the Cranked Arrow Wing Aerodynamics Project International (CAWAPI) project have contributed steady and unsteady viscous simulations of a full-scale, semi-span model of the F-16XL aircraft. Three different categories of flight Reynolds/Mach number combinations were computed and compared with flight-test measurements for the purpose of code validation and improved understanding of the flight physics. Steady-state simulations are done with several turbulence models of different complexity with no topology information required and which overcome Boussinesq-assumption problems in vortical flows. Detached-eddy simulation (DES) and its successor delayed detached-eddy simulation (DDES) have been used to compute the time accurate flow development. Common structured and unstructured grids as well as individually-adapted unstructured grids were used. Although discrepancies are observed in the comparisons, overall reasonable agreement is demonstrated for surface pressure distribution, local skin friction and boundary velocity profiles at subsonic speeds. The physical modeling, steady or unsteady, and the grid resolution both contribute to the discrepancies observed in the comparisons with flight data, but at this time it cannot be determined how much each part contributes to the whole. Overall it can be said that the technology readiness of CFD-simulation technology for the study of vehicle performance has matured since 2001 such that it can be used today with a reasonable level of confidence for complex configurations.

High-speed flight propulsion systems. Progress in Astronautics and Aeronautics. Vol. 137

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

Murthy, S.N.B.; Curran, E.T.

1991-01-01

Various papers on high-speed flight propulsion systems are presented. The topics addressed are: propulsion systems from takeoff to high-speed flight, propulsion system performance and integration for high Mach air-breathing flight, energy analysis of high-speed flight systems, waves and thermodynamics in high Mach number propulsive ducts, turbulent free shear layer mixing and combustion, turbulent mixing in supersonic combustion systems, mixing and mixing enhancement in supersonic reacting flowfields, study of combustion and heat-exchange processes in high-enthalpy short-duration facilities, and facility requirements for hypersonic propulsion system testing.

X-38 Vehicle 131R Free Flights 1 and 2

NASA Technical Reports Server (NTRS)

Munday, Steve

2000-01-01

The X-38 program is using a modern flight control system (FCS) architecture originally developed by Honeywell called MACH. During last year's SAE G&C subcommittee meeting, we outlined the design, implementation and testing of MACH in X-38 Vehicles 132, 131R & 201. During this year's SAE meeting, I'll focus upon the first two free flights of V131R, describing what caused the roll-over in FF1 and how we fixed it for FF2. I only have 30 minutes, so it will be a quick summary including VHS video. X-38 is a NASA JSC/DFRC experimental flight test program developing a series of prototypes for an International Space Station (ISS) Crew Return Vehicle (CRV), often described as an ISS "lifeboat." X-38 Vehicle 132 Free Flight 3 was the first flight test of a modern FCS architecture called Multi-Application ControlH (MACH), developed by the Honeywell Technology Center in Minneapolis and Honeywell's Houston Engineering Center. MACH wraps classical Proportional+integral (P+I) outer attitude loops around modern dynamic inversion attitude rate loops. The presentation at last year's SAE Aerospace Meeting No. 85 focused upon the design and testing of the FCS algorithm and Vehicle 132 Free Flight 3. This presentation will summarize flight control and aerodynamics lessons learned during Free Flights 1 and 2 of Vehicle 131R, a subsonic test vehicle laying the groundwork for the orbital/entry test of Vehicle 201 in 2003.

Hypersonic vehicle simulation model: Winged-cone configuration

NASA Technical Reports Server (NTRS)

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

1990-01-01

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

Instrumentation and Performance Analysis Plans for the HIFiRE Flight 2 Experiment

NASA Technical Reports Server (NTRS)

Gruber, Mark; Barhorst, Todd; Jackson, Kevin; Eklund, Dean; Hass, Neal; Storch, Andrea M.; Liu, Jiwen

2009-01-01

Supersonic combustion performance of a bi-component gaseous hydrocarbon fuel mixture is one of the primary aspects under investigation in the HIFiRE Flight 2 experiment. In-flight instrumentation and post-test analyses will be two key elements used to determine the combustion performance. Pre-flight computational fluid dynamics (CFD) analyses provide valuable information that can be used to optimize the placement of a constrained set of wall pressure instrumentation in the experiment. The simulations also allow pre-flight assessments of performance sensitivities leading to estimates of overall uncertainty in the determination of combustion efficiency. Based on the pre-flight CFD results, 128 wall pressure sensors have been located throughout the isolator/combustor flowpath to minimize the error in determining the wall pressure force at Mach 8 flight conditions. Also, sensitivity analyses show that mass capture and combustor exit stream thrust are the two primary contributors to uncertainty in combustion efficiency.

Development of a short length combustor for a supersonic cruise turbofan engine using a 90 deg sector of a full annulus

NASA Technical Reports Server (NTRS)

Clements, T. R.

1972-01-01

A performance development program has been conducted on a short length, double-annular, ram-induction combustor. The combustor was designed for a large augmented turbofan engine capable of sustained flight speeds up to Mach 3.0. Performance tests were conducted at an inlet temperature and Mach number simulating engine sea level takeoff conditions. At the design temperature rise of 1600 F, combustion efficiency was 100%, pattern factor was 0.20, and combined diffuser-combustor pressure loss was 4.4% or 1.12 times the diffuser inlet velocity head. A temperature rise in excess of 2400 F with a combustion efficiency of 94% was demonstrated.

Space shuttle plume/simulation application: Results and math model supersonic data

NASA Technical Reports Server (NTRS)

Boyle, W.; Conine, B.; Bell, G.

1979-01-01

The analysis of pressure and gage wind tunnel data from space shuttle wind tunnel test IA138 was performed to define the aerodynamic influence of the main propulsion system and solid rocket booster plumes on the total vehicles, elements, and components of the space shuttle vehicle during the supersonic portion of ascent flight. A math model of the plume induced aerodynamic characteristics was developed for a range of Mach numbers to match the forebody aerodynamic math model. The base aerodynamic characteristics are presented in terms of forces and moments versus attitude. Total vehicle base and forebody aerodynamic characteristics are presented in terms of aerodynamic coefficients for Mach numbers from 1.55 to 2.5.

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

NASA Technical Reports Server (NTRS)

Axdahl, Erik L.

2013-01-01

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

Effects of Wing Leading Edge Penetration with Venting and Exhaust Flow from Wheel Well at Mach 24 in Flight

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.

2003-01-01

A baseline solution for CFD Point 1 (Mach 24) in the STS-107 accident investigation was modified to include effects of: (1) holes through the leading edge into a vented cavity; and (2) a scarfed, conical nozzle directed toward the centerline of the vehicle from the forward, inboard corner of the landing gear door. The simulations were generated relatively quickly and early in the investigation because simplifications were made to the leading edge cavity geometry and an existing utility to merge scarfed nozzle grid domains with structured baseline external domains was implemented. These simplifications in the breach simulations enabled: (1) a very quick grid generation procedure; and (2) high fidelity corroboration of jet physics with internal surface impingements ensuing from a breach through the leading edge, fully coupled to the external shock layer flow at flight conditions. These simulations provided early evidence that the flow through a two-inch diameter (or larger) breach enters the cavity with significant retention of external flow directionality. A normal jet directed into the cavity was not an appropriate model for these conditions at CFD Point 1 (Mach 24). The breach diameters were of the same order or larger than the local, external boundary-layer thickness. High impingement heating and pressures on the downstream lip of the breach were computed. It is likely that hole shape would evolve as a slot cut in the direction of the external streamlines. In the case of the six-inch diameter breach the boundary layer is fully ingested. The intent of externally directed jet simulations in the second scenario was to approximately model aerodynamic effects of a relatively large internal wing pressure, fueled by combusting aluminum, which deforms the corner of the landing gear door and directs a jet across the windside surface. These jet interactions, in and of themselves, were not sufficiently large to explain observed aerodynamic behavior.

Numerical Solutions for the CAWAPI Configuration on Structured Grids at NASA LaRC, United States. Chapter 7

NASA Technical Reports Server (NTRS)

Elmiligui, Alaa A.; Abdol-Hamid, Khaled S.; Massey, Steven J.

2009-01-01

In this chapter numerical simulations of the flow around F-16XL are performed as a contribution to the Cranked Arrow Wing Aerodynamic Project International (CAWAPI) using the PAB3D CFD code. Two turbulence models are used in the calculations: a standard k-epsilon model, and the Shih-Zhu-Lumley (SZL) algebraic stress model. Seven flight conditions are simulated for the flow around the F-16XL where the free stream Mach number varies from 0.242 to 0.97. The range of angles of attack varies from 0 deg to 20 deg. Computational results, surface static pressure, boundary layer velocity profiles, and skin friction are presented and compared with flight data. Numerical results are generally in good agreement with flight data, considering that only one grid resolution is utilized for the different flight conditions simulated in this study. The Algebraic Stress Model (ASM) results are closer to the flight data than the k-epsilon model results. The ASM predicted a stronger primary vortex, however, the origin of the vortex and footprint is approximately the same as in the k-epsilon predictions.

Traveling Crossow Instability for HIFiRE-5 in a Quiet Hypersonic Wind Tunnel (Postprint)

DTIC Science & Technology

2013-06-01

scale model of the 2:1 elliptic cone HIFiRE-5 flight vehicle was used to investigate the traveling crossflow instability at Mach 6 in Purdue...Force Research Laboratory, Air Vehicles Directorate 2130 8th St., WPAFB, OH 45433-7542, USA Abstract A scale model of the 2:1 elliptic cone HIFiRE-5...flight vehicle was used to investigate the traveling crossflow instability at Mach 6 in Purdue University’s Mach-6 quiet wind tunnel. Traveling crossflow

Developing and flight testing the HL-10 lifting body: A precursor to the Space Shuttle

NASA Technical Reports Server (NTRS)

Kempel, Robert W.; Painter, Weneth D.; Thompson, Milton O.

1994-01-01

The origins of the lifting-body idea are traced back to the mid-1950's, when the concept of a manned satellite reentering the Earth's atmosphere in the form of a wingless lifting body was first proposed. The advantages of low reentry deceleration loads, range capability, and horizontal landing of a lifting reentry vehicle (as compared with the high deceleration loads and parachute landing of a capsule) are presented. The evolution of the hypersonic HL-10 lifting body is reviewed from the theoretical design and development process to its selection as one of two low-speed flight vehicles for fabrication and piloted flight testing. The design, development, and flight testing of the low-speed, air-launched, rocket-powered HL-10 was part of an unprecedented NASA and contractor effort. NASA Langley Research Center conceived and developed the vehicle shape and conducted numerous theoretical, experimental, and wind-tunnel studies. NASA Flight Research Center (now NASA Dryden Flight Research Center) was responsible for final low-speed (Mach numbers less than 2.0) aerodynamic analysis, piloted simulation, control law development, and flight tests. The prime contractor, Northrop Corp., was responsible for hardware design, fabrication, and integration. Interesting and unusual events in the flight testing are presented with a review of significant problems encountered in the first flight and how they were solved. Impressions by the pilots who flew the HL-10 are included. The HL-10 completed a successful 37-flight program, achieved the highest Mach number and altitude of this class vehicle, and contributed to the technology base used to develop the space shuttle and future generations of lifting bodies.

Preliminary Results of Altitude-Wind-Tunnel Investigation of X24C-4B Turbojet Engine. IV - Performance of Modified Compressor. Part 4; Performance of Modified Compressor

NASA Technical Reports Server (NTRS)

Thorman, H. Carl; Dupree, David T.

1947-01-01

The performance of the 11-stage axial-flow compressor, modified to improve the compressor-outlet velocity, in a revised X24C-4B turbojet engine is presented and compared with the performance of the compressor in the original engine. Performance data were obtained from an investigation of the revised engine in the MACA Cleveland altitude wind tunnel. Compressor performance data were obtained for engine operation with four exhaust nozzles of different outlet area at simulated altitudes from 15,OOO to 45,000 feet, simulated flight Mach numbers from 0.24 to 1.07, and engine speeds from 4000 to 12,500 rpm. The data cover a range of corrected engine speeds from 4100 to 13,500 rpm, which correspond to compressor Mach numbers from 0.30 to 1.00.

Application of Spontaneous Raman Scattering to the Flowfield in a Scramjet Combustor

NASA Astrophysics Data System (ADS)

Sander, T.; Sattelmayer, T.

2002-07-01

For the investigation of the ignition and reaction of fuel injected into the combustor of a Scramjet at a flight Mach number of 8 high temperature test air at supersonic speed is required. One economic possibility to simulate these inlet conditions experimentally is the use of vitiators which preheat the air by the burning of hydrogen. Downstream of the precombustor the flow is accelerated in a Laval nozzle to a Mach number of 2.15 and enters the combustor. For the numerical simulation of a supersonic reacting flow precise information concerning the physical properties during ignition and reaction are required. Optical measurements are best suited for delivering this information as they do not disturb the supersonic flow like probes and as their application is not limited by thermal stress. Raman scattering offers the possibility of measuring the static temperature and the concentration of majority species.

Flight-Test-Determined Aerodynamic Force and Moment Characteristics of the X-43A at Mach 7.0

NASA Technical Reports Server (NTRS)

Davis. Marl C.; White, J. Terry

2006-01-01

The second flight of the Hyper-X program afforded a unique opportunity to determine the aerodynamic force and moment characteristics of an airframe-integrated scramjet-powered aircraft in hypersonic flight. These data were gathered via a repeated series of pitch, yaw, and roll doublets; frequency sweeps; and pushover-pullup maneuvers performed throughout the X-43A cowl-closed descent. Maneuvers were conducted at Mach numbers of 6.80 to 0.95 and altitudes from 92,000 ft msl to sea level. The dynamic pressure varied from 1300 psf to 400 psf with the angle of attack ranging from 0 deg to 14 deg. The flight-extracted aerodynamics were compared with preflight predictions based on wind-tunnel-test data. The X-43A flight-derived axial force was found to be 10 percent to 15 percent higher than prediction. Under-predictions of similar magnitude were observed for the normal force. For Mach numbers above 4.0, the flight-derived stability and control characteristics resulted in larger-than-predicted static margins, with the largest discrepancy approximately 5 in. forward along the x-axis center of gravity at Mach 6.0. This condition would result in less static margin in pitch. The predicted lateral-directional stability and control characteristics matched well with flight data when allowance was made for the high uncertainty in angle of sideslip.

X-43A Flight-Test-Determined Aerodynamic Force and Moment Characteristics at Mach 7.0

NASA Technical Reports Server (NTRS)

Davis, Mark C.; White, J. Terry

2008-01-01

The second flight of the Hyper-X program afforded a unique opportunity to determine the aerodynamic force and moment characteristics of an airframe-integrated scramjet-powered aircraft in hypersonic flight. These data were gathered via a repeated series of pitch, yaw, and roll doublets, frequency sweeps, and pushover-pullup maneuvers performed throughout the X-43A cowl-closed descent. Maneuvers were conducted at Mach numbers of 6.80-0.95 and at altitudes from 92,000 ft mean sea level to sea level. The dynamic pressure varied from 1300 to 400 psf with the angle of attack ranging from 0 to 14 deg. The flight-extracted aerodynamics were compared with preflight predictions based on wind-tunnel test data. The X-43A flight-derived axial force was found to be 10-15%higher than prediction. Underpredictions of similar magnitude were observed for the normal force. For Mach numbers above 4.0, the flight-derived stability and control characteristics resulted in larger-than-predicted static margins, with the largest discrepancy approximately 5 in. forward along the x-axis center of gravity at Mach 6.0. This condition would result in less static margin in pitch. The predicted lateral-directional stability and control characteristics matched well with flight data when allowance was made for the high uncertainty in angle of sideslip.

The Road to Mach 10: A History of the X-43A Hypersonic Flight Test Program at NASA Dryden -- Origins to First Flight

NASA Technical Reports Server (NTRS)

Peebles, Curtis

2006-01-01

The NASA Dryden Flight Research Center, in partnership with the NASA Langley Research Center and industrial contractors, conducted the first flight tests of a supersonic combustion ramjet (scramjet) in 2004. This was a revolutionary airbreathing engine able to operate at speeds above Mach 5, which carries potential for both high-speed atmospheric flight and as a space launcher. For the Dryden engineers, the X-43 program was the culmination of a nearly 60-year history of flight research, going back to the early days of supersonic flight, and to rocket planes such as the X-1, D-558-II Skyrocket, and the X-15. For the propulsion community, it marked a turning point in a quest that had taken nearly as long. The scramjet engine did not arise from the work of a single individual or from a single technological breakthrough. It evolved instead from work under way on ramjets in the early 1950s, and from research programs at the National Advisory Committee for Aeronautics (NACA) Lewis Research Center, at the U.S. Army Aberdeen Proving Ground, and by the U.S. Navy. Studies developed in the course of these disparate projects raised the possibility of supersonic combustion. Many researchers had considered the notion impractical due to the difficulty of stabilizing a flame front in a supersonic airflow. NACA researchers at Lewis attempted to test the idea's feasibility by burning aluminum borohydride in a supersonic wind tunnel. Sustained burning was believed to have been observed at Mach 1.5, Mach 2, and Mach 3 for as long as two seconds.

Numerically simulated comparative performance of a scramjet and shcramjet at Mach 11

NASA Astrophysics Data System (ADS)

Chan, Jonathan

This study investigates the design and aeropropulsive performance of a complete, hydrogen powered, shock-induced combustion ramjet (shcramjet) at a flight Mach number of 11 and altitude of 34.5 km. The design includes a Prandtl-Meyer compression inlet, cantilevered ramp fuel injectors, a shock-inducing wedge and a divergent nozzle. Numerical studies are undertaken using the WARP code that solves the three-dimensional Favre-averaged Navier-Stokes equations closed by the Wilcox k-o turbulence model and the Jachimowski H2/air chemical kinetics model. Studies of fuel injection properties, mixing duct length, combustor wedge and nozzle geometry are completed to maximize the overall performance of the vehicle. The final shcramjet configuration generates a specific impulse of 1110 s. A comparison is undertaken with a scramjet vehicle at identical flight conditions and using many of the same components. The comparable scramjet generates a higher specific impulse of 1450 s although it is significantly larger and therefore heavier.

Development of a Flush Airdata Sensing System on a Sharp-Nosed Vehicle for Flight at Mach 3 to 8

NASA Technical Reports Server (NTRS)

Davis, Mark C.; Pahle, Joseph W.; White, John Terry; Marshall, Laurie A.; Mashburn, Michael J.; Franks, Rick

2000-01-01

NASA Dryden Flight Research Center has developed a flush airdata sensing (FADS) system on a sharp-nosed, wedge-shaped vehicle. This paper details the design and calibration of a real-time angle-of-attack estimation scheme developed to meet the onboard airdata measurement requirements for a research vehicle equipped with a supersonic-combustion ramjet engine. The FADS system has been designed to perform in flights at speeds between Mach 3 and Mach 8 and at angles of attack between -6 deg. and 12 deg. The description of the FADS architecture includes port layout, pneumatic design, and hardware integration. Predictive models of static and dynamic performance are compared with wind-tunnel results across the Mach and angle-of-attack range. Results indicate that static angle-of-attack accuracy and pneumatic lag can be adequately characterized and incorporated into a real-time algorithm.

Preliminary Results of Altitude-Wind-Tunnel Investigation of X24C-4B Turbojet Engine. I - Pressure and Temperature Distributions

NASA Technical Reports Server (NTRS)

Prince, William R.; Hawkins, W. Kent

1947-01-01

Pressures and temperatures throughout the X24C-4B turbojet engine are presented in both tabular and graphical forms to show the effect of altitude, flight Mach number, and engine speed on the internal operation of the engine. These data were obtained in the NACA Cleveland altitude wind tunnel at simulated altitudes from 5000 to 45,000 feet, simulated flight Mach numbers from 0.25 to 1.08, and engine speeds from 4000 to 12,500 rpm. Location and detail drawings of the instrumentation installed at seven survey stations in the engine are shown. Application of generalization factors to pressures and temperatures at each measuring station for the range of altitudes investigated showed that the data did not generalize above an altitude of 25,000 feet. Total-pressure distribution at the compressor outlet varied only with change in engine speed. At altitudes above 35,000 feet and engine speeds above 11,000 rpm, the peak temperature at the turbine-outlet annulus moved inward toward the root of the blade, which is undesirable from blade-stress considerations. The temperature levels at the turbine outlet and the exhaust-nozzle outlet were lowered as the Mach number was increased. The static-pressure measurements obtained at each stator stage of the compressor showed a pressure drop through the inlet guide vanes and the first-stage rotor at high engine speeds. The average values measured by the manufacturer's instrumentation werein close agreement with the average values obtained with NACA instrumentation.

Wind-tunnel/flight correlation study of aerodynamic characteristics of a large flexible supersonic cruise airplane (XB-70-1). 3: A comparison between characteristics predicted from wind-tunnel measurements and those measured in flight

NASA Technical Reports Server (NTRS)

Arnaiz, H. H.; Peterson, J. B., Jr.; Daugherty, J. C.

1980-01-01

A program was undertaken by NASA to evaluate the accuracy of a method for predicting the aerodynamic characteristics of large supersonic cruise airplanes. This program compared predicted and flight-measured lift, drag, angle of attack, and control surface deflection for the XB-70-1 airplane for 14 flight conditions with a Mach number range from 0.76 to 2.56. The predictions were derived from the wind-tunnel test data of a 0.03-scale model of the XB-70-1 airplane fabricated to represent the aeroelastically deformed shape at a 2.5 Mach number cruise condition. Corrections for shape variations at the other Mach numbers were included in the prediction. For most cases, differences between predicted and measured values were within the accuracy of the comparison. However, there were significant differences at transonic Mach numbers. At a Mach number of 1.06 differences were as large as 27 percent in the drag coefficients and 20 deg in the elevator deflections. A brief analysis indicated that a significant part of the difference between drag coefficients was due to the incorrect prediction of the control surface deflection required to trim the airplane.

Development of an advanced pitch active control system for a wide body jet aircraft

NASA Technical Reports Server (NTRS)

Guinn, Wiley A.; Rising, Jerry J.; Davis, Walt J.

1984-01-01

An advanced PACS control law was developed for a commercial wide-body transport (Lockheed L-1011) by using modern control theory. Validity of the control law was demonstrated by piloted flight simulation tests on the NASA Langley visual motion simulator. The PACS design objective was to develop a PACS that would provide good flying qualities to negative 10 percent static stability margins that were equivalent to those of the baseline aircraft at a 15 percent static stability margin which is normal for the L-1011. Also, the PACS was to compensate for high-Mach/high-g instabilities that degrade flying qualities during upset recoveries and maneuvers. The piloted flight simulation tests showed that the PACS met the design objectives. The simulation demonstrated good flying qualities to negative 20 percent static stability margins for hold, cruise and high-speed flight conditions. Analysis and wind tunnel tests performed on other Lockheed programs indicate that the PACS could be used on an advanced transport configuration to provide a 4 percent fuel savings which results from reduced trim drag by flying at negative static stability margins.

Flight measurements of surface pressures on a flexible supercritical research wing

NASA Technical Reports Server (NTRS)

Eckstrom, C. V.

1985-01-01

A flexible supercritical research wing, designated as ARW-1, was flight-tested as part of the NASA Drones for Aerodynamic and Structural Testing Program. Aerodynamic loads, in the form of wing surface pressure measurements, were obtained during flights at altitudes of 15,000, 20,000, and 25,000 feet at Mach numbers from 0.70 to 0.91. Surface pressure coefficients determined from pressure measurements at 80 orifice locations are presented individually as nearly continuous functions of angle of attack for constant values of Mach number. The surface pressure coefficients are also presented individually as a function of Mach number for an angle of attack of 2.0 deg. The nearly continuous values of the pressure coefficient clearly show details of the pressure gradient, which occurred in a rather narrow Mach number range. The effects of changes in angle of attack, Mach number, and dynamic pressure are also shown by chordwise pressure distributions for the range of test conditions experienced. Reynolds numbers for the tests ranged from 5.7 to 8.4 x 1,000,000.

Skylon Aerodynamics and SABRE Plumes

NASA Technical Reports Server (NTRS)

Mehta, Unmeel; Afosmis, Michael; Bowles, Jeffrey; Pandya, Shishir

2015-01-01

An independent partial assessment is provided of the technical viability of the Skylon aerospace plane concept, developed by Reaction Engines Limited (REL). The objectives are to verify REL's engineering estimates of airframe aerodynamics during powered flight and to assess the impact of Synergetic Air-Breathing Rocket Engine (SABRE) plumes on the aft fuselage. Pressure lift and drag coefficients derived from simulations conducted with Euler equations for unpowered flight compare very well with those REL computed with engineering methods. The REL coefficients for powered flight are increasingly less acceptable as the freestream Mach number is increased beyond 8.5, because the engineering estimates did not account for the increasing favorable (in terms of drag and lift coefficients) effect of underexpanded rocket engine plumes on the aft fuselage. At Mach numbers greater than 8.5, the thermal environment around the aft fuselage is a known unknown-a potential design and/or performance risk issue. The adverse effects of shock waves on the aft fuselage and plumeinduced flow separation are other potential risks. The development of an operational reusable launcher from the Skylon concept necessitates the judicious use of a combination of engineering methods, advanced methods based on required physics or analytical fidelity, test data, and independent assessments.

Rocket-Based Combined Cycle Engine Concept Development

NASA Technical Reports Server (NTRS)

Ratekin, G.; Goldman, Allen; Ortwerth, P.; Weisberg, S.; McArthur, J. Craig (Technical Monitor)

2001-01-01

The development of rocket-based combined cycle (RBCC) propulsion systems is part of a 12 year effort under both company funding and contract work. The concept is a fixed geometry integrated rocket, ramjet, scramjet, which is hydrogen fueled and uses hydrogen regenerative cooling. The baseline engine structural configuration uses an integral structure that eliminates panel seals, seal purge gas, and closeout side attachments. Engine A5 is the current configuration for NASA Marshall Space Flight Center (MSFC) for the ART program. Engine A5 models the complete flight engine flowpath of inlet, isolator, airbreathing combustor, and nozzle. High-performance rocket thrusters are integrated into the engine enabling both low speed air-augmented rocket (AAR) and high speed pure rocket operation. Engine A5 was tested in GASL's new Flight Acceleration Simulation Test (FAST) facility in all four operating modes, AAR, RAM, SCRAM, and Rocket. Additionally, transition from AAR to RAM and RAM to SCRAM was also demonstrated. Measured performance demonstrated vision vehicle performance levels for Mach 3 AAR operation and ramjet operation from Mach 3 to 4. SCRAM and rocket mode performance was above predictions. For the first time, testing also demonstrated transition between operating modes.

Effects of the Orion Launch Abort Vehicle Plumes on Aerodynamics and Controllability

NASA Technical Reports Server (NTRS)

Vicker, Darby; Childs, Robert; Rogers,Stuart E.; McMullen, Matthew; Garcia, Joseph; Greathouse, James

2013-01-01

Characterization of the launch abort system of the Multi-purpose Crew Vehicle (MPCV) for control design and accurate simulation has provided a significant challenge to aerodynamicists and design engineers. The design space of the launch abort vehicle (LAV) includes operational altitudes from ground level to approximately 300,000 feet, Mach numbers from 0-9, and peak dynamic pressure near 1300psf during transonic flight. Further complicating the characterization of the aerodynamics and the resultant vehicle controllability is the interaction of the vehicle flowfield with the plumes of the two solid propellant motors that provide attitude control and the main propulsive impulse for the LAV. These interactions are a function of flight parameters such as Mach number, altitude, dynamic pressure, vehicle attitude, as well as parameters relating to the operation of the motors themselves - either as a function of time for the AM, or as a result of the flight control system requests for control torque from the ACM. This paper discusses the computational aerodynamic modeling of the aerodynamic interaction caused by main abort motor and the attitude control motor of the MPCV LAV, showing the effects of these interactions on vehicle controllability.

Effect of Rocket-Motor Operation on the Drag of Three 1/5-Scale Hermes A-3A Models in Free Flight

NASA Technical Reports Server (NTRS)

Jackson, H. Herbert

1954-01-01

Three 1/5-scale models of the Hermes A-3A missile have been flown to determine the effect of rocket-motor operation on the drag corresponding to various altitude and Mach number combinations. The flights covered a Mach number range from 0.5 to 1.8, and ratios of jet-exit static pressure to free-stream static pressure from 0.8 to 1.8. The results indicate that the power-on drag of the missile should be the same as the power-off drag at Mach number 1.3 and slightly less than the power-off drag at Mach number 1.55.

In-flight acoustic test results for the SR-2 and SR-3 advanced-design propellers

NASA Technical Reports Server (NTRS)

Lasagna, P. L.; Mackall, K. G.; Cohn, R. B.

1983-01-01

Several advanced-design propellers, previously tested in the wind tunnel at the Lewis Research Center, have been tested in flight at the Dryden Flight Research Facility. The flight-test propellers were mounted on a pylon on the top of the fuselage of a JetStar airplane. Acoustic data for the advanced-design SR-2 and SR-3 propellers at Mach numbers to 0.8 and helical-tip Mach numbers to 1.15 are presented; maximum blade-passage frequency sound-pressure levels are also compared.

X-38 Experimental Aerothermodynamics

NASA Technical Reports Server (NTRS)

Horvath, Thomas J.; Berry, Scott A.; Merski, N. Ronald; Fitzgerald, Steve M.

2000-01-01

The X-38 program seeks to demonstrate an autonomously returned orbital test flight vehicle to support the development of an operational Crew Return Vehicle for the International Space Station. The test flight, anticipated in 2002, is intended to demonstrate the entire mission profile of returning Space Station crew members safely back to earth in the event of medical or mechanical emergency. Integral to the formulation of the X-38 flight data book and the design of the thermal protection system, the aerothermodynamic environment is being defined through a synergistic combination of ground based testing and computational fluid dynamics. This report provides an overview of the hypersonic aerothermodynamic wind tunnel program conducted at the NASA Langley Research Center in support of the X-38 development. Global and discrete surface heat transfer force and moment, surface streamline patterns, and shock shapes were measured on scaled models of the proposed X-38 configuration in different test gases at Mach 6, 10 and 20. The test parametrics include angle of attack from 0 to 50 degs, unit Reynolds numbers from 0.3 x 10 (exp 6) to 16 x 10 (exp 6)/ ft, rudder deflections of 0, 2, and 5 deg. and body flap deflections from 0 to 30 deg. Results from hypersonic aerodynamic screening studies that were conducted as the configuration evolved to the present shape at, presented. Heavy gas simulation tests have indicated that the primary real gas effects on X-38 aerodynamics at trim conditions are expected to favorably influence flap effectiveness. Comparisons of the experimental heating and force and moment data to prediction and the current aerodynamic data book are highlighted. The effects of discrete roughness elements on boundary layer transition were investigated at Mach 6 and the development of a transition correlation for the X-38 vehicle is described. Extrapolation of ground based heating measurements to flight radiation equilibrium wall temperatures at Mach 6 and 10 were made and generally compared to within 50 deg F of flight prediction.

Skin Friction Measurements at a Mach Number of Three and Momentum Thickness Reynolds Numbers Up to a Half Million.

DTIC Science & Technology

1980-09-01

k ADAOGZ 826 ~~~~~~~~~AIR FORCE FLIGHT DNMC A RGTPTESNABO / / I SKIN FRICTION MEASUREMENTS AT A MACH NUMBER OF THREE AND MOMENT--ETCIU) UNCASSFIE...AT A MACH NUMBER OF THREE AND MOMENTUM THICKNESS REYNOLDS NUMEERS UP TO A HALF MILLION Anthony W. Fiore Aeromechanics Division DTIC September 1980...NOR(&) S. CONTRACT OR GRANT NUMBER (s) 9. PER ORMING ORGANIZATION NAME AND ADDRESS PROR UNT N WU Flight Dynamics Laboratory (AFWAL/FIMG)RA; WOKUNTU

Comparison of wind tunnel and flight test afterbody and nozzle pressures for a twin-jet fighter aircraft at transonic speeds

NASA Technical Reports Server (NTRS)

Nugent, Jack; Pendergraft, Odis C., Jr.

1987-01-01

Afterbody and nozzle pressures measured on a 1/12-scale model and in flight on a twin-jet fighter aircraft were compared as Mach number varied from 0.6 to 1.2, Reynolds number from 17.5 million to 302.5 million, and angle of attack from 1 to 7 deg. At Mach 0.6 and 0.8, nozzle pressure coefficient distributions and nozzle axial force coefficients agreed and showed good recompression. At Mach 0.9 and 1.2, flow complexity caused a loss in recompression for both flight and wind tunnel nozzle data. The flight data exhibited less negative values of pressure coefficient and lower axial force coefficients than did the wind tunnel data. Reynolds number effects were noted only at these Mach numbers. Jet temperature and mass flux ratio did not affect the comparisons of nozzle axial flow coefficient. At subsonic speeds, the levels of pressure coefficient distributions on the upper fuselage and lower nacelle surfaces for flight were less negative than those for the model. The model boundary layer thickness at the aft rake station exceeded that for the forward rake station and increased with increasing angle of attack. The flight boundary layer thickness at the aft rake station was less than that for the forward rake station and decreased with increasing angle of attack.

Longitudinal Stability and Control Characteristics from a Flight Investigation of a Cruciform Canard Missile Configuration Having an Exposed Wing-canard Area Ratio of 16:1

NASA Technical Reports Server (NTRS)

Moul, Martin T; Wineman, Andrew R

1952-01-01

A flight investigation has been made to determine the longitudinal stability and control characteristics of a 60 0 delta-wing-canard missile configuration with an exposed wing-canard area ratio of 16:1. The results presented include the longitudinal stability derivatives, control effectiveness, and drag characteristics for a Mach number range of 0.75 to 1.80 and are compared with the results of a similar configuration having larger 6ontrols. Stability characteristics are also presented from the flights of an interdigitated canard configuration at a Mach number of 2.08 and a wing-body configuration at Mach numbers of 1.25 to 1.45. The stability derivatives varied gradually with Mach number with the exception of the damping-in-pitch derivative. Aerodynamic damping in pitch decreased to a minimum at a Mach number of 1.0 3, then increased to a peak value at a Mach number of 1.26 followed by a gradual decrease at higher Mach numbers. The aerodynamic-center location of the in-line canard configuration shifted rearward 13 percent of the mean aerodynamic chord at transonic speeds. The pitching-moment curve slope was 25 percent greater for the model having no canards than for the in-line configuration. No large effects of interdigitation were noted in the stability derivatives. Pitching effectiveness of the in-line configuration was maintained throughout the Mach number range. A comparison of the stability and control characteristics of two canard configurations having different area controls showed that decreasing the control area 44 percent decreased the pitching effectiveness proportionally, shifted the aerodynamic-center location rearward 9 to 14 percent of the mean aerodynamic chord, and reduced the total hinge moments required for 10 trimmed flight about 50 percent at transonic speeds.

Correlation of the Drag Characteristics of a Typical Pursuit Airplane Obtained from High-Speed Wind-Tunnel and Flight Tests

NASA Technical Reports Server (NTRS)

Nissen, James M; Gadebero, Burnett L; Hamilton, William T

1948-01-01

In order to obtain a correlation of drag data from wind-tunnel and flight tests at high Mach numbers, a typical pursuit airplane, with the propeller removed, was tested in flight at Mach numbers up to 0.755, and the results were compared with wind-tunnel tests of a 1/3-scale model of the airplane. The tests results show that the drag characteristics of the test airplane can be predicted with satisfactory accuracy from tests in the Ames 16-foot high-speed wind tunnel of the Ames Aeronautical Laboratory at both high and low Mach numbers. It is considered that this result is not unique with the airplane.

Space shuttle: Heat transfer rate measurements on Convair booster (B-15B-2) and North American Rockwell orbiter (161B) at nominal Mach number of 8

NASA Technical Reports Server (NTRS)

Warmbrod, J. D.; Martindale, W. R.; Matthews, R. K.

1971-01-01

Plotted and tabulated data from the thin-skin thermocouple phase of an experimental test program are presented. These data are representative of three events of simulated flight and are described as booster-orbiter ascent heating data, booster reentry heating data, and orbiter reentry heating data. The test was conducted in a 50-inch hypersonic tunnel b at a nominal Mach number of 8 and free-stream Reynolds number range of 700,000 to 3,700,000 per foot. The model employed was a 0.009 scale replica of the Convair B-15B-2 booster and North American Rockwell 161B orbiter.

Orbiter BLT Flight Experiment Wind Tunnel Simulations: Nearfield Flowfield Imaging and Surface Thermography

NASA Technical Reports Server (NTRS)

Danehy, Paul M.; Ivey, Christoper B.; Barthel, Brett F.; Inman, Jennifer A.; Jones, Stephen B.; Watkins, Anthony N.; Goodman, Kyle Z.; McCrea, Andrew C.; Leighty, Bradley D.; Lipford, William K.;

Thermochemical Modeling of Nonequilibrium Oxygen Flows

NASA Astrophysics Data System (ADS)

Neitzel, Kevin Joseph

The development of hypersonic vehicles leans heavily on computational simulation due to the high enthalpy flow conditions that are expensive and technically challenging to replicate experimentally. The accuracy of the nonequilibrium modeling in the computer simulations dictates the design margin that is required for the thermal protection system and flight dynamics. Previous hypersonic vehicles, such as Apollo and the Space Shuttle, were primarily concerned with re-entry TPS design. The strong flow conditions of re-entry, involving Mach numbers of 25, quickly dissociate the oxygen molecules in air. Sustained flight, hypersonic vehicles will be designed to operate in Mach number ranges of 5 to 10. The oxygen molecules will not quickly dissociate and will play an important role in the flow field behavior. The development of nonequilibrium models of oxygen is crucial for limiting modeling uncertainty. Thermochemical nonequilibrium modeling is investigated for oxygen flows. Specifically, the vibrational relaxation and dissociation behavior that dominate the nonequilibrium physics in this flight regime are studied in detail. The widely used two-temperature (2T) approach is compared to the higher fidelity and more computationally expensive state-to-state (STS) approach. This dissertation utilizes a wide range of rate sources, including newly available STS rates, to conduct a comprehensive study of modeling approaches for hypersonic nonequilibrium thermochemical modeling. Additionally, the physical accuracy of the computational methods are assessed by comparing the numerical results with available experimental data. The numerical results and experimental measurements present strong nonequilibrium, and even non-Boltzmann behavior in the vibrational energy mode for the sustained hypersonic flight regime. The STS approach is able to better capture the behavior observed in the experimental data, especially for stronger nonequilibrium conditions. Additionally, a reduced order model (ROM) modification to the 2T model is developed to improve the capability of the 2T approach framework.

Rapid Fabrication of Flat Plate Cavity Phosphor Thermography Test Models for Shuttle Return-to-Flight Aero-Heating

NASA Technical Reports Server (NTRS)

Buck, Gregory M.; Powers, Michael A.; Nevins, Stephen C.; Griffith, Mark S.; Wainwright, Gary A.

2006-01-01

Methods, materials and equipment are documented for fabricating flat plate test models at NASA Langley Research Center for Shuttle return-to-flight aeroheating experiments simulating open and closed cavity interactions in Langley s hypersonic 20-Inch Mach 6 air wind tunnel. Approximately 96 silica ceramic flat plate cavity phosphor thermography test models have been fabricated using these methods. On one model, an additional slot is machined through the back of the plate and into the cavity and vented into an evacuated plenum chamber to simulate a further opening in the cavity. After sintering ceramic to 2150 F, and mounting support hardware, a ceramic-based two-color thermographic phosphor coating is applied for global temperature and heat transfer measurements, with fiducial markings for image registration.

Control Design for an Advanced Geared Turbofan Engine

NASA Technical Reports Server (NTRS)

Chapman, Jeffryes W.; Litt, Jonathan S.

2017-01-01

This paper describes the design process for the control system of an advanced geared turbofan engine. This process is applied to a simulation that is representative of a 30,000 lbf thrust class concept engine with two main spools, ultra-high bypass ratio, and a variable area fan nozzle. Control system requirements constrain the non-linear engine model as it operates throughout its flight envelope of sea level to 40,000 ft and from 0 to 0.8 Mach. The control architecture selected for this project was developed from literature and reflects a configuration that utilizes a proportional integral controller integrated with sets of limiters that enable the engine to operate safely throughout its flight envelope. Simulation results show the overall system meets performance requirements without exceeding system operational limits.

Effects of aircraft and flight parameters on energy-efficient profile descents in time-based metered traffic

NASA Technical Reports Server (NTRS)

Dejarnette, F. R.

1984-01-01

Concepts to save fuel while preserving airport capacity by combining time based metering with profile descent procedures were developed. A computer algorithm is developed to provide the flight crew with the information needed to fly from an entry fix to a metering fix and arrive there at a predetermined time, altitude, and airspeed. The flight from the metering fix to an aim point near the airport was calculated. The flight path is divided into several descent and deceleration segments. Descents are performed at constant Mach numbers or calibrated airspeed, whereas decelerations occur at constant altitude. The time and distance associated with each segment are calculated from point mass equations of motion for a clean configuration with idle thrust. Wind and nonstandard atmospheric properties have a large effect on the flight path. It is found that uncertainty in the descent Mach number has a large effect on the predicted flight time. Of the possible combinations of Mach number and calibrated airspeed for a descent, only small changes were observed in the fuel consumed.

Boundary-Layer Edge Conditions and Transition Reynolds Number Data for a Flight Test at Mach 20 (Reentry F)

NASA Technical Reports Server (NTRS)

Johnson, Charles B.; Stainback, P. Calvin; Wicker, Kathleen C.; Boney, Lillian R.

1972-01-01

A flight experiment, designated Reentry F, was conducted to measure heat-transfer rates for laminar, transitional, and turbulent boundary layers on a 5 deg half-angle cone 3.962 m (13 ft) long with a preflight nose radius of 2.54 mm (0.10 in.). Data were obtained over an altitude range from 36.58 to 18.29 km (120 000 to 60 000 ft) at a flight velocity of about 6.096 km/sec (20 000 ft/sec). The nominal values of the free-stream total enthalpy, sharp-cone Mach number, and the wall-to-total enthalpy ratio were 18 MJ/kg (8000 Btu/lb), 15, and 0.03, respectively. Calculated boundary-layer edge conditions that account for effects of the entropy layer and corresponding local transition Reynolds numbers are reported in the present paper. Fully developed turbulent flow occurred with essentially constant boundary-layer edge conditions near the sharp-cone values. Transition data were obtained with local edge Mach numbers ranging from about 5.55 to 15. Transition Reynolds numbers, based on local condition, were as high as 6.6 x 10(exp 7) with an edge Mach number of about 14.4 at an altitude of 24.38 km (80 000 ft). The transition could be correlated with previous flight data taken over a Mach number range from 3 to 12 in terms of parameters including the effects of local unit Reynolds number, boundary-layer wall-to-edge enthalpy ratio, and local Mach number.

High-Speed Wind-Tunnel Tests of a Model of the Lockheed YP-80A Airplane Including Correlation with Flight Tests and Tests of Dive-Recovery Flaps

NASA Technical Reports Server (NTRS)

Cleary, Joseph W.; Gray, Lyle J.

1947-01-01

This report contains the results of tests of a 1/3-scale model of the Lockheed YP-90A "Shooting Star" airplane and a comparison of drag, maximum lift coefficient, and elevator angle required for level flight as measured in the wind tunnel and in flight. Included in the report are the general aerodynamic characteristics of the model and of two types of dive-recovery flaps, one at several positions along the chord on the lower surface of the wing and the other on the lower surface of the fuselage. The results show good agreement between the flight and wind-tunnel measurements at all Mach numbers. The results indicate that the YP-80A is controllable in pitch by the elevators to a Mach number of at least 0.85. The fuselage dive-recovery flaps are effective for producing a climbing moment and increasing the drag at Mach numbers up to at least 0.8. The wing dive-recovery flaps are most effective for producing a climbing moment at 0.75 Mach number. At 0.85 Mach number, their effectiveness is approximately 50 percent of the maximum. The optimum position for the wing dive-recovery flaps to produce a climbing moment is at approximately 35 percent of the chord.

An Inviscid Computational Study of an X-33 Configuration at Hypersonic Speeds

NASA Technical Reports Server (NTRS)

Prabhu, Ramadas K.

1999-01-01

This report documents the results of a study conducted to compute the inviscid longitudinal aerodynamic characteristics of a simplified X-33 configuration. The major components of the X-33 vehicle, namely the body, the canted fin, the vertical fin, and the body-flap, were simulated in the CFD (Computational Fluid Dynamic) model. The rear-ward facing surfaces at the base including the aerospike engine surfaces were not simulated. The FELISA software package consisting of an unstructured surface and volume grid generator and two inviscid flow solvers was used for this study. Computations were made for Mach 4.96, 6.0, and 10.0 with perfect gas air option, and for Mach 10 with equilibrium air option with flow condition of a typical point on the X-33 flight trajectory. Computations were also made with CF4 gas option at Mach 6.0 to simulate the CF4 tunnel flow condition. An angle of attack range of 12 to 48 deg was covered. The CFD results were compared with available wind tunnel data. Comparison was good at low angles of attack; at higher angles of attack (beyond 25 deg) some differences were found in the pitching moment. These differences progressively increased with increase in angle of attack, and are attributed to the viscous effects. However, the computed results showed the trends exhibited by the wind tunnel data.

Lateral and longitudinal stability and control parameters for the space shuttle discovery as determined from flight test data

NASA Technical Reports Server (NTRS)

Suit, William T.; Schiess, James R.

1988-01-01

The Discovery vehicle was found to have longitudinal and lateral aerodynamic characteristics similar to those of the Columbia and Challenger vehicles. The values of the lateral and longitudinal parameters are compared with the preflight data book. The lateral parameters showed the same trends as the data book. With the exception of C sub l sub Beta for Mach numbers greater than 15, C sub n sub delta r for Mach numbers greater than 2 and for Mach numbers less than 1.5, where the variation boundaries were not well defined, ninety percent of the extracted values of the lateral parameters fell within the predicted variations. The longitudinal parameters showed more scatter, but scattered about the preflight predictions. With the exception of the Mach 1.5 to .5 region of the flight envelope, the preflight predictions seem a reasonable representation of the Shuttle aerodynamics. The models determined accounted for ninety percent of the actual flight time histories.

Selected Performance Measurements of the F-15 Active Axisymmetric Thrust-vectoring Nozzle

NASA Technical Reports Server (NTRS)

Orme, John S.; Sims, Robert L.

1998-01-01

Flight tests recently completed at the NASA Dryden Flight Research Center evaluated performance of a hydromechanically vectored axisymmetric nozzle onboard the F-15 ACTIVE. A flight-test technique whereby strain gages installed onto engine mounts provided for the direct measurement of thrust and vector forces has proven to be extremely valuable. Flow turning and thrust efficiency, as well as nozzle static pressure distributions were measured and analyzed. This report presents results from testing at an altitude of 30,000 ft and a speed of Mach 0.9. Flow turning and thrust efficiency were found to be significantly different than predicted, and moreover, varied substantially with power setting and pitch vector angle. Results of an in-flight comparison of the direct thrust measurement technique and an engine simulation fell within the expected uncertainty bands. Overall nozzle performance at this flight condition demonstrated the F100-PW-229 thrust-vectoring nozzles to be highly capable and efficient.

Selected Performance Measurements of the F-15 ACTIVE Axisymmetric Thrust-Vectoring Nozzle

NASA Technical Reports Server (NTRS)

Orme, John S.; Sims, Robert L.

1999-01-01

Flight tests recently completed at the NASA Dryden Flight Research Center evaluated performance of a hydromechanically vectored axisymmetric nozzle onboard the F-15 ACTIVE. A flight-test technique whereby strain gages installed onto engine mounts provided for the direct measurement of thrust and vector forces has proven to be extremely valuable. Flow turning and thrust efficiency, as well as nozzle static pressure distributions were measured and analyzed. This report presents results from testing at an altitude of 30,000 ft and a speed of Mach 0.9. Flow turning and thrust efficiency were found to be significantly different than predicted, and moreover, varied substantially with power setting and pitch vector angle. Results of an in-flight comparison of the direct thrust measurement technique and an engine simulation fell within the expected uncertainty bands. Overall nozzle performance at this flight condition demonstrated the F100-PW-229 thrust-vectoring nozzles to be highly capable and efficient.

Effect of aerodynamic and angle-of-attack uncertainties on the blended entry flight control system of the Space Shuttle from Mach 10 to 2.5

NASA Technical Reports Server (NTRS)

Stone, H. W.; Powell, R. W.

1984-01-01

A six-degree-of-freedom simulation analysis has been performed for the Space Shuttle Orbiter during entry from Mach 10 to 2.5 with realistic off-nominal conditions using the entry flight control system specified in May 1978. The off-nominal conditions included the following: (1) aerodynamic uncertainties, (2) an error in deriving the angle of attack from onboard instrumentation, (3) the failure of two of the four reaction control-system thrusters on each side, and (4) a lateral center-of-gravity offset. With combinations of the above off-nominal conditions, the control system performed satisfactorily with a few exceptions. The cases that did not exhibit satisfactory performance displayed the following main weaknesses. Marginal performance was exhibited at hypersonic speeds with a sensed angle-of-attack error of 4 deg. At supersonic speeds the system tended to be oscillatory, and the system diverged for several cases because of the inability to hold lateral trim. Several system modifications were suggested to help solve these problems and to maximize safety on the first flight: alter the elevon-trim and speed-brake schedules, delay switching to rudder trim until the rudder effectiveness is adequate, and reduce the overall rudder loop gain. These and other modifications were incorporated in a flight-control-system redesign in May 1979.

Comparisons of wing pressure distribution from flight tests of flush and external orifices for Mach numbers from 0.50 to 0.97

NASA Technical Reports Server (NTRS)

Montoya, L. C.; Lux, D. P.

1975-01-01

Wing pressure distributions obtained in flight with flush orifice and external tubing orifice installations for Mach numbers from 0.50 to 0.97 are compared. The procedure used to install the external tubing orifice is discussed. The results indicate that external tubing orifice installations can give useful results.

Force and Pressure Recovery Characteristics at Supersonic Speeds of a Conical Spike Inlet with a Bypass Discharging from the Top or Bottom of the Diffuser in an Axial Direction

NASA Technical Reports Server (NTRS)

Allen, J L; Beke, Andrew

1953-01-01

Force and pressure-recovery characteristics of a nacelle-type conical-spike inlet with a fixed-area bypass located in the top or bottom of the diffuser are presented for flight Mach numbers of 1.6, 1.8, and 2.0 for angles of attack from 0 degrees to 9 degrees. Top or bottom location of the bypass did not have significant effects on diffuser pressure-recovery, bypass mass-flow ratio, or drag coefficient over the range of angles of attack, flight Mach numbers, and stable engine mass-flow ratios investigated. A larger stable subcritical operating range was obtained with the bypass on the bottom at angles of attack from 3 degrees to 9 degrees at a flight Mach number of 2.0. At a flight Mach number of 2.0, the discharge of 14 percent of the critical mass flow of the inlet by means of a bypass increased the drag only one-fifth of the additive drag that would result for equivalent spillage behind an inlet normal shock without significant reductions in diffuser pressure recovery.

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)

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

NASA Astrophysics Data System (ADS)

Huh, Lynn

The objective of the present research was to develop a new method to derive the aerodynamic coefficients and the associated uncertainties for flight vehicles via post- flight inertial navigation analysis using data from the inertial measurement unit. Statistical estimates of vehicle state and aerodynamic coefficients are derived using Monte Carlo simulation. Trajectory reconstruction using the inertial navigation system (INS) is a simple and well used method. However, deriving realistic uncertainties in the reconstructed state and any associated parameters is not so straight forward. Extended Kalman filters, batch minimum variance estimation and other approaches have been used. However, these methods generally depend on assumed physical models, assumed statistical distributions (usually Gaussian) or have convergence issues for non-linear problems. The approach here assumes no physical models, is applicable to any statistical distribution, and does not have any convergence issues. The new approach obtains the statistics directly from a sufficient number of Monte Carlo samples using only the generally well known gyro and accelerometer specifications and could be applied to the systems of non-linear form and non-Gaussian distribution. When redundant data are available, the set of Monte Carlo simulations are constrained to satisfy the redundant data within the uncertainties specified for the additional data. The proposed method was applied to validate the uncertainty in the pre-flight aerodynamic database of the X-43A Hyper-X research vehicle. In addition to gyro and acceleration data, the actual flight data include redundant measurements of position and velocity from the global positioning system (GPS). The criteria derived from the blend of the GPS and INS accuracy was used to select valid trajectories for statistical analysis. The aerodynamic coefficients were derived from the selected trajectories by either direct extraction method based on the equations in dynamics, or by the inquiry of the pre-flight aerodynamic database. After the application of the proposed method to the case of the X-43A Hyper-X research vehicle, it was found that 1) there were consistent differences in the aerodynamic coefficients from the pre-flight aerodynamic database and post-flight analysis, 2) the pre-flight estimation of the pitching moment coefficients was significantly different from the post-flight analysis, 3) the type of distribution of the states from the Monte Carlo simulation were affected by that of the perturbation parameters, 4) the uncertainties in the pre-flight model were overestimated, 5) the range where the aerodynamic coefficients from the pre-flight aerodynamic database and post-flight analysis are in closest agreement is between Mach *.* and *.* and more data points may be needed between Mach * and ** in the pre-flight aerodynamic database, 6) selection criterion for valid trajectories from the Monte Carlo simulations was mostly driven by the horizontal velocity error, 7) the selection criterion must be based on reasonable model to ensure the validity of the statistics from the proposed method, and 8) the results from the proposed method applied to the two different flights with the identical geometry and similar flight profile were consistent.

X-33 Attitude Control System Design for Ascent, Transition, and Entry Flight Regimes

NASA Technical Reports Server (NTRS)

Hall, Charles E.; Gallaher, Michael W.; Hendrix, Neal D.

1998-01-01

The Vehicle Control Systems Team at Marshall Space Flight Center, Systems Dynamics Laboratory, Guidance and Control Systems Division is designing under a cooperative agreement with Lockheed Martin Skunkworks, the Ascent, Transition, and Entry flight attitude control system for the X-33 experimental vehicle. Ascent flight control begins at liftoff and ends at linear aerospike main engine cutoff (NECO) while Transition and Entry flight control begins at MECO and concludes at the terminal area energy management (TAEM) interface. TAEM occurs at approximately Mach 3.0. This task includes not only the design of the vehicle attitude control systems but also the development of requirements for attitude control system components and subsystems. The X-33 attitude control system design is challenged by a short design cycle, the design environment (Mach 0 to about Mach 15), and the X-33 incremental test philosophy. The X-33 design-to-launch cycle of less than 3 years requires a concurrent design approach while the test philosophy requires design adaptation to vehicle variations that are a function of Mach number and mission profile. The flight attitude control system must deal with the mixing of aerosurfaces, reaction control thrusters, and linear aerospike engine control effectors and handle parasitic effects such as vehicle flexibility and propellant sloshing from the uniquely shaped propellant tanks. The attitude control system design is, as usual, closely linked to many other subsystems and must deal with constraints and requirements from these subsystems.

Free-jet Testing of a REST Scramjet at Off-Design Conditions

NASA Technical Reports Server (NTRS)

Smart, Michael K.; Ruf, Edward G.

2006-01-01

Scramjet flowpaths employing elliptical combustors have the potential to improve structural efficiency and performance relative to those using planar geometries. NASA Langley has developed a scramjet flowpath integrated into a lifting body vehicle, while transitioning from a rectangular capture area to both an elliptical throat and combustor. This Rectangular-to-Elliptical Shape Transition (REST) scramjet, has a design point of Mach 7.1, and is intended to operate with fixed-geometry between Mach 4.5 and 8.0. This paper describes initial free-jet testing of the heat-sink REST scramjet engine model at conditions simulating Mach 5.3 flight. Combustion of gaseous hydrogen fuel at equivalence ratios between 0.5 and 1.5 generated robust performance after ignition with a silane-hydrogen pilot. Facility model interactions were experienced for fuel equivalence ratios above 1.1, yet despite this, the flowpath was not unstarted by fuel addition at the Mach 5.3 test condition. Combustion tests at reduced stagnation enthalpy indicated that the engine self-started following termination of the fuel injection. Engine data is presented for the largest fuel equivalence ratio tested without facility interaction. These results indicate that this class of three-dimensional scramjet engine operates successfully at off-design conditions.

Generating an Aerodynamic Model for Projectile Flight Simulation Using Unsteady, Time Accurate Computational Fluid Dynamic Results

DTIC Science & Technology

2006-09-01

Figure 17. Station line center of Magnus force vs. Mach number for spin-stabilized projectile...forces and moments on the projectile. It is also relatively easy to change the wind tunnel model to allow detailed parametric effects to be...such as pitch and roll damping, as well as, Magnus force and moment coefficients, are difficult to obtain in a wind tunnel and require a complex

Fuel-conservation evaluation of US Army helicopters. Part 6. Performance calculator evaluation. Final report for period ending January 1981

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

Dominick, F.; Lockwood, R.A.

1986-07-01

The US Army Aviation Engineering Flight Activity conducted an evaluation of Flight Management Calculator for the UH-1H. The calculator was a Hewlett-Packard HP-41CV. The performance calculator was evaluated for flight planning and in-flight use during 14 mission flights simulating operational conditions. The calculator was much easier to use in-flight than the operator's manual data. The calculator program needs improvement in the areas of pre-flight planning and execution speed. The mission flights demonstrated a 19% fuel saving using optimum over normal flight profiles in warm temperatures (15/sup 0/C above standard). Savings would be greater at colder temperatures because of increasing compressibilitymore » effects. Acceptable accuracy for individual aircraft under operational conditions may require a regressive analog model in which individual aircraft data are used to update the program. The performance data base for the UH-1H was expanded with level flight and hover data to thrust coefficients and Mach numbers to the practical limits of aircraft operation.« less

PSL Icing Facility Upgrade Overview

NASA Technical Reports Server (NTRS)

Griffin, Thomas A.; Dicki, Dennis J.; Lizanich, Paul J.

2014-01-01

The NASA Glenn Research Center Propulsion Systems Lab (PSL) was recently upgraded to perform engine inlet ice crystal testing in an altitude environment. The system installed 10 spray bars in the inlet plenum for ice crystal generation using 222 spray nozzles. As an altitude test chamber, the PSL is capable of simulating icing events at altitude in a groundtest facility. The system was designed to operate at altitudes from 4,000 to 40,000 ft at Mach numbers up to 0.8M and inlet total temperatures from -60 to +15 degF. This paper and presentation will be part of a series of presentations on PSL Icing and will cover the development of the icing capability through design, developmental testing, installation, initial calibration, and validation engine testing. Information will be presented on the design criteria and process, spray bar developmental testing at Cox and Co., system capabilities, and initial calibration and engine validation test. The PSL icing system was designed to provide NASA and the icing community with a facility that could be used for research studies of engine icing by duplicating in-flight events in a controlled ground-test facility. With the system and the altitude chamber we can produce flight conditions and cloud environments to simulate those encountered in flight. The icing system can be controlled to set various cloud uniformities, droplet median volumetric diameter (MVD), and icing water content (IWC) through a wide variety of conditions. The PSL chamber can set altitudes, Mach numbers, and temperatures of interest to the icing community and also has the instrumentation capability of measuring engine performance during icing testing. PSL last year completed the calibration and initial engine validation of the facility utilizing a Honeywell ALF502-R5 engine and has duplicated in-flight roll back conditions experienced during flight testing. This paper will summarize the modifications and buildup of the facility to accomplish these tests.

Minimum time and fuel flight profiles for an F-15 airplane with a Highly Integrated Digital Electronic Control (HIDEC) system

NASA Technical Reports Server (NTRS)

Haering, E. A., Jr.; Burcham, F. W., Jr.

1984-01-01

A simulation study was conducted to optimize minimum time and fuel consumption paths for an F-15 airplane powered by two F100 Engine Model Derivative (EMD) engines. The benefits of using variable stall margin (uptrim) to increase performance were also determined. This study supports the NASA Highly Integrated Digital Electronic Control (HIDEC) program. The basis for this comparison was minimum time and fuel used to reach Mach 2 at 13,716 m (45,000 ft) from the initial conditions of Mach 0.15 at 1524 m (5000 ft). Results were also compared to a pilot's estimated minimum time and fuel trajectory determined from the F-15 flight manual and previous experience. The minimum time trajectory took 15 percent less time than the pilot's estimate for the standard EMD engines, while the minimum fuel trajectory used 1 percent less fuel than the pilot's estimate for the minimum fuel trajectory. The F-15 airplane with EMD engines and uptrim, was 23 percent faster than the pilot's estimate. The minimum fuel used was 5 percent less than the estimate.

NASA's F-15B conducts a local Mach investigation flight over California's Mojave Desert.

NASA Image and Video Library

2004-06-01

NASA's F-15B Research Testbed aircraft flew instrumentation in June 2004 called the Local Mach Investigation (LMI), designed to gather local airflow data for future research projects using the aircraft's Propulsion Flight Test Fixture (PFTF). The PFTF is the black rectangular fixture attached to the aircraft's belly. The LMI package was located in the orange device attached to the PFTF.

NASA's F-15B conducts a local Mach investigation flight over California's Mojave Desert.

NASA Image and Video Library

2004-06-04

NASA's F-15B Research Testbed aircraft flew instrumentation in June 2004 called the Local Mach Investigation (LMI), designed to gather local airflow data for future research projects using the aircraft's Propulsion Flight Test Fixture (PFTF). The PFTF is the black rectangular fixture attached to the aircraft's belly. The LMI package was located in the orange device attached to the PFTF.

An experimental determination in Calspan Ludwieg tube of the base environment of the integrated space shuttle vehicle at simulated Mach 4.5 flight conditions (test IH5 of model 19-OTS)

NASA Technical Reports Server (NTRS)

Drzewiecki, R. F.; Foust, J. W.

1976-01-01

A model test program was conducted to determine heat transfer and pressure distributions in the base region of the space shuttle vehicle during simulated launch trajectory conditions of Mach 4.5 and pressure altitudes between 90,000 and 210,000 feet. Model configurations with and without the solid propellant booster rockets were examined to duplicate pre- and post-staging vehicle geometries. Using short duration flow techniques, a tube wind tunnel provided supersonic flow over the model. Simultaneously, combustion generated exhaust products reproduced the gasdynamic and thermochemical structure of the main vehicle engine plumes. Heat transfer and pressure measurements were made at numerous locations on the base surfaces of the 19-OTS space shuttle model with high response instrumentation. In addition, measurements of base recovery temperature were made indirectly by using dual fine wire and resistance thermometers and by extrapolating heat transfer measurements.

X-2 on Transportation Dolly

NASA Technical Reports Server (NTRS)

1952-01-01

This 1952 photograph shows the X-2 #2 aircraft mounted on a special transportation dolly at Edwards Air Force Base, California. The dolly was steerable and was used for transporting the X-2 around and for towing it off the lakebed at Edwards Air Force Base after a landing. This was the number 2 airplane (46-675), which was lost on May 12, 1953, on a captive flight over Lake Ontario when the airplane exploded during a liquid-oxygen topoff test, killing the pilot, Jean Ziegler, and EB-50A crewman Frank Wolko. Almost no debris was recovered from Lake Ontario, so no cause for the explosion could be determined. Later, however, investigations of similar explosions in the X-1 #3, X-1A, and X-1D traced the problem to Ulmer leather gaskets, which exuded tricresyl phosphate. This substance caused detonations in the supercold atmosphere of the airplanes' liquid oxygen tanks. As the X-2 #2 also had these gaskets, they were probably the cause of the explosion in that aircraft as well. The X-2 was a swept-wing, rocket-powered aircraft designed to fly faster than Mach 3 (three times the speed of sound). It was built for the U.S. Air Force by the Bell Aircraft Company, Buffalo, New York. The X-2 was flown to investigate the problems of aerodynamic heating as well as stability and control effectiveness at high altitudes and high speeds (in excess of Mach 3). Bell aircraft built two X-2 aircraft. These were constructed of K-monel (a copper and nickel alloy) for the fuselage and stainless steel for the swept wings and control surfaces. The aircraft had ejectable nose capsules instead of ejection seats because the development of ejection seats had not reached maturity at the time the X-2 was conceived. The X-2 ejection canopy was successfully tested using a German V-2 rocket. The X-2 used a skid-type landing gear to make room for more fuel. The airplane was air launched from a modified Boeing B-50 Superfortress Bomber. X-2 Number 1 made its first unpowered glide flight on Aug. 5, 1954, and made a total of 17 (4 glide and 13 powered) flights before it was lost Sept. 27, 1956. The pilot on Flight 17, Capt. Milburn Apt, had flown the aircraft to a record speed of Mach 3.2 (2,094 mph), thus becoming the first person to exceed Mach 3. During that last flight, inertial coupling occurred and the pilot was killed. The aircraft suffered little damage in the crash, resulting in proposals (never implemented) from the Langley Memorial Aeronautical Laboratory, Hampton, Virginia, to rebuild it for use in a hypersonic (Mach 5+) test program. In 1953, X-2 Number 2 was lost in an in-flight explosion while at the Bell Aircraft Company during captive flight trials and was jettisoned into Lake Ontario. The Air Force had previously flown the aircraft on three glide flights at Edwards Air Force Base, California, in 1952. Although the NACA's High-Speed Flight Station, Edwards, California, (predecessor of NASA's Dryden Flight Research Center) never actually flew the X-2 aircraft, the NACA did support the program primarily through Langley Memorial Aeronautical Laboratory wind-tunnel tests and Wallops Island, Virginia, rocket-model tests. The NACA High-Speed Flight Station also provided stability-and-control recording instrumentation and simulator support for the Air Force flights. In the latter regard, the NACA worked with the Air Force in using a special computer to extrapolate and predict aircraft behavior from flight data.

Dual-Mode Combustion

NASA Technical Reports Server (NTRS)

Goyne, Christopher P.; McDaniel, James C.

2002-01-01

The Department of Mechanical and Aerospace Engineering at the University of Virginia has conducted an investigation of the mixing and combustion processes in a hydrogen fueled dual-mode scramjet combustor. The experiment essentially consisted of the "direct connect" continuous operation of a Mach 2 rectangular combustor with a single unswept ramp fuel injector. The stagnation enthalpy of the test flow simulated a flight Mach number of 5. Measurements were obtained using conventional wall instrumentation and laser based diagnostics. These diagnostics included, pressure and wall temperature measurements, Fuel Plume Imaging (FPI) and Particle Image Velocimetry (PIV). A schematic of the combustor configuration and a summary of the measurements obtained are presented. The experimental work at UVa was parallel by Computational Fluid Dynamics (CFD) work at NASA Langley. The numerical and experiment results are compared in this document.

Airbreathing Acceleration Toward Earth Orbit

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

Whitehead, J C

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

Scramjet Tests in a Shock Tunnel at Flight Mach 7, 10, and 15 Conditions

NASA Technical Reports Server (NTRS)

Rogers, R. C.; Shih, A. T.; Tsai, C.-Y.; Foelsche, R. O.

2001-01-01

Tests of the Hyper-X scramjet engine flowpath have been conducted in the HYPULSE shock tunnel at conditions duplicating the stagnation enthalpy at flight Mach 7, 10, and 15. For the tests at Mach 7 and 10 HYPULSE was operated as a reflected-shock tunnel; at the Mach 15 condition, HYPULSE was operated as a shock-expansion tunnel. The test conditions matched the stagnation enthalpy of a scramjet engine on an aerospace vehicle accelerating through the atmosphere along a 1000 psf dynamic pressure trajectory. Test parameter variation included fuel equivalence ratios from lean (0.8) to rich (1.5+); fuel composition from pure hydrogen to mixtures of 2% and 5% silane in hydrogen by volume; and inflow pressure and Mach number made by changing the scramjet model mounting angle in the HYPULSE test chamber. Data sources were wall pressures and heat flux distributions and schlieren and fuel plume imaging in the combustor/nozzle sections. Data are presented for calibration of the facility nozzles and the scramjet engine model. Comparisons of pressure distributions and flowpath streamtube performance estimates are made for the three Mach numbers tested.

Optimum flight paths of turbojet aircraft

NASA Technical Reports Server (NTRS)

Miele, Angelo

1955-01-01

The climb of turbojet aircraft is analyzed and discussed including the accelerations. Three particular flight performances are examined: minimum time of climb, climb with minimum fuel consumption, and steepest climb. The theoretical results obtained from a previous study are put in a form that is suitable for application on the following simplifying assumptions: the Mach number is considered an independent variable instead of the velocity; the variations of the airplane mass due to fuel consumption are disregarded; the airplane polar is assumed to be parabolic; the path curvatures and the squares of the path angles are disregarded in the projection of the equation of motion on the normal to the path; lastly, an ideal turbojet with performance independent of the velocity is involved. The optimum Mach number for each flight condition is obtained from the solution of a sixth order equation in which the coefficients are functions of two fundamental parameters: the ratio of minimum drag in level flight to the thrust and the Mach number which represents the flight at constant altitude and maximum lift-drag ratio.

The effect of perforations on the ballistics of a flare-stabilized projectile

NASA Astrophysics Data System (ADS)

Mermagen, W. H.; Yalamanchili, R. J.

Flight tests of two 35/105 mm flare stabilized projectile configurations were conducted. The projectiles were fired from the M68 tank cannon at Mach 4.0 using a standard M735 sabot. Flight data were obtained with a doppler velocimeter. The perforated flare projectiles showed a reduced drag for M greater than 1.7, which increased significantly at velocities below M 1.7, while the solid flare rounds had a slightly higher drag above Mach 1.7. Below Mach 1.7, the solid flare rounds had less drag than the perforated flare round. Both solid and perforated flare projectiles had a maximum range of less than 8.0 kilometers. The effects of the perforations on the flight performance were small. The flight data are compared to previous tests of the German-made 'LKL' projectile. The differences in flight performance between solid flares, flares with perforations, and the LKL-perforated flare rounds were small and of no practical consequence. Dispersion tests of the solid-flare stabilized projectiles were conducted at one, two, and three kilometers with excellent results. Very low dispersions were observed at all ranges.

Techniques for hot structures testing

NASA Technical Reports Server (NTRS)

Deangelis, V. Michael; Fields, Roger A.

1990-01-01

Hot structures testing have been going on since the early 1960's beginning with the Mach 6, X-15 airplane. Early hot structures test programs at NASA-Ames-Dryden focused on operational testing required to support the X-15 flight test program, and early hot structures research projects focused on developing lab test techniques to simulate flight thermal profiles. More recent efforts involved numerous large and small hot structures test programs that served to develop test methods and measurement techniques to provide data that promoted the correlation of test data with results from analytical codes. In Nov. 1988 a workshop was sponsored that focused on the correlation of hot structures test data with analysis. Limited material is drawn from the workshop and a more formal documentation is provided of topics that focus on hot structures test techniques used at NASA-Ames-Dryden. Topics covered include the data acquisition and control of testing, the quartz lamp heater systems, current strain and temperature sensors, and hot structures test techniques used to simulate the flight thermal environment in the lab.

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

NASA Technical Reports Server (NTRS)

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

1997-01-01

A Rocket Based Combined Cycle (RBCC) engine system is designed to combine the high thrust to weight ratio of a rocket along with the high specific impulse of a ramjet in a single, integrated propulsion system. This integrated, combined cycle propulsion system is designed to provide higher vehicle performance than that achievable with a separate rocket and ramjet. The RBCC engine system studied in the current program is the Aerojet strutjet engine concept, which is being developed jointly by a government-industry team as part of the Air Force HyTech program pre-PRDA activity. The strutjet is an ejector-ramjet engine in which small rocket chambers are embedded into the trailing edges of the inlet compression struts. The engine operates as an ejector-ramjet from take-off to slightly above Mach 3. Above Mach 3 the engine operates as a ramjet and transitions to a scramjet at high Mach numbers. For space launch applications the rockets would be re-ignited at a Mach number or altitude beyond which air-breathing propulsion alone becomes impractical. The focus of the present study is to develop and demonstrate a strutjet flowpath using hydrocarbon fuel at up to Mach 7 conditions. Freejet tests of a candidate flowpath for this RBCC engine were conducted at the NASA Lewis Research Center's Hypersonic Tunnel Facility between July and September 1996. This paper describes the engine flowpath and installation, outlines the primary objectives of the program, and describes the overall results of this activity. Through this program 15 full duration tests, including 13 fueled tests were made. The first major achievement was the further demonstration of the HTF capability. The facility operated at conditions up to 1950 K and 7.34 MPa, simulating approximately Mach 6.6 flight. The initial tests were unfueled and focused on verifying both facility and engine starting. During these runs additional aerodynamic appliances were incorporated onto the facility diffuser to enhance starting. Both facility and engine starting were achieved. Further, the static pressure distributions compared well with the results previously obtained in a 40% subscale flowpath study conducted in the LERC 1X1 supersonic wind tunnel (SWT), as well as the results of CFD analysis. Fueled performance results were obtained for the engine at both simulated Mach 6 (1670 K) and Mach 6.6 (1950 K) conditions. For all these tests the primary fuel was liquid JP-10 with gaseous silane (a mixture of 20% SiH4 and 80% H2 by volume) as an ignitor/pilot. These tests verified performance of this engine flowpath in a freejet mode. High combustor pressures were reached and significant changes in axial force were achieved due to combustion. Future test plans include redistributing the fuel to improve mixing, and consequently performance, at higher equivalence ratios.

Pegasus Rocket Booster Being Prepared for X-43A/Hyper-X Flight Test

NASA Image and Video Library

1999-08-25

A close-up view of the front end of a Pegasus rocket booster being prepared by technicians at the Dryden Flight Research Center for flight tests with the X-43A "Hypersonic Experimental Vehicle," or "Hyper-X." The X-43A, which will be attached to the Pegasus booster and drop launched from NASA's B-52 mothership, was developed to research 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).

Flight investigation of an air-cooled plug nozzle with afterburning turbojet

NASA Technical Reports Server (NTRS)

Samanich, N. E.

1972-01-01

A convectively cooled plug nozzle, using 4 percent of the engine air as the coolant, was tested in 1967 K (3540 R) temperature exhaust gas. No significant differences in cooling characteristics existed between flight and static results. At flight speeds above Mach 1.1, nozzle performance was improved by extending the outer shroud. Increasing engine power improved nozzle efficiency considerably more at Mach 1.2 than at 0.9. The effect of nozzle pressure ratio and secondary weight flow on nozzle performance are also presented.

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

NASA Technical Reports Server (NTRS)

2001-01-01

The X-43A hypersonic research aircraft and its modified Pegasus booster rocket are nestled under the wing of NASA's NB-52B carrier aircraft during pre-flight systems testing at the Dryden Flight Research Center, Edwards, Calif. The combined systems test was one of the last major milestones in the Hyper-X research program before the first X-43A flight. The X-43A flights will be the first actual flight tests of an aircraft powered by a revolutionary supersonic-combustion ramjet ('scramjet') engine capable of operating at hypersonic speeds (above Mach 5, or five times the speed of sound). The 12-foot, unpiloted research vehicle was developed and built by MicroCraft Inc., Tullahoma, Tenn., under NASA contract. The booster was built by Orbital Sciences Corp., Dulles, Va. After being air-launched from NASA's venerable NB-52 mothership, the booster will accelerate the X-43A to test speed and altitude. The X-43A will then separate from the rocket and fly a pre-programmed trajectory, conducting aerodynamic and propulsion experiments until it descends into the Pacific Ocean. Three research flights are planned, two at Mach 7 and one at Mach 10.

Analysis of Crossflow Transition Flight Experiment aboard the Pegasus Launch Vehicle

NASA Technical Reports Server (NTRS)

Malik, Mujeeb R.; Li, Fei; Choudhan, Meelan

2007-01-01

The Pegasus wing-glove flight experiment was designed to provide crossflow transition data at high Mach numbers, specifically to help validate stability based predictions for transition onset in a flight environment. This paper provides an analysis of the flight experiment, with emphasis on computational results for crossflow disturbances and the correlation of disturbance growth factors with in-flight transition locations via the e(sup N) method. Implications of the flight data for attachment line stability are also examined. Analysis of the thermocouple data reveals that transition (from turbulent to laminar flow) was first detected during the ascending flight of the rocket when the free stream Mach number exceeded about 4. Therefore, computations have been performed for flight Mach numbers of 4.13, 4.35, 4.56 and 4.99. Due to continually decreasing unit Reynolds number at higher altitudes, the entire wing-glove boundary layer became laminar at the highest flight Mach number computed above. In contrast, the boundary layer flow over the inboard tile region remained transitional up to and somewhat beyond the time of laminarization over the instrumented glove region. Linear stability predictions confirmed that the tile boundary layer is indeed more unstable to crossflow disturbances than the much colder stainless steel glove boundary layer. The transition locations based on thermocouple data from both the glove and the tile regions are found to correlate with stationary-crossflow N-factors within the range of 7 to 12.4 and with traveling mode N-factors between 7.6 and 14.1. Data from the thermocouples and hot film sensors indicates that transition from turbulent to laminar flow (i.e., laminarization) at a fixed point over the glove is generally completed within a flight time interval of 3 seconds. However, the times at which transition begins and ends as inferred from the hot film sensors are found to differ by about 2 seconds from the corresponding estimates based on the thermocouple data.

Flight Test Results from the Rake Airflow Gage Experiment on the F-15B Airplane

NASA Technical Reports Server (NTRS)

Frederick, Michael A.; Ratnayake, Nalin A.

2010-01-01

The Rake Airflow Gage Experiment involves a flow-field survey rake that was flown on the Propulsion Flight Test Fixture at the NASA Dryden Flight Research Center using the Dryden F-15B research test bed airplane. The objective of this flight test was to ascertain the flow-field angularity, local Mach number profile, total pressure distortion, and dynamic pressure at the aerodynamic interface plane of the Channeled Centerbody Inlet Experiment. This new mixed-compression, supersonic inlet is planned for flight test in the near term. Knowledge of the flow-field characteristics at this location underneath the airplane is essential to flight test planning and computational modeling of the new inlet, and it is also applicable for future propulsion systems research that may use the Propulsion Flight Test Fixture. This report describes the flight test preparation and execution, and the local flowfield properties calculated from pressure measurements of the rake. Data from the two Rake Airflow Gage Experiment research flights demonstrate that the F-15B airplane, flying at a free-stream Mach number of 1.65 and a pressure altitude of 40,000 ft, would achieve the desired local Mach number for the future inlet flight test. Interface plane distortion levels of 2 percent and a local angle of attack of 2 were observed at this condition. Alternative flight conditions for future testing and an exploration of certain anomalous data also are provided.

Flight Test Results from the Rake Airflow Gage Experiment on the F-15B Airplane

NASA Technical Reports Server (NTRS)

Frederick, Michael A.; Ratnayake, Nalin A.

2011-01-01

The Rake Airflow Gage Experiment involves a flow-field survey rake that was flown on the Propulsion Flight Test Fixture at the NASA Dryden Flight Research Center using the Dryden F-15B research test bed airplane. The objective of this flight test was to ascertain the flow-field angularity, local Mach number profile, total pressure distortion, and dynamic pressure at the aerodynamic interface plane of the Channeled Centerbody Inlet Experiment. This new mixed-compression, supersonic inlet is planned for flight test in the near term. Knowledge of the flow-field characteristics at this location underneath the airplane is essential to flight test planning and computational modeling of the new inlet, an< it is also applicable for future propulsion systems research that may use the Propulsion Flight Test Fixture. This report describes the flight test preparation and execution, and the local flow-field properties calculated from pressure measurements of the rake. Data from the two Rake Airflow Gage Experiment research flights demonstrate that the F-15B airplane, flying at a free-stream Mach number of 1.65 and a pressure altitude of 40,000 ft, would achieve the desired local Mach number for the future inlet flight test. Interface plane distortion levels of 2 percent and a local angle of attack of -2 deg were observed at this condition. Alternative flight conditions for future testing and an exploration of certain anomalous data also are provided.

Full-Envelope Launch Abort System Performance Analysis Methodology

NASA Technical Reports Server (NTRS)

Aubuchon, Vanessa V.

2014-01-01

The implementation of a new dispersion methodology is described, which dis-perses abort initiation altitude or time along with all other Launch Abort System (LAS) parameters during Monte Carlo simulations. In contrast, the standard methodology assumes that an abort initiation condition is held constant (e.g., aborts initiated at altitude for Mach 1, altitude for maximum dynamic pressure, etc.) while dispersing other LAS parameters. The standard method results in large gaps in performance information due to the discrete nature of initiation conditions, while the full-envelope dispersion method provides a significantly more comprehensive assessment of LAS abort performance for the full launch vehicle ascent flight envelope and identifies performance "pinch-points" that may occur at flight conditions outside of those contained in the discrete set. The new method has significantly increased the fidelity of LAS abort simulations and confidence in the results.

Advanced Computational Techniques for Hypersonic Propulsion

NASA Technical Reports Server (NTRS)

Povinelli, Louis A.

1996-01-01

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

Impact of Fluidic Chevrons on Jet Noise

NASA Technical Reports Server (NTRS)

Henderson, Brenda S.; Kinzie, Kevin W.; Whitmire, Julia; Abeysinghe, Amal

2005-01-01

The impact of alternating fluidic core chevrons on the production of jet noise is investigated. Core nozzles for a representative 1/9th scale, bypass ratio 5 model system were manufactured with slots cut near the trailing edges to allow for air injection into the core and fan streams. The injectors followed an alternating pattern around the nozzle perimeter so that the injection alternated between injection into the core stream and injection into the fan stream. For the takeoff condition and a forward flight Mach number of 0.10, the overall sound pressure levels at the peak jet noise angle decrease with increasing injection pressure. Sound pressure levels increase for observation angles less than 110o at higher injection pressures due to increases in high frequency noise. Greater increases in high frequency noise are observed when the number of injectors increases from 8 to 12. When the forward flight Mach number is increased to 0.28, jet noise reduction (relative to the baseline) is observed at aft angles for increasing injection pressure while significant increases in jet noise are observed at forward observation angles due to substantial acoustic radiation at high frequencies. A comparison between inflow and alternating injectors shows that, for equal mass injection rates, the inflow nozzle produces greater low frequency noise reduction (relative to the baseline) than the alternating injectors at 90o and aft observation angles and a forward flight Mach number of 0.28. Preliminary computational fluid dynamic simulations indicate that the spatial decay rate of the hot potential core flow is less for the inflow nozzle than for the alternating nozzles which indicates that gentle mixing may be preferred over sever mixing when fluidic chevrons are used for jet noise reduction.

Flight tests of external modifications used to reduce blunt base drag

NASA Technical Reports Server (NTRS)

Powers, Sheryll Goecke

1988-01-01

The effectiveness of a trailing disk (the trapped vortex concept) in reducing the blunt base drag of an 8-in diameter body of revolution was studied from measurements made both in flight and in full-scale wind-tunnel tests. The experiment demonstrated the significant base drag reduction capability of the trailing disk to Mach 0.93. The maximum base drag reduction obtained from a cavity tested on the flight body of revolution was not significant. The effectiveness of a splitter plate and a vented-wall cavity in reducing the base drag of a quasi-two-dimensional fuselage closure was studied from base pressure measurements made in flight. The fuselage closure was between the two engines of the F-111 airplane; therefore, the base pressures were in the presence of jet engine exhaust. For Mach numbers from 1.10 to 1.51, significant base drag reduction was provided by the vented-wall cavity configuration. The splitter plate was not considered effective in reducing base drag at any Mach number tested.

The Prediction of Noise Due to Jet Turbulence Convecting Past Flight Vehicle Trailing Edges

NASA Technical Reports Server (NTRS)

Miller, Steven A. E.

2014-01-01

High intensity acoustic radiation occurs when turbulence convects past airframe trailing edges. A mathematical model is developed to predict this acoustic radiation. The model is dependent on the local flow and turbulent statistics above the trailing edge of the flight vehicle airframe. These quantities are dependent on the jet and flight vehicle Mach numbers and jet temperature. A term in the model approximates the turbulent statistics of single-stream heated jet flows and is developed based upon measurement. The developed model is valid for a wide range of jet Mach numbers, jet temperature ratios, and flight vehicle Mach numbers. The model predicts traditional trailing edge noise if the jet is not interacting with the airframe. Predictions of mean-flow quantities and the cross-spectrum of static pressure near the airframe trailing edge are compared with measurement. Finally, predictions of acoustic intensity are compared with measurement and the model is shown to accurately capture the phenomenon.

Assessment of Aerothermodynamic Flight Prediction Tools through Ground and Flight Experimentation (Evaluation des outils aerothermodynamiques de prediction de vol par l’experimentation au sol et en vol)

DTIC Science & Technology

2011-11-01

6.5 Conclusions 6-30 6.6 Acknowledgments 6-31 6.7 References 6-31 Chapter 7 – Experimental Investigation of the Supersonic Wake of a Re-entry 7-1... Noise on the Axial Location of 4-4 Transition for the HIFiRE-1 Cone at Zero Angle of Attack and Mach 6 Figure 4-2 Correlations for Transition...Without Sting) Mach 2 AoA 19 Symmetry Plane Computed 6-14 with LORE: Effect of Sting/ Blade vs. No Sting/ Blade on 8M Cells Mesh Figure 6-12 Mach 2

Measurements of Aerodynamic Heat Transfer and Boundary-Layer Transition on a 10 deg Cone in Free Flight at Supersonic Mach Numbers up to 5.9

NASA Technical Reports Server (NTRS)

Rumsey, Charles B.; Lee, Dorothy B.

1961-01-01

Measurements of aerodynamic heat transfer have been made at six stations on the 40-inch-long 10 deg. total-angle conical nose of a rocket- propelled model which was flight tested at Mach numbers up to 5.9. are presented for a range of local Mach number just outside the bound- ary layer on the cone from 1.57 to 5.50, and a range of local Reynolds number from 6.6 x 10(exp 6) to 55.2 x 10(exp 6) based on length from the nose tip.

Mach Cutoff Analysis and Results from NASA's Farfield Investigation of No-Boom Thresholds

NASA Technical Reports Server (NTRS)

Cliatt, Larry J., II; Hill, Michael A.; Haering, Edward A, Jr.

2016-01-01

In support of the ongoing effort by the National Aeronautics and Space Administration (NASA) to bring supersonic commercial travel to the public, the NASA Armstrong Flight Research Center and the NASA Langley Research Center,in partnership with other industry organizations and academia, conducted a flight research experiment to analyze acoustic propagation in the Mach cutoff shadow zone. The effort was conducted in fall of 2012 and named the Farfield Investigation of No-boom Thresholds (FaINT). The test helped to build a data set that will go toward further understanding of the unique acoustic propagation characteristics below Mach cutoff altitude.

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

NASA Image and Video Library

2001-03-15

The X-43A hypersonic research aircraft and its modified Pegasus® booster rocket are nestled under the wing of NASA's NB-52B carrier aircraft during pre-flight systems testing at the Dryden Flight Research Center, Edwards, Calif. The combined systems test was one of the last major milestones in the Hyper-X research program before the first X-43A flight. The X-43A flights will be the first actual flight tests of an aircraft powered by a revolutionary supersonic-combustion ramjet ("scramjet") engine capable of operating at hypersonic speeds (above Mach 5, or five times the speed of sound). The 12-foot, unpiloted research vehicle was developed and built by MicroCraft Inc., Tullahoma, Tenn., under NASA contract. The booster was built by Orbital Sciences Corp., Dulles, Va. After being air-launched from NASA's venerable NB-52 mothership, the booster will accelerate the X-43A to test speed and altitude. The X-43A will then separate from the rocket and fly a pre-programmed trajectory, conducting aerodynamic and propulsion experiments until it descends into the Pacific Ocean. Three research flights are planned, two at Mach 7 and one at Mach 10.

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.

Stability and Control Estimation Flight Test Results for the SR-71 Aircraft With Externally Mounted Experiments

NASA Technical Reports Server (NTRS)

Moes, Timothy R.; Iliff, Kenneth

2002-01-01

A maximum-likelihood output-error parameter estimation technique is used to obtain stability and control derivatives for the NASA Dryden Flight Research Center SR-71A airplane and for configurations that include experiments externally mounted to the top of the fuselage. This research is being done as part of the envelope clearance for the new experiment configurations. Flight data are obtained at speeds ranging from Mach 0.4 to Mach 3.0, with an extensive amount of test points at approximately Mach 1.0. Pilot-input pitch and yaw-roll doublets are used to obtain the data. This report defines the parameter estimation technique used, presents stability and control derivative results, and compares the derivatives for the three configurations tested. The experimental configurations studied generally show acceptable stability, control, trim, and handling qualities throughout the Mach regimes tested. The reduction of directional stability for the experimental configurations is the most significant aerodynamic effect measured and identified as a design constraint for future experimental configurations. This report also shows the significant effects of aircraft flexibility on the stability and control derivatives.

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.

Mach 5 to 7 RBCC Propulsion System Testing at NASA-LeRC HTF

NASA Technical Reports Server (NTRS)

Perkins, H. Douglas; Thomas, Scott R.; Pack, William D.

1996-01-01

A series of Mach 5 to 7 freejet tests of a Rocket Based Combined Cycle (RBCC) engine were cnducted at the NASA Lewis Research Center (LERC) Hypersonic Tunnel Facility (HTF). This paper describes the configuration and operation of the HTF and the RBCC engine during these tests. A number of facility support systems are described which were added or modified to enhance the HTF test capability for conducting this experiment. The unfueled aerodynamic perfor- mance of the RBCC engine flowpath is also presented and compared to sub-scale test results previously obtained in the NASA LERC I x I Supersonic Wind Tunnel (SWT) and to Computational Fluid Dynamic (CFD) analysis results. This test program demonstrated a successful configuration of the HTF for facility starting and operation with a generic RBCC type engine and an increased range of facility operating conditions. The ability of sub-scale testing and CFD analysis to predict flowpath performance was also shown. The HTF is a freejet, blowdown propulsion test facility that can simulate up to Mach 7 flight conditions with true air composition. Mach 5, 6, and 7 facility nozzles are available, each with an exit diameter of 42 in. This combination of clean air, large scale, and Mach 7 capabilities is unique to the HTF. This RBCC engine study is the first engine test program conducted at the HTF since 1974.

Flight Test of the Engine Fuel Schedules of the X-43A Hyper-X Research Vehicles

NASA Technical Reports Server (NTRS)

Jones, Thomas

2006-01-01

The Hyper-X program flew two X-43A Hyper-X Research Vehicles (HXRVs) in 2004, referred to as Ship 2 and Ship 3. The scramjet engine of the X-43A research vehicle was autonomously controlled in flight to track a predetermined fueling schedule. Ship 2 flew at approximately Mach 7 and Ship 3 flew at approximately Mach 10.

Flight Determination of the Longitudinal Stability Characteristics of a 0.133-Scale Rocket-Powered Model of the Consolidated Vultee XFY-1 Airplane without Propellers at Mach Numbers from 0.73 to 1.19, TED No. NACA DE 369

NASA Technical Reports Server (NTRS)

Hastings, Earl E., Jr.; Mitcham, Grady L.

1954-01-01

A flight test has been conducted to determine the longitudinal stability and control,characteristics of a 0.133-scale model of the Consolidated Vultee XFY-1 airplane without propellers for the Mach number range between 0.73 and 1.19.

Measurements in Flight of the Longitudinal-Stability Characteristics of a Republic YF-84A Airplane (Army Serial No. 45-59488) at High Subsonic Mach Numbers

NASA Technical Reports Server (NTRS)

Turner, Howard L.; Cooper, George E.

1948-01-01

A brief investigation was made of the longitudinal-stability characteristics of a YF-84A airplane (Army Serial No. 45-79488). The airplane developed a pitching-up tendency at approximately 0.80 Mach number which necessitated large push forces and down-elevator deflections for further increases in speed. In steady turns at 35,000 feet with the center of gravity at 28.3 percent mean aerodynamic chord for normal accelerations up to the maximum test value, the control-force gradients were excessive at Mach numbers over 0.78. Airplane buffeting did not present a serious problem in accelerated or unaccelerated flight at 15,000 and 35,000 feet up to the maximum test Mach number of 0.84. It is believed that excessive control force would be the limiting factor in attaining speeds in excess of 0.84 Mach number, especially at altitudes below 35,000 feet.

Flight control systems development and flight test experience with the HiMAT research vehicles

NASA Technical Reports Server (NTRS)

Kempel, Robert W.; Earls, Michael R.

1988-01-01

Two highly maneuverable aircraft technology (HiMAT) remotely piloted vehicles were flown a total of 26 flights. These subscale vehicles were of advanced aerodynamic configuration with advanced technology concepts such as composite and metallic structures, digital integrated propulsion control, and ground (primary) and airborne (backup) relaxed static stability, digital fly-by-wire control systems. Extensive systems development, checkout, and flight qualification were required to conduct the flight test program. The design maneuver goal was to achieve a sustained 8-g turn at Mach 0.9 at an altitude of 25,000 feet. This goal was achieved, along with the acquisition of high-quality flight data at subsonic and supersonic Mach numbers. Control systems were modified in a variety of ways using the flight-determined aerodynamic characteristics. The HiMAT program was successfully completed with approximately 11 hours of total flight time.

Global Infrared Observations of Roughness Induced Transition on the Space Shuttle Orbiter

NASA Technical Reports Server (NTRS)

Horvath, Thomas J.; Zalameda, Joseph N.; Wood, William A.; Berry, Scott A.; Schwartz, Richard J.; Dantowitz, Ronald F.; Spisz, Thomas S.; Taylor, Jeff C.

2012-01-01

High resolution infrared observations made from a mobile ground based optical system captured the laminar-to-turbulent boundary layer transition process as it occurred during Space Shuttle Endeavour's return to earth following its final mission in 2011. The STS-134 imagery was part of a larger effort to demonstrate an emerging and reliable non-intrusive global thermal measurement capability and to complement a series of boundary layer transition flight experiments that were flown on the Shuttle. The STS-134 observations are believed to be the first time that the development and movement of a hypersonic boundary layer transition front has been witnessed in flight over the entire vehicle surface and in particular, at unprecedented spatial resolution. Additionally, benchmark surface temperature maps of the Orbiter lower surface collected over multiple flights and spanning a Mach range of 18 to 6 are now available and represent an opportunity for collaborative comparison with computational techniques focused on hypersonic transition and turbulence modeling. The synergy of the global temperature maps with the companion in-situ thermocouple measurements serve as an example of the effective leveraging of resources to achieve a common goal of advancing our understanding of the complex nature of high Mach number transition. It is shown that quantitative imaging can open the door to a multitude of national and international opportunities for partnership associated with flight-testing and subsequent validation of numerical simulation techniques. The quantitative imaging applications highlighted in this paper offer unique and complementary flight measurement alternatives and suggest collaborative instrumentation opportunities to advance the state of the art in transition prediction and maximize the return on investment in terms of developmental flight tests for future vehicle designs.

In-Flight Capability for Evaluating Skin-Friction Gages and Other Near-Wall Flow Sensors

NASA Technical Reports Server (NTRS)

Bui, Trong T.; Pipitone, Brett J.; Krake, Keith L.; Richwine, Dave (Technical Monitor)

2003-01-01

An 8-in.-square boundary-layer sensor panel has been developed for in-flight evaluation of skin-friction gages and other near-wall flow sensors on the NASA Dryden Flight Research Center F-15B/Flight Test Fixture (FTF). Instrumentation on the sensor panel includes a boundary-layer rake, temperature sensors, static pressure taps, and a Preston tube. Space is also available for skin-friction gages or other near-wall flow sensors. Pretest analysis of previous F-15B/FTF flight data has identified flight conditions suitable for evaluating skin-friction gages. At subsonic Mach numbers, the boundary layer over the sensor panel closely approximates the two-dimensional (2D), law-of-the-wall turbulent boundary layer, and skin-friction estimates from the Preston tube and the rake (using the Clauser plot method) can be used to evaluate skin-friction gages. At supersonic Mach numbers, the boundary layer over the sensor panel becomes complex, and other means of measuring skin friction are needed to evaluate the accuracy of new skin-friction gages. Results from the flight test of a new rubber-damped skin-friction gage confirm that at subsonic Mach numbers, nearly 2D, law-of-the-wall turbulent boundary layers exist over the sensor panel. Sensor panel data also show that this new skin-friction gage prototype does not work in flight.

OVRhyp, Scramjet Test Aircraft

NASA Technical Reports Server (NTRS)

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

1990-01-01

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

Use of Heated Helium to Simulate Surface Pressure Fluctuations on the Launch Abort Vehicle During Abort Motor Firing

NASA Technical Reports Server (NTRS)

Panda, Jayanta; James, George H.; Burnside, Nathan J.; Fong, Robert; Fogt, Vincent A.

2011-01-01

The solid-rocket plumes from the Abort motor of the Multi-Purpose Crew Vehicle (MPCV, also know as Orion) were simulated using hot, high pressure, Helium gas to determine the surface pressure fluctuations on the vehicle in the event of an abort. About 80 different abort situations over a wide Mach number range, (0.3< or =M< or =1.2) and vehicle attitudes (+/-15deg) were simulated inside the NASA Ames Unitary Plan, 11-Foot Transonic Wind Tunnel. For each abort case, typically two different Helium plume and wind tunnel conditions were used to bracket different flow matching critera. This unique, yet cost-effective test used a custom-built hot Helium delivery system, and a 6% scale model of a part of the MPCV, known as the Launch Abort Vehicle. The test confirmed the very high level of pressure fluctuations on the surface of the vehicle expected during an abort. In general, the fluctuations were found to be dominated by the very near-field hydrodynamic fluctuations present in the plume shear-layer. The plumes were found to grow in size for aborts occurring at higher flight Mach number and altitude conditions. This led to an increase in the extent of impingement on the vehicle surfaces; however, unlike some initial expectations, the general trend was a decrease in the level of pressure fluctuations with increasing impingement. In general, the highest levels of fluctuations were found when the outer edges of the plume shear layers grazed the vehicle surface. At non-zero vehicle attitudes the surface pressure distributions were found to become very asymmetric. The data from these wind-tunnel simulations were compared against data collected from the recent Pad Abort 1 flight test. In spite of various differences between the transient flight situation and the steady-state wind tunnel simulations, the hot-Helium data were found to replicate the PA1 data fairly reasonably. The data gathered from this one-of-a-kind wind-tunnel test fills a gap in the manned-space programs, and will be used to establish the acoustic environment for vibro-acoustic qualification testing of the MPCV.

Development of a Flush Airdata Sensing System on a Sharp-Nosed Vehicle for Flight at Mach 3 to 8

NASA Technical Reports Server (NTRS)

Davis, Mark C.; Pahle, Joseph W.; White, John Terry; Marshall, Laurie A.; Mashburn, Michael J.; Franks, Rick

2000-01-01

NASA Dryden Flight Research Center has developed a flush airdata sensing (FADS) system on a sharp-nosed, wedge-shaped vehicle. This paper details the design and calibration of a real-time angle-of-attack estimation scheme developed to meet the onboard airdata measurement requirements for a research vehicle equipped with a supersonic-combustion ramjet engine. The FADS system has been designed to perform in flights at Mach 3-8 and at -6 deg - 12 deg angle of attack. The description of the FADS architecture includes port layout, pneumatic design, and hardware integration. Predictive models of static and dynamic performance are compared with wind-tunnel results across the Mach and angle-of-attack range. Results indicate that static angle-of-attack accuracy and pneumatic lag can be adequately characterized and incorporated into a real-time algorithm.

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

NASA Technical Reports Server (NTRS)

Bobbitt, Percy J.; Maglieri, Domenic J.; Banks, Daniel W.; Frederick, Michael A.; Fuchs, Aaron W.

2012-01-01

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.

Flightweight radiantly and actively cooled panel: Thermal and structural performance

NASA Technical Reports Server (NTRS)

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

1982-01-01

A 2- by 4-ft flightweight panel was subjected to thermal/structural tests representative of design flight conditions for a Mach 6.7 transport and to off-design conditions simulating flight maneuvers and cooling system failures. The panel utilized Rene 41 heat shields backed by a thin layer of insulation to radiate away most of the 12 Btu/ft2-sec incident heating. A solution of ethylene glycol in water circulating through tubes in an aluminum-honeycomb-sandwich panel absorbed the remainder of the incident heating (0.8 Btu/sq ft-sec). The panel successfully withstood (1) 46.7 hr of radiant heating which included 53 thermal cycles and 5000 cycles of uniaxial inplane loading of + or - 1200 lfb/in; (2) simulated 2g-maneuver heating conditions and simulated cooling system failures without excessive temperatures on the structural panel; and (3) the extensive thermal/structural tests and the aerothermal tests reported in NASA TP-1595 without significant damage to the structural panel, coolant leaks, or hot-gas ingress to the structural panel.

Computation of losses in a scramjet combustor

NASA Technical Reports Server (NTRS)

Kamath, Pradeep S.; Mcclinton, Charles R.

1992-01-01

The losses in a conceptual scramjet combustor at flight Mach numbers of 8, 10, 12, 16 and 20 are computed. These losses are extracted from three-dimensional parabolized Navier-Stokes solutions of the turbulent, reacting combustor flow field. A combustor performance index was defined based on the rationale that an efficient scramjet combustor should add heat to the fluid in such a manner as to maximize the stream thrust at the combustor exit while minimizing the losses. This index showed a decrease of more than 40 percent as the flight Mach number increased from 8 to 20, indicative of a drop in the thrust-producing potential of the scramjet at the upper end of the speed regime studied. A breakdown of the losses showed that dissipation, nonequilibrium chemistry and heat diffusion contributed roughly 15 percent, 35 percent, and 50 percent to the irreversible increase in entropy at Mach 8 and 22 percent, 13 and 65 percent at Mach 20.

Turbulent-Spot Growth Characteristics: Wind-Tunnel and Flight Measurements of Natural Transition at High Reynolds and Mach Numbers

NASA Technical Reports Server (NTRS)

Clark, J. P.; Jones, T. V.; LaGraff, J. E.

2007-01-01

A series of experiments are described which examine the growth of turbulent spots on a flat plate at Reynolds and Mach numbers typical of gas-turbine blading. A short-duration piston tunnel is employed and rapid-response miniature surface-heat-transfer gauges are used to asses the state of the boundary layer. The leading- and trailing-edge velocities of spots are reported for different external pressure gradients and Mach numbers. Also, the lateral spreading angle is determined from the heat-transfer signals which demonstrate dramatically the reduction in spot growth associated with favorable pressure gradients. An associated experiment on the development of turbulent wedges is also reported where liquid-crystal heat-transfer techniques are employed in low-speed wind tunnel to visualize and measure the wedge characteristics. Finally, both liquid crystal techniques and hot-film measurements from flight tests at Mach number of 0.6 are presented.

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

NASA Technical Reports Server (NTRS)

2001-01-01

The first of three X-43A hypersonic research aircraft and its modified Pegasus booster rocket recently underwent combined systems testing while mounted to NASA's NB-52B carrier aircraft at the Dryden Flight Research Center, Edwards, Calif. The combined systems test was one of the last major milestones in the Hyper-X research program before the first X-43A flight. The X-43A flights will be the first actual flight tests of an aircraft powered by a revolutionary supersonic-combustion ramjet ('scramjet') engine capable of operating at hypersonic speeds (above Mach 5, or five times the speed of sound). The 12-foot, unpiloted research vehicle was developed and built by MicroCraft Inc., Tullahoma, Tenn., under NASA contract. The booster was built by Orbital Sciences Corp., Dulles, Va.,After being air-launched from NASA's venerable NB-52 mothership, the booster will accelerate the X-43A to test speed and altitude. The X-43A will then separate from the rocket and fly a pre-programmed trajectory, conducting aerodynamic and propulsion experiments until it descends into the Pacific Ocean. Three research flights are planned, two at Mach 7 and one at Mach 10.

Three-Component Force Measurements on a Scramjet in a Reflected-Shock Tunnel

NASA Technical Reports Server (NTRS)

Tsai, C.-Y.; Bakos, R. J.; Mee, D. J.

1998-01-01

A three-component stress-wave force-balance for a large scramjet has been designed, calibrated and tested in the HYPULSE reflected shock tunnel at GASL Inc., New York. The scramjet model is over 3-foot long and weighs in excess of 90 Ibm. The stress-wave force-balance is comprised of three stress bars which are attached to the model. Calibration results indicate that the force balance responds well within about 1 ms and that the sensitivity of the balance to the distribution of load is not large. Results with and without fuel injection were obtained in the tunnel operated for Mach 7 and Mach 10 flight simulation. These tests showed the force-balance can resolve axial force increments due to combustion of about 40 lb in the presence of model lift forces of 500-700 lb.

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.

Control Design for an Advanced Geared Turbofan Engine

NASA Technical Reports Server (NTRS)

Chapman, Jeffryes W.; Litt, Jonathan S.

2017-01-01

This paper describes the design process for the control system of an advanced geared turbofan engine. This process is applied to a simulation that is representative of a 30,000 pound-force thrust class concept engine with two main spools, ultra-high bypass ratio, and a variable area fan nozzle. Control system requirements constrain the non-linear engine model as it operates throughout its flight envelope of sea level to 40,000 feet and from 0 to 0.8 Mach. The purpose of this paper is to review the engine control design process for an advanced turbofan engine configuration. The control architecture selected for this project was developed from literature and reflects a configuration that utilizes a proportional integral controller with sets of limiters that enable the engine to operate safely throughout its flight envelope. Simulation results show the overall system meets performance requirements without exceeding operational limits.

Flight Investigation at High Speeds of Profile Drag of Wing of a P-47D Airplane Having Production Surfaces Covered with Camouflage Paint

NASA Technical Reports Server (NTRS)

Daum, Fred L.; Zalovcik, John A.

1946-01-01

Wing section outboard of flap was tested by wake surveys in Mach range of 0.25 - 0.78 and lift coefficient range 0.06 - 0.69. Results indicated that minimum profile-drag coefficient of 0.0097 was attained for lift coefficients from 0.16 to 0.25 at Mach less than 0.67. Below Mach number at which compressibility shock occurred, variations in Mach of 0.2 had negligible effect on profile drag coefficient. Shock was not evident until critical Mach was exceeded by 0.025.

Flight Calibration of four airspeed systems on a swept-wing airplane at Mach numbers up to 1.04 by the NACA radar-phototheodolite method

NASA Technical Reports Server (NTRS)

Thompson, Jim Rogers; Bray, Richard S; COOPER GEORGE E

1950-01-01

The calibrations of four airspeed systems installed in a North American F-86A airplane have been determined in flight at Mach numbers up to 1.04 by the NACA radar-phototheodolite method. The variation of the static-pressure error per unit indicated impact pressure is presented for three systems typical of those currently in use in flight research, a nose boom and two different wing-tip booms, and for the standard service system installed in the airplane. A limited amount of information on the effect of airplane normal-force coefficient on the static-pressure error is included. The results are compared with available theory and with results from wind-tunnel tests of the airspeed heads alone. Of the systems investigated, a nose-boom installation was found to be most suitable for research use at transonic and low supersonic speeds because it provided the greatest sensitivity of the indicated Mach number to a unit change in true Mach number at very high subsonic speeds, and because it was least sensitive to changes in airplane normal-force coefficient. The static-pressure error of the nose-boom system was small and constant above a Mach number of 1.03 after passage of the fuselage bow shock wave over the airspeed head.

Forebody and Inlet Design for the HIFiRE 2 Flight Test

NASA Technical Reports Server (NTRS)

Ferlemann, Paul G.

2008-01-01

A forebody and inlet have been designed for the HIFiRE 2 scramjet flight test. The test will explore the operating, performance, and stability characteristics of a simple hydrocarbon-fueled scramjet combustor as it transitions from dual-mode to scramjet-mode operation and during supersonic combustion at Mach 8+ flight conditions. Requirements for the compression system were derived from inlet starting and combustor inflow requirements as well as physical size constraints. The design process is described. A planar, fixed geometry, mixed compression concept was used to produce laterally uniform flow at the inlet entrance and a conservative amount of internal contraction with respect to inlet starting. A grid sensitivity study was performed so that important flow physics caused by three-dimensional shock boundary layer interactions could be captured with confidence. Results from low Mach number operability studies, nominal trajectory cases, and high dynamic pressure heat load cases are discussed. The forebody and inlet solutions provide information for on-going combustor calculations, mass capture across the trajectory for fuel system design, and surface heating rates for thermal/structural analysis. The design has a one freestream Mach number margin for inlet starting, exceeds the high Mach number combustor entrance pressure requirement, produces high quality flow at the inlet exit for all Mach numbers and vehicle attitudes in the design space, and fits inside the booster shroud.

Tests and Techniques for Characterizing and Modeling X-43A Electromechanical Actuators

NASA Technical Reports Server (NTRS)

Lin, Yohan; Baumann, Ethan; Bose, David M.; Beck, Roger; Jenney, Gavin

2008-01-01

A series of tests were conducted on the electromechanical actuators of the X-43A research vehicle in preparation for the Mach 7 and 10 hypersonic flights. The tests were required to help validate the actuator models in the simulation and acquire a better understanding of the installed system characteristics. Static and dynamic threshold, multichannel crosstalk, command-to-surface timing, free play, voltage regeneration, calibration, frequency response, compliance, hysteretic damping, and aircraft-in-the-loop tests were performed as part of this effort. This report describes the objectives, configurations, and methods for those tests, as well as the techniques used for developing second-order actuator models from the test results. When the first flight attempt failed because of actuator problems with the launch vehicle, further analysis and model enhancements were performed as part of the return-to-flight activities. High-fidelity models are described, along with the modifications that were required to match measurements taken from the research vehicle. Problems involving the implementation of these models into the X-43A simulation are also discussed. This report emphasizes lessons learned from the actuator testing, simulation modeling, and integration efforts for the X-43A hypersonic research vehicle.

Pegasus Rocket Booster Being Prepared for X-43A/Hyper-X Flight Test

NASA Image and Video Library

1999-08-25

Technicians prepare a Pegasus rocket booster for flight tests with the X-43A "Hypersonic Experimental Vehicle," or "Hyper-X." The X-43A, which will be attached to the Pegasus booster and drop launched from NASA's B-52 mothership, was developed to research 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).

Preliminary Altitude Performance Data of J71-A2 Turbojet Engine Afterburner

NASA Technical Reports Server (NTRS)

Useller, James W.; Mallett, William E.

1954-01-01

The performance and operational characteristics of the J71-A2 turbojet-engine afterburner were investigated for a range of altitudes from 23,000 to 60,000 feet at a flight Mach number of 0,9 and at flight Mach numbers of 0.6, 0.9, and 1.0 at an altitude of 45,000 feet. The combustion performance and altitude operational limits, as well as the altitude starting characteristics have been determined.

Use of a pitot-static probe for determining wing section drag in flight at Mach numbers from 0.5 to approximately 1.0

NASA Technical Reports Server (NTRS)

Montoya, L. C.; Economu, M. A.; Cissell, R. E.

1974-01-01

The use of a pitot-static probe to determine wing section drag at speeds from Mach 0.5 to approximately 1.0 was evaluated in flight. The probe unit is described and operational problems are discussed. Typical wake profiles and wing section drag coefficients are presented. The data indicate that the pitot-static probe gave reliable results up to speeds of approximately 1.0.

Preliminary airborne measurements for the SR-71 sonic boom propagation experiment

NASA Technical Reports Server (NTRS)

Haering, Edward A., Jr.; Ehernberger, L. J.; Whitmore, Stephen A.

1995-01-01

SR-71 sonic boom signatures were measured to validate sonic boom propagation prediction codes. An SR-71 aircraft generated sonic booms from Mach 1.25 to Mach 1.6, at altitudes of 31,000 to 48,000 ft, and at various gross weights. An F-16XL aircraft measured the SR-71 near-field shock waves from close to the aircraft to more than 8,000 ft below, gathering 105 signatures. A YO-3A aircraft measured the SR-71 sonic booms from 21,000 to 38,000 feet below, recording 17 passes. The sonic booms at ground level and atmospheric data were recorded for each flight. Data analysis is underway. Preliminary results show that shock wave patterns and coalescence vary with SR-71 gross weight, Mach number, and altitude. For example, noncoalesced shock wave signatures were measured by the YO-3A at 21,000 ft below the SR-71 aircraft while at a low gross weight, Mach 1.25, and 31,000-ft altitude. This paper describes the design and execution of the flight research experiment. Instrumentation and flight maneuvers of the SR-71, F-16XL, and YO-3A aircraft and sample sonic boom signatures are included.

Free-Flight Tests of 0.11-Scale North American F-100 Airplane Wings to Investigate the Possibility of Flutter in Transonic Speed Range at Varying Angles of Attack

NASA Technical Reports Server (NTRS)

O'Kelly, Burke R.

1954-01-01

Free-flight tests in the transonic speed range utilizing rocketpropelled models have been made on three pairs of 0.11-scale North American F-100 airplane wings having an aspect ratio of 3.47, a taper ratio of 0.308, 45 degree sweepback at the quarter-chord line, and thickness ratios of 31 and 5 percent to investigate the possibility of flutte r. Data from tests of two other rocket-propelled models which accidentally fluttered during a drag investigation of the North American F-100 airplane are also presented. The first set of wings (5 percent thick) was tested on a model which was disturbed in pitch by a moving tail and reached a maximum Mach number of 0.85. The wings encountered mild oscillations near the first - bending frequency at high lift coefficients. The second set of wings 9 percent thick was tested up to a maximum Mach number of 0.95 at (2) angles of attack provided by small rocket motors installed in the nose of the model. No oscillations resembling flutter were encountered during the coasting flight between separation from the booster and sustainer firing (Mach numbers from 0.86 to 0.82) or during the sustainer firing at accelerations of about 8g up to the maximum Mach number of the test (0.95). The third set of wings was similar to the first set and was tested up to a maximum Mach number of 1.24. A mild flutter at frequencies near the first-bending frequency of the wings was encountered between a Mach number of 1.15 and a Mach number of 1.06 during both accelerating and coasting flight. The two drag models, which were 0.ll-scale models of the North American F-100 airplane configuration, reached a maximum Mach number of 1.77. The wings of these models had bending and torsional frequencies which were 40 and 89 percent, respectively, of the calculated scaled frequencies of the full-scale 7-percent-thick wing. Both models experienced flutter of the same type as that experienced-by the third set of wings.

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

NASA Technical Reports Server (NTRS)

Axdahl, Erik; Kumar, Ajay; Wilhite, Alan

2012-01-01

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

Near-Field Acoustic Power Level Analysis of F31/A31 Open Rotor Model at Simulated Cruise Conditions, Technical Report II

NASA Technical Reports Server (NTRS)

Sree, Dave

2015-01-01

Near-field acoustic power level analysis of F31A31 open rotor model has been performed to determine its noise characteristics at simulated cruise flight conditions. The non-proprietary parts of the test data obtained from experiments in the 8x6 supersonic wind tunnel were provided by NASA-Glenn Research Center. The tone and broadband components of total noise have been separated from raw test data by using a new data analysis tool. Results in terms of sound pressure levels, acoustic power levels, and their variations with rotor speed, freestream Mach number, and input shaft power, with different blade-pitch setting angles at simulated cruise flight conditions, are presented and discussed. Empirical equations relating models acoustic power level and input shaft power have been developed. The near-field acoustic efficiency of the model at simulated cruise conditions is also determined. It is hoped that the results presented in this work will serve as a database for comparison and improvement of other open rotor blade designs and also for validating open rotor noise prediction codes.

X-33 Computational Aeroheating/Aerodynamic Predictions and Comparisons With Experimental Data

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Thompson, Richard A.; Berry, Scott A.; Horvath, Thomas J.; Murphy, Kelly J.; Nowak, Robert J.; Alter, Stephen J.

2003-01-01

This report details a computational fluid dynamics study conducted in support of the phase II development of the X-33 vehicle. Aerodynamic and aeroheating predictions were generated for the X-33 vehicle at both flight and wind-tunnel test conditions using two finite-volume, Navier-Stokes solvers. Aerodynamic computations were performed at Mach 6 and Mach 10 wind-tunnel conditions for angles of attack from 10 to 50 with body-flap deflections of 0 to 20. Additional aerodynamic computations were performed over a parametric range of free-stream conditions at Mach numbers of 4 to 10 and angles of attack from 10 to 50. Laminar and turbulent wind-tunnel aeroheating computations were performed at Mach 6 for angles of attack of 20 to 40 with body-flap deflections of 0 to 20. Aeroheating computations were performed at four flight conditions with Mach numbers of 6.6 to 8.9 and angles of attack of 10 to 40. Surface heating and pressure distributions, surface streamlines, flow field information, and aerodynamic coefficients from these computations are presented, and comparisons are made with wind-tunnel data.

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.

Implementation of a Single-Stage-To-Orbit (SSTO) model for stability and control analysis

NASA Astrophysics Data System (ADS)

Ingalls, Stephen A.

1995-07-01

Three NASA centers: Marshall Space Flight Center (MSFC), Langley Research Center (LaRC), and Johnson Space Center (JSC) are currently involved in studying a family of single-stage- and two-stage-to-orbit (SSTO/TSTO) vehicles to serve as the next generation space transportation system (STS). A rocketed winged-body is the current focus. The configuration (WB001) is a vertically-launched, horizontally-landing system with circular cross-section. Preliminary aerodynamic data was generated by LaRC and is a combination of wind-tunnel data, empirical methods, and Aerodynamic Preliminary Analysis System-(APAS) generated values. JSC's efforts involve descent trajectory design, stability analysis, and flight control system synthesis. Analysis of WB001's static stability indicates instability in 'tuck' (C(sub mu) less than 0: Mach = 0.30, alpha greater than 3.25 deg; Mach = 0.60, alpha greater than 8.04), an unstable dihedral effects (C(sub l(beta)) greater than 0: Mach = 30,alpha less than 12 deg.; Mach = 0.60, alpha less than 10.00 deg.), and, most significantly, an unstable weathercock stability derivative, C(sub n(beta)), at all angles of attack and subsonic Mach numbers. Longitudinal trim solutions for Mach = 0.30 and 0.60 indicate flight path angle possibilities ranging from around 12 (M = 0.30) to slightly over 20 degrees at Mach = 0.60. Trim angles of attack increase from 6.24 at Mach 0.60 and 10,000 feet to 17.7 deg. at Mach 0.30, sea-level. Lateral trim was attempted for a design cross-wind of 25.0 knots. The current vehicle aerodynamic and geometric characteristics will only yield a lateral trim solution at impractical tip-fin deflections (approximately equal to 43 deg.) and bank angles (21 deg.). A study of the lateral control surfaces, tip-fin controllers for WB001, indicate increased surface area would help address these instabilities, particularly the deficiency in C(sub n(beta)), but obviously at the expense of increased vehicle weight. Growth factors of approximately 7 were determined using a design C(sub n(beta)) of 0.100/radian (approximate subsonic values for the orbiter).

Implementation of a Single-Stage-To-Orbit (SSTO) model for stability and control analysis

NASA Technical Reports Server (NTRS)

Ingalls, Stephen A.

1995-01-01

Three NASA centers: Marshall Space Flight Center (MSFC), Langley Research Center (LaRC), and Johnson Space Center (JSC) are currently involved in studying a family of single-stage- and two-stage-to-orbit (SSTO/TSTO) vehicles to serve as the next generation space transportation system (STS). A rocketed winged-body is the current focus. The configuration (WB001) is a vertically-launched, horizontally-landing system with circular cross-section. Preliminary aerodynamic data was generated by LaRC and is a combination of wind-tunnel data, empirical methods, and Aerodynamic Preliminary Analysis System-(APAS) generated values. JSC's efforts involve descent trajectory design, stability analysis, and flight control system synthesis. Analysis of WB001's static stability indicates instability in 'tuck' (C(sub mu) less than 0: Mach = 0.30, alpha greater than 3.25 deg; Mach = 0.60, alpha greater than 8.04), an unstable dihedral effects (C(sub l(beta)) greater than 0: Mach = 30,alpha less than 12 deg.; Mach = 0.60, alpha less than 10.00 deg.), and, most significantly, an unstable weathercock stability derivative, C(sub n(beta)), at all angles of attack and subsonic Mach numbers. Longitudinal trim solutions for Mach = 0.30 and 0.60 indicate flight path angle possibilities ranging from around 12 (M = 0.30) to slightly over 20 degrees at Mach = 0.60. Trim angles of attack increase from 6.24 at Mach 0.60 and 10,000 feet to 17.7 deg. at Mach 0.30, sea-level. Lateral trim was attempted for a design cross-wind of 25.0 knots. The current vehicle aerodynamic and geometric characteristics will only yield a lateral trim solution at impractical tip-fin deflections (approximately equal to 43 deg.) and bank angles (21 deg.). A study of the lateral control surfaces, tip-fin controllers for WB001, indicate increased surface area would help address these instabilities, particularly the deficiency in C(sub n(beta)), but obviously at the expense of increased vehicle weight. Growth factors of approximately 7 were determined using a design C(sub n(beta)) of 0.100/radian (approximate subsonic values for the orbiter).

APEX 3D Propeller Test Preliminary Design

NASA Technical Reports Server (NTRS)

Colozza, Anthony J.

2002-01-01

A low Reynolds number, high subsonic mach number flight regime is fairly uncommon in aeronautics. Most flight vehicles do not fly under these aerodynamic conditions. However, recently there have been a number of proposed aircraft applications (such as high altitude observation platforms and Mars aircraft) that require flight within this regime. One of the main obstacles to flight under these conditions is the ability to reliably generate sufficient thrust for the aircraft. For a conventional propulsion system, the operation and design of the propeller is the key aspect to its operation. Due to the difficulty in experimentally modeling the flight conditions in ground-based facilities, it has been proposed to conduct propeller experiments from a high altitude gliding platform (APEX). A preliminary design of a propeller experiment under the low Reynolds number, high mach number flight conditions has been devised. The details of the design are described as well as the potential data that will be collected.

Women in Flight Research at NASA Dryden Flight Research Center from 1946 to 1995. Number 6; Monographs in Aerospace History

NASA Technical Reports Server (NTRS)

Powers, Sheryll Goecke

1997-01-01

This monograph discusses the working and living environment of women involved with flight research at NASA Dryden Flight Research Center during the late 1940s and early 1950s. The women engineers, their work and the airplanes they worked on from 1960 to December 1995 are highlighted. The labor intensive data gathering and analysis procedures and instrumentation used before the age of digital computers are explained by showing and describing typical instrumentation found on the X-series aircraft from the X-1 through the X-15. The data reduction technique used to obtain the Mach number position error curve for the X-1 aircraft and which documents the historic first flight to exceed the speed of sound is described and a Mach number and altitude plot from an X-15 flight is shown.

Heat Flux and Wall Temperature Estimates for the NASA Langley HIFiRE Direct Connect Rig

NASA Technical Reports Server (NTRS)

Cuda, Vincent, Jr.; Hass, Neal E.

2010-01-01

An objective of the Hypersonic International Flight Research Experimentation (HIFiRE) Program Flight 2 is to provide validation data for high enthalpy scramjet prediction tools through a single flight test and accompanying ground tests of the HIFiRE Direct Connect Rig (HDCR) tested in the NASA LaRC Arc Heated Scramjet Test Facility (AHSTF). The HDCR is a full-scale, copper heat sink structure designed to simulate the isolator entrance conditions and isolator, pilot, and combustor section of the HIFiRE flight test experiment flowpath and is fully instrumented to assess combustion performance over a range of operating conditions simulating flight from Mach 5.5 to 8.5 and for various fueling schemes. As part of the instrumentation package, temperature and heat flux sensors were provided along the flowpath surface and also imbedded in the structure. The purpose of this paper is to demonstrate that the surface heat flux and wall temperature of the Zirconia coated copper wall can be obtained with a water-cooled heat flux gage and a sub-surface temperature measurement. An algorithm was developed which used these two measurements to reconstruct the surface conditions along the flowpath. Determinations of the surface conditions of the Zirconia coating were conducted for a variety of conditions.

Dual-Mode Combustor

NASA Technical Reports Server (NTRS)

Trefny, Charles J (Inventor); Dippold, Vance F (Inventor)

2013-01-01

A new dual-mode ramjet combustor used for operation over a wide flight Mach number range is described. Subsonic combustion mode is usable to lower flight Mach numbers than current dual-mode scramjets. High speed mode is characterized by supersonic combustion in a free-jet that traverses the subsonic combustion chamber to a variable nozzle throat. Although a variable combustor exit aperture is required, the need for fuel staging to accommodate the combustion process is eliminated. Local heating from shock-boundary-layer interactions on combustor walls is also eliminated.

E-959

NASA Image and Video Library

1953-04-27

The aircraft in this 1953 photo of the National Advisory Committee for Aeronautics (NACA) hangar at South Base of Edwards Air Force Base showed the wide range of research activities being undertaken. On the left side of the hanger are the three D-558-2 research aircraft. These were designed to test swept wings at supersonic speeds approaching Mach 2. The front D-558-2 is the third built (NACA 145/Navy 37975). It has been modified with a leading-edge chord extension. This was one of a number of wing modifications, using different configurations of slats and/or wing fences, to ease the airplane's tendency to pitch-up. NACA 145 had both a jet and a rocket engine. The middle aircraft is NACA 144 (Navy 37974), the second built. It was all-rocket powered, and Scott Crossfield made the first Mach 2 flight in this aircraft on November 20, 1953. The aircraft in the back is D-558-2 number 1. NACA 143 (Navy 37973) was also carried both a jet and a rocket engine in 1953. It had been used for the Douglas contractor flights, then was turned over to the NACA. The aircraft was not converted to all-rocket power until June 1954. It made only a single NACA flight before NACA's D-558-2 program ended in 1956. Beside the three D-558-2s is the third D-558-1. Unlike the supersonic D-558-2s, it was designed for flight research at transonic speeds, up to Mach 1. The D-558-1 was jet-powered, and took off from the ground. The D-558-1's handling was poor as it approached Mach 1. Given the designation NACA 142 (Navy 37972), it made a total of 78 research flights, with the last in June 1953. In the back of the hangar is the X-4 (Air Force 46-677). This was a Northrop-built research aircraft which tested a swept wing design without horizontal stabilizers. The aircraft proved unstable in flight at speeds above Mach 0.88. The aircraft showed combined pitching, rolling, and yawing motions, and the design was considered unsuitable. The aircraft, the second X-4 built, was then used as a pilot traine

Performance of a hypersonic hot fuselage structure with a carbon dioxide frost projected, nonintegral cryogenic tank

NASA Technical Reports Server (NTRS)

Sharpe, E. L.; Jackson, L. R.

1975-01-01

A model which consisted of a hot structure and a nonintegral tank protected by a carbon dioxide frost thermal protection system was tested under the following conditions: (1) room temperature loading and (2) heating and loading corresponding to the Mach 8 flight of an air-breathing launch vehicle. In the simulated flight tests, liquid nitrogen inside the tank was withdrawn at the rate fuel would be consumed. Prior to each simulated flight test, carbon dioxide was cryodeposited in the insulation surrounding the tank; during the tests, subliming CO2 frost absorbed heat and provided a purge gas for the space between the tank and the structure. A method of flame spraying the joints between panels with a nickel-aluminum material was developed to prevent excessive leakage of the purge gas through the outer structure. The tests indicated that the hot structure (with a joint repaired by riveting), the nonintegral tank and suspension system, and the carbon dioxide frost thermal protection system provide a workable concept with predictable performance.

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.

Design and Flight Evaluation of a New Force-Based Flow Angle Probe

NASA Technical Reports Server (NTRS)

Corda, Stephen; Vachon, Michael Jacob

2006-01-01

A novel force-based flow angle probe was designed and flight tested on the NASA F-15B Research Testbed aircraft at NASA Dryden Flight Research Center. The prototype flow angle probe is a small, aerodynamic fin that has no moving parts. Forces on the prototype flow angle probe are measured with strain gages and correlated with the local flow angle. The flow angle probe may provide greater simplicity, greater robustness, and better access to flow measurements in confined areas relative to conventional moving vane-type flow angle probes. Flight test data were obtained at subsonic, transonic, and supersonic Mach numbers to a maximum of Mach 1.70. Flight conditions included takeoff, landing, straight and level flight, flight at higher aircraft angles of attack, and flight at elevated g-loadings. Flight test maneuvers included angle-of-attack and angle-of-sideslip sweeps. The flow angle probe-derived flow angles are compared with those obtained with a conventional moving vane probe. The flight tests validated the feasibility of a force-based flow angle measurement system.

Flight and Static Exhaust Flow Properties of an F110-GE-129 Engine in an F-16XL Airplane During Acoustic Tests

NASA Technical Reports Server (NTRS)

Holzman, Jon K.; Webb, Lannie D.; Burcham, Frank W., Jr.

1996-01-01

The exhaust flow properties (mass flow, pressure, temperature, velocity, and Mach number) of the F110-GE-129 engine in an F-16XL airplane were determined from a series of flight tests flown at NASA Dryden Flight Research Center, Edwards, California. These tests were performed in conjunction with NASA Langley Research Center, Hampton, Virginia (LARC) as part of a study to investigate the acoustic characteristics of jet engines operating at high nozzle pressure conditions. The range of interest for both objectives was from Mach 0.3 to Mach 0.9. NASA Dryden flew the airplane and acquired and analyzed the engine data to determine the exhaust characteristics. NASA Langley collected the flyover acoustic measurements and correlated these results with their current predictive codes. This paper describes the airplane, tests, and methods used to determine the exhaust flow properties and presents the exhaust flow properties. No acoustics results are presented.

Unsteady Aerodynamic Model Tuning for Precise Flutter Prediction

NASA Technical Reports Server (NTRS)

Pak, Chan-Gi

2011-01-01

A simple method for an unsteady aerodynamic model tuning is proposed in this study. This method is based on the direct modification of the aerodynamic influence coefficient matrices. The aerostructures test wing 2 flight-test data is used to demonstrate the proposed model tuning method. The flutter speed margin computed using only the test validated structural dynamic model can be improved using the additional unsteady aerodynamic model tuning, and then the flutter speed margin requirement of 15 % in military specifications can apply towards the test validated aeroelastic model. In this study, unsteady aerodynamic model tunings are performed at two time invariant flight conditions, at Mach numbers of 0.390 and 0.456. When the Mach number for the unsteady model tuning approaches to the measured fluttering Mach number, 0.502, at the flight altitude of 9,837 ft, the estimated flutter speed is approached to the measured flutter speed at this altitude. The minimum flutter speed difference between the estimated and measured flutter speed is -.14 %.

Unsteady Aerodynamic Model Tuning for Precise Flutter Prediction

NASA Technical Reports Server (NTRS)

Pak, Chan-gi

2011-01-01

A simple method for an unsteady aerodynamic model tuning is proposed in this study. This method is based on the direct modification of the aerodynamic influence coefficient matrices. The aerostructures test wing 2 flight-test data is used to demonstrate the proposed model tuning method. The flutter speed margin computed using only the test validated structural dynamic model can be improved using the additional unsteady aerodynamic model tuning, and then the flutter speed margin requirement of 15 percent in military specifications can apply towards the test validated aeroelastic model. In this study, unsteady aerodynamic model tunings are performed at two time invariant flight conditions, at Mach numbers of 0.390 and 0.456. When the Mach number for the unsteady aerodynamic model tuning approaches to the measured fluttering Mach number, 0.502, at the flight altitude of 9,837 ft, the estimated flutter speed is approached to the measured flutter speed at this altitude. The minimum flutter speed difference between the estimated and measured flutter speed is -0.14 percent.

Transition Analysis for the HIFiRE-5 Vehicle

NASA Technical Reports Server (NTRS)

Choudhari, Meelan M.; Chang, Chau-Lyan; Li, Fei; Berger, Karen T.; Candler, Graham V.; Kimmel, Roger

2009-01-01

The Hypersonic International Flight Research and Experimentation (HIFiRE) 5 flight experiment by Air Force Research Laboratories and Australian Defense Science and Technology Organization is designed to provide in-flight boundary-layer transition data for a canonical 3D configuration at hypersonic Mach numbers. This paper outlines the progress, to date, on boundary layer stability analysis for the HIFiRE-5 flight configuration, as well as for selected test conditions from the wind tunnel experiments supporting the flight test. At flow conditions corresponding to the end of the test window, rather large values of linear amplification factor are predicted for both second mode (N>40) and crossflow (N>20) instabilities, strongly supporting the feasibility of first in-flight measurements of natural transition on a fully three-dimensional hypersonic configuration. Additional results highlight the rich mixture of instability mechanisms relevant to a large segment of the flight trajectory, as well as the effects of angle of attack and yaw angle on the predicted transition fronts for ground facility experiments at Mach 6.

X-43D Conceptual Design and Feasibility Study

NASA Technical Reports Server (NTRS)

Johnson, Donald B.; Robinson, Jeffrey S.

2005-01-01

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

The X-43A hypersonic research aircraft and its modified Pegasus® booster rocket recently underwent combined systems testing while mounted to NASA's NB-52B carrier aircraft

NASA Image and Video Library

2001-03-15

The first of three X-43A hypersonic research aircraft and its modified Pegasus® booster rocket recently underwent combined systems testing while mounted to NASA's NB-52B carrier aircraft at the Dryden Flight Research Center, Edwards, Calif. The combined systems test was one of the last major milestones in the Hyper-X research program before the first X-43A flight. The X-43A flights will be the first actual flight tests of an aircraft powered by a revolutionary supersonic-combustion ramjet ("scramjet") engine capable of operating at hypersonic speeds (above Mach 5, or five times the speed of sound). The 12-foot, unpiloted research vehicle was developed and built by MicroCraft Inc., Tullahoma, Tenn., under NASA contract. The booster was built by Orbital Sciences Corp., Dulles, Va.,After being air-launched from NASA's venerable NB-52 mothership, the booster will accelerate the X-43A to test speed and altitude. The X-43A will then separate from the rocket and fly a pre-programmed trajectory, conducting aerodynamic and propulsion experiments until it descends into the Pacific Ocean. Three research flights are planned, two at Mach 7 and one at Mach 10.

Wind Tunnel Measurements of Shuttle Orbiter Global Heating with Comparisons to Flight

NASA Technical Reports Server (NTRS)

Berry, Scott A.; Merski, N. Ronald; Blanchard, Robert C.

2002-01-01

An aerothermodynamic database of global heating images was acquired of the Shuttle Orbiter in the NASA Langley Research Center 20-Inch Mach 6 Air Tunnel. These results were obtained for comparison to the global infrared images of the Orbiter in flight from the infrared sensing aeroheating flight experiment (ISAFE). The most recent ISAFE results from STS-103, consisted of port side images, at hypersonic conditions, of the surface features that result from the strake vortex scrubbing along the side of the vehicle. The wind tunnel results were obtained with the phosphor thermography system, which also provides global information and thus is ideally suited for comparison to the global flight results. The aerothermodynamic database includes both windward and port side heating images of the Orbiter for a range of angles of attack (20 to 40 deg), freestream unit Reynolds number (1 x 10(exp 6))/ft to 8 x 10(exp 6)/ft, body flap deflections (0, 5, and 10 deg), speed brake deflections (0 and 45 deg), as well as with boundary layer trips for forced transition to turbulence heating results. Sample global wind tunnel heat transfer images were extrapolated to flight conditions for comparison to Orbiter flight data. A windward laminar case for an angle of attack of 40 deg was extrapolated to Mach 11.6 flight conditions for comparison to STS-2 flight thermocouple results. A portside wind tunnel image for an angle of attack of 25 deg was extrapolated for Mach 5 flight conditions for comparison to STS-103 global surface temperatures. The comparisons showed excellent qualitative agreement, however the extrapolated wind tunnel results over-predicted the flight surface temperatures on the order of 5% on the windward surface and slightly higher on the portside.

High Altitude Flight Test of a Reefed 12.2 Meter Diameter Disk-Gap-Band Parachute with Deployment at Mach Number of 2.58

NASA Technical Reports Server (NTRS)

Grow, R. Bruce; Preisser, John S.

1971-01-01

A reefed 12.2-meter nominal-diameter (40-ft) disk-gap-band parachute was flight tested as part of the NASA Supersonic High Altitude Parachute Experiment (SHAPE) program. A three-stage rocket was used to drive the instrumented payload to an altitude of 43.6 km (143,000 ft), a Mach number of 2.58, and a dynamic pressure of 972 N/m(exp 2) (20.3 lb/ft(exp 2)) where the parachute was deployed by means of a mortar. The parachute deployed satisfactorily and reached a partially inflated condition characterized by irregular variations in parachute projected area. A full, stable reefed inflation was achieved when the system had decelerated to a Mach number of about 1.5. The steady, reefed projected area was 49 percent of the steady, unreefed area and the average drag coefficient was 0.30. Disreefing occurred at a Mach number of 0.99 and a dynamic pressure of 81 N/m(exp 2) (1.7 lb/ft(exp 2)). The parachute maintained a steady inflated shape for the remainder of the deceleration portion of the flight and throughout descent. During descent, the average effective drag coefficient was 0.57. There was little, if any, coning motion, and the amplitude of planar oscillations was generally less than 10 degrees. The film also shows a wind tunnel test of a 1.7-meter-diameter parachute inflating at Mach number 2.0.

Calibration of HYPULSE for hypervelocity air flows corresponding to flight Mach numbers 13.5, 15, and 17

NASA Technical Reports Server (NTRS)

Calleja, John; Tamagno, Jose

1993-01-01

A series of air calibration tests were performed in GASL's HYPULSE facility in order to more accurately determine test section flow conditions for flows simulating total enthalpies in the Mach 13 to 17 range. Present calibration data supplements previous data and includes direct measurement of test section pitot and static pressure, acceleration tube wall pressure and heat transfer, and primary and secondary incident shock velocities. Useful test core diameters along with the corresponding free-stream conditions and usable testing times were determined. For the M13.5 condition, in-stream static pressure surveys showed the temporal and spacial uniformity of this quantity across the useful test core. In addition, finite fringe interferograms taken of the free-stream flow at the test section did not indicate the presence of any 'strong' wave system for any of the conditions investigated.

Comparison of the Aeroacoustics of Two Small-Scale Supersonic Inlets

NASA Technical Reports Server (NTRS)

Ng, Wing

1996-01-01

An aerodynamic and acoustic investigation was performed on two small-scale supersonic inlets to determine which inlet would be more suitable for a High Speed Civil Transport (HSCT) aircraft during approach and takeoff flight conditions. The comparison was made between an axisymmetric supersonic P inlet and a bifurcated two-dimensional supersonic inlet. The 1/14 scale model supersonic inlets were used in conjunction with a 4.1 in (10.4 cm) turbofan engine simulator. A bellmouth was utilized on each inlet to eliminate lip separation commonly associated with airplane engine inlets that are tested under static conditions. Steady state measurements of the aerodynamic flowfield and acoustic farfield were made in order to evaluate the aeroacoustic performance of the inlets. The aerodynamic results show the total pressure recovery of the two inlets to be nearly identical, 99% at the approach condition and 98% at the takeoff condition. At the approach fan speed (60% design speed), there was no appreciable difference in the acoustic performance of either inlet over the entire 0 deg to 110 deg farfield measurement sector. The inlet flow field results at the takeoff fan speed (88% design speed), show the average inlet throat Mach number for the P inlet (Mach 0.52) to be approximately 2 times that of the 2D inlet (Mach 0.26). The difference in the throat Mach number is a result of the smaller throughflow area of the P inlet. This reduced area resulted in a 'soft choking' of the P inlet which lowered the tone and overall sound pressure levels of the simulator in the forward sector by an average of 9 dB and 3 dB, respectively, when compared to the 2D inlet.

Variable Mach number design approach for a parallel waverider with a wide-speed range based on the osculating cone theory

NASA Astrophysics Data System (ADS)

Zhao, Zhen-tao; Huang, Wei; Li, Shi-Bin; Zhang, Tian-Tian; Yan, Li

2018-06-01

In the current study, a variable Mach number waverider design approach has been proposed based on the osculating cone theory. The design Mach number of the osculating cone constant Mach number waverider with the same volumetric efficiency of the osculating cone variable Mach number waverider has been determined by writing a program for calculating the volumetric efficiencies of waveriders. The CFD approach has been utilized to verify the effectiveness of the proposed approach. At the same time, through the comparative analysis of the aerodynamic performance, the performance advantage of the osculating cone variable Mach number waverider is studied. The obtained results show that the osculating cone variable Mach number waverider owns higher lift-to-drag ratio throughout the flight profile when compared with the osculating cone constant Mach number waverider, and it has superior low-speed aerodynamic performance while maintaining nearly the same high-speed aerodynamic performance.

Simulation model of the F/A-18 high angle-of-attack research vehicle utilized for the design of advanced control laws

NASA Technical Reports Server (NTRS)

Strickland, Mark E.; Bundick, W. Thomas; Messina, Michael D.; Hoffler, Keith D.; Carzoo, Susan W.; Yeager, Jessie C.; Beissner, Fred L., Jr.

1996-01-01

The 'f18harv' six degree-of-freedom nonlinear batch simulation used to support research in advanced control laws and flight dynamics issues as part of NASA's High Alpha Technology Program is described in this report. This simulation models an F/A-18 airplane modified to incorporate a multi-axis thrust-vectoring system for augmented pitch and yaw control power and actuated forebody strakes for enhanced aerodynamic yaw control power. The modified configuration is known as the High Alpha Research Vehicle (HARV). The 'f18harv' simulation was an outgrowth of the 'f18bas' simulation which modeled the basic F/A-18 with a preliminary version of a thrust-vectoring system designed for the HARV. The preliminary version consisted of two thrust-vectoring vanes per engine nozzle compared with the three vanes per engine actually employed on the F/A-18 HARV. The modeled flight envelope is extensive in that the aerodynamic database covers an angle-of-attack range of -10 degrees to +90 degrees, sideslip range of -20 degrees to +20 degrees, a Mach Number range between 0.0 and 2.0, and an altitude range between 0 and 60,000 feet.

Analytical and Experimental Verification of a Flight Article for a Mach-8 Boundary-Layer Experiment

NASA Technical Reports Server (NTRS)

Richards, W. Lance; Monaghan, Richard C.

1996-01-01

Preparations for a boundary-layer transition experiment to be conducted on a future flight mission of the air-launched Pegasus(TM) rocket are underway. The experiment requires a flight-test article called a glove to be attached to the wing of the Mach-8 first-stage booster. A three-dimensional, nonlinear finite-element analysis has been performed and significant small-scale laboratory testing has been accomplished to ensure the glove design integrity and quality of the experiment. Reliance on both the analysis and experiment activities has been instrumental in the success of the flight-article design. Results obtained from the structural analysis and laboratory testing show that all glove components are well within the allowable thermal stress and deformation requirements to satisfy the experiment objectives.

NASA Administrator Sean O'Keefe, left, learned about the Mach 10 X-43 research vehicle from manager

NASA Technical Reports Server (NTRS)

2002-01-01

NASA Administrator Sean O'Keefe left, learned about the Mach 10 X-43 research vehicle from manager, Joel Sitz during O'Keefe's visit to the NASA Dryden Flight Research Center, Edwards, California, January 31, 2002.

Blade Tip Pressure Measurements Using Pressure Sensitive Paint

NASA Technical Reports Server (NTRS)

Wong, Oliver D.; Watkins, Anthony Neal; Goodman, Kyle Z.; Crafton, James; Forlines, Alan; Goss, Larry; Gregory, James W.; Juliano, Thomas J.

2012-01-01

This paper discusses the application of pressure sensitive paint using laser-based excitation for measurement of the upper surface pressure distribution on the tips of rotor blades in hover and simulated forward flight. The testing was conducted in the Rotor Test Cell and the 14- by 22-ft Subsonic Tunnel at the NASA Langley Research Center on the General Rotor Model System (GRMS) test stand. The Mach-scaled rotor contained three chordwise rows of dynamic pressure transducers for comparison with PSP measurements. The rotor had an 11 ft 1 in. diameter, 5.45 in. main chord and a swept, tapered tip. Three thrust conditions were examined in hover, C(sub T) = 0.004, 0.006 and 0.008. In forward flight, an additional thrust condition, C(sub T) = 0.010 was also examined. All four thrust conditions in forward flight were conducted at an advance ratio of 0.35.

Thermodynamic Cycle Analysis of Magnetohydrodynamic-Bypass Hypersonic Airbreathing Engines

NASA Technical Reports Server (NTRS)

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

2000-01-01

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

Flight Test Measurements From The Tu-144LL Structure/Cabin Noise Follow-On Experiment

NASA Technical Reports Server (NTRS)

Rizzi, Stephen A.; Rackl, Robert G.; Andrianov, Eduard V.

2000-01-01

This follow-on flight experiment on the TU-144LL Supersonic Flying Laboratory, conducted during the period September 1998 to April 1999, was a continuation of previous Structure/Cabin Noise Experiment 2.1. Data was obtained over a wide range of altitudes and Mach numbers. Measured were: turbulent boundary layer pressure fluctuations on the fuselage over its length; structural response on skin panels using accelerometers; and flow direction over three windows using 'flow cones'. The effect of steps in the flow was also measured using two window blank pairs; each pair bridged by a plate which created small sharp forward and aft facing steps. The effect of transducer flushness with the exterior surface was also measured during flight. Height test points were chosen to cover much of the TU-144's flight envelope, as well as to obtain as large a unit Reynolds number range as possible at various Mach numbers: takeoff, subsonic, transonic, and supersonic cruise conditions up to Mach 2. Data on engine runups and background noise were acquired on the ground. The data in the form of time histories of the acoustic signals, together with auxiliary data and basic MATLAB processing modules, are available on CD-R disks.

Advanced and Adaptable Military Propulsion

DTIC Science & Technology

2008-01-22

turbine. The accelerating flow in the turbine environment would mitigate this somewhat (i.e. (favorable) axial pressure gradient vs radial pressure...For instance in a low bypass ratio (0.8) turbofan engine operating at flight Mach numbers ranging from 0.85 to 2.5, the specific fuel consumption...ratio, T8/PT2 - Fan inlet axial Mach No, M2, capability - Compressor exit axial Mach No, M3 - Compressor pressure ratio, PT3/PT2 - Turbine nozzle area

Shock-tunnel combustor testing for hypersonic vehicles

NASA Technical Reports Server (NTRS)

Loomis, Mark P.

1994-01-01

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

Dual-Pump CARS Development and Application to Supersonic Combustion

NASA Astrophysics Data System (ADS)

Magnotti, Gaetano

Successful design of hypersonic air-breathing engines requires new computational fluid dynamics (CFD) models for turbulence and turbulence-chemistry interaction in supersonic combustion. Unfortunately, not enough data are available to the modelers to develop and validate their codes, due to difficulties in taking measurements in such a harsh environment. Dual-pump coherent anti-Stokes Raman spectroscopy (CARS) is a non-intrusive, non-linear, laser-based technique that provides temporally and spatially resolved measurements of temperature and absolute mole fractions of N2, O2 and H2 in H2-air flames. A dual-pump CARS instrument has been developed to obtain measurements in supersonic combustion and generate databases for the CFD community. Issues that compromised previous attempts, such as beam steering and high irradiance perturbation effects, have been alleviated or avoided. Improvements in instrument precision and accuracy have been achieved. An axis-symmetric supersonic combusting coaxial jet facility has been developed to provide a simple, yet suitable flow to CFD modelers. The facility provides a central jet of hot "vitiated air" simulating the hot air entering the engine of a hypersonic vehicle flying at Mach numbers between 5 and 7. Three different silicon carbide nozzles, with exit Mach number 1, 1.6 and 2, are used to provide flows with the effects of varying compressibility. H2 co-flow is available in order to generate a supersonic combusting free jet. Dual-pump CARS measurements have been obtained for varying values of flight and exit Mach numbers at several locations. Approximately one million Dual-pump CARS single shots have been collected in the supersonic jet for varying values of flight and exit Mach numbers at several locations. Data have been acquired with a H2 co-flow (combustion case) or a N 2 co-flow (mixing case). Results are presented and the effects of the compressibility and of the heat release are discussed.

Supersonic Disk Gap Band Parachute Performance in the Wake of a Viking-Type Aeroshell from Mach 2 to 2.5

NASA Technical Reports Server (NTRS)

Sengupta, Anita; Roeder, James; Kelsch, Richard; Wernet, Mark; Machalick, Walt; Reuter, James; Witkowski, Al

2008-01-01

Supersonic wind tunnel testing of 0.813 m diameter Disk-Gap-Band parachutes is being conducted in the NASA Glenn Research Center (GRC) 10' x 10' wind-tunnel. The tests are conducted in support of the Mars Science Laboratory Parachute Decelerator System development and qualification. Four percent of full-scale parachutes were constructed similarly to the flight-article in material and construction techniques. The parachutes are attached to a 4% scale MSL entry-vehicle to simulate the free-flight configuration. The parachutes are tested from Mach 2 to 2.5 over a Reynolds number (Re) range of 1 to 3 x 10(exp 6), representative of the MSL deployment envelope. Constrained and unconstrained test configurations are investigated to quantify the effects of parachute trim, suspension line interaction, and alignment with the capsule wake. The parachute is constrained horizontally through the vent region, to measure canopy breathing and wake interaction for fixed trim angles of 0 and 10 degrees from the velocity vector. In the unconstrained configuration the parachute is permitted to trim and cone, similar to the free-flight varying its alignment relative to the entry-vehicle wake. Test diagnostics were chosen to quantify parachute performance and to provide insight into the flow field structure. An in-line load cell provided measurement of unsteady and mean drag as a function of Mach and Re. High-speed shadowgraph video of the upstream parachute flow field was used to capture bow-shock motion and stand of distance. Particle image velocimetry of the upstream parachute flow field provides spatially and temporally resolved measurement velocity and turbulent statistics. Multiple high speed video views of targets placed in the interior of the canopy enable photo-grammetric measurement of the fabric motion in time and space from reflective. High speed video is also used to document the supersonic inflation and measure trim angle, projected area, and frequency of area oscillations.

Flight Investigation at Low Angles of Attack to Determine the Longitudinal Stability and Control Characteristics of a Cruciform Canard Missile Configuration with a Low-Aspect-Ratio Wing and Blunt Nose at Mach Numbers from 1.2 to 2.1

NASA Technical Reports Server (NTRS)

Brown, Clarence A , Jr

1957-01-01

A full- scale rocket-powered model of a cruciform canard missile configuration with a low- aspect - ratio wing and blunt nose has been flight tested by the Langley Pilotless Aircraft Research Division. Static and dynamic longitudinal stability and control derivatives of this interdigitated canard-wing missile configuration were determined by using the pulsed- control technique at low angles of attack and for a Mach number range of 1.2 to 2.1. The lift - curve slope showed only small nonlinearities with changes in control deflection or angle of attack but indicated a difference in lift- .curve slope of approximately 7 percent for the two control deflections of delta = 3.0 deg and delta= -0.3 deg . The large tail length of the missile tested was effective in producing damping in pitch throughout the Mach number range tested. The aerodynamic- center location was nearly constant with Mach number for the two control deflections but was shown to be less stable with the larger control deflection. The increment of lift produced by the controls was small and positive throughout the Mach number range tested, whereas the pitching moment produced by the controls exhibited a normal trend of reduced effectiveness with increasing Mach number.The effectiveness of the controls in producing angle of attack, lift, and pitching moment was good at all Mach numbers tested.

Flight Investigation at Low Angles of Attack to Determine the Longitudinal Stability and Control Characteristics of a Cruciform Canard Missile Configuration with a Low-Aspect-Ratio Wing and Blunt Nose at Mach Numbers from 1.2 to 2.1

NASA Technical Reports Server (NTRS)

Brown, C. A., Jr.

1957-01-01

A full-scale rocket-powered model of a cruciform canard missile configuration with a low-aspect-ratio wing and blunt nose has been flight tested by the Langley Pilotless Aircraft Research Division. Static and dynamic longitudinal stability and control derivatives of this interdigitated canard-wing missile configuration were determined by using the pulsed-control technique at low angles of attack and for a Mach number range of 1.2 to 2.1. The lift-curve slope showed only small nonlinearities with changes in control deflection or angle of attack but indicated a difference in lift-curve slope of approximately 7 percent for the two control deflections of delta = 3.0 deg and delta = -0.3 deg. The large tail length of the missile tested was effective in producing damping in pitch throughout the Mach number range tested. The aerodynamic-center location was nearly constant with Mach number for the two control deflections but was shown to be less stable with the larger control deflection. The increment of lift produced by the controls was small and positive throughout the Mach number range tested, whereas the pitching moment produced by the controls exhibited a normal trend of reduced effectiveness with increasing Mach number. The effectiveness of the controls in producing angle of attack, lift, and pitching moment was good at all Mach numbers tested.

An investigation of bleed configurations and their effect on shock wave/boundary layer interactions

NASA Technical Reports Server (NTRS)

Hamed, Awatef

1995-01-01

The design of high efficiency supersonic inlets is a complex task involving the optimization of a number of performance parameters such as pressure recovery, spillage, drag, and exit distortion profile, over the flight Mach number range. Computational techniques must be capable of accurately simulating the physics of shock/boundary layer interactions, secondary corner flows, flow separation, and bleed if they are to be useful in the design. In particular, bleed and flow separation, play an important role in inlet unstart, and the associated pressure oscillations. Numerical simulations were conducted to investigate some of the basic physical phenomena associated with bleed in oblique shock wave boundary layer interactions that affect the inlet performance.

Flight test measurements and analysis of sonic boom phenomena near the shock wave extremity

NASA Technical Reports Server (NTRS)

Haglund, G. T.; Kane, E. J.

1973-01-01

The sonic boom flight test program conducted at Jackass Flats, Nevada, during the summer and fall of 1970 consisted of 121 sonic-boom-generating flights over the 1500 ft instrumented BREN tower. This test program was designed to provide information on several aspects of sonic boom, including caustics produced by longitudinal accelerations, caustics produced by steady flight near the threshold Mach number, sonic boom characteristics near lateral cutoff, and the vertical extent of shock waves attached to near-sonic airplanes. The measured test data, except for the near-sonic flight data, were analyzed in detail to determine sonic boom characteristics for these flight conditions and to determine the accuracy and the range of validity of linear sonic boom theory. The caustic phenomena observed during the threshold Mach number flights and during the transonic acceleration flights are documented and analyzed in detail. The theory of geometric acoustics is shown to be capable of predicting shock wave-ground intersections, and current methods for calculating sonic boom pressure signature away from caustics are shown to be reasonably accurate.

Dispersion development program. [development of a fin stabilized submissile and ejection system for the Little John warhead

NASA Technical Reports Server (NTRS)

Carlson, D. J.; Lusardi, R. J.; Phillips, W. H.

1975-01-01

The requirement for the predictable dispersion of small munitions over large areas from ground support missile systems has resulted in the development of a fin stabilized submissile and sling ejection system for the Little John warhead. The progressive development of this system is traced including a comparison of simulator, sled test, and flight test results. The results indicate that it is not only necessary but also possible to eject long slender bodies, from a missile warhead at Mach 1, in a stable, uniform and predictable manner.

Flight and Preflight Tests of a Ram Jet Burning Magnesium Slurry Fuel and Utilizing a Solid-propellant Gas Generator for Fuel Expulsion

NASA Technical Reports Server (NTRS)

Bartlett, Walter, A , jr; Hagginbotham, William K , Jr

1955-01-01

Data obtained from the first flight test of a ram jet utilizing a magnesium slurry fuel are presented. The ram jet accelerated from a Mach number of 1.75 to a Mach number of 3.48 in 15.5 seconds. During this period a maximum values of air specific impulse and gross thrust coefficient were calculated to be 151 seconds and 0.658, respectively. The rocket gas generator used as a fuel-pumping system operated successfully.

Process for Operating a Dual-Mode Combustor

NASA Technical Reports Server (NTRS)

Trefny, Charles J. (Inventor); Dippold, Vance F. (Inventor)

2017-01-01

A new dual-mode ramjet combustor used for operation over a wide flight Mach number range is described. Subsonic combustion mode is usable to lower flight Mach numbers than current dual-mode scramjets. High speed mode is characterized by supersonic combustion in a free-jet that traverses the subsonic combustion chamber to a variable nozzle throat. Although a variable combustor exit aperture is required, the need for fuel staging to accommodate the combustion process is eliminated. Local heating from shock-boundary-layer interactions on combustor walls is also eliminated.

Flight-determined engine exhaust characteristics of an F404 engine in an F-18 airplane

NASA Technical Reports Server (NTRS)

Ennix, Kimberly A.; Burcham, Frank W., Jr.; Webb, Lannie D.

1993-01-01

Personnel at the NASA Langley Research Center (NASA-Langley) and the NASA Dryden Flight Research Facility (NASA-Dryden) recently completed a joint acoustic flight test program. Several types of aircraft with high nozzle pressure ratio engines were flown to satisfy a twofold objective. First, assessments were made of subsonic climb-to-cruise noise from flights conducted at varying altitudes in a Mach 0.30 to 0.90 range. Second, using data from flights conducted at constant altitude in a Mach 0.30 to 0.95 range, engineers obtained a high quality noise database. This database was desired to validate the Aircraft Noise Prediction Program and other system noise prediction codes. NASA-Dryden personnel analyzed the engine data from several aircraft that were flown in the test program to determine the exhaust characteristics. The analysis of the exhaust characteristics from the F-18 aircraft are reported. An overview of the flight test planning, instrumentation, test procedures, data analysis, engine modeling codes, and results are presented.

Inverse Flush Air Data System (FADS) for Real Time Simulations

NASA Astrophysics Data System (ADS)

Madhavanpillai, Jayakumar; Dhoaya, Jayanta; Balakrishnan, Vidya Saraswathi; Narayanan, Remesh; Chacko, Finitha Kallely; Narayanan, Shyam Mohan

2017-12-01

Flush Air Data Sensing System (FADS) forms a mission critical sub system in future reentry vehicles. FADS makes use of surface pressure measurements from the nose cap of the vehicle for deriving the air data parameters of the vehicle such as angle of attack, angle of sideslip, Mach number, etc. These parameters find use in the flight control and guidance systems, and also assist in the overall mission management. The FADS under consideration in this paper makes use of nine pressure ports located in the nose cap of a technology demonstrator vehicle. In flight, the air data parameters are obtained from the FADS estimation algorithm using the pressure data at the nine pressure ports. But, these pressure data will not be available, for testing the FADS package during ground simulation. So, an inverse software to FADS which estimates the pressure data at the pressure ports for a given flight condition is developed. These pressure data at the nine ports will go as input to the FADS package during ground simulation. The software is run to generate the pressure data for the descent phase trajectory of the technology demonstrator. This data is used again to generate the air data parameters from FADS algorithm. The computed results from FADS algorithm match well with the trajectory data.

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.

Experimental Hypersonic Aerodynamic Characteristics of the Space Shuttle Orbiter for a Range of Damage Scenarios

NASA Technical Reports Server (NTRS)

Brauckman, Gregory J.; Scallion, William I.

2003-01-01

Aerodynamic tests in support of the Columbia accident investigation were conducted in two hypersonic wind tunnels at the NASA Langley Research Center, the 20-Inch Mach 6 Air Tunnel and the 20-Inch Mach 6 CF4 Tunnel. The primary purpose of these tests was to measure the forces and moments generated by a variety of outer mold line alterations (damage scenarios) using 0.0075-scale models of the Space Shuttle Orbiter (approximately 10 inches in length). Simultaneously acquired global heat transfer mappings were obtained for a majority of the configurations tested. Test parameters include angles of attack from 38 to 42 deg, unit Reynolds numbers from 0.26 to 3.0 x10^6 per foot, and normal shock density ratios of 5 (Mach 6 air) and 12 (Mach 6 CF4). The damage scenarios evaluated included asymmetric boundary layer transition, gouges in the windward surface acreage thermal protection system tiles, wing leading edge damage (partially and fully missing reinforced carbon-carbon (RCC) panels), holes through the wing from the windward surface to the leeside, deformation of the wing windward surface, and main landing gear door and/or gear deployment. The aerodynamic data were compared to the magnitudes and directions observed in flight, and the heating images were evaluated in terms of the location of the generated disturbances and how these disturbance might relate to the response of discrete gages on the Columbia Orbiter vehicle during entry. The measured aerodynamic increments were generally small in magnitude, as were the flight-derived values during most of the entry. Asymmetric boundary layer transition (ABLT) results were consistent with the flight-derived Shuttle ABLT model, but not with the observed flight trends for STS-107. The partially missing leading edge panel results best matched both the early aerodynamic and heating trends observed in flight. A progressive damage scenario is presented that qualitatively matches the flight observations for the full entry.

Aerodynamic Models for the Low Density Supersonic Declerator (LDSD) Supersonic Flight Dynamics Test (SFDT)

NASA Technical Reports Server (NTRS)

Van Norman, John W.; Dyakonov, Artem; Schoenenberger, Mark; Davis, Jody; Muppidi, Suman; Tang, Chun; Bose, Deepak; Mobley, Brandon; Clark, Ian

2015-01-01

An overview of pre-flight aerodynamic models for the Low Density Supersonic Decelerator (LDSD) Supersonic Flight Dynamics Test (SFDT) campaign is presented, with comparisons to reconstructed flight data and discussion of model updates. The SFDT campaign objective is to test Supersonic Inflatable Aerodynamic Decelerator (SIAD) and large supersonic parachute technologies at high altitude Earth conditions relevant to entry, descent, and landing (EDL) at Mars. Nominal SIAD test conditions are attained by lifting a test vehicle (TV) to 36 km altitude with a large helium balloon, then accelerating the TV to Mach 4 and and 53 km altitude with a solid rocket motor. The first flight test (SFDT-1) delivered a 6 meter diameter robotic mission class decelerator (SIAD-R) to several seconds of flight on June 28, 2014, and was successful in demonstrating the SFDT flight system concept and SIAD-R. The trajectory was off-nominal, however, lofting to over 8 km higher than predicted in flight simulations. Comparisons between reconstructed flight data and aerodynamic models show that SIAD-R aerodynamic performance was in good agreement with pre-flight predictions. Similar comparisons of powered ascent phase aerodynamics show that the pre-flight model overpredicted TV pitch stability, leading to underprediction of trajectory peak altitude. Comparisons between pre-flight aerodynamic models and reconstructed flight data are shown, and changes to aerodynamic models using improved fidelity and knowledge gained from SFDT-1 are discussed.

The Effect of Blunt-Trailing-Edge Modifications on the High-Speed Stability and Control Characteristics of a Swept-Wing Fighter Airplane

NASA Technical Reports Server (NTRS)

Sadoff, Melvin; Matteson, Frederick H.; Van Dyke, Rudolph D., Jr.

1954-01-01

An investigation was conducted on a 35 deg swept-wing fighter airplane to determine the effects of several blunt-trailing-edge modifications to the wing and tail on the high-speed stability and control characteristics and tracking performance. The results indicated significant improvement in the pitch-up characteristics for the blunt-aileron configuration at Mach numbers around 0.90. As a result of increased effectiveness of the blunt-trailing-edge aileron, the roll-off, customarily experienced with the unmodified airplane in wings-level flight between Mach numbers of about 0.9 and 1.0 was eliminated, The results also indicated that the increased effectiveness of the blunt aileron more than offset the large associated aileron hinge moment, resulting in significant improvement in the rolling performance at Mach numbers between 0.85 and 1.0. It appeared from these results that the tracking performance with the blunt-aileron configuration in the pitch-up and buffeting flight region at high Mach numbers was considerably improved over that of the unmodified airplane; however, the tracking errors of 8 to 15 mils were definitely unsatisfactory. A drag increment of about O.OOl5 due to the blunt ailerons was noted at Mach numbers to about 0.85. The drag increment was 0 at Mach numbers above 0.90.

Preliminary sonic boom correlation of predicted and measured levels for STS-1 entry

NASA Technical Reports Server (NTRS)

Garcia, F., Jr.; Morrison, K. M.; Jones, J. H.; Henderson, H. R.

1982-01-01

A preliminary analysis correlating peaks from sonic boom pressure signatures recorded during the descent trajectory of the Orbiter Columbia, which landed in the dry lake bed at Edwards Air Force Base (EAFB), California, with measured wind tunnel signatures extrapolated from flight altitudes to the ground has been made for Mach numbers ranging from 1.3 to 6. The flight pressure signatures were recorded by microphones positioned at ground level near the groundtrack, whereas the wind tunnel signatures were measured during a test of a 0.0041-scale model Orbiter. The agreement between overpressure estimates based on wind tunnel data using preliminary flight trajectory data and oscillograph traces from ground measurements appears reasonable at this time for the range of Mach numbers considered. More detailed studies using final flight trajectory data and digitized ground measured data will be performed.

Helicopter far-field acoustic levels as a function of reduced rotor speeds

NASA Technical Reports Server (NTRS)

Mueller, Arnold W.; Lemasurier, Philip; Smith, Charles D.

1990-01-01

This paper will present far-field measured noise levels relative to tests conducted with a model S-76A helicopter. The project was designed to provide supplemental experimental flight data which may be used to further study reduced helicopter rotor speeds (and thus, advancing blade-tip Mach number) effects on far-field acoustic levels. The aircraft was flown in straight and level flight while operating with both the rotor speed and flight speed as test variables. The rotor speed was varied over the range of 107 percent of the main-rotor speed (NR) to 90 percent NR and with the forward flight speed varied over the range of 155 to 35 knots indicated air speed. These conditions produced a wide range of advancing blade-tip Mach numbers to which the noise data are related.

Computational analysis of forebody tangential slot blowing on the high alpha research vehicle

NASA Technical Reports Server (NTRS)

Gee, Ken

1994-01-01

Current and future fighter aircraft can maneuver in the high-angle-of-attack flight regime while flying at low subsonic and transonic freestream Mach numbers. However, at any flight speed, the ability of the vertical tails to generate yawing moment is limited in high-angle-of-attack flight. Thus, any system designed to provide the pilot with additional side force and yawing moment must work in both low subsonic and transonic flight. However, previous investigations of the effectiveness of forebody tangential slot blowing in generating the desired control forces and moments have been limited to the low subsonic freestream flow regime. In order to investigate the effectiveness of tangential slot blowing in transonic flight, a computational fluid dynamics analysis was carried out during the grant period. Computational solutions were obtained at three different freestream Mach numbers and at various jet mass flow ratios. All results were obtained using the isolated F/A-18 forebody grid geometry at 30.3 degrees angle of attack. One goal of the research was to determine the effect of freestream Mach number on the effectiveness of forebody tangential slot blowing in generating yawing moment. The second part of the research studied the force onset time lag associated with blowing. The time required for the yawing moment to reach a steady-state value from the onset of blowing may have an impact on the implementation of a pneumatic system on a flight vehicle.

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.

Supersonic Free-Jet Combustion in a Ramjet Burner

NASA Technical Reports Server (NTRS)

Trefny, Charles J.; Dippold, Vance F., III

2010-01-01

A new dual-mode ramjet combustor concept intended for operation over a wide flight Mach number range is described. Subsonic combustion mode is similar to that of a traditional ram combustor which allows operation at higher efficiency, and to lower flight Mach numbers than current dual-mode scramjets. High speed mode is characterized by supersonic combustion in a free-jet that traverses the subsonic combustion chamber to a variable nozzle. The maximum flight Mach number of this scheme is governed largely by the same physics as its classical counterpart. Although a variable combustor exit aperture is required, the need for fuel staging to accommodate the combustion process is eliminated. Local heating from shock-boundary-layer interactions on combustor walls is also eliminated. Given the parallel nature of the present scheme, overall flowpath length is less than that of present dual-mode configurations. Cycle analysis was done to define the flowpath geometry for computational fluid dynamics (CFD) analysis, and then to determine performance based on the CFD results. CFD results for Mach 5, 8, and 12 flight conditions indicate stable supersonic free-jet formation and nozzle reattachment, thereby establishing the basic feasibility of the concept. These results also reveal the structure of, and interactions between the free-jet and recirculating combustion chamber flows. Performance based on these CFD results is slightly less than that of the constant-pressure-combustion cycle analysis primarily due to these interactions. These differences are quantified and discussed. Additional CFD results at the Mach 8 flight condition show the effects of nozzle throat area variation on combustion chamber pressure, flow structure, and performance. Calculations with constant temperature walls were also done to evaluate heat flux and overall heat loads. Aspects of the concept that warrant further study are outlined. These include diffuser design, ramjet operation, mode transition, loss mechanisms, and the effects of secondary flow for wall cooling and combustion chamber pressurization. Also recommended is an examination of system-level aspects such as weight, thermal management and rocket integration as well as alternate geometries and variable geometry schemes.

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

NASA Image and Video Library

2001-03-15

As part of a combined systems test conducted by NASA Dryden Flight Research Center, NASA's NB-52B carrier aircraft rolls down a taxiway at Edwards Air Force Base with the X-43A hypersonic research aircraft and its modified Pegasus® booster rocket attached to a pylon under its right wing. The taxi test was one of the last major milestones in the Hyper-X research program before the first X-43A flight. The X-43A flights will be the first actual flight tests of an aircraft powered by a revolutionary supersonic-combustion ramjet ("scramjet") engine capable of operating at hypersonic speeds (above Mach 5, or five times the speed of sound). The 12-foot, unpiloted research vehicle was developed and built by MicroCraft Inc., Tullahoma, Tenn., under NASA contract. The booster was built by Orbital Sciences Corp., Dulles, Va. After being air-launched from NASA's venerable NB-52 mothership, the booster will accelerate the X-43A to test speed and altitude. The X-43A will then separate from the rocket and fly a pre-programmed trajectory, conducting aerodynamic and propulsion experiments until it descends into the Pacific Ocean. Three research flights are planned, two at Mach 7 and one at Mach 10.

Preliminary results from a subsonic high angle-of-attack flush airdata sensing (HI-FADS) system: Design, calibration, and flight test evaluation

NASA Technical Reports Server (NTRS)

Whitmore, Stephen A.; Moes, Timothy R.; Larson, Terry J.

1990-01-01

A nonintrusive high angle-of-attack flush airdata sensing (HI-FADS) system was installed and flight-tested on the F-18 high alpha research flight vehicle. The system is a matrix of 25 pressure orifices in concentric circles on the nose of the vehicle. The orifices determine angles of attack and sideslip, Mach number, and pressure altitude. Pressure was transmitted from the orifices to an electronically scanned pressure module by lines of pneumatic tubing. The HI-FADS system was calibrated and demonstrated using dutch roll flight maneuvers covering large Mach, angle-of-attack, and sideslip ranges. Reference airdata for system calibration were generated by a minimum variance estimation technique blending measurements from two wingtip airdata booms with inertial velocities, aircraft angular rates and attitudes, precision radar tracking, and meteorological analyses. The pressure orifice calibration was based on identifying empirical adjustments to modified Newtonian flow on a hemisphere. Calibration results are presented. Flight test results used all 25 orifices or used a subset of 9 orifices. Under moderate maneuvering conditions, the HI-FADS system gave excellent results over the entire subsonic Mach number range up to 55 deg angle of attack. The internal pneumatic frequency response of the system is accurate to beyond 10 Hz. Aerodynamic lags in the aircraft flow field caused some performance degradation during heavy maneuvering.

Noise Prediction of NASA SR2 Propeller in Transonic Conditions

NASA Astrophysics Data System (ADS)

Gennaro, Michele De; Caridi, Domenico; Nicola, Carlo De

2010-09-01

In this paper we propose a numerical approach for noise prediction of high-speed propellers for Turboprop applications. It is based on a RANS approach for aerodynamic simulation coupled with Ffowcs Williams-Hawkings (FW-H) Acoustic Analogy for propeller noise prediction. The test-case geometry adopted for this study is the 8-bladed NASA SR2 transonic cruise propeller, and simulated Sound Pressure Levels (SPL) have been compared with experimental data available from Wind Tunnel and Flight Tests for different microphone locations in a range of Mach numbers between 0.78 and 0.85 and rotational velocities between 7000 and 9000 rpm. Results show the ability of this approach to predict noise to within a few dB of experimental data. Moreover corrections are provided to be applied to acoustic numerical results in order for them to be compared with Wind Tunnel and Flight Test experimental data, as well computational grid requirements and guidelines in order to perform complete aerodynamic and aeroacoustic calculations with highly competitive computational cost.

Structural and Optical Properties Studies Of Ar2+ Ion Implanted Mn Deposited GaAs

NASA Astrophysics Data System (ADS)

De Gennaro, Michele; Caridi, Domenico; de Nicola, Carlo

2010-09-01

In this paper we propose a numerical approach for noise prediction of high-speed propellers for Turboprop applications. It is based on a RANS approach for aerodynamic simulation coupled with Ffowcs Williams-Hawkings (FW-H) Acoustic Analogy for propeller noise prediction. The test-case geometry adopted for this study is the 8-bladed NASA SR2 transonic cruise propeller, and simulated Sound Pressure Levels (SPL) have been compared with experimental data available from Wind Tunnel and Flight Tests for different microphone locations in a range of Mach numbers between 0.78 and 0.85 and rotational velocities between 7000 and 9000 rpm. Results show the ability of this approach to predict noise to within a few dB of experimental data. Moreover corrections are provided to be applied to acoustic numerical results in order for them to be compared with Wind Tunnel and Flight Test experimental data, as well computational grid requirements and guidelines in order to perform complete aerodynamic and aeroacoustic calculations with highly competitive computational cost.

Comparison of the Drag of a Fin-Stabilized Body of Revolution and of a Complete Airplane Configuration as Obtained at Transonic Speeds in a Slotted Wind Tunnel and in Free Flight

NASA Technical Reports Server (NTRS)

Howell, Robert R.; Braslow, Albert L.

1955-01-01

A comparison of the zero-lift drag coefficients at Mach numbers from 0.81 to 1.41 of a fin-stabilized parabolic body of revolution as measured in the Langley transonic blowdown tunnel has been made with measurements obtained in free-flight on a larger but geometrically similar model. The absolute values of drag coefficient obtained in the slotted wind tunnel were equivalent to the free-flight drag-coefficient values up to a Mach number of 1.4 when adjustments were made for the effect on viscous drag of differences in Reynolds number between the two test conditions. Excellent agreement was obtained between the two tests for the pressure-drag variation with Mach number, regardless of whether the scale effect on skin friction was considered. Favorable agreement was also obtained between the pressure-drag increments due t o the presence of the stabilizing fins as determined in the wine tunnel from fins-on and fins-off tests and as obtained by a different method in free flight. Tests of a specific airplane configuration to obtain an indication of the problems involved in the construction and tests of small-scale (approximately 7-inch span) complete airplane configuration with internal air flow indicated that reliable zero-lift drag-coefficient measurements at Mach numbers up to 1.4 can be attained with such models, provided the model is constructed with a high but not an unreasonable degree of accuracy.

Experimental Aerodynamic Characteristics of the Pegasus Air-Launched Booster and Comparisons with Predicted and Flight Results

NASA Technical Reports Server (NTRS)

Rhode, M. N.; Engelund, Walter C.; Mendenhall, Michael R.

1995-01-01

Experimental longitudinal and lateral-directional aerodynamic characteristics were obtained for the Pegasus and Pegasus XL configurations over a Mach number range from 1.6 to 6 and angles of attack from -4 to +24 degrees. Angle of sideslip was varied from -6 to +6 degrees, and control surfaces were deflected to obtain elevon, aileron, and rudder effectiveness. Experimental data for the Pegasus configuration are compared with engineering code predictions performed by Nielsen Engineering & Research, Inc. (NEAR) in the aerodynamic design of the Pegasus vehicle, and with results from the Aerodynamic Preliminary Analysis System (APAS) code. Comparisons of experimental results are also made with longitudinal flight data from Flight #2 of the Pegasus vehicle. Results show that the longitudinal aerodynamic characteristics of the Pegasus and Pegasus XL configurations are similar, having the same lift-curve slope and drag levels across the Mach number range. Both configurations are longitudinally stable, with stability decreasing towards neutral levels as Mach number increases. Directional stability is negative at moderate to high angles of attack due to separated flow over the vertical tail. Dihedral effect is positive for both configurations, but is reduced 30-50 percent for the Pegasus XL configuration because of the horizontal tail anhedral. Predicted longitudinal characteristics and both longitudinal and lateral-directional control effectiveness are generally in good agreement with experiment. Due to the complex leeside flowfield, lateral-directional characteristics are not as well predicted by the engineering codes. Experiment and flight data are in good agreement across the Mach number range.

Qualification flight tests of the Viking decelerator system.

NASA Technical Reports Server (NTRS)

Moog, R. D.; Bendura, R. J.; Timmons, J. D.; Lau, R. A.

1973-01-01

The Balloon Launched Decelerator Test (BLDT) series conducted at White Sands Missile Range (WSMR) during July and August of 1972 flight qualified the NASA Viking '75 decelerator system at conditions bracketing those expected for Mars. This paper discusses the decelerator system design requiremnts, compares the test results with prior work, and discusses significant considerations leading to successful qualification in earth's atmosphere. The Viking decelerator system consists of a single-stage mortar-deployed 53-foot nominal diameter disk-gap-band parachute. Full-scale parachutes were deployed behind a full-scale simulated Viking vehicle at Mach numbers from 0.47 to 2.18 and dynamic pressures from 6.9 to 14.6 psf. Analyses show that the system is qualified with sufficient margin to perform successfully for the Viking mission.

Prediction of Aerodynamic Coefficient using Genetic Algorithm Optimized Neural Network for Sparse Data

NASA Technical Reports Server (NTRS)

Rajkumar, T.; Bardina, Jorge; Clancy, Daniel (Technical Monitor)

2002-01-01

Wind tunnels use scale models to characterize aerodynamic coefficients, Wind tunnel testing can be slow and costly due to high personnel overhead and intensive power utilization. Although manual curve fitting can be done, it is highly efficient to use a neural network to define the complex relationship between variables. Numerical simulation of complex vehicles on the wide range of conditions required for flight simulation requires static and dynamic data. Static data at low Mach numbers and angles of attack may be obtained with simpler Euler codes. Static data of stalled vehicles where zones of flow separation are usually present at higher angles of attack require Navier-Stokes simulations which are costly due to the large processing time required to attain convergence. Preliminary dynamic data may be obtained with simpler methods based on correlations and vortex methods; however, accurate prediction of the dynamic coefficients requires complex and costly numerical simulations. A reliable and fast method of predicting complex aerodynamic coefficients for flight simulation I'S presented using a neural network. The training data for the neural network are derived from numerical simulations and wind-tunnel experiments. The aerodynamic coefficients are modeled as functions of the flow characteristics and the control surfaces of the vehicle. The basic coefficients of lift, drag and pitching moment are expressed as functions of angles of attack and Mach number. The modeled and training aerodynamic coefficients show good agreement. This method shows excellent potential for rapid development of aerodynamic models for flight simulation. Genetic Algorithms (GA) are used to optimize a previously built Artificial Neural Network (ANN) that reliably predicts aerodynamic coefficients. Results indicate that the GA provided an efficient method of optimizing the ANN model to predict aerodynamic coefficients. The reliability of the ANN using the GA includes prediction of aerodynamic coefficients to an accuracy of 110% . In our problem, we would like to get an optimized neural network architecture and minimum data set. This has been accomplished within 500 training cycles of a neural network. After removing training pairs (outliers), the GA has produced much better results. The neural network constructed is a feed forward neural network with a back propagation learning mechanism. The main goal has been to free the network design process from constraints of human biases, and to discover better forms of neural network architectures. The automation of the network architecture search by genetic algorithms seems to have been the best way to achieve this goal.

Experimental and Computational Investigation of a Translating-Throat Single-Expansion-Ramp Nozzle

NASA Technical Reports Server (NTRS)

Deere, Karen A.; Asbury, Scott C.

1999-01-01

An experimental and computational study was conducted on a high-speed, single-expansion-ramp nozzle (SERN) concept designed for efficient off-design performance. The translating-throat SERN concept adjusts the axial location of the throat to provide a variable expansion ratio and allow a more optimum jet exhaust expansion at various flight conditions in an effort to maximize nozzle performance. Three design points (throat locations) were investigated to simulate the operation of this concept at subsonic-transonic, low supersonic, and high supersonic flight conditions. The experimental study was conducted in the jet exit test facility at the Langley Research Center. Internal nozzle performance was obtained at nozzle pressure ratios (NPR's) up to 13 for six nozzles with design nozzle pressure ratios near 9, 42, and 102. Two expansion-ramp surfaces, one concave and one convex, were tested for each design point. Paint-oil flow and focusing schlieren flow visualization techniques were utilized to acquire additional flow data at selected NPR'S. The Navier-Stokes code, PAB3D, was used with a two-equation k-e turbulence model for the computational study. Nozzle performance characteristics were predicted at nozzle pressure ratios of 5, 9, and 13 for the concave ramp, low Mach number nozzle and at 10, 13, and 102 for the concave ramp, high Mach number nozzle.

Hyper-X Research Vehicle (HXRV) Experimental Aerodynamics Test Program Overview

NASA Technical Reports Server (NTRS)

Holland, Scott D.; Woods, William C.; Engelund, Walter C.

2000-01-01

This paper provides an overview of the experimental aerodynamics test program to ensure mission success for the autonomous flight of the Hyper-X Research Vehicle (HXRV). The HXRV is a 12-ft long, 2700 lb lifting body technology demonstrator designed to flight demonstrate for the first time a fully airframe integrated scramjet propulsion system. Three flights are currently planned, two at Mach 7 and one at Mach 10, beginning in the fall of 2000. The research vehicles will be boosted to the prescribed scramjet engine test point where they will separate from the booster, stabilize. and initiate engine test. Following 5+ seconds of powered flight and 15 seconds of cowl-open tares, the cowl will close and the vehicle will fly a controlled deceleration trajectory which includes numerous control doublets for in-flight aerodynamic parameter identification. This paper reviews the preflight testing activities, wind tunnel models, test rationale. risk reduction activities, and sample results from wind tunnel tests supporting the flight trajectory of the HXRV from hypersonic engine test point through subsonic flight termination.

Hyper-X Research Vehicle (HXRV) Experimental Aerodynamics Test Program Overview

NASA Technical Reports Server (NTRS)

Holland, Scott D.; Woods, William C.; Engelund, Walter C.

2000-01-01

This paper provides an overview of the experimental aerodynamics test program to ensure mission success for the autonomous flight of the Hyper-X Research Vehicle (HXRV). The HXRV is a 12-ft long, 2700 lb lifting body technology demonstrator designed to flight demonstrate for the first time a fully airframe integrated scramjet propulsion system. Three flights are currently planned, two at Mach 7 and one at Mach 10, beginning in the fall of 2000. The research vehicles will be boosted to the prescribed scramjet engine test point where they will separate from the booster, stabilize, and initiate engine test. Following 5+ seconds of powered flight and 15 seconds of cow-open tares, the cowl will close and the vehicle will fly a controlled deceleration trajectory which includes numerous control doublets for in-flight aerodynamic parameter identification. This paper reviews the preflight testing activities, wind tunnel models, test rationale, risk reduction activities, and sample results from wind tunnel tests supporting the flight trajectory of the HXRV from hypersonic engine test point through subsonic flight termination.

Further considerations of engine emissions from subsonic aircraft at cruise altitude

NASA Astrophysics Data System (ADS)

Lee, S. H.; Le Dilosquer, M.; Singh, R.; Rycroft, M. J.

The most significant man-made sources of pollution of the higher troposphere and lower stratosphere are exhaust emissions from civil subsonic aircraft at cruise altitude (8-12 km). This paper examines such issues by computational modelling of Boeing 747-400 flights during their cruise phase between selected city pairs, for example London to Tokyo. The engine performance, exhaust pollutant prediction, and detailed flight history analysis effects of different Mach numbers and of increasing the cruise altitude from 9.8 to 12.1 km during the flight rather than staying at a constant cruise altitude of 10.5 km are studied in detail. To minimise the overall effects of atmospheric pollution, a Mach number of 0.85 and increasing altitude is the favoured cruise technique.

Flow Field Investigations of a Simulated Weapons Cavity at Mach 3.

DTIC Science & Technology

1981-12-01

AD-Alll 843 ARMO" ENGINEERING. DEVELOPMENT CENTER ARNOLD AFS TN F/9 14/2 FLOW FIELD INVESTIGATIONS OF A SIMULATED WEAPONS CAVITY AT MACN--ETC(U) DEC...TEST CHART .. AEDC-TSR-81-V37 FLOW FIELD INVESTIGATIONS OF A _____SIMULATED WEAPONS CAVITY AT MACH 3 _~W. A. Crosby Calspan Field Services, Inc...TYPE OF REPORT & PERIOD COVERED Final Report FLOW FIELD INVESTIGATIONS OF A SIMULATED WEAPONS 27 October 1981 CAVITY AT MACH 3 6. PERFORMING O1G

A hypersonic research vehicle to develop scramjet engines

NASA Technical Reports Server (NTRS)

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

1990-01-01

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

Effects of Wing Sweep on In-flight Boundary-layer Transition for a Laminar Flow Wing at Mach Numbers from 0.60 to 0.79

NASA Technical Reports Server (NTRS)

Anderson, Bianca Trujillo; Meyer, Robert R., Jr.

1990-01-01

The variable sweep transition flight experiment (VSTFE) was conducted on an F-14A variable sweep wing fighter to examine the effect of wing sweep on natural boundary layer transition. Nearly full span upper surface gloves, extending to 60 percent chord, were attached to the F-14 aircraft's wings. The results are presented of the glove 2 flight tests. Glove 2 had an airfoil shape designed for natural laminar flow at a wing sweep of 20 deg. Sample pressure distributions and transition locations are presented with the complete results tabulated in a database. Data were obtained at wing sweeps of 15, 20, 25, 30, and 35 deg, at Mach numbers ranging from 0.60 to 0.79, and at altitudes ranging from 10,000 to 35,000 ft. Results show that a substantial amount of laminar flow was maintained at all the wing sweeps evaluated. The maximum transition Reynolds number obtained was 18.6 x 10(exp 6) at 15 deg of wing sweep, Mach 0.75, and at an altitude of 10,000 ft.

B-70 Aircraft Study. Volume 2

NASA Technical Reports Server (NTRS)

1972-01-01

Volume 2 of the final report on the B-70 aircraft study is presented here. The B-70 Program, at the onset, was a full weapon system capable of sustained Mach 3 flight for the major portion of its design missions. The weapon system was to enter the SAC inventory as an RS-70 with the first intercontinental resonnaissance/bomber wing scheduled to go operational in July, 1964. After several redirections, a two XB-70 air vehicle program emerged with its prime objective being to demonstrate the technical feasibility of sustained Mach 3 flight. This section describes the original Weapon System 110A concepts, the evolution of the RS-70 design, and the XB-70 air vehicles which demonstrated the design, fabrication, and technical feasibility of long range Mach 3 flights at high altitude. The data presented shows that a very large step forward in the state-of-the-art of manned aircraft design was achieved during the B-70 development program and that advances were made and incorporated in every area, including design, materials application, and manufacturing techniques.

77 FR 12158 - Airworthiness Directives; Bombardier, Inc. Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2012-02-29

... limited to not more than 0.78 Mach. (6) The climb ceiling obtained from the Flight Planning and Cruise... damage to airplane structure, which could adversely affect the airplane's continued safe flight and... requirements.'' Under that section, Congress charges the FAA with promoting safe flight of civil aircraft in...

Potential benefits of propulsion and flight control integration for supersonic cruise vehicles

NASA Technical Reports Server (NTRS)

Berry, D. T.; Schweikhard, W. G.

1976-01-01

Typical airframe/propulsion interactions such as Mach/altitude excursions and inlet unstarts are reviewed. The improvements in airplane performance and flight control that can be achieved by improving the interfaces between propulsion and flight control are estimated. A research program to determine the feasibility of integrating propulsion and flight control is described. This program includes analytical studies and YF-12 flight tests.

Ram-Jet off Design Performances

NASA Astrophysics Data System (ADS)

Andriani, Roberto; Ghezzi, Umberto

2002-01-01

In this work it is intended to study the off-design performances of a ram jet engine. To this purpouse it has been analyzed in a first time the behaviour of an ideal engine, that means to not consider the losses in the various components, or, under a thermodynamic point of view, to consider the fluid transformation through the air intake and exhaust nozzle, remembering that in a ram jet there are not rotating components as compressor and turbine, isentropic. Referring to the ram-jet scheme of fig.1. we can say, neglecting the fuel introduced, that the air mass flow rate throughout the engine is constant. If we consider the two control sections 4 and 8, respectively the throat section of the converging-diverging supersonic inlet and the throat section of the discharge nozzle, the condition of constant mass flow leads to the relation: m4 =f (M 4 ) m8 = m 4 = m8 We can imaging that the throat section # 4 is always choked for any value of the flight Mach number M0. This means that the throat section 4 is adjusted at any value of M0 so that the flow Mach number in 4 is equal to unity. In this it follows: R. Andriani, U. Ghezzi1 Since in an ideal case T t8 The relation [1] allows to determine the T8 temperature, that represent the maximum cycle temperature, for different operating conditions, as flight Mach number and altitude. We then have two cases: the first is A8 (nozzle throat section) fixed, and the second is A8 variable. In the first case the maximum temperature T8 is univocally determined by the operating condition. In the second case A8 can be varied so to maintain T8 at a chosen value. The graphic of fig.2 shows the first case. In particular it has been considered as design point an altitude of 15000 meters and a flight Mach number equal to 2. In this condition it has been evaluated the section A8 for unity mass flow rate. At the same altitude, varying the flight Mach number, with the section A4 always choked, the graphic shows the variation of the maximum cycle temperature according to the equation *. We notice as the temperature raises to unacceptable levels as we accelerate from the design point. For instance we can see as if we move to flight Mach number 2.2 the maximum temperature raises from 1800 Kelvin degrees (design point) to more than 2900 [K], a value practically unreachable. This shows as the ramjet with fixed discharge nozzle allows very little variation of the operating conditions around the design point. If we consider instead the possibility to vary the nozzle section A8 we can imagine to adjust it so to maintain constant the temperature T8, according to the relation *. The graphics of fig.3 show the area variation required to the section A8 to maintain constant the temperature T8 at the same altitude for different values of the flight Mach number. Considering a circular section, on the same R. Andriani, U. Ghezzi2 picture is reported the corresponding variation of the diameter. On the Y-axis it is reported the ratio between the entity (section and diameter) at the considered condition and at the design point. For the case of A8 section variable it has also been evaluated the behavior of the specific fuel consumption, the mass flow rate and the thrust. In fig.4 they are reported the behavior TSFC and Fuel/air ratio at different flight Mach numbers, and the ratio between them and the design value. In fig .5 it is reported the ratio between the thrust and mass flow rate at different flight conditions and their value at design point. R. Andriani, U. Ghezzi3

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.

Some effects of wing and body geometry on the aerodynamic characteristics of configurations designed for high supersonic Mach numbers

NASA Technical Reports Server (NTRS)

Spearman, M. L.; Tice, David C.; Braswell, Dorothy O.

1992-01-01

Experimental and theoretical results are presented for a family of aerodynamic configurations for flight Mach numbers as high as Mach 8. All of these generic configurations involved 70-deg sweep delta planform wings of three different areas and three fuselage shapes with circular-to-elliptical cross sections. It is noted that fuselage ellipticity enhances lift-curve slope and maximum L/D, while decreasing static longitudinal stability (especially with smaller wing areas).

Pressure distribution for the wing of the YAV-8B airplane; with and without pylons

NASA Technical Reports Server (NTRS)

Saltzman, Edwin J.; Delfrate, John H.; Sabsay, Catherine M.; Yarger, Jill M.

1992-01-01

Pressure distribution data have been obtained in flight at four span stations on the wing panel of the YAV-8B airplane. Data obtained for the supercritical profiled wing, with and without pylons installed, ranged from Mach 0.46 to 0.88. The altitude ranged from approximately 20,000 to 40,000 ft and the resultant Reynolds numbers varied from approximately 7.2 million to 28.7 million based on the mean aerodynamic chord. Pressure distribution data and flow visualization results show that the full-scale flight wing performance is compromised because the lower surface cusp region experiences flow separation for some important transonic flight conditions. This condition is aggravated when local shocks occur on the lower surface of the wing (mostly between 20 and 35 percent chord) when the pylons are installed for Mach 0.8 and above. There is evidence that convex fairings, which cover the pylon attachment flanges, cause these local shocks. Pressure coefficients significantly more negative than those for sonic flow also occur farther aft on the lower surface (near 60 percent chord) whether or not the pylons are installed for Mach numbers greater than or equal to 0.8. These negative pressure coefficient peaks and associated local shocks would be expected to cause increasing wave and separation drag at transonic Mach number increases.

On the attitude control and flight result of winged reentry test vehicle

NASA Astrophysics Data System (ADS)

Kawaguchi, Jun'ichiro; Inatani, Yoshifumi; Yonemoto, Koichi; Hinada, Motoki

The Institute of Space and Astronautical Science (ISAS) has been studying the unmanned winged space vehicle HIMES (HIghly Maneuverable Engineering Space vehicle) for a decade and successfully carried out sub-sonic Gliding Flight Experiments several years ago, which was followed by Reentry Flight Experiment, utilizing so called 'Rockoon' method, in September of 1988, which failed due to the unexpected burst of the balloon. ISAS conducted it again making use of refined 'Rockoon' scheme in February of 1992. In spite of its small bulk property, it was equipped with not only a reaction control system (RCS) but a surface control system (SCS) capability as well, which enabled it to make a successful flight under both vacuum and atmospheric circumstances. The highest Mach number exceeded 3.5 and the highest altitude was a bit lower to 67 km. Switching from reaction control to surface control was one of the essential engineering interests in the flight like this. Supersonic autonomous flight control with high angle of attack was also what should be established through this, since in general it inevitably carries inherent lateral instability. A flight test this time revealed those features and characteristics quite well. This paper deals with the attitude control strategy with three-axis Motion Simulation Test as well as the flight results.

KC-135 wing and winglet flight pressure distributions, loads, and wing deflection results with some wind tunnel comparisons

NASA Technical Reports Server (NTRS)

Montoya, L. C.; Jacobs, P.; Flechner, S.; Sims, R.

1982-01-01

A full-scale winglet flight test on a KC-135 airplane with an upper winglet was conducted. Data were taken at Mach numbers from 0.70 to 0.82 at altitudes from 34,000 feet to 39,000 feet at stabilized flight conditions for wing/winglet configurations of basic wing tip, 15/-4 deg, 15/-2 deg, and 0/-4 deg winglet cant/incidence. An analysis of selected pressure distribution and data showed that with the basic wing tip, the flight and wind tunnel wing pressure distribution data showed good agreement. With winglets installed, the effects on the wing pressure distribution were mainly near the tip. Also, the flight and wind tunnel winglet pressure distributions had some significant differences primarily due to the oilcanning in flight. However, in general, the agreement was good. For the winglet cant and incidence configuration presented, the incidence had the largest effect on the winglet pressure distributions. The incremental flight wing deflection data showed that the semispan wind tunnel model did a reasonable job of simulating the aeroelastic effects at the wing tip. The flight loads data showed good agreement with predictions at the design point and also substantiated the predicted structural penalty (load increase) of the 15 deg cant/-2 deg incidence winglet configuration.

Ares I and Ares I-X Stage Separation Aerodynamic Testing

NASA Technical Reports Server (NTRS)

Pinier, Jeremy T.; Niskey, Charles J.

2011-01-01

The aerodynamics of the Ares I crew launch vehicle (CLV) and Ares I-X flight test vehicle (FTV) during stage separation was characterized by testing 1%-scale models at the Arnold Engineering Development Center s (AEDC) von Karman Gas Dynamics Facility (VKF) Tunnel A at Mach numbers of 4.5 and 5.5. To fill a large matrix of data points in an efficient manner, an injection system supported the upper stage and a captive trajectory system (CTS) was utilized as a support system for the first stage located downstream of the upper stage. In an overall extremely successful test, this complex experimental setup associated with advanced postprocessing of the wind tunnel data has enabled the construction of a multi-dimensional aerodynamic database for the analysis and simulation of the critical phase of stage separation at high supersonic Mach numbers. Additionally, an extensive set of data from repeated wind tunnel runs was gathered purposefully to ensure that the experimental uncertainty would be accurately quantified in this type of flow where few historical data is available for comparison on this type of vehicle and where Reynolds-averaged Navier-Stokes (RANS) computational simulations remain far from being a reliable source of static aerodynamic data.

An inventory of aeronautical ground research facilities. Volume 2: Air breathing engine test facilities

NASA Technical Reports Server (NTRS)

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

1971-01-01

The inventory covers free jet and direct connect altitude cells, sea level static thrust stands, sea level test cells with ram air, and propulsion wind tunnels. Free jet altitude cells and propulsion wind tunnels are used for evaluation of complete inlet-engine-exhaust nozzle propulsion systems under simulated flight conditions. These facilities are similar in principal of operation and differ primarily in test section concept. The propulsion wind tunnel provides a closed test section and restrains the flow around the test specimen while the free jet is allowed to expand freely. A chamber of large diameter about the free jet is provided in which desired operating pressure levels may be maintained. Sea level test cells with ram air provide controlled, conditioned air directly to the engine face for performance evaluation at low altitude flight conditions. Direct connect altitude cells provide a means of performance evaluation at simulated conditions of Mach number and altitude with air supplied to the flight altitude conditions. Sea level static thrust stands simply provide an instrumented engine mounting for measuring thrust at zero airspeed. While all of these facilities are used for integrated engine testing, a few provide engine component test capability.

Navier-Stokes computations with finite-rate chemistry for LO2/LH2 rocket engine plume flow studies

NASA Technical Reports Server (NTRS)

Dougherty, N. Sam; Liu, Baw-Lin

1991-01-01

Computational fluid dynamics methods have been developed and applied to Space Shuttle Main Engine LO2/LH2 plume flow simulation/analysis of airloading and convective base heating effects on the vehicle at high flight velocities and altitudes. New methods are described which were applied to the simulation of a Return-to-Launch-Site abort where the vehicle would fly briefly at negative angles of attack into its own plume. A simplified two-perfect-gases-mixing approach is used where one gas is the plume and the other is air at 180-deg and 135-deg flight angle of attack. Related research has resulted in real gas multiple-plume interaction methods with finite-rate chemistry described herein which are applied to the same high-altitude-flight conditions of 0 deg angle of attack. Continuing research plans are to study Orbiter wake/plume flows at several Mach numbers and altitudes during ascent and then to merge this model with the Shuttle 'nose-to-tail' aerodynamic and SRB plume models for an overall 'nose-to-plume' capability. These new methods are also applicable to future launch vehicles using clustered-engine LO2/LH2 propulsion.

In-flight transition measurement on a 10 deg cone at Mach numbers from 0.5 to 2.0

NASA Technical Reports Server (NTRS)

Fisher, D. F.; Dougherty, N. S., Jr.

1982-01-01

Boundary layer transition measurements were made in flight on a 10 deg transition cone tested previously in 23 wind tunnels. The cone was mounted on the nose of an F-15 aircraft and flown at Mach numbers room 0.5 to 2.0 and altitudes from 1500 meters (5000 feet) to 15,000 meters (50,000 feet), overlapping the Mach number/Reynolds number envelope of the wind tunnel tests. Transition was detected using a traversing pitot probe in contact with the surface. Data were obtained near zero cone incidence and adiabatic wall temperature. Transition Reynolds number was found to be a function of Mach number and of the ratio of wall temperature to adiabatic all temperature. Microphones mounted flush with the cone surface measured free-stream disturbances imposed on the laminar boundary layer and identified Tollmien-Schlichting waves as the probable cause of transition. Transition Reynolds number also correlated with the disturbance levels as measured by the cone surface microphones under a laminar boundary layer as well as the free-stream impact.

Effects of Inlet Modification and Rocket-Rack Extension on the Longitudinal Trim and Low-Lift Drag of the Douglas F5D-1 Airplane as Obtained with a 0.125-Scale Rocket-Boosted Model between Mach Numbers of 0.81 and 1.64, TED No. NACA AD 399

NASA Technical Reports Server (NTRS)

Hastings, Earl C., Jr.; Dickens, Waldo L.

1957-01-01

A flight investigation was conducted to determine the effects of an inlet modification and rocket-rack extension on the longitudinal trim and low-lift drag of the Douglas F5D-1 airplane. The investigation was conducted with a 0.125-scale rocket-boosted model which was flight tested at the Langley Pilotless Aircraft Research Station at Wallops Island, Va. Results indicate that the combined effects of the modified inlet and fully extended rocket racks on the trim lift coefficient and trim angle of attack were small between Mach numbers of 0.94 and 1.57. Between Mach numbers of 1.10 and 1.57 there was an average increase in drag coefficient of about o,005 for the model with modified inlet and extended rocket racks. The change in drag coefficient due to the inlet modification alone is small between Mach numbers of 1.59 and 1.64

A Transonic Wind-Tunnel Investigation of a Seaplane Configuration having a 40 Deg Sweptback Wing, TED No. NACA DE 387

NASA Technical Reports Server (NTRS)

Hieser, Gerald; Kudlacik, Louis; Gray, W. H.

1956-01-01

During the course of an aerodynamic loads investigation of a model of the Martin XP6M-1 flying boat in the.Langley 16-foot transonic tunnel, longitudinal-aerodynamic-performance information was obtained. Data were obtained at speeds up to and exceeding those anticipated for the seaplane in level flight and included the Mach number range from 0.84. to 1.09. The angle of attack was varied from -2deg to 6deg and the average Reynolds number, based on wing mean aerodyn&ic chord, was about 3.7 x 10(exp 6). This seaplane, although not designed to maintain level flight at Mach numbers beyond the force break, was found to have a transonic drag-rise coefficient of 0.0728, with an accompanying drag-rise Mach number of about 0.85. A large portion of the.drag rise and the relatively low value of drag-rise Mach number result from the axial coincidence of the maximum areas of the principal airplane components.

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.

X-15A-2 with full scale ablative and external tanks installed parked in front of hangar

NASA Image and Video Library

1967-08-04

X-15A-2 with full scale ablative and external tanks installed parked in front of hangar. In June 1967, the X-15A-2 rocket-powered research aircraft received a full-scale ablative coating to protect the craft from the high temperatures associated with hypersonic flight (above Mach 5). This pink eraser-like substance, applied to the X-15A-2 aircraft (56-6671), was then covered with a white sealant coat before flight. This coating would help the #2 aircraft reach the record speed of 4,520 mph (Mach 6.7).

An Investigation of Transonic Resonance in a Mach 2.2 Round Convergent-Divergent Nozzle

NASA Technical Reports Server (NTRS)

Dippold, Vance F., III; Zaman, Khairul B. M. Q.

2015-01-01

Hot-wire and acoustic measurements were taken for a round convergent nozzle and a round convergent-divergent (C-D) nozzle at a jet Mach number of 0.61. The C-D nozzle had a design Mach number of 2.2. Compared to the convergent nozzle jet flow, the Mach 2.2 nozzle jet flow produced excess broadband noise (EBBN). It also produced a transonic resonance tone at 1200 Herz. Computational simulations were performed for both nozzle flows. A steady Reynolds-Averaged Navier-Stokes simulation was performed for the convergent nozzle jet flow. For the Mach 2.2 nozzle flow, a steady RANS simulation, an unsteady RANS (URANS) simulation, and an unsteady Detached Eddy Simulation (DES) were performed. The RANS simulation of the convergent nozzle showed good agreement with the hot-wire velocity and turbulence measurements, though the decay of the potential core was over-predicted. The RANS simulation of the Mach 2.2 nozzle showed poor agreement with the experimental data, and more closely resembled an ideally-expanded jet. The URANS simulation also showed qualitative agreement with the hot-wire data, but predicted a transonic resonance at 1145 Herz. The DES showed good agreement with the hot-wire velocity and turbulence data. The DES also produced a transonic tone at 1135 Herz. The DES solution showed that the destabilization of the shock-induced separation region inside the nozzle produced increased levels of turbulence intensity. This is likely the source of the EBBN.

Measured far-field flight noise of a counterrotation turboprop at cruise conditions

NASA Technical Reports Server (NTRS)

Woodward, Richard P.; Loeffler, Irvin J.; Dittmar, James H.

1989-01-01

Modern high speed propeller (advanced turboprop) aircraft are expected to operate on 50 to 60 percent less fuel than the 1980 vintage turbofan fleet while at the same time matching the flight speed and performance of those aircraft. Counterrotation turboprop engines offer additional fuel savings by means of upstream propeller swirl recovery. This paper presents acoustic sideline results for a full-scale counterrotation turboprop engine at cruise conditions. The engine was installed on a Boeing 727 aircraft in place of the right-side turbofan engine. Acoustic data were taken from an instrumented Learjet chase plane. Sideline acoustic results are presented for 0.50 and 0.72 Mach cruise conditions. A scale model of the engine propeller was tested in a wind tunnel at 0.72 Mach cruise conditions. The model data were adjusted to flight acquisition conditions and were in general agreement with the flight results.

Results of wind tunnel tests of an ASRM configured 0.03 scale Space Shuttle integrated vehicle model (47-OTS) in the AEDC 16-foot transonic wind tunnel, volume 2

NASA Technical Reports Server (NTRS)

Marroquin, J.; Lemoine, P.

1992-01-01

An experimental Aerodynamic and Aero-Acoustic loads data base was obtained at transonic Mach numbers for the Space Shuttle Launch Vehicle configured with the ASRM Solid Rocket Boosters as an increment to the current flight configuration (RSRB). These data were obtained during transonic wind tunnel tests (IA 613A) conducted in the Arnold Engineering Development Center 16-Foot transonic propulsion wind tunnel from March 27, 1991 through April 12, 1991. This test is the first of a series of two tests covering the Mach range from 0.6 to 3.5. Steady state surface static and fluctuating pressure distributions over the Orbiter, External Tank and Solid Rocket Boosters of the Shuttle Integrated Vehicle were measured. Total Orbiter forces, Wing forces and Elevon hinge moments were directly measured as well from force balances. Two configurations of Solid Rocket Boosters were tested, the Redesigned Solid Rocket Booster (RSRB) and the Advanced Solid Rocket Motor (ASRM). The effects of the position (i.e., top, bottom, top and bottom) of the Integrated Electronics Assembly (IEA) box, mounted on the SRB attach ring, were obtained on the ASRM configured model. These data were obtained with and without Solid Plume Simulators which, when used, matched as close as possible the flight derived pressures on the Orbiter and External Tank base. Data were obtained at Mach numbers ranging from 0.6 to 1.55 at a Unit Reynolds Number of 2.5 million per foot through model angles of attack from -8 to +4 degrees at sideslip angles of 0, +4 and -4 degrees.

Results of wind tunnel tests of an ASRM configured 0.03 scale Space Shuttle integrated vehicle model (47-OTS) in the AEDC 16-foot Transonic wind tunnel (IA613A), volume 1

NASA Technical Reports Server (NTRS)

Marroquin, J.; Lemoine, P.

1992-01-01

An experimental Aerodynamic and Aero-Acoustic loads data base was obtained at transonic Mach numbers for the Space Shuttle Launch Vehicle configured with the ASRM Solid Rocket Boosters as an increment to the current flight configuration (RSRB). These data were obtained during transonic wind tunnel tests (IA 613A) conducted in the Arnold Engineering Development Center 16-Foot transonic propulsion wind tunnel from March 27, 1991 through April 12, 1991. This test is the first of a series of two tests covering the Mach range from 0.6 to 3.5. Steady state surface static and fluctuating pressure distributions over the Orbiter, External Tank and Solid Rocket Boosters of the Shuttle Integrated Vehicle were measured. Total Orbiter forces, Wing forces and Elevon hinge moments were directly measured as well from force balances. Two configurations of Solid Rocket Boosters were tested, the Redesigned Solid Rocket Booster (RSRB) and the Advanced Solid Rocket Motor (ASRM). The effects of the position (i.e. top, bottom, top and bottom) of the Integrated Electronics Assembly (IEA) box, mounted on the SRB attach ring, were obtained on the ASRM configured model. These data were obtained with and without Solid Plume Simulators which, when used, matched as close as possible the flight derived pressures on the Orbiter and External Tank base. Data were obtained at Mach numbers ranging from 0.6 to 1.55 at a Unit Reynolds Number of 2.5 million per foot through model angles of attack from -8 to +4 degrees at sideslip angles of 0, +4 and -4 degrees.

Flight Test Results of an Axisymmetric Channeled Center Body Supersonic Inlet at Off-Design Conditions

NASA Technical Reports Server (NTRS)

St. John, Clinton W.; Frederick, Michael Alan

2013-01-01

Flight-testing of a channeled center-body axisymmetric supersonic inlet design concept was conducted at the National Aeronautics and Space Administration (NASA) Dryden Flight Research Center in collaboration with the NASA Glenn Research Center (Cleveland, Ohio) and TechLand Research, Inc. (North Olmsted, Ohio). This testing utilized the Propulsion Flight Test Fixture, flown on the NASA F-15B research test bed airplane (NASA tail number 836) at local experiment Mach numbers up to 1.50. The translating channeled center-body inlet was designed by TechLand Research, Inc. (U.S. Patent No. 6,276,632 B1) to allow for a novel method of off-design flow matching, with original test planning conducted under a NASA Small Business Innovative Research study. Data were collected in flight at various off-design Mach numbers for fixed-geometry representations of both the channeled center-body design and an equivalent area smooth center-body design for direct comparison of total pressure recovery and limited distortion measurements.

Computational Analyses of the LIMX TBCC Inlet High-Speed Flowpath

NASA Technical Reports Server (NTRS)

Dippold, Vance F., III

2012-01-01

Reynolds-Averaged Navier-Stokes (RANS) simulations were performed for the high-speed flowpath and isolator of a dual-flowpath Turbine-Based Combined-Cycle (TBCC) inlet using the Wind-US code. The RANS simulations were performed in preparation for the Large-scale Inlet for Mode Transition (LIMX) model tests in the NASA Glenn Research Center (GRC) 10- by 10-ft Supersonic Wind Tunnel. The LIMX inlet has a low-speed flowpath that is coupled to a turbine engine and a high-speed flowpath designed to be coupled to a Dual-Mode Scramjet (DMSJ) combustor. These RANS simulations were conducted at a simulated freestream Mach number of 4.0, which is the nominal Mach number for the planned wind tunnel testing with the LIMX model. For the simulation results presented in this paper, the back pressure, cowl angles, and freestream Mach number were each varied to assess the performance and robustness of the high-speed inlet and isolator. Under simulated wind tunnel conditions at maximum inlet mass flow rates, the high-speed flowpath pressure rise was found to be greater than a factor of four. Furthermore, at a simulated freestream Mach number of 4.0, the high-speed flowpath and isolator showed stability for freestream Mach number that drops 0.1 Mach below the design point. The RANS simulations indicate the yet-untested highspeed inlet and isolator flowpath should operate as designed. The RANS simulation results also provided important insight to researchers as they developed test plans for the LIMX experiment in GRC s 10- by 10-ft Supersonic Wind Tunnel.

Aerodynamic Models for the Low Density Supersonic Decelerator (LDSD) Test Vehicles

NASA Technical Reports Server (NTRS)

Van Norman, John W.; Dyakonov, Artem; Schoenenberger, Mark; Davis, Jody; Muppidi, Suman; Tang, Chun; Bose, Deepak; Mobley, Brandon; Clark, Ian

2016-01-01

An overview of aerodynamic models for the Low Density Supersonic Decelerator (LDSD) Supersonic Flight Dynamics Test (SFDT) campaign test vehicle is presented, with comparisons to reconstructed flight data and discussion of model updates. The SFDT campaign objective is to test Supersonic Inflatable Aerodynamic Decelerator (SIAD) and large supersonic parachute technologies at high altitude Earth conditions relevant to entry, descent, and landing (EDL) at Mars. Nominal SIAD test conditions are attained by lifting a test vehicle (TV) to 36 km altitude with a helium balloon, then accelerating the TV to Mach 4 and 53 km altitude with a solid rocket motor. Test flights conducted in June of 2014 (SFDT-1) and 2015 (SFDT-2) each successfully delivered a 6 meter diameter decelerator (SIAD-R) to test conditions and several seconds of flight, and were successful in demonstrating the SFDT flight system concept and SIAD-R technology. Aerodynamic models and uncertainties developed for the SFDT campaign are presented, including the methods used to generate them and their implementation within an aerodynamic database (ADB) routine for flight simulations. Pre- and post-flight aerodynamic models are compared against reconstructed flight data and model changes based upon knowledge gained from the flights are discussed. The pre-flight powered phase model is shown to have a significant contribution to off-nominal SFDT trajectory lofting, while coast and SIAD phase models behaved much as predicted.

Comparison of concurrent strain gage- and pressure transducer-measured flight loads on a lifting reentry vehicle and correlation with wind tunnel predictions

NASA Technical Reports Server (NTRS)

Tang, M. H.; Sefic, W. J.; Sheldon, R. G.

1978-01-01

Concurrent strain gage and pressure transducer measured flight loads on a lifting reentry vehicle are compared and correlated with wind tunnel-predicted loads. Subsonic, transonic, and supersonic aerodynamic loads are presented for the left fin and control surfaces of the X-24B lifting reentry vehicle. Typical left fin pressure distributions are shown. The effects of variations in angle of attack, angle of sideslip, and Mach number on the left fin loads and rudder hinge moments are presented in coefficient form. Also presented are the effects of variations in angle of attack and Mach number on the upper flap, lower flap, and aileron hinge-moment coefficients. The effects of variations in lower flap hinge moments due to changes in lower flap deflection and Mach number are presented in terms of coefficient slopes.

Heat flux and shock shape measurements on an Aeroassist Flight Experiment model in a high enthalpy free piston shock tunnel

NASA Technical Reports Server (NTRS)

Gai, S. L.; Mudford, N. R.; Hackett, C.

1992-01-01

This paper describes measurements of heat flux and shock shapes made on a 2.08 percent scale model of the proposed Aeroassist Flight Experiment model in a high enthalpy free piston shock tunnel T3 at the Australian National University in Canberra, Australia. The enthalpy and Reynolds number range covered were 7.5 MJ/kg to 20 MJ/kg and 150,000 to 270,000 per meter respectively. The test Mach number varied between 7.5 and 8. Two test gases, air and nitrogen, were used and the model angle of attack varied from -10 deg to +10 deg to the free stream. The results are discussed and compared to the Mach 10 cold hypersonic air data as obtained in the Langley 31 inch Mach 10 Facility as well as the perfect gas CFD calculations of NASA LaRC.

Altitude test of several afterburner configurations on a turbofan engine with a hydrogen heater to simulate an elevated turbine discharge temperature

NASA Technical Reports Server (NTRS)

Johnsen, R. L.; Cullom, R. R.

1977-01-01

A performance test of several experimental afterburner configurations was conducted with a mixed-flow turbofan engine in an altitude facility. The simulated flight conditions were for Mach 1.4 at two altitudes, 12,190 and 14,630 meters. Turbine discharge temperatures of 889 and 1056 K were used. A production afterburner was tested for comparison. The research afterburners included partial forced mixers with V-gutter flameholders, a carburetted V-gutter flameholder, and a triple ring V-gutter flameholder with four swirl-can fuel mixers. Fuel injection variations were included. Performance data shown include augmented thrust ratio, thrust specific fuel consumption, combustion efficiency, and total pressure drop across the afterburner.

Trajectory Simulation of Meteors Assuming Mass Loss and Fragmentation

NASA Technical Reports Server (NTRS)

Allen, Gary A., Jr.; Prabhu, Dinesh K.; Saunders, David A

2015-01-01

Program used to simulate atmospheric flight trajectories of entry capsules [1] Includes models of atmospheres of different planetary destinations - Earth, Mars, Venus, Jupiter, Saturn, Uranus, Titan, ... Solves 3-­-degrees of freedom (3DoF) equations for a single body treated as a point mass. Also supports 6-DoF trajectory simula4on and Monte Carlo analyses. Uses Fehlberg-­-Runge-­-Kuna (4th-5th order) time integraion with automaic step size control. Includes rotating spheroidal planet with gravitational field having a J2 harmonic. Includes a variety of engineering aerodynamic and heat flux models. Capable of specifying events - heatshield jettison, parachute deployment, etc. - at predefined altitudes or Mach number. Has material thermal response models of typical aerospace materials integrated.

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

NASA Image and Video Library

2001-03-13

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

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.

Rocket-Model Investigation of the Longitudinal Stability, Drag, and Duct Performance Characteristics of the North American MX-770 (X-10) Missile at Mach Numbers from 0.80 to 1.70

NASA Technical Reports Server (NTRS)

Bond, Aleck C.; Swanson, Andrew G.

1953-01-01

A free-flight 0.12-scale rocket-boosted model of the North American MX-770 (X-10) missile has been tested in flight by the Pilotless Aircraft Research Division of the Langley Aeronautical Laboratory. Drag, longitudinal stability, and duct performance data were obtained at Mach numbers from 0.8 to 1.7 covering a Reynolds number range of about 9 x 10(exp 6) to 24 x 10(exp 6) based on wing mean aerodynamic chord. The lift-curve slope, static stability, and damping-in-pitch derivatives showed similar variations with Mach number, the parameters increasing from subsonic values in the transonic region and decreasing in the supersonic region. The variations were for the most part fairly smooth. The aerodynamic center of the configuration shifted rearward in the transonic region and moved forward gradually in the supersonic region. The pitching effectiveness of the canard control surfaces was maintained throughout the flight speed range, the supersonic values being somewhat greater than the subsonic. Trim values of angle of attack and lift coefficient changed abruptly in the transonic region, the change being associated with variations in the out-of-trim pitching moment, control effectiveness, and aerodynamic-center travel in this speed range. Duct total-pressure recovery decreased with increase in free-stream Mach number and the values were somewhat less than normal-shock recovery. Minimum drag data indicated a supersonic drag coefficient about twice the subsonic drag coefficient and a drag-rise Mach number of approximately 0.90. Base drag was small subsonically but was about 25 percent of the minimum drag of the configuration supersonically.

Design of a Mach-15 Total-Enthalpy Nozzle With Non-uniform Inflow Using Rotational MOC

NASA Technical Reports Server (NTRS)

Gaffney, Richard L., Jr.

2004-01-01

A new computer program to design nozzles with non-uniform inflow has been developed using the rotational method of characteristics (MOC). This program has been used to design a nozzle for the NASA's HYPULSE shock-expansion tunnel for use in scramjet engine tests at a Mach-15 flight-enthalpy condition. The nozzle has an area ratio of 9.5:1 that expands the inflow from Mach 6 along the centerline to Mach 8.7. Although the density and Mach number vary radially at the exit due to the non-uniformities of the inflow, the MOC procedure produces exit flow that is parallel and has uniform static pressure. The design has been verified with CFD which compares favorably with the MOC solution.

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

NASA Technical Reports Server (NTRS)

1997-01-01

This computational fluid dynamics (CFD) image shows the Hyper-X vehicle at a Mach 7 test condition with the engine operating. The solution includes both internal (scramjet engine) and external flow fields, including the interaction between the engine exhaust and vehicle aerodynamics. The image illustrates surface heat transfer on the vehicle surface (red is highest heating) and flowfield contours at local Mach number. The last contour illustrates the engine exhaust plume shape. This solution approach is one method of predicting the vehicle performance, and the best method for determination of vehicle structural, pressure and thermal design loads. The Hyper-X program is an ambitious series of experimental flights to expand the boundaries of high-speed aeronautics and develop new technologies for space access. When the first of three aircraft flies, it will be the first time a non-rocket engine has powered a vehicle in flight at hypersonic speeds--speeds above Mach 5, equivalent to about one mile per second or approximately 3,600 miles per hour at sea level. Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

Flight Wing Surface Pressure and Boundary-Layer Data Report from the F-111 Smooth Variable-Camber Supercritical Mission Adaptive Wing

NASA Technical Reports Server (NTRS)

Powers, Sheryll Goecke; Webb, Lannie D.

1997-01-01

Flight tests were conducted using the advanced fighter technology integration F-111 (AFTI/F-111) aircraft modified with a variable-sweep supercritical mission adaptive wing (MAW). The MAW leading- and trailing-edge variable-camber surfaces were deflected in flight to provide a near-ideal wing camber shape for the flight condition. The MAW features smooth, flexible upper surfaces and fully enclosed lower surfaces, which distinguishes it from conventional flaps that have discontinuous surfaces and exposed or semi-exposed mechanisms. Upper and lower surface wing pressure distributions were measured along four streamwise rows on the right wing for cruise, maneuvering, and landing configurations. Boundary-layer measurements were obtained near the trailing edge for one of the rows. Cruise and maneuvering wing leading-edge sweeps were 26 deg for Mach numbers less than 1 and 45 deg or 58 deg for Mach numbers greater than 1. The landing wing sweep was 9 deg or 16 deg. Mach numbers ranged from 0.27 to 1.41, angles of attack from 2 deg to 13 deg, and Reynolds number per unit foot from 1.4 x 10(exp 6) to 6.5 x 10(exp 6). Leading-edge cambers ranged from O deg to 20 deg down, and trailing-edge cambers ranged from 1 deg up to 19 deg down. Wing deflection data for a Mach number of 0.85 are shown for three cambers. Wing pressure and boundary-layer data are given. Selected data comparisons are shown. Measured wing coordinates are given for three streamwise semispan locations for cruise camber and one spanwise location for maneuver camber.

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

NASA Image and Video Library

2001-03-15

NASA's NB-52B carrier aircraft rolls down a taxiway at Edwards Air Force Base with the X-43A hypersonic research aircraft and its modified Pegasus® booster rocket slung from a pylon under its right wing. Part of a combined systems test conducted by NASA's Dryden Flight Research Center at Edwards, the taxi test was one of the last major milestones in the Hyper-X research program before the first X-43A flight. The X-43A flights will be the first actual flight tests of an aircraft powered by a revolutionary supersonic-combustion ramjet ("scramjet") engine capable of operating at hypersonic speeds (above Mach 5, or five times the speed of sound). The 12-foot, unpiloted research vehicle was developed and built by MicroCraft Inc., Tullahoma, Tenn., under NASA contract. The booster was built by Orbital Sciences Corp., Dulles, Va.,After being air-launched from NASA's venerable NB-52 mothership, the booster will accelerate the X-43A to test speed and altitude. The X-43A will then separate from the rocket and fly a pre-programmed trajectory, conducting aerodynamic and propulsion experiments until it descends into the Pacific Ocean. Three research flights are planned, two at Mach 7 and one at Mach 10, with the first tentatively scheduled for late spring to early summer, 2001.

X-1A in flight with flight data superimposed

NASA Image and Video Library

1953-12-12

This photo of the X-1A includes graphs of the flight data from Maj. Charles E. Yeager's Mach 2.44 flight on December 12, 1953. (This was only a few days short of the 50th anniversary of the Wright brothers' first powered flight.) After reaching Mach 2.44, then the highest speed ever reached by a piloted aircraft, the X-1A tumbled completely out of control. The motions were so violent that Yeager cracked the plastic canopy with his helmet. He finally recovered from a inverted spin and landed on Rogers Dry Lakebed. Among the data shown are Mach number and altitude (the two top graphs). The speed and altitude changes due to the tumble are visible as jagged lines. The third graph from the bottom shows the G-forces on the airplane. During the tumble, these twice reached 8 Gs or 8 times the normal pull of gravity at sea level. (At these G forces, a 200-pound human would, in effect, weigh 1,600 pounds if a scale were placed under him in the direction of the force vector.) Producing these graphs was a slow, difficult process. The raw data from on-board instrumentation recorded on oscillograph film. Human computers then reduced the data and recorded it on data sheets, correcting for such factors as temperature and instrument errors. They used adding machines or slide rules for their calculations, pocket calculators being 20 years in the future.

Orion Multi-Purpose Crew Vehicle (MPCV) Capsule Parachute Assembly System (CPAS) Wake Deficit Wind Tunnel Testing

NASA Technical Reports Server (NTRS)

Ross, James C.; Schuster, David M.

2014-01-01

During descent after re-entry into the Earth's atmosphere, the Orion CM deploys its drogue parachutes at approximately Mach 0.7. Accurately predicting the dynamic pressure experienced by the drogue parachutes at deployment is critical to properly designing the parachutes. This NASA Engineering and Safety Center assessment was designed to provide a complete set of flowfield measurements on and around an idealized Orion Crew Module shape with the most appropriate wind tunnel simulation of the Orion flight conditions prior to parachute deployment. This document contains the details of testing and the outcome of the assessment.

Effect of simulated in-flight thrust reversing on vertical-tail loads of F-18 and F-15 airplane models. [conducted in the Langley 16-foot transonic tunnel

NASA Technical Reports Server (NTRS)

Bare, E. A.; Berrier, B. L.; Capone, F. J.

1981-01-01

Investigations were conducted in the Langley 16-Foot Transonic Tunnel to provide data on a 0.10-scale model of the prototype F-18 airplane and a 0.047-scale model of the F-15 three-surface configuration (canard, wing, and horizontal tails). Test data were obtained at static conditions and at Mach numbers from 0.6 to 1.2 over an angle-of-attack range from 2 deg to 15 deg. Nozzle pressure ratio was varied from jet off to about 8.0.

NASA Glenn Propulsion Systems Lab (PSL) Icing Facility Update

NASA Technical Reports Server (NTRS)

Thomas, Queito P.

2015-01-01

The NASA Glenn Research Center Propulsion Systems Lab (PSL) was recently upgraded to perform engine inlet ice crystal testing in an altitude environment. The system installed 10 spray bars in the inlet plenum for ice crystal generation using 222 spray nozzles. As an altitude test chamber, PSL is capable of simulation of in-flight icing events in a ground test facility. The system was designed to operate at altitudes from 4,000 ft. to 40,000 ft. at Mach numbers up to 0.8M and inlet total temperatures from -60F to +15F.

Aerosciences, Aero-Propulsion and Flight Mechanics Technology Development for NASA's Next Generation Launch Technology Program

NASA Technical Reports Server (NTRS)

Cockrell, Charles E., Jr.

2003-01-01

The Next Generation Launch Technology (NGLT) program, Vehicle Systems Research and Technology (VSR&T) project is pursuing technology advancements in aerothermodynamics, aeropropulsion and flight mechanics to enable development of future reusable launch vehicle (RLV) systems. The current design trade space includes rocket-propelled, hypersonic airbreathing and hybrid systems in two-stage and single-stage configurations. Aerothermodynamics technologies include experimental and computational databases to evaluate stage separation of two-stage vehicles as well as computational and trajectory simulation tools for this problem. Additionally, advancements in high-fidelity computational tools and measurement techniques are being pursued along with the study of flow physics phenomena, such as boundary-layer transition. Aero-propulsion technology development includes scramjet flowpath development and integration, with a current emphasis on hypervelocity (Mach 10 and above) operation, as well as the study of aero-propulsive interactions and the impact on overall vehicle performance. Flight mechanics technology development is focused on advanced guidance, navigation and control (GN&C) algorithms and adaptive flight control systems for both rocket-propelled and airbreathing vehicles.

Comparison of shock structure solutions using independent continuum and kinetic theory approaches

NASA Technical Reports Server (NTRS)

Fiscko, Kurt A.; Chapman, Dean R.

1988-01-01

A vehicle traversing the atmosphere will experience flight regimes at high altitudes in which the thickness of a hypersonic shock wave is not small compared to the shock standoff distance from the hard body. When this occurs, it is essential to compute accurate flow field solutions within the shock structure. In this paper, one-dimensional shock structure is investigated for various monatomic gases from Mach 1.4 to Mach 35. Kinetic theory solutions are computed using the Direct Simulation Monte Carlo method. Steady-state solutions of the Navier-Stokes equations and of a slightly truncated form of the Burnett equations are determined by relaxation to a steady state of the time-dependent continuum equations. Monte Carlo results are in excellent agreement with published experimental data and are used as bases of comparison for continuum solutions. For a Maxwellian gas, the truncated Burnett equations are shown to produce far more accurate solutions of shock structure than the Navier-Stokes equations.

Progress in Payload Separation Risk Mitigation for a Deployable Venus Heat Shield

NASA Technical Reports Server (NTRS)

Smith, Brandon P.; Yount, Bryan C.; Venkatapathy, Ethiraj; Stern, Eric C.; Prabhu, Dinesh K.; Litton, Daniel K.

2013-01-01

A deployable decelerator known as the Adaptive Deployable Entry and Placement Technology (ADEPT) offers substantial science and mass savings for the Venus In Situ Explorer (VISE) mission. The lander and science payload must be separated from ADEPT during atmospheric entry. This paper presents a trade study of the separation system concept of operations and provides a conceptual design of the baseline: aft-separation with a subsonic parachute. Viability of the separation system depends on the vehicle's dynamic stability characteristics during deceleration from supersonic to subsonic speeds. A trajectory sensitivity study presented shows that pitch damping and Venusian winds drive stability prior to parachute deployment, while entry spin rate is not a driver of stability below Mach 5. Additionally, progress in free-flight CFD techniques capable of computing aerodynamic damping parameters is presented. Exploratory simulations of ADEPT at a constant speed of Mach number of 0.8 suggest the vehicle may have an oscillation limit cycle near 5 angle-of-attack. The proposed separation system conceptual design is thought to be viable.

Hypersonic shock tunnel heat transfer tests of the Space Shuttle SILTS pod configuration

NASA Technical Reports Server (NTRS)

Wittliff, C. E.

1983-01-01

Heat transfer measurements have been made on a 0.0175-scale NASA Space Shuttle orbiter model having a simulated SILTS (Shuttle Infrared Leeside Temperature Sensor) pod on top of the vertical tail. Heat transfer distributions were measured both on the pod and on the vertical tail. The test program covered Mach numbers of 8, 11 and 16 in air, at Reynolds numbers from 100,000 to 18 million, based on model length. The angle of attack ranged from 30 deg to 40 deg at sideslip angles from -2 to +2 deg. Data were obtained with 92 thin film assistance thermometers located on the SILTS pod and on the upper 30 percent of the vertical tail. Heat transfer rates measured on the vertical tail show good agreement with flight data obtained from missions STS-1, -2 and -3. The variation of heat transfer to the pod with Reynolds number, Mach number and angle of attack is discussed.

STS-1 operational flight profile. Volume 5: Descent cycle 3. Appendix D: GRTLS six degree of freedom Monte Carlo dispersion analysis

NASA Technical Reports Server (NTRS)

Montez, M. N.

1980-01-01

The results of a six degree of freedom (6-DOF) nonlinear Monte Carlo dispersion analysis for the latest glide return to landing site (GRTLS) abort trajectory for the Space Transportation System 1 Flight are presented. For this GRTLS, the number two main engine fails at 262.5 seconds ground elapsed time. Fifty randomly selected simulations, initialized at external tank separation, are analyzed. The initial covariance matrix is a 20 x 20 matrix and includes navigation errors and dispersions in position and velocity, time, accelerometer bias, and inertial platform misalinements. In all 50 samples, speedbrake, rudder, elevon, and body flap hinge moments are acceptable. Transitions to autoland begin before 9,000 feet and there are no tailscrapes. Navigation derived dynamic pressure accuracies exceed the flight control system constraints above Mach 2.5. Three out of 50 landings exceeded tire specification limit speed of 222 knots. Pilot manual landings are expected to reduce landing speed by landing farther downrange.

Pressure distributions obtained on a 0.10-scale model of the space shuttle Orbiter's forebody in the AEDC 16T propulsion wind tunnel

NASA Technical Reports Server (NTRS)

Siemers, P. M., III; Henry, M. W.

1986-01-01

Pressure distribution test data obtained on a 0.10-scale model of the forward fuselage of the Space Shuttle Orbiter are presented without analysis. The tests were completed in the AEDC 16T Propulsion Wind Tunnel. The 0.10-scale model was tested at angles of attack from -2 deg to 18 deg and angles of side slip from -6 to 6 deg at Mach numbers from 0.25 to 1/5 deg. The tests were conducted in support of the development of the Shuttle Entry Air Data System (SEADS). In addition to modeling the 20 SEADS orifices, the wind-tunnel model was also instrumented with orifices to match Development Flight Instrumentation (DFI) port locations that existed on the Space Shuttle Orbiter Columbia (OV-102) during the Orbiter Flight Test program. This DFI simulation has provided a means of comparisons between reentry flight pressure data and wind-tunnel and computational data.

Measurements of Aerodynamic Heat Transfer and Boundary-Layer Transition on a 15 deg. Cone in Free Flight at Supersonic Mach Numbers up to 5.2

NASA Technical Reports Server (NTRS)

Rumsey, Charles B.; Lee, Dorothy B.

1961-01-01

Measurements of aerodynamic heat transfer have been made at several stations on the 15 deg total-angle conical nose of a rocket-propelled model in free flight at Mach numbers up to 5.2. Data are presented for a range of local Mach number just outside the boundary layer from 1.40 to 4.65 and a range of local Reynolds number from 3.8 x 10(exp 6) to 46.5 x 10(exp 6), based on length from the nose tip to a measurement station. Laminar, transitional, and turbulent heat-transfer coefficients were measured. The laminar data were in agreement with laminar theory for cones, and the turbulent data agreed well with turbulent theory for cones using Reynolds number based on length from the nose tip. At a nearly constant ratio of wall to local static temperature of 1.2 the Reynolds number of transition increased from 14 x 10(exp 6) to 30 x 10(exp 6) as Mach number increased from 1.4 to 2.9 and then decreased to 17 x 10(exp 6) as Mach number increased to 3.7. At Mach numbers near 3.5, transition Reynolds numbers appeared to be independent of skin temperature at skin temperatures very cold with respect to adiabatic wall temperature. The transition Reynolds number was 17.7 x 10(exp 6) at a condition of Mach number and ratio of wall to local static temperature near that for which three-dimensional disturbance theory has been evaluated and has predicted laminar boundary-layer stability to very high Reynolds numbers (approximately 10(exp 12)).

NASA Test Flights Examine Effect of Atmospheric Turbulence on Sonic Booms

NASA Image and Video Library

2016-07-20

NASA pilot Nils Larson, and flight test engineer and pilot Wayne Ringelberg, head for a mission debrief after flying a NASA F/A-18 at Mach 1.38 to create sonic booms as part of the SonicBAT flight series at NASA’s Armstrong Flight Research Center in California, to study sonic boom signatures with and without the element of atmospheric turbulence.

Hyper-X Engine Design and Ground Test Program

NASA Technical Reports Server (NTRS)

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

1998-01-01

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

Acoustic Radiation from a Mach 14 Turbulent Boundary layer

NASA Astrophysics Data System (ADS)

Duan, Lian; Choudhari, Meelan

2014-11-01

Direct numerical simulations (DNS) are used to examine the pressure fluctuations generated by a high-speed turbulent boundary layer with a nominal freestream Mach number of 14 and wall temperature of 0.15 times the recovery temperature. The emphasis is on characterizing the acoustic radiation from the turbulent boundary layer and comparing it with previous simulations at Mach 2.5 and Mach 6 to assess the Mach-number dependence of the freestream pressure fluctuations. In particular, the numerical database is used to provide insights into the pressure disturbance spectrum and amplitude scaling with respect to the freestream Mach number as well as to understand the acoustic source mechanisms at very high Mach numbers. Such information is important for characterizing the freestream disturbance environment in conventional (i.e., noisy) hypersonic wind tunnels. Spectral characteristics of pressure fluctuations at the surface are also investigated. Supported by AFOSR and NASA Langley Research Center.

Instability of Poiseuille flow at extreme Mach numbers: linear analysis and simulations.

PubMed

Xie, Zhimin; Girimaji, Sharath S

2014-04-01

We develop the perturbation equations to describe instability evolution in Poiseuille flow at the limit of very high Mach numbers. At this limit the equation governing the flow is the pressure-released Navier-Stokes equation. The ensuing semianalytical solution is compared against simulations performed using the gas-kinetic method (GKM), resulting in excellent agreement. A similar comparison between analytical and computational results of small perturbation growth is performed at the incompressible (zero Mach number) limit, again leading to excellent agreement. The study accomplishes two important goals: it (i) contrasts the small perturbation evolution in Poiseuille flows at extreme Mach numbers and (ii) provides important verification of the GKM simulation scheme.

A wind tunnel investigation of circular and straked cylinders in transonic cross flow

NASA Technical Reports Server (NTRS)

Macha, J.

1976-01-01

Pressure distributions around circular and circular/strake cylinders were measured in a wind tunnel at Mach numbers from 0.6 to 1.2 with Reynolds number independently variable from 10,000 to 100,000. The local pressures are integrated over the cylinder surface to determine the variation of drag coefficient with both Mach number and Reynolds number. Effects of tunnel blockage are evaluated by comparing results from circular cylinders of various diameters at common Mach and Reynolds number conditions. Compressibility effects are concluded to be responsible for a flight reduction of the drag coefficient near Mach 0.7. Drag increases with strake height, presumably approaching a maximum drag corresponding to a flat plate configuration.

Longitudinal Stability Characteristics of the Consolidated Vultee XFY-1 Airplane with Windmilling Propellers as Obtained from Flight of 0.133-Scale Rocket-Propelled Model at Mach Numbers from 0.70 to 1.13

NASA Technical Reports Server (NTRS)

Hastings, Earl C.; Mitcham, Grady L.

1954-01-01

A flight test has been conducted to determine the longitudinal stability and control characteristics of a 0.133-scale model of the Consolidated Vultee XFY-1 airplane with windmilling propellers for the Mach number range between 0.70 and 1.13. The variation of lift-curve slope C(sub L(sub alpha) with Mach number was gradual with a maximum value of 0.074 occurring at a Mach number of 0.97. Propellers had little effect upon the values of lift-curve slope or the linearity of lift coefficient with angle of attack. At lift coefficients between approximately 0.25 and 0.45 with an elevon angle of approximately -l0 deg, there was a region of neutral longitudinal stability at Mach numbers below 0.93 introduced by the addition of windmilling propellers. Below a lift coefficient of 0.10 and above a lift coefficient of 0.45, the model was longitudinally stable throughout the Mach number range of the test. There was a forward shift in the aerodynamic center of about 3-percent mean aerodynamic chord introduced by the addition of propellers. The aerodynamic center as determined at low lift moved gradually from a value of 28.5-percent mean aerodynamic chord at a Mach number of 0.75 to a value of 47-percent mean aerodynamic chord at a Mach number of 1.10. There was an abrupt decrease in pitch damping between Mach numbers of 0.88 and 0.99 followed by a rapid increase in damping to a Mach number of 1.06. The propellers had little effect upon the pitch damping characteristics . The transonic trim change was a large pitching-down tendency with and without windmilling propellers. The elevons were effective pitch controls throughout the speed range; however, their effectiveness was reduced about 50 percent at supersonic speeds. The propellers had no appreciable effect upon the control effectiveness.

A comparative study of scramjet injection strategies for high Mach numbers flows

NASA Technical Reports Server (NTRS)

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

1992-01-01

A simple method for predicting the axial distribution of supersonic combustor thrust potential is described. A complementary technique for illustrating the spatial evolution and distribution of thrust potential and loss mechanisms in reacting flows is developed. Wall jet cases and swept ramp injector cases for Mach 17 and Mach 13.5 flight enthalpy inflow conditions are numerically modeled and analyzed using these techniques. The visualization of thrust potential in the combustor for the various cases examined provides a unique tool for increasing understanding of supersonic combustor performance potential.

Operating capability and current status of the reactivated NASA Lewis Research Center Hypersonic Tunnel Facility

NASA Technical Reports Server (NTRS)

Thomas, Scott R.; Trefny, Charles J.; Pack, William D.

1995-01-01

The NASA Lewis Research Center's Hypersonic Tunnel Facility (HTF) is a free-jet, blowdown propulsion test facility that can simulate up to Mach-7 flight conditions with true air composition. Mach-5, -6, and -7 nozzles, each with a 42 inch exit diameter, are available. Previously obtained calibration data indicate that the test flow uniformity of the HTF is good. The facility, without modifications, can accommodate models approximately 10 feet long. The test gas is heated using a graphite core induction heater that generates a nonvitiated flow. The combination of clean-air, large-scale, and Mach-7 capabilities is unique to the HTF and enables an accurate propulsion performance determination. The reactivation of the HTF, in progress since 1990, includes refurbishing the graphite heater, the steam generation plant, the gaseous oxygen system, and all control systems. All systems were checked out and recertified, and environmental systems were upgraded to meet current standards. The data systems were also upgraded to current standards and a communication link with NASA-wide computers was added. In May 1994, the reactivation was complete, and an integrated systems test was conducted to verify facility operability. This paper describes the reactivation, the facility status, the operating capabilities, and specific applications of the HTF.

X-43A departs NASA Dryden Flight Research Center for first free-flight attempt.

NASA Technical Reports Server (NTRS)

2001-01-01

The first X-43A hypersonic research aircraft and its modified Pegasus booster rocket were carried aloft by NASA's NB-52B carrier aircraft from Dryden Flight Research Center at Edwards Air Force Base, Calif., on June 2, 2001 for the first of three high-speed free flight attempts. About an hour and 15 minutes later the Pegasus booster was released from the B-52 to accelerate the X-43A to its intended speed of Mach 7. Before this could be achieved, the combined Pegasus and X-43A 'stack' lost control about eight seconds after ignition of the Pegasus rocket motor. The mission was terminated and explosive charges ensured the Pegasus and X-43A fell into the Pacific Ocean in a cleared Navy range area. A NASA investigation board is being assembled to determine the cause of the incident. Work continues on two other X-43A vehicles, the first of which could fly by late 2001. Central to the X-43A program is its integration of an air-breathing 'scramjet' engine that could enable a variety of high-speed aerospace craft, and promote cost-effective access to space. The 12-foot, unpiloted research vehicle was developed and built for NASA by MicroCraft Inc., Tullahoma, Tenn. The booster was built by Orbital Sciences Corp. at Chandler, Ariz. The X-43A flights are the first actual flight tests of an aircraft powered by a scramjet engine capable of operating at hypersonic speeds (above Mach 5, or five times the speed of sound). Some 90 minutes after takeoff, the Pegasus will launch from a B-52, rocketing the X-43A to Mach 7 at 95,000 feet altitude, or Mach 10 at 105,000 feet altitude. The X-43A will be powered by its revolutionary air-breathing supersonic-combustion ramjet or 'scramjet' engine. The X-43A will then fly a pre-programmed trajectory, conducting aerodynamic and propulsion experiments as it descends until it splashes into the Pacific Ocean.

Altitude Performance Characteristics of Turbojet-engine Tail-pipe Burner with Variable-area Exhaust Nozzle Using Several Fuel Systems and Flame Holders

NASA Technical Reports Server (NTRS)

Johnson, Lavern A; Meyer, Carl L

1950-01-01

A tail-pipe burner with a variable-area exhaust nozzle was investigated. From five configurations a fuel-distribution system and a flame holder were selected. The best configuration was investigated over a range of altitudes and flight Mach numbers. For the best configuration, an increase in altitude lowered the augmented thrust ratio, exhaust-gas total temperature, and tail-pipe combustion efficiency, and raised the specific fuel consumption. An increase in flight Mach number raised the augmented thrust ratio but had no apparent effect on exhaust-gas total temperature, tail-pipe combustion efficiency, or specific fuel consumption.

Reuse fo a Cold War Surveillance Drone to Flight Test a NASA Rocket Based Combined Cycle Engine

NASA Technical Reports Server (NTRS)

Brown, T. M.; Smith, Norm

1999-01-01

Plans for and early feasibility investigations into the modification of a Lockheed D21B drone to flight test the DRACO Rocket Based Combined Cycle (RBCC) engine are discussed. Modifications include the addition of oxidizer tanks, modern avionics systems, actuators, and a vehicle recovery system. Current study results indicate that the D21B is a suitable candidate for this application and will allow demonstrations of all DRACO engine operating modes at Mach numbers between 0.8 and 4.0. Higher Mach numbers may be achieved with more extensive modification. Possible project risks include low speed stability and control, and recovery techniques.

Experimental and predicted pressure and heating distributions for an Aeroassist Flight Experiment vehicle in air at Mach 10

NASA Technical Reports Server (NTRS)

Micol, John R.

1989-01-01

The Aeroassisted Flight Experiment vehicle for whose scale model pressure and heat-transfer rate distributions have been measured in air at Mach 10 is a 60-deg elliptic cone, raked off at a 73-percent angle, with an ellipsoid nose and a skirt added to the base of the rake plane to reduce heating. The predictions of both an inviscid flow-field code and a Navier-Stokes solver are compared with measured values. Good agreement is obtained in the case of pressure distributions; the effect of Reynolds number on heat-transfer distributions is noted to be small.

X-15A-2 is rolled out of the paint shop after having the full scale ablative applied

NASA Image and Video Library

1967-06-23

X-15A-2 is rolled out of the paint shop after having the full scale ablative applied. In June 1967, the X-15A-2 rocket-powered research aircraft received a full-scale ablative coating to protect the craft from the high temperatures associated with hypersonic flight (above Mach 5). This pink eraser-like substance, applied to the X-15A-2 aircraft (56-6671), was then covered with a white sealant coat before flight. This coating would help the #2 aircraft reach the record speed of 4,520 mph (Mach 6.7).

In-flight measurement of propeller noise on the fuselage of an airplane

NASA Technical Reports Server (NTRS)

Pla, Frederic G.; Ranaudo, Richard; Woodward, Richard P.

1989-01-01

In-flight measurements of propeller noise on the fuselage of an OV-10A aircraft were obtained using a horizontal and a vertical microphone array. A wide range of flight conditions were tested including changes in angle of attack, sideslip angle, power coefficient, helical tip Mach number and advance ratio, and propeller direction of rotation. Results show a dependence of the level and directivity of the tones on the angle of attack and on the sideslip angle with the propeller direction of rotation, which is similar to results obtained in wind tunnel tests with advanced propeller designs. The level of the tones at each microphone increases with increasing angle of attack for inboard-down propeller rotation and decreases for inboard-up rotation. The level also increases with increasing slideslip angle for both propeller directions of rotation. Increasing the power coefficient results in a slight increase in the level of the tones. A strong shock wave is generated by the propeller blades even at relatively low helical tip Mach numbers resulting in high harmonic levels. As the helical tip Mach number and the advance ratio are increased, the level of the higher harmonics increases much faster than the level of the blade passage frequency.

Preliminary Results of Stability and Control Investigation of the Bell X-5 Research Airplane

NASA Technical Reports Server (NTRS)

Finch, Thomas W; Briggs, Donald W

1953-01-01

During the acceptance tests of the Bell X-5 airplane, measurements of the static stability and control characteristics and horizontal-tail loads were obtained by the NACA High-Speed Flight Research Station. The results of the stability and control measurements are presented in this paper. A change in sweep angle between 20 deg and 59 deg had a minor effect on the longitudinal trim, with a maximum change of about 2.5 deg in elevator deflection being required at a Mach number near 0.85; however, sweeping the wings produced a total stick-force change of about 40 pounds. At low Mach numbers there was a rapid increase in stability at high normal-force coefficients for both 20 0 and 1100 sweepback, whereas a condition of neutral stability existed for 58 0 sweepback at high normal-force coefficients. At Mach numbers near 0.8 there was an instability at normal-force coefficients above 0.5 for all sweep angles tested. In the low normal-force-coefficient range a high degree of stability resulted in high stick forces which limited the maximum load factors attainable in the demonstration flights to values under 5g for all sweep angles at a Mach number near 0.8 and an altitude of 12,000 feet. The aileron effectiveness at 200 sweepback was found to be low over the Mach number range tested.

Shock Generation and Control Using DBD Plasma Actuators

NASA Technical Reports Server (NTRS)

Patel, Mehul P.; Cain, Alan B.; Nelson, Christopher C.; Corke, Thomas C.; Matlis, Eric H.

2012-01-01

This report is the final report of a NASA Phase I SBIR contract, with some revisions to remove company proprietary data. The Shock Boundary Layer Interaction (SBLI) phenomena in a supersonic inlet involve mutual interaction of oblique shocks with boundary layers, forcing the boundary layer to separate from the inlet wall. To improve the inlet efficiency, it is desired to prevent or delay shock-induced boundary layer separation. In this effort, Innovative Technology Applications Company (ITAC), LLC and the University of Notre Dame (UND) jointly investigated the use of dielectric-barrier-discharge (DBD) plasma actuators for control of SBLI in a supersonic inlet. The research investigated the potential for DBD plasma actuators to suppress flow separation caused by a shock in a turbulent boundary layer. The research involved both numerical and experimental investigations of plasma flow control for a few different SBLI configurations: (a) a 12 wedge flow test case at Mach 1.5 (numerical and experimental), (b) an impinging shock test case at Mach 1.5 using an airfoil as a shock generator (numerical and experimental), and (c) a Mach 2.0 nozzle flow case in a simulated 15 X 15 cm wind tunnel with a shock generator (numerical). Numerical studies were performed for all three test cases to examine the feasibility of plasma flow control concepts. These results were used to guide the wind tunnel experiments conducted on the Mach 1.5 12 degree wedge flow (case a) and the Mach 1.5 impinging shock test case (case b) which were at similar flow conditions as the corresponding numerical studies to obtain experimental evidence of plasma control effects for SBLI control. The experiments also generated data that were used in validating the numerical studies for the baseline cases (without plasma actuators). The experiments were conducted in a Mach 1.5 test section in the University of Notre Dame Hessert Laboratory. The simulation results from cases a and b indicated that multiple spanwise actuators in series and at a voltage of 75 kVp-p could fully suppress the flow separation downstream of the shock. The simulation results from case c showed that the streamwise plasma actuators are highly effective in creating pairs of counter-rotating vortices, much like the mechanical vortex generators, and could also potentially have beneficial effects for SBLI control. However, to achieve these effects, the positioning and the quantity of the DBD actuators used must be optimized. The wind tunnel experiments mapped the baseline flow with good agreement to the numerical simulations. The experimental results were conducted with spanwise actuators for cases a and b, but were limited by the inability to generate a sufficiently high voltage due to arcing in the wind-tunnel test-section. The static pressure in the tunnel was lower than the static pressure in an inlet at flight conditions, promoting arching and degrading the actuator performance.

Propagation of propeller tone noise through a fuselage boundary layer

NASA Technical Reports Server (NTRS)

Hanson, D. B.; Magliozzi, B.

1984-01-01

In earlier experimental and analytical studies, it was found that the boundary layer on an aircraft could provide significant shielding from propeller noise at typical transport airplane cruise Mach numbers. In this paper a new three-dimensional theory is described that treats the combined effects of refraction and scattering by the fuselage and boundary layer. The complete wave field is solved by matching analytical expressions for the incident and scattered waves in the outer flow to a numerical solution in the boundary layer flow. The model for the incident waves is a near-field frequency-domain propeller source theory developed previously for free field studies. Calculations for an advanced turboprop (Prop-Fan) model flight test at 0.8 Mach number show a much smaller than expected pressure amplification at the noise directivity peak, strong boundary layer shielding in the forward quadrant, and shadowing around the fuselage. Results are presented showing the difference between fuselage surface and free-space noise predictions as a function of frequency and Mach number. Comparison of calculated and measured effects obtained in a Prop-Fan model flight test show good agreement, particularly near and aft of the plane of rotation at high cruise Mach number.

Comparison of Engine Cycle Codes for Rocket-Based Combined Cycle Engines

NASA Technical Reports Server (NTRS)

Waltrup, Paul J.; Auslender, Aaron H.; Bradford, John E.; Carreiro, Louis R.; Gettinger, Christopher; Komar, D. R.; McDonald, J.; Snyder, Christopher A.

2002-01-01

This paper summarizes the results from a one day workshop on Rocket-Based Combined Cycle (RBCC) Engine Cycle Codes held in Monterey CA in November of 2000 at the 2000 JANNAF JPM with the authors as primary participants. The objectives of the workshop were to discuss and compare the merits of existing Rocket-Based Combined Cycle (RBCC) engine cycle codes being used by government and industry to predict RBCC engine performance and interpret experimental results. These merits included physical and chemical modeling, accuracy and user friendliness. The ultimate purpose of the workshop was to identify the best codes for analyzing RBCC engines and to document any potential shortcomings, not to demonstrate the merits or deficiencies of any particular engine design. Five cases representative of the operating regimes of typical RBCC engines were used as the basis of these comparisons. These included Mach 0 sea level static and Mach 1.0 and Mach 2.5 Air-Augmented-Rocket (AAR), Mach 4 subsonic combustion ramjet or dual-mode scramjet, and Mach 8 scramjet operating modes. Specification of a generic RBCC engine geometry and concomitant component operating efficiencies, bypass ratios, fuel/oxidizer/air equivalence ratios and flight dynamic pressures were provided. The engine included an air inlet, isolator duct, axial rocket motor/injector, axial wall fuel injectors, diverging combustor, and exit nozzle. Gaseous hydrogen was used as the fuel with the rocket portion of the system using a gaseous H2/O2 propellant system to avoid cryogenic issues. The results of the workshop, even after post-workshop adjudication of differences, were surprising. They showed that the codes predicted essentially the same performance at the Mach 0 and I conditions, but progressively diverged from a common value (for example, for fuel specific impulse, Isp) as the flight Mach number increased, with the largest differences at Mach 8. The example cases and results are compared and discussed in this paper.

Study of heat sink thermal protection systems for hypersonic research aircraft

NASA Technical Reports Server (NTRS)

Vahl, W. A.; Edwards, C. L. W.

1978-01-01

The feasibility of using a single metallic heat sink thermal protection system (TPS) over a projected flight test program for a hypersonic research vehicle was studied using transient thermal analyses and mission performance calculations. Four materials, aluminum, titanium, Lockalloy, and beryllium, as well as several combinations, were evaluated. Influence of trajectory parameters were considered on TPS and mission performance for both the clean vehicle configuration as well as with an experimental scramjet mounted. From this study it was concluded that a metallic heat sink TPS can be effectively employed for a hypersonic research airplane flight envelope which includes dash missions in excess of Mach 8 and 60 seconds of cruise at Mach numbers greater than 6. For best heat sink TPS match over the flight envelope, Lockalloy and titanium appear to be the most promising candidates

Experience with high performance V/STOL fighter projects at MBB

NASA Technical Reports Server (NTRS)

Aulehla, F.; Kissel, G. K.

1981-01-01

Flight control systems and aerodynamic aspects of experimental V/STOL aircraft are discussed. The VJ 101 C featured tilting engines for increased thrust, reheat for takeoff, simple translation, triangular decentralization of the engines for thrust modulation, and moderate ground effects. Two experimental aircraft were built, with and without reheat, capable of Mach 2 and Mach 1.04, respectively. The mechanical flight control system and tests are outlined, both for hover rig and flight configurations. Ground suction, acoustic and thermal loading, sodium silicate coatings to avoid ground corrosion, and recirculation are considered. Results of the follow-on project to the VJ 101 C, the AVS, which was developed by NASA, are reviewed, and it is noted that trends toward thrust-to-weight ratios exceeding one, in concert with low wing loading, favor the development of V/STOL aircraft.

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.

A modified Pegasus rocket ignites moments after release from the B-52B, beginning the acceleration of the X-43A over the Pacific Ocean on March 27, 2004

NASA Image and Video Library

2004-03-27

The second X-43A hypersonic research aircraft and its modified Pegasus booster rocket accelerate after launch from NASA's B-52B launch aircraft over the Pacific Ocean on March 27, 2004. The mission originated from the NASA Dryden Flight Research Center at Edwards Air Force Base, Calif. Minutes later the X-43A separated from the Pegasus booster and accelerated to its intended speed of Mach 7. In a combined research effort involving Dryden, Langley, and several industry partners, NASA demonstrated the value of its X-43A hypersonic research aircraft, as it became the first air-breathing, unpiloted, scramjet-powered plane to fly freely by itself. The March 27 flight, originating from NASA's Dryden Flight Research Center, began with the Agency's B-52B launch aircraft carrying the X-43A out to the test range over the Pacific Ocean off the California coast. The X-43A was boosted up to its test altitude of about 95,000 feet, where it separated from its modified Pegasus booster and flew freely under its own power. Two very significant aviation milestones occurred during this test flight: first, controlled accelerating flight at Mach 7 under scramjet power, and second, the successful stage separation at high dynamic pressure of two non-axisymmetric vehicles. To top it all off, the flight resulted in the setting of a new aeronautical speed record. The X-43A reached a speed of over Mach 7, or about 5,000 miles per hour faster than any known aircraft powered by an air-breathing engine has ever flown.

NASA's B-52B launch aircraft cruises to a test range over the Pacific Ocean carrying the second X-43A vehicle attached to a Pegasus rocket on March 27, 2004

NASA Image and Video Library

2004-03-27

The second X-43A hypersonic research aircraft, attached to a modified Pegasus booster rocket and followed by a chase F-18, was taken to launch altitude by NASA's B-52B launch aircraft from the NASA Dryden Flight Research Center at Edwards Air Force Base, Calif., on March 27, 2004. About an hour later the Pegasus booster was released from the B-52 to accelerate the X-43A to its intended speed of Mach 7. In a combined research effort involving Dryden, Langley, and several industry partners, NASA demonstrated the value of its X-43A hypersonic research aircraft, as it became the first air-breathing, unpiloted, scramjet-powered plane to fly freely by itself. The March 27 flight, originating from NASA's Dryden Flight Research Center, began with the Agency's B-52B launch aircraft carrying the X-43A out to the test range over the Pacific Ocean off the California coast. The X-43A was boosted up to its test altitude of about 95,000 feet, where it separated from its modified Pegasus booster and flew freely under its own power. Two very significant aviation milestones occurred during this test flight: first, controlled accelerating flight at Mach 7 under scramjet power, and second, the successful stage separation at high dynamic pressure of two non-axisymmetric vehicles. To top it all off, the flight resulted in the setting of a new aeronautical speed record. The X-43A reached a speed of over Mach 7, or about 5,000 miles per hour faster than any known aircraft powered by an air-breathing engine has ever flown.

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.

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.

Development of the Orion Crew Module Static Aerodynamic Database. Part 1; Hypersonic

NASA Technical Reports Server (NTRS)

Bibb, Karen L.; Walker, Eric L.; Robinson, Philip E.

2011-01-01

The Orion aerodynamic database provides force and moment coefficients given the velocity, attitude, configuration, etc. of the Crew Exploration Vehicle (CEV). The database is developed and maintained by the NASA CEV Aerosciences Project team from computational and experimental aerodynamic simulations. The database is used primarily by the Guidance, Navigation, and Control (GNC) team to design vehicle trajectories and assess flight performance. The initial hypersonic re-entry portion of the Crew Module (CM) database was developed in 2006. Updates incorporating additional data and improvements to the database formulation and uncertainty methodologies have been made since then. This paper details the process used to develop the CM database, including nominal values and uncertainties, for Mach numbers greater than 8 and angles of attack between 140deg and 180deg. The primary available data are more than 1000 viscous, reacting gas chemistry computational simulations using both the Laura and Dplr codes, over a range of Mach numbers from 2 to 37 and a range of angles of attack from 147deg to 172deg. Uncertainties were based on grid convergence, laminar-turbulent solution variations, combined altitude and code-to-code variations, and expected heatshield asymmetry. A radial basis function response surface tool, NEAR-RS, was used to fit the coefficient data smoothly in a velocity-angle-of-attack space. The resulting database is presented and includes some data comparisons and a discussion of the predicted variation of trim angle of attack and lift-to-drag ratio. The database provides a variation in trim angle of attack on the order of +/-2deg, and a range in lift-to-drag ratio of +/-0.035 for typical vehicle flight conditions.

Flight and wind-tunnel calibrations of a flush airdata sensor at high angles of attack and sideslip and at supersonic Mach numbers

NASA Technical Reports Server (NTRS)

Moes, Timothy R.; Whitmore, Stephen A.; Jordan, Frank L., Jr.

1993-01-01

A nonintrusive airdata-sensing system was calibrated in flight and wind-tunnel experiments to an angle of attack of 70 deg and to angles of sideslip of +/- 15 deg. Flight-calibration data have also been obtained to Mach 1.2. The sensor, known as the flush airdata sensor, was installed on the nosecap of an F-18 aircraft for flight tests and on a full-scale F-18 forebody for wind-tunnel tests. Flight tests occurred at the NASA Dryden Flight Research Facility, Edwards, California, using the F-18 High Alpha Research Vehicle. Wind-tunnel tests were conducted in the 30- by 60-ft wind tunnel at the NASA LaRC, Hampton, Virginia. The sensor consisted of 23 flush-mounted pressure ports arranged in concentric circles and located within 1.75 in. of the tip of the nosecap. An overdetermined mathematical model was used to relate the pressure measurements to the local airdata quantities. The mathematical model was based on potential flow over a sphere and was empirically adjusted based on flight and wind-tunnel data. For quasi-steady maneuvering, the mathematical model worked well throughout the subsonic, transonic, and low supersonic flight regimes. The model also worked well throughout the angle-of-attack and sideslip regions studied.

Flight and wind-tunnel calibrations of a flush airdata sensor at high angles of attack and sideslip and at supersonic Mach numbers

NASA Technical Reports Server (NTRS)

Moes, Timothy R.; Whitmore, Stephen A.; Jordan, Frank L., Jr.

1993-01-01

A nonintrusive airdata-sensing system was calibrated in flight and wind-tunnel experiments to an angle of attack of 70 deg and to angles of sideslip of +/- 15 deg. Flight-calibration data have also been obtained to Mach 1.2. The sensor, known as the flush airdata sensor, was installed on the nosecap of an F-18 aircraft for flight tests and on a full-scale F-18 forebody for wind-tunnel tests. Flight tests occurred at the NASA Dryden Flight Research Facility, Edwards, California, using the F-18 High Alpha Research Vehicle. Wind-tunnel tests were conducted in the 30- by 60-ft wind tunnel at the NASA LaRC, Hampton, Virginia. The sensor consisted of 23 flush-mounted pressure ports arranged in concentric circles and located within 1.75 in. of the tip of the nosecap. An overdetermined mathematical model was used to relate the pressure measurements to the local airdata quantities. The mathematical model was based on potential flow over a sphere and was empirically adjusted based on flight and wind-tunnel data. For quasi-steady maneuvering, the mathematical model worked well throughout the subsonic, transonic, and low supersonic flight regimes. The model also worked well throughout the angles-of-attack and -sideslip regions studied.

SHEFEX II Flight Instrumentation And Preparation Of Post Flight Analysis

NASA Astrophysics Data System (ADS)

Thiele, Thomas; Siebe, Frank; Gulhan, Ali

2011-05-01

A main disadvantage of modern TPS systems for re- entry vehicles is the expensive manufacturing and maintenance process due to the complex geometry of these blunt nose configurations. To reduce the costs and to improve the aerodynamic performance the German Aerospace Center (DLR) is following a different approach using TPS structures consisting of flat ceramic tiles. To test these new sharp edged TPS structures the SHEFEX I flight experiment was designed and successfully performed by DLR in 2005. To further improve the reliability of the sharp edged TPS design at even higher Mach numbers, a second flight experiment SHEFEX II will be performed in September 2011. In comparison to SHEFEX I the second flight experiment has a fully symmetrical shape and will reach a maximum Mach number of about 11. Furthermore the vehicle has an active steering system using four canards to control the flight attitude during re-entry, e.g. roll angle, angle of attack and sideslip. After a successful flight the evaluation of the flight data will be performed using a combination of numerical and experimental tools. The data will be used for the improvement of the present numerical analysis tools and to get a better understanding of the aerothermal behaviour of sharp TPS structures. This paper presents the flight instrumentation of the SHEFEX II TPS. In addition the concept of the post flight analysis is presented.

Aerodynamic Database Development for the Hyper-X Airframe Integrated Scramjet Propulsion Experiments

NASA Technical Reports Server (NTRS)

Engelund, Walter C.; Holland, Scott D.; Cockrell, Charles E., Jr.; Bittner, Robert D.

2000-01-01

This paper provides an overview of the activities associated with the aerodynamic database which is being developed in support of NASA's Hyper-X scramjet flight experiments. Three flight tests are planned as part of the Hyper-X program. Each will utilize a small, nonrecoverable research vehicle with an airframe integrated scramjet propulsion engine. The research vehicles will be individually rocket boosted to the scramjet engine test points at Mach 7 and Mach 10. The research vehicles will then separate from the first stage booster vehicle and the scramjet engine test will be conducted prior to the terminal decent phase of the flight. An overview is provided of the activities associated with the development of the Hyper-X aerodynamic database, including wind tunnel test activities and parallel CFD analysis efforts for all phases of the Hyper-X flight tests. A brief summary of the Hyper-X research vehicle aerodynamic characteristics is provided, including the direct and indirect effects of the airframe integrated scramjet propulsion system operation on the basic airframe stability and control characteristics. Brief comments on the planned post flight data analysis efforts are also included.

Flight and wind-tunnel correlation of boundary-layer transition on the AEDC transition cone

NASA Technical Reports Server (NTRS)

Fisher, D. L.; Dougherty, N. S., Jr.

1982-01-01

Transition and fluctuating surface pressure data were acquired on a 10 deg included angle cone, using the same instrumentation and technique over a wide range of Mach and Reynolds numbers in 23 wind tunnels and in flight. Transition was detected with a traversing pitot-pressure probe in contact with the surface. The surface pressure fluctuations were measured with microphones set flush in the cone surface. Good correlation of end of transition Reynolds number RE(T) was obtained between data from the lower disturbance wind tunnels and flight up to a boundary layer edge Mach number, M(e) = 1.2. Above M(e) = 1.2, however, this correlation deteriorates, with the flight Re(T) being 25 to 30% higher than the wind tunnel Re(T) at M(e) = 1.6. The end of transition Reynolds number correlated within + or - 20% with the surface pressure fluctuations, according to the equation used. Broad peaks in the power spectral density distributions indicated that Tollmien-Schlichting waves were the probable cause of transition in flight and in some of the wind tunnels.

In-flight lift and drag measurements on a first generation jet transport equipped with winglets

NASA Technical Reports Server (NTRS)

Lux, D. P.

1982-01-01

A KC-135A aircraft equipped with wing tip winglets was flight tested to demonstrate and validate the potential performance gain of the winglet concept as predicted from analytical and wind tunnel data. Flight data were obtained at cruise conditions for Mach numbers of 0.70, 0.75, and 0.80 at a nominal altitude of 36,000 ft. and winglet configurations of 15 deg cant/-4 deg incidence, 0 deg cant/-4 deg incidence, and baseline. For the Mach numbers tested the data show that the addition of winglets did not affect the lifting characteristics of the wing. However, both winglet configurations showed a drag reduction over the baseline configuration, with the best winglet configuration being the 15 deg cant/-4 deg incidence configuration. This drag reduction due to winglets also increased with increasing lift coefficient. It was also shown that a small difference exists between the 15 deg cant/-4 deg incidence flight and wind tunnel predicted data. This difference was attributed to the pillowing of the winglet skins in flight which would decrease the winglet performance.

The dynamic-response characteristics of a 35 degree swept-wing airplane as determined from flight measurements

NASA Technical Reports Server (NTRS)

Triplett, William C; Brown, Stuart C; Smith, G Allan

1955-01-01

The longitudinal and lateral-directional dynamic-response characteristics of a 35 degree swept-wing fighter-type airplane determined from flight measurements are presented and compared with predictions based on theoretical studies and wind-tunnel data. Flights were made at an altitude of 35,000 feet covering the Mach number range of 0.50 to 1.04. A limited amount of lateral-directional data were also obtained at 10,000 feet. The flight consisted essentially of recording transient responses to pilot-applied pulsed motions of each of the three primary control surfaces. These transient data were converted into frequency-response form by means of the Fourier transformation and compared with predicted responses calculated from the basic equations. Experimentally determined transfer functions were used for the evaluation of the stability derivatives that have the greatest effect on the dynamic response of the airplane. The values of these derivatives, in most cases, agreed favorably with predictions over the Mach number range of the test.

HAWK MACH-III Intelligent Maintenance Tutor Design Development Report

DTIC Science & Technology

1986-12-01

objective can best be achieved by designing the MACH-IIl to provide augmented hands-on experience in troubleshooting in a setting which will emphasize...artificial intelligence supporting the development activity will focus on development of a strategy for effective and efficient hierarchical simulation of...main components of such a system are the system simulation and problem-solving expertise, the student model, and the tutorial strategies . In the MACH

A preliminary investigation of inlet unstart effects on a high-speed civil transport concept

NASA Technical Reports Server (NTRS)

Domack, Christopher S.

1991-01-01

Vehicle motions resulting from a supersonic mixed-compression inlet unstart were examined to determine if the unstart constituted a hazard severe enough to warrant rejection of mixed-compression inlets on high-speed civil transport (HSCT) concepts. A simple kinematic analysis of an inlet unstart during cruise was performed for a Mach 2, 4, 250-passenger HSCT concept using data from a wind-tunnel test of a representative configuration with unstarted inlets simulated. A survey of previously published research on inlet unstart effects, including simulation and flight test data for the YF-12, XB-70, and Concorde aircraft, was conducted to validate the calculated results. It was concluded that, when countered by suitable automatic propulsion and flight control systems, the vehicle dynamics induced by an inlet unstart are not severe enough to preclude the use of mixed-compression inlets on an HSCT from a passenger safety standpoint. The ability to provide suitable automatic controls appears to be within the current state of the art. However, the passenger startle and discomfort caused by the noise, vibration, and cabin motions associated with an inlet unstart remain a concern.

Design and Implementation of the Boundary Layer Transition Flight Experiment on Space Shuttle Discovery

NASA Technical Reports Server (NTRS)

Spanos, Theodoros A.; Micklos, Ann

2010-01-01

In an effort to better the understanding of high speed aerodynamics, a series of flight experiments were installed on Space Shuttle Discovery during the STS-119 and STS-128 missions. This experiment, known as the Boundary Layer Transition Flight Experiment (BLTFE), provided the technical community with actual entry flight data from a known height protuberance at Mach numbers at and above Mach 15. Any such data above Mach 15 is irreproducible in a laboratory setting. Years of effort have been invested in obtaining this valuable data, and many obstacles had to be overcome in order to ensure the success of implementing an Orbiter modification. Many Space Shuttle systems were involved in the installation of appropriate components that revealed 'concurrent engineering' was a key integration tool. This allowed the coordination of all various parts and pieces which had to be sequenced appropriately and installed at the right time. Several issues encountered include Orbiter configuration and access, design requirements versus current layout, implementing the modification versus typical processing timelines, and optimizing the engineering design cycles and changes. Open lines of communication within the entire modification team were essential to project success as the team was spread out across the United States, from NASA Kennedy Space Center in Florida, to NASA Johnson Space Center in Texas, to Boeing Huntington Beach, California among others. The forum permits the discussion of processing concerns from the design phase to the implementation phase, which eventually saw the successful flights and data acquisition on STS-119 in March 2009 and on STS-128 in September 2009.

Investigation of wing upper surface flow-field disturbance due to NASA DC-8-72 in-flight inboard thrust-reverser deployment

NASA Technical Reports Server (NTRS)

Hamid, Hedayat U.; Margason, Richard J.; Hardy, Gordon

1995-01-01

An investigation of the wing upper surface flow-field disturbance due to in-flight inboard thrust reverser deployment on the NASA DC-8-72, which was conducted cooperatively by NASA Ames, the Federal Aviation Administration (FAA), McDonnell Douglas, and the Aerospace Industry Association (AIA), is outlined and discussed in detail. The purpose of this flight test was to obtain tufted flow visualization data which demonstrates the effect of thrust reverser deployment on the wing upper surface flow field to determine if the disturbed flow regions could be modeled by computational methods. A total of six symmetric thrust reversals of the two inboard engines were performed to monitor tuft and flow cone patterns as well as the character of their movement at the nominal Mach numbers of 0.55, 0.70, and 0.85. The tufts and flow cones were photographed and video-taped to determine the type of flow field that occurs with and without the thrust reversers deployed. In addition, the normal NASA DC-8 onboard Data Acquisition Distribution System (DADS) was used to synchronize the cameras. Results of this flight test will be presented in two parts. First, three distinct flow patterns associated with the above Mach numbers were sketched from the motion videos and discussed in detail. Second, other relevant aircraft parameters, such as aircraft's angular orientation, altitude, Mach number, and vertical descent, are discussed. The flight test participants' comments were recorded on the videos and the interested reader is referred to the video supplement section of this report for that information.

Study of boundary-layer transition using transonic cone Preston tube data

NASA Technical Reports Server (NTRS)

Reed, T. D.; Abu-Mostafa, A.

1982-01-01

Laminar layer Preston tube data on a sharp nose, ten degree cone obtained in the Ames 11 ft TWT and in flight tests are analyzed. During analyses of the laminar-boundary layer data, errors were discovered in both the wind tunnel and the flight data. A correction procedure for errors in the flight data is recommended which forces the flight data to exhibit some of the orderly characteristics of the wind tunnel data. From corrected wind tunnel data, a correlation is developed between Preston tube pressures and the corresponding values of theoretical laminar skin friction. Because of the uncertainty in correcting the flight data, a correlation for the unmodified data is developed, and, in addition, three other correlations are developed based on different correction procedures. Each of these correlations are used in conjunction with the wind tunnel correlation to define effective freestream unit Reynolds numbers for the 11 ft TWT over a Mach number range of 0.30 to 0.95. The maximum effective Reynolds numbers are approximately 6.5% higher than the normal values. These maximum values occur between freestream Mach numbers of 0.60 and 0.80. Smaller values are found outside this Mach number range. These results indicate wind tunnel noise affects the average laminar skin friction much less than it affects boundary layer transition. Data on the onset, extent, and end of boundary layer transition are summarized. Application of a procedure for studying the relative effects of varying nose radius on a ten degree cone at supercritical speeds indicates that increasing nose radius promotes boundary layer transition and separation of laminar boundary layers.

Effects of reaction control system jet flow field interactions on the aerodynamic characteristics of a 0.010-scale space shuttle orbiter model in the Langley Research Center 31 inch CFHT (OA85)

NASA Technical Reports Server (NTRS)

Daileda, J. J.; Marroquin, J.

1974-01-01

An experimental investigation was conducted to obtain detailed effects on supersonic vehicle hypersonic aerodynamic and stability and control characteristics of reaction control system jet flow field interactions with the local vehicle flow field. A 0.010-scale model was used. Six-component force data and wing, elevon, and body flap surface pressure data were obtained through an angle-of-attack range of -10 to +35 degrees 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 and reaction control system plume simulation of flight dynamic pressures of 5, 10 and 20 PSF.

Transonic Aerodynamic Characteristics of a Model of a Proposed Six-Engine Hull-Type Seaplane Designed for Supersonic Flight

NASA Technical Reports Server (NTRS)

Wornom, Dewey E.

1960-01-01

Force tests of a model of a proposed six-engine hull-type seaplane were performed in the Langley 8-foot transonic pressure tunnel. The results of these tests have indicated that the model had a subsonic zero-lift drag coefficient of 0.0240 with the highest zero-lift drag coefficient slightly greater than twice the subsonic drag level. Pitchup tendencies were noted for subsonic Mach numbers at relatively high lift coefficients. Wing leading-edge droop increased the maximum lift-drag ratio approximately 8 percent at a Mach number of 0.80 but this effect was negligible at a Mach number of 0.90 and above. The configuration exhibited stable lateral characteristics over the test Mach number range.

STS 2000: Structural design of the airbreathing launcher

NASA Astrophysics Data System (ADS)

Boyeldieu, E.

This paper presents a description of the structural design and the choice of materials of the different parts of the Space Transportation System 2000 (STS 2000). This launcher is one of the different concepts studied by AEROSPATIALE to evaluate its feasibility and its performance. The STS 2000 Single-Stage-To-Orbit (SSTO) is a reusable single stage launcher using airbreathing propulsion till Mach 6. This SSTO takes off horizontally using an undercarriage It takes off with a speed of 150 m/s and with an incidence angle of 12 deg. The STS 2000 flights from Mach 0.4 to Mach 3.6 using four turbo-rockets engines, from Mach 3.6 to Mach 6 using four ramjets-rockets engines and from Mach 6 to Mach 25 using four rockets engines. During its reentry, it glides from orbit to earth and it horizontally lands at the same base (KOUROU in French Guiana). The initial take-off mass is 338 metric tons. The ascent phase specification are: a maximum axial acceleration of 4 g's and a maximum dynamic pressure of 70 kPa.

Measurements in Flight of the Pressure Distribution on the Right Wing of a Pursuit-Type Airplane at Several Values of Mach Number

NASA Technical Reports Server (NTRS)

Clousing, Lawrence A; Turner, William N; Rolls, L Stewart

1946-01-01

Pressure-distribution measurements were made on the right wing of a pursuit-type airplane at values of Mach number up to 0.80. The results showed that a considerable portion of the lift was carried by components of the airplane other than the wings, and that the proportion of lift carried by the wings may vary considerably with Mach number, thus changing the bending moment at the wing root whether or not there is a shift in the lateral position of the center of pressure. It was also shown that the center of pressure does not necessarily move outward at high Mach numbers, even though the wing-thickness ratio decreases toward the wing tip. The wing pitching-moment coefficient increased sharply in a negative direction at a Mach lift-curve slope increased with Mach number up to values of above the critical value. Pressures inside the wing were small and negative.

Finite-Element Analysis of a Mach-8 Flight Test Article Using Nonlinear Contact Elements

NASA Technical Reports Server (NTRS)

Richards, W. Lance

1997-01-01

A flight test article, called a glove, is required for a Mach-8 boundary-layer experiment to be conducted on a flight mission of the air-launched Pegasus(reg) space booster. The glove is required to provide a smooth, three-dimensional, structurally stable, aerodynamic surface and includes instrumentation to determine when and where boundary-layer transition occurs during the hypersonic flight trajectory. A restraint mechanism has been invented to attach the glove to the wing of the space booster. The restraint mechanism securely attaches the glove to the wing in directions normal to the wing/glove interface surface, but allows the glove to thermally expand and contract to alleviate stresses in directions parallel to the interface surface. A finite-element analysis has been performed using nonlinear contact elements to model the complex behavior of the sliding restraint mechanism. This paper provides an overview of the glove design and presents details of the analysis that were essential to demonstrate the flight worthiness of the wing-glove test article. Results show that all glove components are well within the allowable stress and deformation requirements to satisfy the objectives of the flight research experiment.

Ares I-X Test Flight Reference Trajectory Development

NASA Technical Reports Server (NTRS)

Starr, Brett R.; Gumbert, Clyde R.; Tartabini, Paul V.

2011-01-01

Ares I-X was the first test flight of NASA's Constellation Program's Ares I crew launch vehicle. Ares I is a two stage to orbit launch vehicle that provides crew access to low Earth orbit for NASA's future manned exploration missions. The Ares I first stage consists of a Shuttle solid rocket motor (SRM) modified to include an additional propellant segment and a liquid propellant upper stage with an Apollo J2X engine modified to increase its thrust capability. The modified propulsion systems were not available for the first test flight, thus the test had to be conducted with an existing Shuttle 4 segment reusable solid rocket motor (RSRM) and an inert Upper Stage. The test flight's primary objective was to demonstrate controllability of an Ares I vehicle during first stage boost and the ability to perform a successful separation. In order to demonstrate controllability, the Ares I-X ascent control algorithms had to maintain stable flight throughout a flight environment equivalent to Ares I. The goal of the test flight reference trajectory development was to design a boost trajectory using the existing RSRM that results in a flight environment equivalent to Ares I. A trajectory similarity metric was defined as the integrated difference between the Ares I and Ares I-X Mach versus dynamic pressure relationships. Optimization analyses were performed that minimized the metric by adjusting the inert upper stage weight and the ascent steering profile. The sensitivity of the optimal upper stage weight and steering profile to launch month was also investigated. A response surface approach was used to verify the optimization results. The analyses successfully defined monthly ascent trajectories that matched the Ares I reference trajectory dynamic pressure versus Mach number relationship to within 10% through Mach 3.5. The upper stage weight required to achieve the match was found to be feasible and varied less than 5% throughout the year. The paper will discuss the flight test requirements, provide Ares I-X vehicle background, discuss the optimization analyses used to meet the requirements, present analysis results, and compare the reference trajectory to the reconstructed flight trajectory.

Numerical exploration of mixing and combustion in ethylene fueled scramjet combustor

NASA Astrophysics Data System (ADS)

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

2015-12-01

Numerical simulations are performed for full scale scramjet combustor of a hypersonic airbreathing vehicle with ethylene fuel at ground test conditions corresponding to flight Mach number, altitude and stagnation enthalpy of 6.0, 30 km and 1.61 MJ/kg respectively. Three dimensional RANS equations are solved along with species transport equations and SST-kω turbulence model using Commercial CFD software CFX-11. Both nonreacting (with fuel injection) and reacting flow simulations [using a single step global reaction of ethylene-air with combined combustion model (CCM)] are carried out. The computational methodology is first validated against experimental results available in the literature and the performance parameters of full scale combustor in terms of thrust, combustion efficiency and total pressure loss are estimated from the simulation results. Parametric studies are conducted to study the effect of fuel equivalence ratio on the mixing and combustion behavior of the combustor.

Verification of a Constraint Force Equation Methodology for Modeling Multi-Body Stage Separation

NASA Technical Reports Server (NTRS)

Tartabini, Paul V.; Roithmayr, Carlos; Toniolo, Matthew D.; Karlgaard, Christopher; Pamadi, Bandu N.

2008-01-01

This paper discusses the verification of the Constraint Force Equation (CFE) methodology and its implementation in the Program to Optimize Simulated Trajectories II (POST2) for multibody separation problems using three specially designed test cases. The first test case involves two rigid bodies connected by a fixed joint; the second case involves two rigid bodies connected with a universal joint; and the third test case is that of Mach 7 separation of the Hyper-X vehicle. For the first two cases, the POST2/CFE solutions compared well with those obtained using industry standard benchmark codes, namely AUTOLEV and ADAMS. For the Hyper-X case, the POST2/CFE solutions were in reasonable agreement with the flight test data. The CFE implementation in POST2 facilitates the analysis and simulation of stage separation as an integral part of POST2 for seamless end-to-end simulations of launch vehicle trajectories.

Prospects for Nonlinear Laser Diagnostics in the Jet Noise Laboratory

NASA Technical Reports Server (NTRS)

Herring, Gregory C.; Hart, Roger C.; Fletcher, mark T.; Balla, R. Jeffrey; Henderson, Brenda S.

2007-01-01

Two experiments were conducted to test whether optical methods, which rely on laser beam coherence, would be viable for off-body flow measurement in high-density, compressible-flow wind tunnels. These tests measured the effects of large, unsteady density gradients on laser diagnostics like laser-induced thermal acoustics (LITA). The first test was performed in the Low Speed Aeroacoustics Wind Tunnel (LSAWT) of NASA Langley Research Center's Jet Noise Laboratory (JNL). This flow facility consists of a dual-stream jet engine simulator (with electric heat and propane burners) exhausting into a simulated flight stream, reaching Mach numbers up to 0.32. A laser beam transited the LSAWT flow field and was imaged with a high-speed gated camera to measure beam steering and transverse mode distortion. A second, independent test was performed on a smaller laboratory jet (Mach number < 1.2 and mass flow rate < 0.1 kg/sec). In this test, time-averaged LITA velocimetry and thermometry were performed at the jet exit plane, where the effect of unsteady density gradients is observed on the LITA signal. Both experiments show that LITA (and other diagnostics relying on beam overlap or coherence) faces significant hurdles in the high-density, compressible, and turbulent flow environments similar to those of the JNL.

Local flow measurements at the inlet spike tip of a Mach 3 supersonic cruise airplane

NASA Technical Reports Server (NTRS)

Johnson, H. J.; Montoya, E. J.

1973-01-01

The flow field at the left inlet spike tip of a YF-12A airplane was examined using at 26 deg included angle conical flow sensor to obtain measurements at free-stream Mach numbers from 1.6 to 3.0. Local flow angularity, Mach number, impact pressure, and mass flow were determined and compared with free-stream values. Local flow changes occurred at the same time as free-stream changes. The local flow usually approached the spike centerline from the upper outboard side because of spike cant and toe-in. Free-stream Mach number influenced the local flow angularity; as Mach number increased above 2.2, local angle of attack increased and local sideslip angle decreased. Local Mach number was generally 3 percent less than free-stream Mach number. Impact-pressure ratio and mass flow ratio increased as free-stream Mach number increased above 2.2, indicating a beneficial forebody compression effect. No degradation of the spike tip instrumentation was observed after more than 40 flights in the high-speed thermal environment encountered by the airplane. The sensor is rugged, simple, and sensitive to small flow changes. It can provide accurate imputs necessary to control an inlet.

Hydrocarbon-Fueled Scramjet Research at Hypersonic Mach Numbers

DTIC Science & Technology

2005-03-31

oxide O atomic oxygen 02 molecular oxygen OH hydroxyl radical ppm parts per million PD photodiode PLLF planar laser-induced fluorescence PMT...photomultiplier tube RAM random access memory RANS Reynolds-averaged Navier-Stokes RET rotational energy transfer TDLAS tunable diode laser absorption...here extend this knowledge base to flight at Mach 11.5. Griffiths (2004) used a tunable diode laser absorption spectroscopy ( TDLAS ) system to measure

Role of Air-Breathing Pulse Detonation Engines in High Speed Propulsion

NASA Technical Reports Server (NTRS)

Povinelli, Louis A.; Lee, Jin-Ho; Anderberg, Michael O.

2001-01-01

In this paper, the effect of flight Mach number on the relative performance of pulse detonation engines and gas turbine engines is investigated. The effect of ram and mechanical compression on combustion inlet temperature and the subsequent sensible heat release is determined. Comparison of specific thrust, fuel consumption and impulse for the two engines show the relative benefits over the Mach number range.

Thermal Analysis Methods and Basic Heat-Transfer Data for a Turbulent Heating Flight Experiment at Mach 20 (Reentry F)

NASA Technical Reports Server (NTRS)

Howard, Floyd G.

1971-01-01

A heat-transfer experiment was flight conducted on a 5 deg half-angle cone, 396.2 cm (13 ft) in length, as it entered the sensible atmosphere under laminar, transitional, and turbulent boundary-layer conditions at a free-stream Mach number of about 20. Accurate turbulent-heat-transfer data with natural transition were obtained for correlation with theories in regions of simultaneous high Mach number, Reynolds number, enthalpy, and total-to-wall temperature ratio. Temperatures were measured at four depths through the 15.24-mm-thick (0.600-in.) beryllium wall. Experimental heating rates at 20 stations on the cone were determined independently from the outermost temperature measurement and from the temperature measurement at the second depth by a single-thermocouple inverse method and also from the temperature histories at all four depths by an integral method. The thermal data analysis procedure, associated problems, and results are presented herein.

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

NASA Technical Reports Server (NTRS)

1999-01-01

An artist's conception of the X-43A Hypersonic Experimental Vehicle, or 'Hyper-X' in flight. The X-43A was developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

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

NASA Technical Reports Server (NTRS)

1998-01-01

An artist's conception of the X-43A Hypersonic Experimental Vehicle, or 'Hyper-X' in flight. The X-43A was developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

Hyper-X Research Vehicle - Artist Concept in Flight

NASA Technical Reports Server (NTRS)

1997-01-01

An artist's conception of the X-43A Hypersonic Experimental Vehicle, or 'Hyper-X' in flight. The X-43A was developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

Flight-determined characteristics of an air intake system on an F-111A airplane

NASA Technical Reports Server (NTRS)

Hughes, D. L.; Johnson, H. J.

1972-01-01

Flow phenomena of the F-111A air intake system were investigated over a large range of Mach number, altitude, and angle of attack. Boundary-layer variations are shown for the fuselage splitter plate and inlet entrance stations. Inlet performance is shown in terms of pressure recovery, airflow, mass-flow ratio, turbulence factor, distortion factor, and power spectral density. The fuselage boundary layer was found to be not completely removed from the upper portion of the splitter plate at all Mach numbers investigated. Inlet boundary-layer ingestion started at approximately Mach 1.6 near the translating spike and cone. Pressure-recovery distribution at the compressor face showed increasing distortion with increasing angle of attack and increasing Mach number. The time-averaged distortion-factor value approached 1300, which is near the distortion tolerance of the engine at Mach numbers above 2.1.

Sonic-boom measurements for SR-71 aircraft operating at Mach numbers to 3.0 and altitudes to 24384 meters

NASA Technical Reports Server (NTRS)

Maglieri, D. J.; Huckel, V.; Henderson, H. R.

1972-01-01

Sonic-boom pressure signatures produced by the SR-71 aircraft at altitudes from 10,668 to 24,384 meters and Mach numbers 1.35 to 3.0 were obtained as an adjunct to the sonic boom evaluation program relating to structural and subjective response which was conducted in 1966-1967 time period. Approximately 2000 sonic-boom signatures from 33 flights of the SR-71 vehicle and two flights of the F-12 vehicle were recorded. Measured ground-pressure signatures for both on-track and lateral measuring station locations are presented and the statistical variations of the overpressure, positive impulse, wave duration, and shock-wave rise time are illustrated.

Surface flow and heating distributions on a cylinder in near wake of Aeroassist Flight Experiment (AFE) configuration at incidence in Mach 10 Air

NASA Technical Reports Server (NTRS)

Wells, William L.

1990-01-01

Experimental heat transfer distributions and surface streamline directions are presented for a cylinder in the near wake of the Aeroassist Flight Experiment forebody configuration. Tests were conducted in air at a nominal free stream Mach number of 10, with post shock Reynolds numbers based on model base height of 6,450 to 50,770, and angles of attack of 5, 0, -5, and -10 degrees. Heat transfer data were obtained with thin film resistance gage and surface streamline directions by the oil flow technique. Comparisons between measured values and predicted values were made by using a Navier-Stokes computer code.

Wind Tunnel Test of Mach 5 Class Hypersonic Airplane

NASA Astrophysics Data System (ADS)

Nakatani, Hiroki; Taguchi, Hideyuki; Fujita, Kazuhisa; Shindo, Shigemi; Honami, Shinji

JAXA is currently performing studies on a Hypersonic Turbojet Experimental Vehicle, which involve a hypersonic flight test of a Small Pre-cooled Turbojet Engine. The aerodynamic performance of this airplane was examined at the JAXA hypersonic, supersonic, and transonic wind tunnel facilities. The 6-degrees-of-freedom forces and pressure distribution around the model were measured and evaluated. This airplane satisfies the lift-to-drag ratio requirement for a flight test at Mach 5. In addition, the results indicate that this airplane has longitudinal and directional static stability if the moment reference point is x/l smaller than 0.35. A separation occurs at the external expanding nozzle. Therefore, a redesign is necessary to solve these problems.

Cruise performance and range prediction reconsidered

NASA Astrophysics Data System (ADS)

Torenbeek, Egbert

1997-05-01

A unified analytical treatment of the cruise performance of subsonic transport aircraft is derived, valid for gas turbine powerplant installations: turboprop, turbojet and turbofan powered aircraft. Different from the classical treatment the present article deals with compressibility effects on the aerodynamic characteristics. Analytical criteria are derived for optimum cruise lift coefficient and Mach number, with and without constraints on the altitude and engine rating. A simple alternative to the Bréguet range equation is presented which applies to several practical cruising flight techniques: flight at constant altitude and Mach number and stepped cruise/climb. A practical non-iterative procedure for computing mission and reserve fuel loads in the preliminary design stage is proposed.

Preliminary Results Obtained from Flight Test of a Rocket Model Having the Tail Only of the Grumman XF10F Airplane Configuration, TED No. NACA DE 354

NASA Technical Reports Server (NTRS)

Gardner, William N.; Edmondson, James L.

1950-01-01

A flight test was made to determine the servoplane effectiveness and stability characteristics of the free-floating horizontal stabilizer to be used on the XF10F airplane. The results of this test indicate that servoplane effectiveness is practically constant through the speed range up to a Mach number of 1.15, and the stabilizer static stability is satisfactory. A loss of damping occurs over a narrow Mach number range near M = 1.0, resulting in dynamic instability of the stabilizer in this narrow range. Above M = 1.0 there is a gradual positive trim change of the stabilizer.

Wind-tunnel investigation of longitudinal and lateral-directional stability and control characteristics of a 0.237-scale model of a remotely piloted research vehicle with a thick, high-aspect-ratio supercritical wing

NASA Technical Reports Server (NTRS)

Byrdsong, T. A.; Brooks, C. W., Jr.

1980-01-01

A 0.237-scale model of a remotely piloted research vehicle equipped with a thick, high-aspect-ratio supercritical wing was tested in the Langley 8-foot transonic tunnel to provide experimental data for a prediction of the static stability and control characteristics of the research vehicle as well as to provide an estimate of vehicle flight characteristics for a computer simulation program used in the planning and execution of specific flight-research mission. Data were obtained at a Reynolds number of 16.5 x 10 to the 6th power per meter for Mach numbers up to 0.92. The results indicate regions of longitudinal instability; however, an adequate margin of longitudinal stability exists at a selected cruise condition. Satisfactory effectiveness of pitch, roll, and yaw control was also demonstrated.

Computed and Experimental Flutter/LCO Onset for the Boeing Truss-Braced Wing Wind-Tunnel Model

NASA Technical Reports Server (NTRS)

Bartels, Robert E.; Scott, Robert C.; Funk, Christie J.; Allen, Timothy J.; Sexton, Bradley W.

2014-01-01

This paper presents high fidelity Navier-Stokes simulations of the Boeing Subsonic Ultra Green Aircraft Research truss-braced wing wind-tunnel model and compares the results to linear MSC. Nastran flutter analysis and preliminary data from a recent wind-tunnel test of that model at the NASA Langley Research Center Transonic Dynamics Tunnel. The simulated conditions under consideration are zero angle of attack, so that structural nonlinearity can be neglected. It is found that, for Mach number greater than 0.78, the linear flutter analysis predicts flutter onset dynamic pressure below the wind-tunnel test and that predicted by the Navier-Stokes analysis. Furthermore, the wind-tunnel test revealed that the majority of the high structural dynamics cases were wing limit cycle oscillation (LCO) rather than flutter. Most Navier-Stokes simulated cases were also LCO rather than hard flutter. There is dip in the wind-tunnel test flutter/LCO onset in the Mach 0.76-0.80 range. Conditions tested above that Mach number exhibited no aeroelastic instability at the dynamic pressures reached in the tunnel. The linear flutter analyses do not show a flutter/LCO dip. The Navier-Stokes simulations also do not reveal a dip; however, the flutter/LCO onset is at a significantly higher dynamic pressure at Mach 0.90 than at lower Mach numbers. The Navier-Stokes simulations indicate a mild LCO onset at Mach 0.82, then a more rapidly growing instability at Mach 0.86 and 0.90. Finally, the modeling issues and their solution related to the use of a beam and pod finite element model to generate the Navier-Stokes structure mode shapes are discussed.

Performance of high mach number scramjets - Tunnel vs flight

NASA Astrophysics Data System (ADS)

Landsberg, Will O.; Wheatley, Vincent; Smart, Michael K.; Veeraragavan, Ananthanarayanan

2018-05-01

While typically analysed through ground-based impulse facilities, scramjets experience significant heating loads in flight, raising engine wall temperatures and the fuel used to cool them beyond standard laboratory conditions. Hence, the present work numerically compares an access-to-space scramjet's performance at both these conditions. The Mach 12 Rectangular-to-Elliptical Shape-Transitioning scramjet flow path is examined via three-dimensional and chemically reacting Reynolds-averaged Navier-Stokes solutions. Flight operation is modelled through 800 K and 1800 K inlet and combustor walls respectively, while fuel is injected at both inlet- and combustor-based stations at 1000 K stagnation temperature. Room temperature walls and fuel plena model shock tunnel conditions. Mixing and combustion performance indicates that while flight conditions promote rapid mixing, high combustor temperatures inhibit the completion of reaction pathways, with reactant dissociation reducing chemical heat release by 16%. However, the heated walls in flight ensured 28% less energy was absorbed by the walls. While inlet fuel injection promotes robust burning of combustor-injected fuel, premature ignition upon the inlet in flight suggests these injectors should be moved further downstream. Coupled with counteracting differences in heat release and loss to the walls, the optimal engine design for flight may differ considerably from that which gives the best performance in the tunnel.

YF-22 in flight (US AF photo)

NASA Technical Reports Server (NTRS)

1990-01-01

The YF-22, prototype aircraft for the Air Force's F-22 fighter, cruises over the desert on a flight for the Air Force. It was never involved in any programs with Dryden. The United States Air Force announced the demonstration/validation phase contractors selection for the Advanced Tactical Fighter (ATF) program October 31, 1986. These contractor programs were the Lockheed YF-22 and the Northrop YF-23; each produced two prototypes and ground-based avionics testbeds. First flights of all four prototypes occured in 1990. The YF-22 was first flown on Sept. 29, 1990. The YF-22 was powered by two General Electric YF120-GE-100 engines. The final design, the F-22, was flown sometime in May 1997. The F-22 is capable of efficient supersonic operation without afterburner use (supercruise). Lockheed teamed with General Dynamics (Fort Worth) and Boeing Military Airplanes to produce two YF-22 prototypes, civil registrations N22YF (with GE YF120) and N22YX (P&W YF119). N22YF rolled out at Palmdale August 29, 1990; first flight/ferry to Edwards AFB September 29, 1990; first air refuelling (11th sortie) October 26, 1990; thrust vectoring in flight November 15, 1990; achieved Mach 1.8 December 26, 1990. Flight test demonstrations included `supercruise' flight in excess of Mach 1.58 without afterburner.

Preliminary Results of British Nene II Engine Altitude-Chamber Performance Investigation. I - Altitude Performance Using Standard 18.75-Inch-Diameter Jet Nozzle. 1; Altitude Performance Using Standard 18.75-Inch-Diameter Jet Nozzle

NASA Technical Reports Server (NTRS)

Barson, Zelmar; Wilsted, H. D.

1948-01-01

An investigation is being conducted to determine the altitude performance characteristics of the British Nene II engine and its components. The present paper presents the preliminary results obtained using a standard jet nozzle. The test results presented are for conditions simulating altitudes from sea level to 60,000 feet and ram pressure ratios from 1.0 to 2.3. These ram pressure ratios correspond to flight Mach numbers between zero and 1.16 assuming a 100 percent ram recovery.

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.

Interior noise considerations for advanced high-speed turboprop aircraft

NASA Technical Reports Server (NTRS)

Mixson, J. S.; Farassat, F.; Leatherwood, J. D.; Prydz, R.; Revell, J. D.

1982-01-01

This paper describes recent research on noise generated by high-speed propellers, on noise transmission through acoustically treated aircraft sidewalls and on subjective response to simulated turboprop noise. Propeller noise discussion focuses on theoretical prediction methods for complex blade shapes designed for low noise at Mach = 0.8 flight and on comparisons with experimental test results. Noise transmission experiments using a 168 cm. diameter aircraft fuselage model and scaled heavy-double-wall treatments indicate that the treatments perform well and that the predictions are usually conservative. Studies of subjective comfort response in an anechoic environment are described for noise signatures having combinations of broadband and propeller-type tone components.

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.

A preliminary study of the use of intercooling and reheat in conjunction with regeneration for aircraft turbine engines

NASA Technical Reports Server (NTRS)

Eisenberg, J. D.

1977-01-01

The effect on fuel consumption of turbofans with intercooled, regenerative cycles and with intercooled, regenerative, reheat cycles was studied. The technology level for both engine and aircraft was that projected for 1985. The simulated mission was a 5556 km flight carrying 200 passengers at Mach 0.8 at 11582 min. Results indicate that these relatively complex cycles offer little, if any, fuel savings potential relative to a conventional turbofan cycle of comparable advanced technology. The intercooled, regenerative cycle yields about the same fuel economy as a conventional cycle at close to the same overall pressure ratio.

S.S.T.O. performance assessment with in-flight lox collection

NASA Astrophysics Data System (ADS)

Vandenkerckhove, J.; Czysz, P.

1995-10-01

Much attention has recently been given, up to harware development to in-flight oxygen collection as a means to improve considerably the performance of both TSTO & SSTO vehicles. A first assessment suggests that it permits simultaneously to improve much both gross take-off weight (by more than 30%) & dry weight (by more than 15%) of an SSTO and to lower significantly the Mach number of transition scramjet → rocket, from 15 down below 10, thereby reducing dramatically the programmatic development risks. After having compared in-flight lox collection with other SSTO concepts, this paper provides a tentative assessment of the performance of SSTO vehicles taking advantage of it, in particular their sensitivity to changes in system characteristics such as transition Mach number, vehicle slenderness (i.e. Küchemann's parameter τ) or planform loading at take-off and in collection characteristics, in particular collection ratio & specific collection plant weight.

Measurement and Empirical Correlation of Transpiration-Cooling Parameters on a 25 degree Cone in a Turbulent Boundary Layer in Both Free Flight and a Hot-Gas Jet

NASA Technical Reports Server (NTRS)

Walton, Thomas E., Jr.; Rashis, Bernard

1961-01-01

Transpiration-cooling parameters are presented for a turbulent boundary layer on a cone configuration with a total angle of 250 which was tested in both free flight and in an ethylene-heated high-temperature jet at a Mach number of 2.0. The flight-tested cone was flown to a maximum Mach number of 4.08 and the jet tests were conducted at stagnation temperatures ranging from 937 R to 1,850 R. In general, the experimental heat transfer was in good agreement with the theoretical values. Inclusion of the ratio of local stream temperature to wall temperature in the nondimensional flow rate parameter enabled good correlation of both sets of transpiration data. The measured pressure at the forward station coincided with the theoretical pressure over a sharp cone; however, the measured pressure increased with distance from the nose tip.

In-flight demonstration of a Real-Time Flush Airdata Sensing (RT-FADS) system

NASA Technical Reports Server (NTRS)

Whitmore, Stephen A.; Davis, Roy J.; Fife, John Michael

1995-01-01

A prototype real-time flush airdata sensing (RT-FADS) system has been developed and flight tested at the NASA Dryden Flight Research Center. This system uses a matrix of pressure orifices on the vehicle nose to estimate airdata parameters in real time using nonlinear regression. The algorithm is robust to sensor failures and noise in the measured pressures. The RT-FADS system has been calibrated using inertial trajectory measurements that were bootstrapped for atmospheric conditions using meteorological data. Mach numbers as high as 1.6 and angles of attack greater than 45 deg have been tested. The system performance has been evaluated by comparing the RT-FADS to the ship system airdata computer measurements to give a quantitative evaluation relative to an accepted measurement standard. Nominal agreements of approximately 0.003 in Mach number and 0.20 deg in angle of attack and angle of sideslip have been achieved.

A shadowgraph study of the National Launch System's 1 1/2 stage vehicle configuration and Heavy Lift Launch Vehicle configuration. [Using the Marshall Space Flight Center's 14-Inch Trisonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Pokora, Darlene C.; Springer, Anthony M.

1994-01-01

A shadowgraph study of the National Launch System's (NLS's) 1 1/2 stage and heavy lift launch vehicle (HLLV) configurations is presented. Shadowgraphs are shown for the range of Mach numbers from Mach 0.6 to 5.0 at various angles-of-attack and roll angles. Since the 1 1/2 stage configuration is generally symmetric, no shadowgraphs of any roll angle are shown for this configuration. The major flow field phenomena over the NLS 1 1/2 stage and HLLV configurations are shown in the shadowgraphs. These shadowgraphs are used in the aerothermodynamic analysis of the external flow conditions the launch vehicle would encounter during the ascent stage of flight. The shadowgraphs presented in this study were obtained from configurations tested in the Marshall Space Flight Center's 14-Inch Trisonic Wind Tunnel during 1992.

Langley Aerothermodynamic Facilities Complex: Enhancements and Testing Capabilities

NASA Technical Reports Server (NTRS)

Micol, J. R.

1998-01-01

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

Heat-Transfer Measurements on a 5.5- Inch-Diameter Hemispherical Concave Nose in Free Flight at Mach Numbers up to 6.6

NASA Technical Reports Server (NTRS)

Levine, Jack; Rumsey, Charles B.

1958-01-01

The aerodynamic heat transfer to a hemispherical concave nose has been measured in free flight at Mach numbers from 3.5 to 6.6 with corresponding Reynolds numbers based on nose diameter from 7.4 x 10(exp 6) to 14 x 10(exp 6). Over the test Mach number range the heating on the cup nose, expressed as a ratio to the theoretical stagnation-point heating on a hemisphere nose of the same diameter, varied from 0.05 to 0.13 at the stagnation point of the cup, was approximately 0.1 at other locations within 40 deg of the stagnation point, and varied from 0.6 to 0.8 just inside the lip where the highest heating rates occurred. At a Mach number of 5 the total heat input integrated over the surface of the cup nose including the lip was 0.55 times the theoretical value for a hemisphere nose with laminar boundary layer and 0.76 times that for a flat face. The heating at the stagnation point was approximately 1/5 as great as steady-flow tunnel results. Extremely high heating rates at the stagnation point (on the order of 30 times the stagnation-point values of the present test), which have occurred in conjunction with unsteady oscillatory flow around cup noses in wind-tunnel tests at Mach and Reynolds numbers within the present test range, were not observed.

Computational Fluid Dynamics Analysis Method Developed for Rocket-Based Combined Cycle Engine Inlet

NASA Technical Reports Server (NTRS)

1997-01-01

Renewed interest in hypersonic propulsion systems has led to research programs investigating combined cycle engines that are designed to operate efficiently across the flight regime. The Rocket-Based Combined Cycle Engine is a propulsion system under development at the NASA Lewis Research Center. This engine integrates a high specific impulse, low thrust-to-weight, airbreathing engine with a low-impulse, high thrust-to-weight rocket. From takeoff to Mach 2.5, the engine operates as an air-augmented rocket. At Mach 2.5, the engine becomes a dual-mode ramjet; and beyond Mach 8, the rocket is turned back on. One Rocket-Based Combined Cycle Engine variation known as the "Strut-Jet" concept is being investigated jointly by NASA Lewis, the U.S. Air Force, Gencorp Aerojet, General Applied Science Labs (GASL), and Lockheed Martin Corporation. Work thus far has included wind tunnel experiments and computational fluid dynamics (CFD) investigations with the NPARC code. The CFD method was initiated by modeling the geometry of the Strut-Jet with the GRIDGEN structured grid generator. Grids representing a subscale inlet model and the full-scale demonstrator geometry were constructed. These grids modeled one-half of the symmetric inlet flow path, including the precompression plate, diverter, center duct, side duct, and combustor. After the grid generation, full Navier-Stokes flow simulations were conducted with the NPARC Navier-Stokes code. The Chien low-Reynolds-number k-e turbulence model was employed to simulate the high-speed turbulent flow. Finally, the CFD solutions were postprocessed with a Fortran code. This code provided wall static pressure distributions, pitot pressure distributions, mass flow rates, and internal drag. These results were compared with experimental data from a subscale inlet test for code validation; then they were used to help evaluate the demonstrator engine net thrust.

Rocket Based Combined Cycle Exchange Inlet Performance Estimation at Supersonic Speeds

NASA Astrophysics Data System (ADS)

Murzionak, Aliaksandr

A method to estimate the performance of an exchange inlet for a Rocket Based Combined Cycle engine is developed. This method is to be used for exchange inlet geometry optimization and as such should be able to predict properties that can be used in the design process within a reasonable amount of time to allow multiple configurations to be evaluated. The method is based on a curve fit of the shocks developed around the major components of the inlet using solutions for shocks around sharp cones and 2D estimations of the shocks around wedges with blunt leading edges. The total pressure drop across the estimated shocks as well as the mass flow rate through the exchange inlet are calculated. The estimations for a selected range of free-stream Mach numbers between 1.1 and 7 are compared against numerical finite volume method simulations which were performed using available commercial software (Ansys-CFX). The total pressure difference between the two methods is within 10% for the tested Mach numbers of 5 and below, while for the Mach 7 test case the difference is 30%. The mass flow rate on average differs by less than 5% for all tested cases with the maximum difference not exceeding 10%. The estimation method takes less than 3 seconds on 3.0 GHz single core processor to complete the calculations for a single flight condition as oppose to over 5 days on 8 cores at 2.4 GHz system while using 3D finite volume method simulation with 1.5 million elements mesh. This makes the estimation method suitable for the use with exchange inlet geometry optimization algorithm.

In-Flight Boundary-Layer Transition of a Large Flat Plate at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Banks, D. W.; Frederick, M. A.; Tracy, R. R.; Matisheck, J. R.; Vanecek, N. D.

2012-01-01

A flight experiment was conducted to investigate the pressure distribution, local-flow conditions, and boundary-layer transition characteristics on a large flat plate in flight at supersonic speeds up to Mach 2.00. The tests used a NASA testbed aircraft with a bottom centerline mounted test fixture. The primary objective of the test was to characterize the local flow field in preparation for future tests of a high Reynolds number natural laminar flow test article. A second objective was to determine the boundary-layer transition characteristics on the flat plate and the effectiveness of using a simplified surface coating. Boundary-layer transition was captured in both analog and digital formats using an onboard infrared imaging system. Surface pressures were measured on the surface of the flat plate. Flow field measurements near the leading edge of the test fixture revealed the local flow characteristics including downwash, sidewash, and local Mach number. Results also indicated that the simplified surface coating did not provide sufficient insulation from the metallic structure, which likely had a substantial effect on boundary-layer transition compared with that of an adiabatic surface. Cold wall conditions were predominant during the acceleration to maximum Mach number, and warm wall conditions were evident during the subsequent deceleration.

D-558-2 in flight with F-86 chase

NASA Technical Reports Server (NTRS)

1950-01-01

This 1950s photograph shows the Douglas D-558-2 and the North American F-86 Sabre chase aircraft in-flight. Both aircraft display early examples of sweptwing airfoils. The Douglas D-558-2 'Skyrockets' were among the early transonic research airplanes like the X-1, X-4, X-5, and X-92A. Three of the single-seat, swept-wing aircraft flew from 1948 to 1956 in a joint program involving the National Advisory Committee for Aeronautics (NACA), with its flight research done at the NACA's Muroc Flight Test Unit in Calif., redesignated in 1949 the High-Speed Flight Research Station (HSFRS); the Navy-Marine Corps; and the Douglas Aircraft Co. The HSFRS became the High-Speed Flight Station in 1954 and is now known as the NASA Dryden Flight Research Center. The Skyrocket made aviation history when it became the first airplane to fly twice the speed of sound. The 2 in the aircraft's designation referred to the fact that the Skyrocket was the phase-two version of what had originally been conceived as a three-phase program, with the phase-one aircraft having straight wings. The third phase, which never came to fruition, would have involved constructing a mock-up of a combat-type aircraft embodying the results from the testing of the phase one and two aircraft. Douglas pilot John F. Martin made the first flight at Muroc Army Airfield (later renamed Edwards Air Force Base) in Calif. on February 4, 1948. The goals of the program were to investigate the characteristics of swept-wing aircraft at transonic and supersonic speeds with particular attention to pitch-up (uncommanded rotation of the nose of the airplane upwards)--a problem prevalent in high-speed service aircraft of that era, particularly at low speeds during take-off and landing and in tight turns. The three aircraft gathered a great deal of data about pitch-up and the coupling of lateral (yaw) and longitudinal (pitch) motions; wing and tail loads, lift, drag, and buffeting characteristics of swept-wing aircraft at transonic and supersonic speeds; and the effects of the rocket exhaust plume on lateral dynamic stability throughout the speed range. (Plume effects were a new experience for aircraft.) The number three aircraft also gathered information about the effects of external stores (bomb shapes, drop tanks) upon the aircraft's behavior in the transonic region (roughly 0.7 to 1.3 times the speed of sound). In correlation with data from other early transonic research aircraft such as the XF-92A, this information contributed to solutions to the pitch-up problem in swept-wing aircraft. The three airplanes flew a total of 313 times--123 by the number one aircraft (Bureau No. 37973--NACA 143), 103 by the second Skyrocket (Bureau No. 37974--NACA 144), and 87 by airplane number three (Bureau No. 37975--NACA 145). Skyrocket 143 flew all but one of its missions as part of the Douglas contractor program to test the airplane's performance. NACA aircraft 143 was initially powered by a Westinghouse J-34-40 turbojet engine configured only for ground take-offs, but in 1954-55 the contractor modified it to an all-rocket air-launch capability featuring an LR8-RM-6, 4-chamber Reaction Motors engine rated at 6,000 pounds of thrust at sea level (the Navy designation for the Air Force's LR-11 used in the X-1). In this configuration, NACA research pilot John McKay flew the airplane only once for familiarization on September 17, 1956. The 123 flights of NACA 143 served to validate wind-tunnel predictions of the airplane's performance, except for the fact that the airplane experienced less drag above Mach 0.85 than the wind tunnels had indicated. NACA 144 also began its flight program with a turbojet powerplant. NACA pilots Robert A. Champine and John H. Griffith flew 21 times in this configuration to test airspeed calibrations and to research longitudinal and lateral stability and control. In the process, during August of 1949 they encountered pitch-up problems, which NACA engineers recognized as serious because they could produce a limiting and dangerous restriction on flight performance. Hence, they determined to make a complete investigation of the problem. In 1950, Douglas replaced the turbojet with an LR-8 rocket engine, and its pilot, William B. Bridgeman, flew the aircraft seven times up to a speed of Mach 1.88 (1.88 times the speed of sound) and an altitude of 79,494 feet (the latter an unofficial world's altitude record at the time, achieved on August 15, 1951). In the rocket configuration, a Navy P2B (Navy version of the B-29) launched the airplane at approximately 30,000 feet after taking off from the ground with the Skyrocket attached beneath its bomb bay. During Bridgeman's supersonic flights, he encountered a violent rolling motion known as lateral instability that was less pronounced on the Mach 1.88 flight on August 7, 1951, than on a Mach 1.85 flight in June when he pushed over to a low angle of attack (angle of the fuselage or wing to the prevailing wind direction). The NACA engineers studied the behavior of the aircraft before beginning their own flight research in the airplane in September 1951. Over the next couple of years, NACA pilot A. Scott Crossfield flew the airplane 20 times to gather data on longitudinal and lateral stability and control, wing and tail loads, and lift, drag, and buffeting characteristics at speeds up to Mach 1.878. At that point, Marine Lt. Col. Marion Carl flew the airplane to a new (unofficial) altitude record of 83,235 feet on August 21, 1953, and to a maximum speed of Mach 1.728. Following Carl's completion of these flights for the Navy, NACA technicians at the High-Speed Flight Research Station (HSFRS) near Mojave, Calif., outfitted the LR-8 engine's cylinders with nozzle extensions to prevent the exhaust gas from affecting the rudders at supersonic speeds. This addition also increased the engine's thrust by 6.5 percent at Mach 1.7 and 70,000 feet. Even before Marion Carl had flown the Skyrocket, HSFRS Chief Walter C. Williams had petitioned NACA headquarters unsuccessfully to fly the aircraft to Mach 2 to garner the research data at that speed. Finally, after Crossfield had secured the agreement of the Navy's Bureau of Aeronautics, NACA director Hugh L. Dryden relaxed the organization's usual practice of leaving record setting to others and consented to attempting a flight to Mach 2. In addition to adding the nozzle extensions, the NACA flight team at the HSFRS chilled the fuel (alcohol) so more could be poured into the tank and waxed the fuselage to reduce drag. With these preparations and employing a flight plan devised by project engineer Herman O. Ankenbruck to fly to approximately 72,000 feet and push over into a slight dive, Crossfield made aviation history on November 20, 1953, when he flew to Mach 2.005 (1,291 miles per hour). He became the first pilot to reach Mach 2 in this, the only flight in which the Skyrocket flew that fast. Following this flight, Crossfield and NACA pilots Joseph A. Walker and John B. McKay flew the airplane for such purposes as to gather data on pressure distribution, structural loads, and structural heating, with the last flight in the program occurring on December 20, 1956, when McKay obtained dynamic stability data and sound-pressure levels at transonic speeds and above. Meanwhile, NACA 145 had completed 21 contractor flights by Douglas pilots Eugene F. May and Bill Bridgeman in November 1950. In this jet-and-rocket-propelled craft, Scott Crossfield and Walter Jones began the NACA's investigation of pitch-up lasting from September 1951 well into the summer of 1953. They flew the Skyrocket with a variety of wing-fence, wing-slat, and leading-edge chord extension configurations, performing various maneuvers as well as straight-and-level flying at transonic speeds. While fences significantly aided recovery from pitch-up conditions, leading edge chord extensions did not, disproving wind-tunnel tests to the contrary. Slats (long, narrow auxiliary airfoils) in the fully open position eliminated pitch-up except in the speed range around Mach 0.8 to 0.85. In June 1954, Crossfield began an investigation of the effects of external stores (bomb shapes and fuel tanks) upon the aircraft's transonic behavior. McKay and Stanley Butchart completed the NACA's investigation of this issue, with McKay flying the final mission on August 28, 1956. Besides setting several records, the Skyrocket pilots had gathered important data and understanding about what would and would not work to provide stable, controlled flight of a swept-wing aircraft in the transonic and supersonic flight regimes. The data they gathered also helped to enable a better correlation of wind-tunnel test results with actual flight values, enhancing the abilities of designers to produce more capable aircraft for the armed services, especially those with swept wings. Moreover, data on such matters as stability and control from this and other early research airplanes aided in the design of the century series of fighter airplanes, all of which featured the movable horizontal stabilizers first employed on the X-1 and D-558 series.

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

NASA Technical Reports Server (NTRS)

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

1990-01-01

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

Preliminary Design of a Ramjet for Integration with Ground-Based Launch Assist

NASA Technical Reports Server (NTRS)

Sayles, Emily L.

2008-01-01

This viewgraph presentation reviews the preliminary design of a ramjet for integration with a ground based launch assist. The reasons for the use of ground-based launch assist and the proposed mechanism for a system are reviewed. The use of a Optimal Trajectory by Implicit Simulation (OTIS), to model the flight and comparison with an actual rocket trajectory is given. The OTIS system is reviewed, The benefits of a launch assist system are analyzed concluding that a launch assist can provide supersonic speeds thus allowing ignition of ramjet without an onboard compressor. This means a further reduction in total launch weight. The Ramjet study is reviewed next. This included a review of the ONX simulations, the verification of the ONX results with the use of Holloman Sled experiment data as derived from the Feasibility of Ramjet Engine Test Capability on The Holloman AFB Sled Track. The conclusion was that the ONX system was not sufficient to meet the needs for the modeling required. The GECAT (Graphical Engine Cycle Analysis Tool) is examined. The results of the GECAT simulations was verified with data from Stataltex and D21 flights. The Next steps are: to create a GECAT Model of a launch assist ramjet, to adjust the geometry to produce the desired thrust, and to survey the ramjet's performance over a range of Mach numbers. The assumptions and requirements of a launch assist ramjet are given, and the acceptable flight regimes are reviewed.

Turbulence Model Effects on RANS Simulations of the HIFiRE Flight 2 Ground Test Configurations

NASA Technical Reports Server (NTRS)

Georgiadis, Nicholas J.; Mankbadi, Mina R.; Vyas, Manan A.

2014-01-01

The Wind-US Reynolds-averaged Navier-Stokes solver was applied to the Hypersonic International Flight Research Experimentation (HIFiRE) Flight 2 scramjet ground test configuration. Two test points corresponding to flight Mach numbers of 5.9 and 8.9 were examined. The emphasis was examining turbulence model effects on the prediction of flow path pressures. Three variants of the Menter k-omega turbulence model family were investigated. These include the baseline (BSL) and shear stress transport (SST) as well as a modified SST model where the shear stress limiter was altered. Variations in the turbulent Schmidt number were also considered. Choice of turbulence model had a substantial effect on prediction of the flow path pressures. The BSL model produced the highest pressures and the SST model produced the lowest pressures. As expected, the settings for the turbulent Schmidt number also had significant effects on predicted pressures. Small values for the turbulent Schmidt number enabled more rapid mass transfer, faster combustion, and in turn higher flowpath pressures. Optimal settings for turbulence model and turbulent Schmidt number were found to be rather case dependent, as has been concluded in other scramjet investigations.

Flush Airdata Sensing (FADS) System Calibration Procedures and Results for Blunt Forebodies

NASA Technical Reports Server (NTRS)

Cobleigh, Brent R.; Whitmore, Stephen A.; Haering, Edward A., Jr.; Borrer, Jerry; Roback, V. Eric

1999-01-01

Blunt-forebody pressure data are used to study the behavior of the NASA Dryden Flight Research Center flush airdata sensing (FADS) pressure model and solution algorithm. The model relates surface pressure measurements to the airdata state. Spliced from the potential flow solution for uniform flow over a sphere and the modified Newtonian impact theory, the model was shown to apply to a wide range of blunt-forebody shapes and Mach numbers. Calibrations of a sphere, spherical cones, a Rankine half body, and the F-14, F/A-18, X-33, X-34, and X-38 configurations are shown. The three calibration parameters are well-behaved from Mach 0.25 to Mach 5.0, an angle-of-attack range extending to greater than 30 deg, and an angle-of-sideslip range extending to greater than 15 deg. Contrary to the sharp calibration changes found on traditional pitot-static systems at transonic speeds, the FADS calibrations are smooth, monotonic functions of Mach number and effective angles of attack and sideslip. Because the FADS calibration is sensitive to pressure port location, detailed measurements of the actual pressure port locations on the flight vehicle are required and the wind-tunnel calibration model should have pressure ports in similar locations. The procedure for calibrating a FADS system is outlined.

Free-Flight Investigation of Heat Transfer to an Unswept Cylinder Subjected to an Incident Shock and Flow Interference from an Upstream Body at Mach Numbers up to 5.50

NASA Technical Reports Server (NTRS)

Carter, Howard S.; Carr, Robert E.

1961-01-01

Heat-transfer rates have been measured in free flight along the stagnation line of an unswept cylinder mounted transversely on an axial cylinder so that the shock wave from the hemispherical nose of the axial cylinder intersected the bow shock of the unswept transverse cylinder. Data were obtained at Mach numbers from 2.53 to 5.50 and at Reynolds numbers based on the transverse cylinder diameter from 1.00 x 10(exp 6) to 1.87 x 10(exp 6). Shadowgraph pictures made in a wind tunnel showed that the flow field was influenced by boundary-layer separation on the axial cylinder and by end effects on the transverse cylinder as well as by the intersecting shocks. Under these conditions, the measured heat-transfer rates had inconsistent variations both in magnitude and distribution which precluded separating the effects of these disturbances. The general magnitude of the measured heating rates at Mach numbers up to 3 was from 0.1 to 0.5 of the theoretical laminar heating rates along the stagnation line for an infinite unswept cylinder in undisturbed flow. At Mach numbers above 4 the measured heating rates were from 1.5 to 2 times the theoretical rates.

Rocketdyne Development of RBCC Engine for Low Cost Access to Space

NASA Technical Reports Server (NTRS)

Ortwerth, P.; Ratekin, G.; Goldman, A.; Emanuel, M.; Ketchum, A.; Horn, M.

1997-01-01

Rocketdyne is pursuing the conceptual design and development of a Rocket Based Combined Cycle (RBCC) engine for booster and SSTO, advanced reusable space transportation ARTT systems under contract with NASA Marshall Space Flight Center. The Rocketdyne concept is fixed geometry integrated Rocket, Ram Scramjet which is Hydrogen fueled and uses Hydrogen regenerative cooling. Vision vehicle integration studies have determined that scramjet operation to Mach 12 has high payoff for low cost reusable space transportation. Rocketdyne is internally developing versions of the concept for other applications in high speed aircraft and missiles with Hydrocarbon fuel systems. Subscale engine ground testing is underway for all modes of operation from takeoff to Mach 8. High altitude Rocket only mode tests will be completed as part of the ground test program to validate high expansion ratio performance. A unique feature of the ground test series is the inclusion of dynamic trajectory simulation with real time Mach number, altitude, engine throttling, and RBCC mode changes in a specially modified freejet test facility at GASL. Preliminary cold flow Air Augmented Rocket mode test results and Short Combustor tests have met program goals and have been used to integrate all modes of operation in a single combustor design with a fixed geometry inlet for design confirmation tests. A water cooled subscale engine is being fabricated and installed for test beginning the last quarter of 1997.

Low-speed wind tunnel performance of high-speed counterrotation propellers at angle-of-attack

NASA Technical Reports Server (NTRS)

Hughes, Christopher E.; Gazzaniga, John A.

1989-01-01

The low-speed aerodynamic performance characteristics of two advanced counterrotation pusher-propeller configurations with cruise design Mach numbers of 0.72 were investigated in the NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel. The tests were conducted at Mach number 0.20, which is representative of the aircraft take-off/landing flight regime. The investigation determined the effect of nonuniform inflow on the propeller performance characteristics for several blade angle settings and a range of rotational speeds. The inflow was varied by yawing the propeller model to angle-of-attack by as much as plus or minus 16 degrees and by installing on the counterrotation propeller test rig near the propeller rotors a model simulator of an aircraft engine support pylon and fuselage. The results of the investigation indicated that the low-speed performance of the counterrotation propeller configurations near the take-off target operating points were reasonable and were fairly insensitive to changes in model angle-of-attack without the aircraft pylon/fuselage simulators installed on the propeller test rig. When the aircraft pylon/fuselage simulators were installed, small changes in propeller performance were seen at zero angle-of-attack, but fairly large changes in total power coefficient and very large changes of aft-to-forward-rotor torque ratio were produced when the propeller model was taken to angle-of-attack. The propeller net efficiency, though, was fairly insensitive to any changes in the propeller flowfield conditions near the take-off target operating points.

Heat-Transfer Measurements in Free Flight at Mach Numbers up to 14.6 on a Flat-Faced Conical Nose with a Total Angle of 29 Degrees

NASA Technical Reports Server (NTRS)

Rumsey, Charles B; Lee, Dorothy B

1958-01-01

Skin-temperature measurements have been made at several locations on a flat-faced cone-cylinder nose which was flight tested on a fivestage rocket-propeller model to a Mach number of 14.64 and a free-stream Reynolds number of 2.0 x 10(exp 6), based on flat-face diameter, at an altitude of 66,300 feet. The copper nose had a 29 deg total-angle conical section which was 1.6 flat-face diameters long. The aerodynamic-heating rates determined from the temperature measurements reached 1,440 Btu/( sec) (sq ft) on the flat face. The heating rates near the center of the flat face agreed well at Mach numbers up to 13.6 with those obtained by a theory for laminar stagnation-point heating in equilibrium dissociated air (Avco Res. Rep. 1). At Mach numbers above 13.6, the heating rates at locations near the center of the flat face became progressively lower than stagnation-point theory and. were 29 percent lower at Mach number 14.6 at the end. of the test. The reason for this behavior of the heating on the central part of the flat face was not determined. Excluding the relatively low heating rates that occurred on the central part of the nose at the highest Mach numbers, the distribution of experimental heating along the innermost 0.79 of the flat-face radius, expressed as a percentage of stagnation-point heating, was in fair agreement with the distribution predicted by laminar theory. At a location of 0.71 radii from the stagnation point, the experimental heating was very near 130 percent of the theoretical stagnation-point rate at Mach numbers from 11 to 14.5. The experimental beating rates on the conical section of the nose were in good agreement with laminar-cone theory using the assumption of theoretical sharp-cone static pressure on the conical section.

Parametric study of shock-induced combustion in a hydrogen air system

NASA Technical Reports Server (NTRS)

Ahuja, J. K.; Tiwari, Surendra N.

1994-01-01

A numerical parametric study is conducted to simulate shock-induced combustion under various free-stream conditions and varying blunt body diameter. A steady combustion front is established if the free-stream Mach number is above the Chapman-Jouguet speed of the mixture, whereas an unsteady reaction front is established if the free-stream Mach number is below or at the Chapman-Jouguet speed of the mixture. The above two cases have been simulated for Mach 5.11 and Mach 6.46 with a projectile diameter of 15 mm. Mach 5.11, which is an underdriven case, shows an unsteady reaction front, whereas Mach 6.46, which is an overdriven case, shows a steady reaction front. Next for Mach 5. 11 reducing the diameter to 2.5 mm causes the instabilities to disappear, whereas, for Mach 6.46 increasing the diameter of the projectile to 225 mm causes the instabilities to reappear, indicating that Chapman-Jouguet speed is not the only deciding factor for these instabilities to trigger. The other key parameters are the projectile diameter, induction time, activation energy and the heat release. The appearance and disappearance of the instabilities have been explained by the one-dimensional wave interaction model.

PHYSX Glove Test

NASA Technical Reports Server (NTRS)

1995-01-01

A mock-up of the stainless-steel Pegasus Hypersonic Experiment (PHYSX) Projects experimental 'glove' undergoes hot-loads tests at NASA's Dryden Flight Research Center, Edwards, California. The thermal ground test simulates heats and pressures the wing glove will experience at hypersonic speeds. Quartz heat lamps subject this model of a Pegasus booster rocket's right wing glove to the extreme heats it will experience at speeds approaching Mach 8. The glove has a highly reflective surface, underneath which are hundreds of temperature and pressure sensors that will send hypersonic flight data to ground tracking facilities during the experimental flight. Pegasus is an air-launched space booster produced by Orbital Sciences Corporation and Hercules Aerospace Company (initially; later, Alliant Tech Systems) to provide small satellite users with a cost-effective, flexible, and reliable method for placing payloads into low earth orbit. Pegasus has been used to launch a number of satellites and the PHYSX experiment. That experiment consisted of a smooth glove installed on the first-stage delta wing of the Pegasus. The glove was used to gather data at speeds of up to Mach 8 and at altitudes approaching 200,000 feet. The flight took place on October 22, 1998. The PHYSX experiment focused on determining where boundary-layer transition occurs on the glove and on identifying the flow mechanism causing transition over the glove. Data from this flight-research effort included temperature, heat transfer, pressure measurements, airflow, and trajectory reconstruction. Hypersonic flight-research programs are an approach to validate design methods for hypersonic vehicles (those that fly more than five times the speed of sound, or Mach 5). Dryden Flight Research Center, Edwards, California, provided overall management of the glove experiment, glove design, and buildup. Dryden also was responsible for conducting the flight tests. Langley Research Center, Hampton, Virginia, was responsible for the design of the aerodynamic glove as well as development of sensor and instrumentation systems for the glove. Other participating NASA centers included Ames Research Center, Mountain View, California; Goddard Space Flight Center, Greenbelt, Maryland; and Kennedy Space Center, Florida. Orbital Sciences Corporation, Dulles, Virginia, is the manufacturer of the Pegasus vehicle, while Vandenberg Air Force Base served as a pre-launch assembly facility for the launch that included the PHYSX experiment. NASA used data from Pegasus launches to obtain considerable data on aerodynamics. By conducting experiments in a piggyback mode on Pegasus, some critical and secondary design and development issues were addressed at hypersonic speeds. The vehicle was also used to develop hypersonic flight instrumentation and test techniques. NASA's B-52 carrier-launch vehicle was used to get the Pegasus airborne during six launches from 1990 to 1994. Thereafter, an Orbital Sciences L-1011 aircraft launched the Pegasus. The Pegasus launch vehicle itself has a 400- to 600-pound payload capacity in a 61-cubic-foot payload space at the front of the vehicle. The vehicle is capable of placing a payload into low earth orbit. This vehicle is 49 feet long and 50 inches in diameter. It has a wing span of 22 feet. (There is also a Pegasus XL vehicle that was introduced in 1994. Dryden has never launched one of these vehicles, but they have greater thrust and are 56 feet long.)

Very High-Speed Digital Video Capability for In-Flight Use

NASA Technical Reports Server (NTRS)

Corda, Stephen; Tseng, Ting; Reaves, Matthew; Mauldin, Kendall; Whiteman, Donald

2006-01-01

digital video camera system has been qualified for use in flight on the NASA supersonic F-15B Research Testbed aircraft. This system is capable of very-high-speed color digital imaging at flight speeds up to Mach 2. The components of this system have been ruggedized and shock-mounted in the aircraft to survive the severe pressure, temperature, and vibration of the flight environment. The system includes two synchronized camera subsystems installed in fuselage-mounted camera pods (see Figure 1). Each camera subsystem comprises a camera controller/recorder unit and a camera head. The two camera subsystems are synchronized by use of an MHub(TradeMark) synchronization unit. Each camera subsystem is capable of recording at a rate up to 10,000 pictures per second (pps). A state-of-the-art complementary metal oxide/semiconductor (CMOS) sensor in the camera head has a maximum resolution of 1,280 1,024 pixels at 1,000 pps. Exposure times of the electronic shutter of the camera range from 1/200,000 of a second to full open. The recorded images are captured in a dynamic random-access memory (DRAM) and can be downloaded directly to a personal computer or saved on a compact flash memory card. In addition to the high-rate recording of images, the system can display images in real time at 30 pps. Inter Range Instrumentation Group (IRIG) time code can be inserted into the individual camera controllers or into the M-Hub unit. The video data could also be used to obtain quantitative, three-dimensional trajectory information. The first use of this system was in support of the Space Shuttle Return to Flight effort. Data were needed to help in understanding how thermally insulating foam is shed from a space shuttle external fuel tank during launch. The cameras captured images of simulated external tank debris ejected from a fixture mounted under the centerline of the F-15B aircraft. Digital video was obtained at subsonic and supersonic flight conditions, including speeds up to Mach 2 and altitudes up to 50,000 ft (15.24 km). The digital video was used to determine the structural survivability of the debris in a real flight environment and quantify the aerodynamic trajectories of the debris.

A study of the transmission characteristics of suppressor nozzles

NASA Technical Reports Server (NTRS)

Ahuja, K. K.; Salikuddin, M.; Burrin, R. H.; Plumbee, H. E., Jr.

1980-01-01

The internal noise radiation characteristics for a single stream 12 lobe 24 tube suppressor nozzle, and for a dual stream 36 chute suppressor nozzle were investigated. An equivalent single round conical nozzle and an equivalent coannular nozzle system were also tested to provide a reference for the two suppressors. The technique utilized a high voltage spark discharge as a noise source within the test duct which permitted separation of the incident, reflected and transmitted signals in the time domain. These signals were then Fourier transformed to obtain the nozzle transmission coefficient and the power transfer function. These transmission parameters for the 12 lobe, 24 tube suppressor nozzle and the reference conical nozzle are presented as a function of jet Mach number, duct Mach number polar angle and temperature. Effects of simulated forward flight are also considered for this nozzle. For the dual stream, 36 chute suppressor, the transmission parameters are presented as a function of velocity ratios and temperature ratios. Possible data for the equivalent coaxial nozzle is also presented. Jet noise suppression by these nozzles is also discussed.

LeRC NATR Free-Jet Development

NASA Technical Reports Server (NTRS)

Long-Davis, M.; Cooper, B. A.

1999-01-01

The Nozzle Acoustic Test Rig (NATR) was developed to provide additional test capabilities at Lewis needed to meet HSR program goals. The NATR is a large f ree-jet facility (free-jet diameter = 53 in.) with a design Mach number of 0.3. It is located inside a geodesic dome, adjacent to the existing Powered Lift Facility (PLF). The NATR allows nozzle concepts to be acoustically assessed for far-field (approximately 50 feet) noise characteristics under conditions simulating forward flight. An ejector concept was identified as a means of supplying the required airflow for this free-jet facility. The primary stream is supplied through a circular array of choked nozzles and the resulting low pressure in the constant, annular- area mixing section causes a "pumping" action that entrains the secondary stream. The mixed flow expands through an annular diffuser and into a plenum chamber. Once inside the plenum, the flow passes over a honeycomb/screen combination intended to remove large disturbances and provide uniform flow. The flow accelerates through an elliptical contraction section where it achieves a free-jet Mach number of up to 0.3.

High-Speed Boundary-Layer Transition Induced by an Isolated Roughness Element

NASA Technical Reports Server (NTRS)

Kegerise, Michael A.; Owens, Lewis R.; King, Rudolph A.

2010-01-01

Progress on an experimental effort to quantify the instability mechanisms associated with roughness-induced transition in a high-speed boundary layer is reported in this paper. To simulate the low-disturbance environment encountered during high-altitude flight, the experimental study was performed in the NASA-Langley Mach 3.5 Supersonic Low-Disturbance Tunnel. A flat plate trip sizing study was performed first to identify the roughness height required to force transition. That study, which included transition onset measurements under both quiet and noisy freestream conditions, confirmed the sensitivity of roughness-induced transition to freestream disturbance levels. Surveys of the laminar boundary layer on a 7deg half-angle sharp-tipped cone were performed via hot-wire anemometry and pitot-pressure measurements. The measured mean mass-flux and Mach-number profiles agreed very well with computed mean-flow profiles. Finally, surveys of the boundary layer developing downstream of an isolated roughness element on the cone were performed. The measurements revealed an instability in the far wake of the roughness element that grows exponentially and has peak frequencies in the 150 to 250 kHz range.

Effect of fuel zoning and fuel nozzle design on pollution emissions at ground idle conditions for a double-annular ram-induction combustor

NASA Technical Reports Server (NTRS)

Clements, T. R.

1973-01-01

An exhaust emission survey was conducted on a double-annular ram induction combustor at simulated ground idle conditions. The combustor was designed for a large augmented turbofan engine capable of sustained flight speeds up to Mach 3.0. The emission levels of total hydrocarbon (THC), carbon monoxide, carbon dioxide, and nitric oxide were measured. The effects of fuel zoning, fuel nozzle design, and operating conditions (inlet temperature and reference Mach number) on the level of these emissions were determined. At an overall combustor fuel/air ratio of 0.007, fuel zoning reduced THC emissions by a factor of 5 to 1. The reduction in THC emissions is attributed to the increase in local fuel/air ratio provided by the fuel zoning. An alternative method of increasing fuel/air ratio would be to operate with larger-than-normal compressor overboard bleed; however, analysis on this method indicated an increase in idle fuel consumption of 20 percent. The use of air-atomizing nozzles reduced the THC emissions by 2 to 1.

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

NASA Technical Reports Server (NTRS)

Sanders, Bobby W.; Weir, Lois J.

2008-01-01

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

Computational fluid dynamics and frequency-dependent finite-difference time-domain method coupling for the interaction between microwaves and plasma in rocket plumes

NASA Astrophysics Data System (ADS)

Kinefuchi, K.; Funaki, I.; Shimada, T.; Abe, T.

2012-10-01

Under certain conditions during rocket flights, ionized exhaust plumes from solid rocket motors may interfere with radio frequency transmissions. To understand the relevant physical processes involved in this phenomenon and establish a prediction process for in-flight attenuation levels, we attempted to measure microwave attenuation caused by rocket exhaust plumes in a sea-level static firing test for a full-scale solid propellant rocket motor. The microwave attenuation level was calculated by a coupling simulation of the inviscid-frozen-flow computational fluid dynamics of an exhaust plume and detailed analysis of microwave transmissions by applying a frequency-dependent finite-difference time-domain method with the Drude dispersion model. The calculated microwave attenuation level agreed well with the experimental results, except in the case of interference downstream the Mach disk in the exhaust plume. It was concluded that the coupling estimation method based on the physics of the frozen plasma flow with Drude dispersion would be suitable for actual flight conditions, although the mixing and afterburning in the plume should be considered depending on the flow condition.

Influence of the Angle of Attack on the Aerothermodynamics of the Mars Science Laboratory

NASA Technical Reports Server (NTRS)

Dyakonov, Artem A.; Edquist, Karl T.; Schoenenberger, Mark

2006-01-01

An investigation of the effects of the incidence angle on the aerothermodynamic environments of the Mars Science Laboratory has been conducted. Flight conditions of peak heating, peak deceleration and chute deploy are selected and the effects of the angle of attack on the aerodynamics and aerothermodynamics are analyzed. The investigation found that static aerodynamics are well behaved within the considered range of incidence angles. Leeside laminar and turbulent computed heating rates decrease with incidence, despite the increase in the leeside running length. Stagnation point was found to stay on the conical flank at all angles of attack, and this is linked to the rapid flow expansion around the shoulder. Hypersonic lift to drag ratio is limited by the heating rates in the region of the windside shoulder. The effects of the high angle of incidence on the dynamic aero at low Mach remains to be determined. Influence of the angle of attack on the smooth-wall transition parameter indicates, that higher angle of attack flight may result in delayed turbulence onset, however, a coupled analysis, involving flight trajectory simulation is necessary.

Longitudinal Aerodynamic Characteristics and Effect of Rocket Jet on Drag of Models of the Hermes A-3A and A-3B Missiles in Free Flight at Mach Numbers From 0.6 to 2.0

NASA Technical Reports Server (NTRS)

Jackson, H. Herbert

1955-01-01

A free-flight investigation over a Mach number range from 0.6 to 2.0 has been conducted to determine the longitudinal aerodynamic characteristics and effect of rocket jet on zero-lift drag of 1/5-scale models of two ballistic-type missiles, the Hermes A-3A and A-3B. Models of both types of missiles exhibited very nearly linear normal forces and pitching moments over the angle-of-attack range of 8 deg to -4 deg and Mach number range tested. The centers of pressure for both missiles were not appreciably affected by Mach number over the subsonic range; however, between a Mach number of 1.02 and 1.50 the center of pressure for the A-3A model moved forward 0.34 caliber with increasing Mach number. At a trim angle-of-attack of approximately 30 deg, the A-3A model indicated a total drag coefficient 30% higher than the power-off zero-lift drag over the subsonic Mach number range and 10% higher over the supersonic range. Under the conditions of the present test, and excluding the effect of the jet on base drag, there was no indicated effect of the propulsive jet on the total drag of the A-3A model. The propulsive jet operating at a jet pressure ratio p(sub j)/p(sub o) of 0.8 caused approximately 100% increase in base drag over the Mach number range M = 0.6 to 1.0. This increase in base drag amounts to 15% of the total drag. An underexpanded jet operating at jet pressure ratios corresponding approximately to those of the full-scale missile caused a 22% reduction in base drag at M = 1.55 (p(sub j)/p(sub o) = 1.76) but indicated no change at M = 1.30 (p(sub j)/p(sub o) = 1.43). At M = 1.1 and p(sub j)/p(sub o) = 1.55, the jet caused a 50% increase in base drag.

The Effect of Altitude Conditions on the Particle Emissions of a J85-GE-5L Turbojet Engine

NASA Technical Reports Server (NTRS)

Rickey, June Elizabeth

1995-01-01

Particles from a J85-GE-5L turbojet engine were measured over a range of engine speeds at simulated altitude conditions ranging from near sea level to 45,000 ft and at flight Mach numbers of 0.5 and 0.8. Samples were collected from the engine by using a specially designed probe positioned several inches behind the exhaust nozzle. A differential mobility particle sizing system was used to determine particle size. Particle data measured at near sea-level conditions were compared with Navy Aircraft Environmental Support Office (AESO) particle data taken from a GE-J85-4A engine at a sea-level static condition. Particle data from the J85 engine were also compared with particle data from a J85 combustor at three different simulated altitudes.

Altitude Performance of AN-F-58 Fuels in British Rolls-Royce Nene Single Combustor

NASA Technical Reports Server (NTRS)

Cook, William P.; Koch, Richard G.

1949-01-01

An investigation was conducted with a single combustor from a British Rolls-Royce Nene turbojet engine to determine the altitude performance characteristics of AN-F-58 fuels. Three fuel blends conforming to AN-F-58 specifications were prepared in order to determine the influence of fuel boiling temperatures and aromatic content on combustion efficiencies and altitude operational limits. The performance of the three AN-F-58 fuels was compared in the range of altitudes from sea level to 65,000 feet, engine speeds from 40- to 100- percent normal rated, and flight Mach numbers of 0.0 and 0.6. Similar information was obtained for AN-F-32 fuel at a flight Mach number of 0.0

Ballistic range experiments on the superboom generated at increasing flight Mach numbers

NASA Technical Reports Server (NTRS)

Sanai, M.; Toong, T.-Y.; Pierce, A. D.

1976-01-01

Ballistic range experiments for the study of the propagation of converging shocks are described and the similarity between the observed phenomenon and that expected for superbooms created by accelerating supersonic aircraft is discussed. For weak shocks (shock Mach numbers of about 1.03), a structure resembling that of a folded shock predicted by geometrical acoustics theory is observed while for stronger shocks, a concave front with enhanced overpressure is recorded. Other results are in general accord with the basic concepts of shock propagation and in conjunction with some theoretical scaling laws indicate that the peak magnification of sonic booms due to aircraft flight acceleration in the real atmosphere should be in the range of 6 to 13.

F-16XL Wing Pressure Distributions and Shock Fence Results from Mach 1.4 to Mach 2.0

NASA Technical Reports Server (NTRS)

Landers, Stephen F.; Saltzman, John A.; Bjarke, Lisa J.

1997-01-01

Chordwise pressure distributions were obtained in-flight on the upper and lower surfaces of the F-16XL ship 2 aircraft wing between Mach 1.4 and Mach 2.0. This experiment was conducted to determine the location of shock waves which could compromise or invalidate a follow-on test of a large chord laminar flow control suction panel. On the upper surface, the canopy closure shock crossed an area which would be covered by a proposed laminar flow suction panel. At the laminar flow experiment design Mach number of 1.9, 91 percent of the suction panel area would be forward of the shock. At Mach 1.4, that value reduces to 65 percent. On the lower surface, a shock from the inlet diverter would impinge on the proposed suction panel leading edge. A chordwise plate mounted vertically to deflect shock waves, called a shock fence, was installed between the inlet diverter and the leading edge. This plate was effective in reducing the pressure gradients caused by the inlet shock system.

Preliminary Investigation of a New Type of Supersonic Inlet

NASA Technical Reports Server (NTRS)

Ferri, Antonio; Nucci, Louis M

1946-01-01

A supersonic inlet with supersonic deceleration of the flow entirely outside of the inlet is considered. A particular arrangement with fixed geometry having a central body with a circular annular intake is analyzed, and it is shown theoretically that this arrangement gives high pressure recovery for a large range of Mach number and mass flow and therefore is practical for use on supersonic airplanes and missiles. For some Mach numbers the drag coefficient for this type of inlet is larger than the drag coefficient for the type of inlet with supersonic compression entirely inside, but the pressure recovery is larger for all flight conditions. The differences in drag can be eliminated for the design Mach number. Experimental results confirm the results of the theoretical analysis and show that pressure recoveries of 95 percent for Mach numbers of 1.33 and 1.52, 92 percent for a Mach number of 1.72, and 86 percent for a Mach number oof 2.10 are possible with the configurations considered. If the mass flow decreases, the total drag coefficient increases gradually and the pressure recovery does not change appreciably.

Flight-determined stability and control characteristics of the M2-F3 lifting body vehicle

NASA Technical Reports Server (NTRS)

Sim, A. G.

1973-01-01

Flight data were obtained over a Mach number range from 0.4 to 1.55 and an angle-of-attack range from -2 deg to 16 deg. Lateral-directional and longitudinal derivatives, reaction control rocket effectiveness, and longitudinal trim information obtained from flight data and wind-tunnel predictions are compared. The effects of power, configuration change, and speed brake are discussed.

Thermodynamic Cycle Analysis of Magnetohydrodynamic-Bypass Airbreathing Hypersonic Engines

NASA Technical Reports Server (NTRS)

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

1999-01-01

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

Induced Moment Effects of Formation Flight Using Two F/A-18 Aircraft

NASA Technical Reports Server (NTRS)

Hansen, Jennifer L.; Cobleigh, Brent R.

2002-01-01

Previous investigations into formation flight have shown the possibility for significant fuel savings through drag reduction. Using two F/A-18 aircraft, NASA Dryden Flight Research Center has investigated flying aircraft in autonomous formation. Positioning the trailing airplane for best drag reduction requires investigation of the wingtip vortex effects induced by the leading airplane. A full accounting of the vortex effect on the trailing airplane is desired to validate vortex-effect prediction methods and provide a database for the design of a formation flight autopilot. A recent flight phase has mapped the complete wingtip vortex effects at two flight conditions with the trailing airplane at varying distances behind the leading one. Force and moment data at Mach 0.56 and an altitude of 25,000 ft and Mach 0.86 and an altitude of 36,000 ft have been obtained with 20, 55, 110, and 190 ft of longitudinal distance between the aircraft. The moments induced by the vortex on the trailing airplane were well within the pilot's ability to control. This report discusses the data analysis methods and vortex-induced effects on moments and side force. An assessment of the impact of the nonlinear vortex effects on the design of a formation autopilot is offered.

Measured Sonic Boom Signatures Above and Below the XB-70 Airplane Flying at Mach 1.5 and 37,000 Feet

NASA Technical Reports Server (NTRS)

Maglieri, Domenic J.; Henderson, Herbert R.; Tinetti, Ana F.

2011-01-01

During the 1966-67 Edwards Air Force Base (EAFB) National Sonic Boom Evaluation Program, a series of in-flight flow-field measurements were made above and below the USAF XB-70 using an instrumented NASA F-104 aircraft with a specially designed nose probe. These were accomplished in the three XB-70 flights at about Mach 1.5 at about 37,000 ft. and gross weights of about 350,000 lbs. Six supersonic passes with the F-104 probe aircraft were made through the XB-70 shock flow-field; one above and five below the XB-70. Separation distances ranged from about 3000 ft. above and 7000 ft. to the side of the XB-70 and about 2000 ft. and 5000 ft. below the XB-70. Complex near-field "sawtooth-type" signatures were observed in all cases. At ground level, the XB-70 shock waves had not coalesced into the two-shock classical sonic boom N-wave signature, but contained three shocks. Included in this report is a description of the generating and probe airplanes, the in-flight and ground pressure measuring instrumentation, the flight test procedure and aircraft positioning, surface and upper air weather observations, and the six in-flight pressure signatures from the three flights.

Real-Time Simulation of Aeroheating of the Hyper-X Airplane

NASA Technical Reports Server (NTRS)

Gong, Les

2005-01-01

A capability for real-time computational simulation of aeroheating has been developed in support of the Hyper-X program, which is directed toward demonstrating the feasibility of operating an air-breathing ramjet/scramjet engine at mach 5, mach 7, and mach 10. The simulation software will serve as a valuable design tool for initial trajectory studies in which aerodynamic heating is expected to exert a major influence in the design of the Hyper-X airplane; this tool will aid in the selection of materials, sizing of structural skin thicknesses, and selection of components of a thermal-protection system (TPS) for structures that must be insulated against aeroheating.

Wedge and Conical Probes for the Instantaneous Measurement of Free-Stream Flow Quantities at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Bobbitt, Percy J.; Maglieri, Domenic J.; Banks, Daniel W.; Fuchs, Aaron W.

2011-01-01

Wedge and conical shaped probes for the measurement of free-stream flow quantities at supersonic speeds have been tested in both wind tunnel and flight. These probes have improved capabilities over similar ones used in the past. Through the use of miniature pressure sensors, that are located inside the probes, they are able to provide instantaneous measurements of a time-varying environment. Detailed herein are the results of the tests in NASA Langley Researcher Center s Unitary Plan Wind Tunnel (UPWT) at Mach numbers of 1.6, 1.8 and 2.0, as well as flight tests carried out at the NASA Dryden Flight Research Center (DFRC) on its F-15 aircraft up to Mach numbers of 1.9. In the flight tests the probes were attached to a fixture on the underside of the F-15 fuselage. 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 impact the accuracy of the measurements.

The SR-71 Test Bed Aircraft: A Facility for High-Speed Flight Research

NASA Technical Reports Server (NTRS)

Corda, Stephen; Moes, Timothy R.; Mizukami, Masashi; Hass, Neal E.; Jones, Daniel; Monaghan, Richard C.; Ray, Ronald J.; Jarvis, Michele L.; Palumbo, Nathan

2000-01-01

The SR-71 test bed aircraft is shown to be a unique platform to flight-test large experiments to supersonic Mach numbers. The test bed hardware mounted on the SR-71 upper fuselage is described. This test bed hardware is composed of a fairing structure called the "canoe" and a large "reflection plane" flat plate for mounting experiments. Total experiment weights, including the canoe and reflection plane, as heavy as 14,500 lb can be mounted on the aircraft and flight-tested to speeds as fast as Mach 3.2 and altitudes as high as 80,000 ft. A brief description of the SR-71 aircraft is given, including details of the structural modifications to the fuselage, modifications to the J58 engines to provide increased thrust, and the addition of a research instrumentation system. Information is presented based on flight data that describes the SR-71 test bed aerodynamics, stability and control, structural and thermal loads, the canoe internal environment, and reflection plane flow quality. Guidelines for designing SR-71 test bed experiments are also provided.

Flight test results of riblets at supersonic speeds

NASA Technical Reports Server (NTRS)

Zuniga, Fanny A.; Anderson, Bianca T.; Bertelrud, Arild

1992-01-01

A flight experiment to test and evaluate the skin friction drag characteristics of a riblet surface in turbulent flow at supersonic speeds was conducted at NASA Dryden. Riblets of groove sizes 0.0030 and 0.0013 in. were mounted on the F-104G flight test fixture. The test surfaces were surveyed with boundary layer rakes and pressure orifices to examine the boundary layer profiles and pressure distributions of the flow. Skin friction reductions caused by the riblet surface were reported based on measured differences of momentum thickness between the smooth and riblet surfaces obtained from the boundary layer data. Flight test results for the 0.0030 in. riblet show skin friction reductions of 4 to 8 % for Mach numbers ranging from 1.2 to 1.6 and Reynolds numbers ranging from 2 to 3.4 million per unit foot. The results from the 0.0013 in. riblets show skin friction reductions of 4 to 15 % for Mach 1.2 to 1.4 and Reynolds numbers ranging from 3.6 to 6 million per unit foot.

Aerodynamic Simulation of Runback Ice Accretion

NASA Technical Reports Server (NTRS)

Broeren, Andy P.; Whalen, Edward A.; Busch, Greg T.; Bragg, Michael B.

2010-01-01

This report presents the results of recent investigations into the aerodynamics of simulated runback ice accretion on airfoils. Aerodynamic tests were performed on a full-scale model using a high-fidelity, ice-casting simulation at near-flight Reynolds (Re) number. The ice-casting simulation was attached to the leading edge of a 72-in. (1828.8-mm ) chord NACA 23012 airfoil model. Aerodynamic performance tests were conducted at the ONERA F1 pressurized wind tunnel over a Reynolds number range of 4.7?10(exp 6) to 16.0?10(exp 6) and a Mach (M) number ran ge of 0.10 to 0.28. For Re = 16.0?10(exp 6) and M = 0.20, the simulated runback ice accretion on the airfoil decreased the maximum lift coe fficient from 1.82 to 1.51 and decreased the stalling angle of attack from 18.1deg to 15.0deg. The pitching-moment slope was also increased and the drag coefficient was increased by more than a factor of two. In general, the performance effects were insensitive to Reynolds numb er and Mach number changes over the range tested. Follow-on, subscale aerodynamic tests were conducted on a quarter-scale NACA 23012 model (18-in. (457.2-mm) chord) at Re = 1.8?10(exp 6) and M = 0.18, using low-fidelity, geometrically scaled simulations of the full-scale castin g. It was found that simple, two-dimensional simulations of the upper- and lower-surface runback ridges provided the best representation of the full-scale, high Reynolds number iced-airfoil aerodynamics, whereas higher-fidelity simulations resulted in larger performance degrada tions. The experimental results were used to define a new subclassification of spanwise ridge ice that distinguishes between short and tall ridges. This subclassification is based upon the flow field and resulting aerodynamic characteristics, regardless of the physical size of the ridge and the ice-accretion mechanism.

Shock Tube and Ballistic Range Facilities at NASA Ames Research Center

DTIC Science & Technology

2010-04-01

in gun performance and launching techniques enabled Mach numbers over 10 to be realized at flight-relevant Reynolds numbers. Although the SSFF...contoured nozzle produced an accelerating supersonic flow with a concomitant exponential decay in density. A two-stage light gas gun fired models upstream...km/s, or Mach 50. Such extreme conditions could not be met by a counterflow facility consisting of the highest performance light gas gun and a

Conventional Prompt Global Strike: Capabilities Today While Planning for Tomorrow

DTIC Science & Technology

2012-03-19

sound. Scramjet powered vehicles are envisioned to operate at speeds up to at least Mach 15.”18 It is beyond the scope of this paper to recommend...launch, the WaveRider “completed the longest supersonic combustion scramjet - powered flight in history, flying at approximately Mach 5 for 143...seconds. This technology is setting the foundation for hypersonic application”17 and could power a hypersonic CPGS delivery vehicle . NASA describes a

Combustion-Powered Actuation for Dynamic Stall Suppression - Simulations and Low-Mach Experiments

NASA Technical Reports Server (NTRS)

Matalanis, Claude G.; Min, Byung-Young; Bowles, Patrick O.; Jee, Solkeun; Wake, Brian E.; Crittenden, Tom; Woo, George; Glezer, Ari

2014-01-01

An investigation on dynamic-stall suppression capabilities of combustion-powered actuation (COMPACT) applied to a tabbed VR-12 airfoil is presented. In the first section, results from computational fluid dynamics (CFD) simulations carried out at Mach numbers from 0.3 to 0.5 are presented. Several geometric parameters are varied including the slot chordwise location and angle. Actuation pulse amplitude, frequency, and timing are also varied. The simulations suggest that cycle-averaged lift increases of approximately 4% and 8% with respect to the baseline airfoil are possible at Mach numbers of 0.4 and 0.3 for deep and near-deep dynamic-stall conditions. In the second section, static-stall results from low-speed wind-tunnel experiments are presented. Low-speed experiments and high-speed CFD suggest that slots oriented tangential to the airfoil surface produce stronger benefits than slots oriented normal to the chordline. Low-speed experiments confirm that chordwise slot locations suitable for Mach 0.3-0.4 stall suppression (based on CFD) will also be effective at lower Mach numbers.

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

NASA Technical Reports Server (NTRS)

1997-01-01

This is an artist's depiction of a Hyper-X research vehicle under scramjet power in free-flight following separation from its booster rocket. The X-43A was developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

Numerical Modeling of Active Flow Control in a Boundary Layer Ingesting Offset Inlet

NASA Technical Reports Server (NTRS)

Allan, Brian G.; Owens, Lewis R.; Berrier, Bobby L.

2004-01-01

This investigation evaluates the numerical prediction of flow distortion and pressure recovery for a boundary layer ingesting offset inlet with active flow control devices. The numerical simulations are computed using a Reynolds averaged Navier-Stokes code developed at NASA. The numerical results are validated by comparison to experimental wind tunnel tests conducted at NASA Langley Research Center at both low and high Mach numbers. Baseline comparisons showed good agreement between numerical and experimental results. Numerical simulations for the inlet with passive and active flow control also showed good agreement at low Mach numbers where experimental data has already been acquired. Numerical simulations of the inlet at high Mach numbers with flow control jets showed an improvement of the flow distortion. Studies on the location of the jet actuators, for the high Mach number case, were conducted to provide guidance for the design of a future experimental wind tunnel test.

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

NASA Technical Reports Server (NTRS)

Benyo, Theresa L.

2010-01-01

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

Procedure for noise prediction and optimization of advanced technology propellers

NASA Technical Reports Server (NTRS)

Jou, W. H.; Bernstein, S.

1979-01-01

The sound field due to a propeller operating at supersonic tip speed in a uniform flow was investigated. Using the fact that the wave front in a uniform stream is a convected sphere, the fundamental solution to the convected wave equation was easily obtained. The Fourier coefficients of the pressure signature were obtained by a far field approximation, and are expressed as an integral over the blade platform. It is shown that cones of silence exist fore and aft the propeller plane. The semiapex angles are shown. These angles are independent of the individual Mach components such as the flight Mach number and the rotation Mach number. The result is confirmed by the computation of the ray path of the emitted Mach waves. The Doppler amplification factor strengthens the signal behind the propeller while it weakens that upstream.

OPACITY BROADENING OF {sup 13}CO LINEWIDTHS AND ITS EFFECT ON THE VARIANCE-SONIC MACH NUMBER RELATION

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

Correia, C.; De Medeiros, J. R.; Burkhart, B.

2014-04-10

We study how the estimation of the sonic Mach number (M{sub s} ) from {sup 13}CO linewidths relates to the actual three-dimensional sonic Mach number. For this purpose we analyze MHD simulations that include post-processing to take radiative transfer effects into account. As expected, we find very good agreement between the linewidth estimated sonic Mach number and the actual sonic Mach number of the simulations for optically thin tracers. However, we find that opacity broadening causes M{sub s} to be overestimated by a factor of ≈1.16-1.3 when calculated from optically thick {sup 13}CO lines. We also find that there ismore » a dependence on the magnetic field: super-Alfvénic turbulence shows increased line broadening compared with sub-Alfvénic turbulence for all values of optical depth for supersonic turbulence. Our results have implications for the observationally derived sonic Mach number-density standard deviation (σ{sub ρ/(ρ)}) relationship, σ{sub ρ/〈ρ〉}{sup 2}=b{sup 2}M{sub s}{sup 2}, and the related column density standard deviation (σ {sub N/(N)}) sonic Mach number relationship. In particular, we find that the parameter b, as an indicator of solenoidal versus compressive driving, will be underestimated as a result of opacity broadening. We compare the σ {sub N/(N)}-M{sub s} relation derived from synthetic dust extinction maps and {sup 13}CO linewidths with recent observational studies and find that solenoidally driven MHD turbulence simulations have values of σ {sub N/(N)}which are lower than real molecular clouds. This may be due to the influence of self-gravity which should be included in simulations of molecular cloud dynamics.« less

Simulation of transient flow in a shock tunnel and a high Mach number nozzle

NASA Technical Reports Server (NTRS)

Jacobs, P. A.

1991-01-01

A finite volume Navier-Stokes code was used to simulate the shock reflection and nozzle starting processes in an axisymmetric shock tube and a high Mach number nozzle. The simulated nozzle starting processes were found to match the classical quasi-1-D theory and some features of the experimental measurements. The shock reflection simulation illustrated a new mechanism for the driver gas contamination of the stagnated test gas.

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

NASA Technical Reports Server (NTRS)

Ko, William L.; Gong, Leslie

2001-01-01

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

Visualization of Flow Separation Around an Atmospheric Entry Capsule at Low-Subsonic Mach Number Using Background-Oriented Schlieren (BOS)

NASA Technical Reports Server (NTRS)

Mizukaki, Toshiharu; Borg, Stephen E.; Danehy, Paul M.; Murman, Scott M.

2014-01-01

This paper presents the results of visualization of separated flow around a generic entry capsule that resembles the Apollo Command Module (CM) and the Orion Multi-Purpose Crew Vehicle (MPCV). The model was tested at flow speeds up to Mach 0.4 at a single angle of attack of 28 degrees. For manned spacecraft using capsule-shaped vehicles, certain flight operations such as emergency abort maneuvers soon after launch and flight just prior to parachute deployment during the final stages of entry, the command module may fly at low Mach number. Under these flow conditions, the separated flow generated from the heat-shield surface on both windward and leeward sides of the capsule dominates the wake flow downstream of the capsule. In this paper, flow visualization of the separated flow was conducted using the background-oriented schlieren (BOS) method, which has the capability of visualizing significantly separated wake flows without the particle seeding required by other techniques. Experimental results herein show that BOS has detection capability of density changes on the order of 10(sup-5).

Helicopter far-field acoustic levels as a function of reduced main-rotor advancing blade-tip Mach number

NASA Technical Reports Server (NTRS)

Mueller, Arnold W.; Smith, Charles D.; Lemasurier, Philip

1990-01-01

During the design of a helicopter, the weight, engine, rotor speed, and rotor geometry are given significant attention when considering the specific operations for which the helicopter will be used. However, the noise radiated from the helicopter and its relationship to the design variables is currently not well modeled with only a limited set of full-scale field test data to study. In general, limited field data have shown that reduced main-rotor advancing blade-tip Mach numbers result in reduced far-field noise levels. The status of a recent helicopter noise research project is reviewed. It is designed to provide flight experimental data which may be used to further understand helicopter main-rotor advancing blade-tip Mach number effects on far-field acoustic levels. Preliminary results are presented relative to tests conducted with a Sikorsky S-76A helicopter operating with both the rotor speed and the flight speed as the control variable. The rotor speed was operated within the range of 107 to 90 percent NR at nominal forward speeds of 35, 100, and 155 knots.

Ares I Aerodynamic Testing at the Boeing Polysonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Pinier, Jeremy T.; Niskey, Charles J.; Hanke, Jeremy L.; Tomek, William G.

2011-01-01

Throughout three full design analysis cycles, the Ares I project within the Constellation program has consistently relied on the Boeing Polysonic Wind Tunnel (PSWT) for aerodynamic testing of the subsonic, transonic and supersonic portions of the atmospheric flight envelope (Mach=0.5 to 4.5). Each design cycle required the development of aerodynamic databases for the 6 degree-of-freedom (DOF) forces and moments, as well as distributed line-loads databases covering the full range of Mach number, total angle-of-attack, and aerodynamic roll angle. The high fidelity data collected in this facility has been consistent with the data collected in NASA Langley s Unitary Plan Wind Tunnel (UPWT) at the overlapping condition ofMach=1.6. Much insight into the aerodynamic behavior of the launch vehicle during all phases of flight was gained through wind tunnel testing. Important knowledge pertaining to slender launch vehicle aerodynamics in particular was accumulated. In conducting these wind tunnel tests and developing experimental aerodynamic databases, some challenges were encountered and are reported as lessons learned in this paper for the benefit of future crew launch vehicle aerodynamic developments.

Hyper-X Vehicle Model - Side View

NASA Technical Reports Server (NTRS)

1996-01-01

A side-view of an early desk-top model of NASA's X-43A 'Hyper-X,' or Hypersonic Experimental Vehicle, which has been developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

Hyper-X Vehicle Model - Front View

NASA Technical Reports Server (NTRS)

1996-01-01

A front view of an early desk-top model of NASA's X-43A 'Hyper-X,' or Hypersonic Experimental Vehicle, which has been developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

X-43A Vehicle During Ground Testing

NASA Technical Reports Server (NTRS)

1999-01-01

The X-43A Hypersonic Experimental Vehicle, or 'Hyper-X' is seen here undergoing ground testing at NASA's Dryden Flight Research Center, Edwards, California. The X-43A was developed to research a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

Hyper-X Vehicle Model - Side View

NASA Technical Reports Server (NTRS)

1996-01-01

Sleek lines are apparent in this side-view of an early desk-top model of NASA's X-43A 'Hyper-X,' or Hypersonic Experimental Vehicle, which has been developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

Hyper-X Vehicle Model - Top Rear View

NASA Technical Reports Server (NTRS)

1996-01-01

This aft-quarter model view of NASA's X-43A 'Hyper-X' or Hypersonic Experimental Vehicle shows its sleek, geometric design. The X-43A was developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.

Hyper-X Vehicle Model - Top Front View

NASA Technical Reports Server (NTRS)

1996-01-01

A top front view of an early desk-top model of NASA's X-43A 'Hyper-X,' or Hypersonic Experimental Vehicle, developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration. The body of the vehicle itself forms critical elements of the engine. The forebody acts as part of the intake for airflow and the aft section serves as the nozzle. The X-43A vehicles were manufactured by Micro Craft, Inc., Tullahoma, Tennessee. Orbital Sciences Corporation, Chandler, Arizona, built the Pegasus rocket booster used to launch the X-43 vehicles. For the Dryden research flights, the Pegasus rocket booster and attached X-43 will be air launched by Dryden's B-52 'Mothership.' After release from the B-52, the booster will accelerate the X-43A vehicle to the established test conditions (Mach 7 to 10) at an altitude of approximately 100,000 feet where the X-43 will separate from the booster and fly under its own power and preprogrammed control.