Science.gov

Sample records for 9-by 7-foot supersonic

  1. Supersonic Retropropulsion Experimental Results from the NASA Ames 9- x 7-Foot Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Berry, Scott A.; Rhode, Matthew N.; Edquist, Karl T.

    2012-01-01

    Supersonic retropropulsion was experimentally examined in the Ames Research Center 9x7-Foot Supersonic Wind Tunnel at Mach 1.8 and 2.4. The experimental model, previously designed for and tested in the Langley Research Center Unitary Plan Wind Tunnel at Mach 2.4, 3.5 and 4.6, was a 5-in diameter 70-deg sphere-cone forebody with a 9.55-in long cylindrical aftbody. The forebody was designed to accommodate up to four 4:1 area ratio nozzles, one on the model centerline and the other three on the half radius spaced 120-deg apart. Surface pressure and flow visualization were the primary measurements, including high-speed data to investigate the dynamics of the interactions between the bow and nozzle shocks. Three blowing configurations were tested with thrust coefficients up to 10 and angles of attack up to 20-deg. Preliminary results and observations from the test are provided

  2. Results of a FRSI material test under Space Shuttle ascent conditions in the Ames Research Center 9x7 foot supersonic wind tunnel (OS13). Space Shuttle aerothermodynamic data report

    NASA Technical Reports Server (NTRS)

    Lemoine, P. L.; Collette, J. G. R.

    1992-01-01

    A test was conducted in the NASA/ARC 9 x 7 foot supersonic wind tunnel to verify the integrity of Felt Reusable Surface Insulation (FRSI) material in a panel flutter environment. A FRSI sample panel was subjected to the shocks, pressure gradients, and turbulence characteristics encountered at dynamic pressure 1.5 times the 3(sigma) dispersed trajectory flight conditions of the Space Shuttle. Static and fluctuating pressure data were obtained for Mach numbers ranging from 1.55 to 2.5 with dynamic pressures of 625 to 1250 psf. The FRSI panel suffered no appreciable damage as a result of the test.

  3. Results of a FRSI material test under Space Shuttle ascent conditions in the Ames Research Center 9x7 foot supersonic wind tunnel (OS13). Space Shuttle aerothermodynamic data report

    NASA Astrophysics Data System (ADS)

    Lemoine, P. L.; Collette, J. G. R.

    1992-07-01

    A test was conducted in the NASA/ARC 9 x 7 foot supersonic wind tunnel to verify the integrity of Felt Reusable Surface Insulation (FRSI) material in a panel flutter environment. A FRSI sample panel was subjected to the shocks, pressure gradients, and turbulence characteristics encountered at dynamic pressure 1.5 times the 3(sigma) dispersed trajectory flight conditions of the Space Shuttle. Static and fluctuating pressure data were obtained for Mach numbers ranging from 1.55 to 2.5 with dynamic pressures of 625 to 1250 psf. The FRSI panel suffered no appreciable damage as a result of the test.

  4. Space Shuttle AFRSI OMS pod environment test using model 81-0 test fixture in the Ames Research Center 9x7-foot supersonic wind tunnel (OS-314A/B/C)

    NASA Technical Reports Server (NTRS)

    Collette, J. G. R.

    1984-01-01

    A test was conducted in the NASA/Ames Research Center 9x7-foot Supersonic Wind Tunnel to help resolve an anomaly that developed during the STS-6 orbiter flight wherein sections of the Advanced Flexible Reusable Surface Insulation (AFRSI) covering the OMS pods suffered some damage. A one-third scale two-dimensional shell structure model of an OMS pod cross-section was employed to support the test articles. These consisted of 15 AFRSI blanket panels form-fitted over the shell structures for exposure to simulated flight conditions. Of six baseline blankets, two were treated with special surface coatings. Two other panels were configured with AFRSI sections removed from the OV099 orbiter vehicle after the STS-6 flight. Seven additional specimens incorporated alternative designs and repairs. Following a series of surface pressure calibration runs, the specimens were exposed to simulated ascent and entry dynamic pressure profiles. Entry conditions included the use of a vortex generator to evaluate the effect of shed vortices on the AFRSI located in the area of concern.

  5. Model Deformation Measurements of Sonic Boom Models in the NASA Ames 9- by 7-Ft Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Schairer, Edward T.; Kushner, Laura K.; Garbeff, Theodore J.; Heineck, James T.

    2015-01-01

    The deformations of two sonic-boom models were measured by stereo photogrammetry during tests in the 9- by 7-Ft Supersonic Wind Tunnel at NASA Ames Research Center. The models were geometrically similar but one was 2.75 times as large as the other. Deformation measurements were made by simultaneously imaging the upper surfaces of the models from two directions by calibrated cameras that were mounted behind windows of the test section. Bending and twist were measured at discrete points using conventional circular targets that had been marked along the leading and trailing edges of the wings and tails. In addition, continuous distributions of bending and twist were measured from ink speckles that had been applied to the upper surfaces of the model. Measurements were made at wind-on (M = 1.6) and wind-off conditions over a range of angles of attack between 2.5 deg. and 5.0 deg. At each condition, model deformation was determined by comparing the wind-off and wind-on coordinates of each measurement point after transforming the coordinates to reference coordinates tied to the model. The necessary transformations were determined by measuring the positions of a set of targets on the rigid center-body of the models whose model-axes coordinates were known. Smoothly varying bending and twist measurements were obtained at all conditions. Bending displacements increased in proportion to the square of the distance to the centerline. Maximum deflection of the wingtip of the larger model was about 5 mm (2% of the semispan) and that of the smaller model was 0.9 mm (1% of the semispan). The change in wing twist due to bending increased in direct proportion to distance from the centerline and reached a (absolute) maximum of about -1? at the highest angle of attack for both models. The measurements easily resolved bending displacements as small as 0.05 mm and bending-induced changes in twist as small as 0.05 deg.

  6. Investigations on an 0.030-scale space shuttle vehicle configuration 140A/B orbiter model in the Ames Research Center unitary plan 8 by 7-foot supersonic wind tunnel (0A53C)

    NASA Technical Reports Server (NTRS)

    Nichols, M. E.

    1974-01-01

    A wind tunnel test was conducted of an 0.030 scale model of the space shuttle orbiter in a supersonic wind tunnel. Tests were conducted at Mach numbers of 2.5, 3.0, and 3.5. Reynolds numbers ranged from 0.75 million per foot to 4.00 million per foot. The objective of the test was to establish and verify longitudinal and lateral-directional aerodynamic performance, stability, and control characteristics for the configuration 140 A/B SSV Orbiter. Six-component force and moment data, base and cavity pressures, body-flap, elevon, speedbrake, and rudder hinge moments, and vertical tail forces and moments were measured.

  7. Flow quality studies of the NASA Lewis Research Center 8- by 6-foot supersonic/9- by 15-foot low speed wind tunnel

    NASA Technical Reports Server (NTRS)

    Arrington, E. Allen; Pickett, Mark T.

    1992-01-01

    A series of studies were conducted to determine the existing flow quality in the NASA Lewis 8 by 6 Foot Supersonic/9 by 15 Foot Low speed Wind Tunnel. The information gathered from these studies was used to determine the types and designs of flow manipulators which can be installed to improve overall tunnel flow quality and efficiency. Such manipulators include honeycomb flow straighteners, turbulence reduction screens, corner turning vanes, and acoustic treatments. The types of measurements, instrumentation, and results obtained from experiments conducted at several locations throughout the tunnel loop are described.

  8. Drive Motor Improved for 8- by 6-Foot Supersonic Wind Tunnel/9- by 15-Foot Low-Speed Wind Tunnel Complex

    NASA Technical Reports Server (NTRS)

    2005-01-01

    An operational change made recently in the drive motor system for the 8- by 6-Foot Supersonic Wind Tunnel (8x6 SWT)/9- by 15-Foot Low-Speed Wind Tunnel (9x15 LSWT) complex resulted in dramatic power savings and expanded operating range. The 8x6 SWT/9x15 LSWT complex offers a unique combination of wind tunnel conditions for both high- and low-speed testing. Prior to the work discussed in this article, the 8- by 6-ft test section offered airflows ranging from Mach 0.36 to 2.0. Subsonic testing was done in the 9-ft high, 15-ft wide test area in the return leg of the facility. The air speed in this test section can range from 0 to 175 mph (Mach 0.23). In the past, we varied the air speed by using a combination of the compressor speed and the position of the tunnel flow-control doors. When very slow speeds were required in the 9x15 LSWT, these large tunnel flow control doors might be very nearly full open, bleeding off large quantities of air, even with the drive system operating at its previous minimum speed of about 510 rpm. Power drawn during this mode of operation varied between 15 and 18 MW/hr, but clearly much of this power was not being used to provide air that would be used for testing in the test section. The air exiting these large doors represented wasted power. Early this year, the facility's tunnel drive system was run on one motor instead of three to see if lower drive speeds could be achieved that would, in turn, result in large power savings because unnecessary air would not be blown out of the flow-control doors unnecessarily. In addition, if the drive could be run slower, then slower speeds would also be possible in the 8x6 SWT test section as an added benefit. Results of the first tests performed early last year showed that in fact the drive, when operating on only one motor, actually reached a steady-state speed of only 337 rpm and drew an amazingly small 6 MW/hr of electrical power. During daytime operation of the drive, this meant that it would be

  9. Preliminary Computational Study for Future Tests in the NASA Ames 9 foot' x 7 foot Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Pearl, Jason M.; Carter, Melissa B.; Elmiligui, Alaa A.; WInski, Courtney S.; Nayani, Sudheer N.

    2016-01-01

    The NASA Advanced Air Vehicles Program, Commercial Supersonics Technology Project seeks to advance tools and techniques to make over-land supersonic flight feasible. In this study, preliminary computational results are presented for future tests in the NASA Ames 9 foot x 7 foot supersonic wind tunnel to be conducted in early 2016. Shock-plume interactions and their effect on pressure signature are examined for six model geometries. Near- field pressure signatures are assessed using the CFD code USM3D to model the proposed test geometries in free-air. Additionally, results obtained using the commercial grid generation software Pointwise Reigistered Trademark are compared to results using VGRID, the NASA Langley Research Center in-house mesh generation program.

  10. Design and testing of an oblique all-wing supersonic transport

    NASA Technical Reports Server (NTRS)

    Lee, Christopher A.

    1994-01-01

    This report describes the preliminary design of an Oblique All-Wing (OAW) supersonic transport and a corresponding wind-tunnel model that was tested in the NASA Ames 9- by 7-Foot supersonic wind tunnel. The main goal was the determination of the cruise performance (lift/drag ratio) of a realistically configured OAW. To achieve an acceptable level of realism, it was necessary to consider many issues of design practicality such as the need for a viable propulsion system, adequate control surfaces, landing gear, provisions for 450 passengers, and fuel to fly 5,000 nautical miles. The aircraft had to be stable, structurally sound, and needed to fit into airports across the world. Support was directed primarily towards integration of the propulsion system, however, there were notable contributions to many aspects of the configuration design, wind tunnel model, and wind tunnel test.

  11. Supersonic compressor

    DOEpatents

    Lawlor, Shawn P.; Novaresi, Mark A.; Cornelius, Charles C.

    2008-02-26

    A gas compressor based on the use of a driven rotor having an axially oriented compression ramp traveling at a local supersonic inlet velocity (based on the combination of inlet gas velocity and tangential speed of the ramp) which forms a supersonic shockwave axially, between adjacent strakes. In using this method to compress inlet gas, the supersonic compressor efficiently achieves high compression ratios while utilizing a compact, stabilized gasdynamic flow path. Operated at supersonic speeds, the inlet stabilizes an oblique/normal shock system in the gasdyanamic flow path formed between the gas compression ramp on a strake, the shock capture lip on the adjacent strake, and captures the resultant pressure within the stationary external housing while providing a diffuser downstream of the compression ramp.

  12. Supersonic compressor

    DOEpatents

    Roberts, II, William Byron; Lawlor, Shawn P.; Breidenthal, Robert E.

    2016-04-12

    A supersonic compressor including a rotor to deliver a gas at supersonic conditions to a diffuser. The diffuser includes a plurality of aerodynamic ducts that have converging and diverging portions, for deceleration of gas to subsonic conditions and then for expansion of subsonic gas, to change kinetic energy of the gas to static pressure. The aerodynamic ducts include vortex generating structures for controlling boundary layer, and structures for changing the effective contraction ratio to enable starting even when the aerodynamic ducts are designed for high pressure ratios, and structures for boundary layer control. In an embodiment, aerodynamic ducts are provided having an aspect ratio of in excess of two to one, when viewed in cross-section orthogonal to flow direction at an entrance to the aerodynamic duct.

  13. Supersonic combustor modeling

    NASA Technical Reports Server (NTRS)

    Riggins, David W.; Drummond, J. Philip; Carpenter, Mark H.

    1990-01-01

    The physical phenomena involved when a supersonic flow undergoes chemical reaction are discussed. Detailed physical models of convective and diffusive mixing, and finite rate chemical reaction in supersonic flow are presented. Numerical algorithms used to solve the equations governing these processes are introduced. Computer programs using these algorithms are used to analyze the structure of the reacting mixing layer. It is concluded that, as in subsonic flow, exothermic heat release in unconfined supersonic flows retards fuel/air mixing. Non mixing is shown to be a potential problem in reducing the efficiency of supersonic as well as subsonic combustion. Techniques for enhancing fuel/air mixing and combustion are described.

  14. Spectroscopy with Supersonic Jets.

    ERIC Educational Resources Information Center

    Skinner, Anne R.; Chandler, Dean W.

    1980-01-01

    Discusses a new technique that enables spectroscopists to study gas phase molecules at temperatures below 1 K, without traditional cryogenic apparatus. This technique uses supersonic jets as samples for gas molecular spectroscopy. Highlighted are points in the theory of supersonic flow which are important for applications in molecular…

  15. Supersonic Elliptical Ramp Inlet

    NASA Technical Reports Server (NTRS)

    Adamson, Eric E. (Inventor); Fink, Lawrence E. (Inventor); Fugal, Spencer R. (Inventor)

    2016-01-01

    A supersonic inlet includes a supersonic section including a cowl which is at least partially elliptical, a ramp disposed within the cowl, and a flow inlet disposed between the cowl and the ramp. The ramp may also be at least partially elliptical.

  16. A corporate supersonic transport

    NASA Technical Reports Server (NTRS)

    Greene, Randall; Seebass, Richard

    1996-01-01

    This talk address the market and technology for a corporate supersonic transport. It describes a candidate configuration. There seems to be a sufficient market for such an aircraft, even if restricted to supersonic operation over water. The candidate configuration's sonic boom overpressure may be small enough to allow overland operation as well.

  17. Supersonic Cruise Technology

    NASA Technical Reports Server (NTRS)

    Mclean, F. Edward

    1985-01-01

    The history and status of supersonic cruise research is covered. The early research efforts of the National Advisory Committee for Aeronautics and efforts during the B-70 and SST phase are included. Technological progress made during the NASA Supersonic Cruise Research and Variable Cycle Engine programs are presented. While emphasis is on NASA's contributions to supersonic cruise research in the U.S., also noted are developments in England, France, and Russia. Written in nontechnical language, this book presents the most critical technology issues and research findings.

  18. Supersonic laser propulsion.

    PubMed

    Rezunkov, Yurii; Schmidt, Alexander

    2014-11-01

    To produce supersonic laser propulsion, a new technique based on the interaction of a laser-ablated jet with supersonic gas flow in a nozzle is proposed. It is shown that such parameters of the jet, such as gas-plasma pressure and temperature in the ablation region as well as the mass consumption rate of the ablated solid propellant, are characteristic in this respect. The results of numerical simulations of the supersonic laser propulsion are presented for two types of nozzle configuration. The feasibility to achieve the momentum coupling coefficient of C(m)∼10(-3) N/W is shown.

  19. Supersonic Combustion Ramjet Research

    DTIC Science & Technology

    2012-08-01

    4.2 Ignition, Flameholding, and Flame Propagation in Supersonic Flows ......................... 18 4.2.1 Plasma -Assisted Ignition and Flameholding...high- speed flows), plasma -assisted combustion, flameholding (particularly in a high-speed flow), and development and application of diagnostic...Flameholding, and Flame Propagation in Supersonic Flows 4.2.1 Plasma -Assisted Ignition and Flameholding Key questions that have guided this

  20. Supersonic gas compressor

    DOEpatents

    Lawlor, Shawn P.; Novaresi, Mark A.; Cornelius, Charles C.

    2007-11-13

    A gas compressor based on the use of a driven rotor having a compression ramp traveling at a local supersonic inlet velocity (based on the combination of inlet gas velocity and tangential speed of the ramp) which compresses inlet gas against a stationary sidewall. In using this method to compress inlet gas, the supersonic compressor efficiently achieves high compression ratios while utilizing a compact, stabilized gasdynamic flow path. Operated at supersonic speeds, the inlet stabilizes an oblique/normal shock system in the gasdyanamic flow path formed between the rim of the rotor, the strakes, and a stationary external housing. Part load efficiency is enhanced by the use of a pre-swirl compressor, and using a bypass stream to bleed a portion of the intermediate pressure gas after passing through the pre-swirl compressor back to the inlet of the pre-swirl compressor. Inlet guide vanes to the compression ramp enhance overall efficiency.

  1. Supersonic nonlinear potential analysis

    NASA Technical Reports Server (NTRS)

    Siclari, M. J.

    1984-01-01

    The NCOREL computer code was established to compute supersonic flow fields of wings and bodies. The method encompasses an implicit finite difference transonic relaxation method to solve the full potential equation in a spherical coordinate system. Two basic topic to broaden the applicability and usefulness of the present method which is encompassed within the computer code NCOREL for the treatment of supersonic flow problems were studied. The first topic is that of computing efficiency. Accelerated schemes are in use for transonic flow problems. One such scheme is the approximate factorization (AF) method and an AF scheme to the supersonic flow problem is developed. The second topic is the computation of wake flows. The proper modeling of wake flows is important for multicomponent configurations such as wing-body and multiple lifting surfaces where the wake of one lifting surface has a pronounced effect on a downstream body or other lifting surfaces.

  2. Supersonic MHD generator system

    SciTech Connect

    Rahman, M.A.

    1983-11-29

    An improved MHD electrical power generating system of the type having a MHD topping cycle and a steam generating bottoming cycle is disclosed. The system typically includes a combustion system, a conventional MHD generator and a first diffuser radiant boiler. The improvement comprises a first supersonic MHD generator and ramjet engine configuration operatively connected in series with each other and with the conventional MHD generator. The first supersonic MHD generator and ramjet engine configuration increase the power output and improve the operating efficiency of the electrical generating system. A diffuser system is also disclosed which is in fluid communication with the supersonic MHD generator and the ramjet engine for collecting bypass plasma gas to be used for heating a second radiant boiler adapted for powering a steam turbine generator.

  3. The Trojan. [supersonic transport

    NASA Technical Reports Server (NTRS)

    1992-01-01

    The Trojan is the culmination of thousands of engineering person-hours by the Cones of Silence Design Team. The goal was to design an economically and technologically viable supersonic transport. The Trojan is the embodiment of the latest engineering tools and technology necessary for such an advanced aircraft. The efficient design of the Trojan allows for supersonic cruise of Mach 2.0 for 5,200 nautical miles, carrying 250 passengers. The per aircraft price is placed at $200 million, making the Trojan a very realistic solution for tomorrows transportation needs. The following is a detailed study of the driving factors that determined the Trojan's super design.

  4. Infinitesimal Conical Supersonic Flow

    NASA Technical Reports Server (NTRS)

    Busemann, Adolf

    1947-01-01

    The calculation of infinitesimal conical supersonic flow has been applied first to the simplest examples that have also been calculated in another way. Except for the discovery of a miscalculation in an older report, there was found the expected conformity. The new method of calculation is limited more definitely to the conical case.

  5. Supersonic Leading Edge Receptivity

    NASA Technical Reports Server (NTRS)

    Maslov, Anatoly A.

    1998-01-01

    This paper describes experimental studies of leading edge boundary layer receptivity for imposed stream disturbances. Studies were conducted in the supersonic T-325 facility at ITAM and include data for both sharp and blunt leading edges. The data are in agreement with existing theory and should provide guidance for the development of more complete theories and numerical computations of this phenomena.

  6. Supersonic laminar flow control research

    NASA Technical Reports Server (NTRS)

    Lo, Ching F.

    1994-01-01

    The objective of the research is to understand supersonic laminar flow stability, transition, and active control. Some prediction techniques will be developed or modified to analyze laminar flow stability. The effects of supersonic laminar flow with distributed heating and cooling on active control will be studied. The primary tasks of the research applying to the NASA/Ames Proof of Concept (POC) Supersonic Wind Tunnel and Laminar Flow Supersonic Wind Tunnel (LFSWT) nozzle design with laminar flow control are as follows: (1) predictions of supersonic laminar boundary layer stability and transition, (2) effects of wall heating and cooling for supersonic laminar flow control, and (3) performance evaluation of POC and LFSWT nozzles design with wall heating and cooling effects applying at different locations and various length.

  7. Supersonic laminar flow control research

    NASA Technical Reports Server (NTRS)

    Lo, C. F.

    1994-01-01

    The objective of the research is to understand supersonic laminar flow stability, transition and active control. Some prediction techniques will be developed or modified to analyze laminar stability. The effects of supersonic laminar flow with distributed heating and cooling on active control will be studied. The primary tasks of the research applying to the NASA/Ames POC and LFSWT's nozzle design with laminar flow control are as follows: (1) supersonic laminar boundary layer stability and transition prediction; (2) effects of heating and cooling for supersonic laminar flow control; and (3) POC and LFSWT nozzle design with heating and cooling effects combining wall contour and length changes.

  8. Supersonic Wind Tunnel Test Section

    NASA Technical Reports Server (NTRS)

    1957-01-01

    8ft x 6ft Supersonic Wind Tunnel Test-Section showing changes made in Stainless Steel walls with 17 inch inlet model installation. The model is the ACN Nozzle model used for aircraft engines. The Supersonic Wind Tunnel is located in the Lewis Flight Propulsion Laboratory, now John H. Glenn Research Center

  9. The Edge supersonic transport

    NASA Technical Reports Server (NTRS)

    Agosta, Roxana; Bilbija, Dushan; Deutsch, Marc; Gallant, David; Rose, Don; Shreve, Gene; Smario, David; Suffredini, Brian

    1992-01-01

    As intercontinental business and tourism volumes continue their rapid expansion, the need to reduce travel times becomes increasingly acute. The Edge Supersonic Transport Aircraft is designed to meet this demand by the year 2015. With a maximum range of 5750 nm, a payload of 294 passengers and a cruising speed of M = 2.4, The Edge will cut current international flight durations in half, while maintaining competitive first class, business class, and economy class comfort levels. Moreover, this transport will render a minimal impact upon the environment, and will meet all Federal Aviation Administration Part 36, Stage III noise requirements. The cornerstone of The Edge's superior flight performance is its aerodynamically efficient, dual-configuration design incorporating variable-geometry wingtips. This arrangement combines the benefits of a high aspect ratio wing at takeoff and low cruising speeds with the high performance of an arrow-wing in supersonic cruise. And while the structural weight concerns relating to swinging wingtips are substantial, The Edge looks to ever-advancing material technologies to further increase its viability. Heeding well the lessons of the past, The Edge design holds economic feasibility as its primary focus. Therefore, in addition to its inherently superior aerodynamic performance, The Edge uses a lightweight, largely windowless configuration, relying on a synthetic vision system for outside viewing by both pilot and passengers. Additionally, a fly-by-light flight control system is incorporated to address aircraft supersonic cruise instability. The Edge will be produced at an estimated volume of 400 aircraft and will be offered to airlines in 2015 at $167 million per transport (1992 dollars).

  10. Progress in supersonic cruise technology

    NASA Technical Reports Server (NTRS)

    Driver, C.

    1983-01-01

    The Supersonic Cruise Research (SCR) program identified significant improvements in the technology areas of aerodynamics, structures, propulsion, noise reduction, takeoff and landing procedures, and advanced configuration concepts. These improvements, when combined in a large supersonic cruise vehicle, offer a far greater technology advance than generally realized. They offer the promise of an advanced commercial family of aircraft which are environmentally acceptable, have flexible range-payload capability, and are economically viable. These same areas of technology have direct application to smaller advanced military aircraft and to supersonic executive aircraft. Several possible applications will be addressed.

  11. Wind tunnel test of the 0.019 (2A configuration) jet plume space shuttle integrated vehicle in the ARC 9- by 7-foot unitary wind tunnel (IA12B)

    NASA Technical Reports Server (NTRS)

    Hardin, R. B.; Burrows, R. R.

    1974-01-01

    The wind tunnel test of the 0.019 jet plume space shuttle integrated vehicle in the Ames 9 ft by 7 ft unitary wind tunnel was conducted at Mach numbers of 1.55 and 2.0 over a Reynolds number range from 3.5 million to 4.1 million/ft. Data were obtained at angles of attack from minus 8 deg to plus 8 deg at 0 deg sideslip and at angles of sideslip from minus 9 deg to plus 8 deg at 0 deg angle of attack. The basic configuration tested was the 2A vehicle with the orbiter at 0 deg angle of incidence with respect to the external tank. The other deviations to the 2A configuration were the solid rocket motor shrouds, which were designed to vehicle '3' lines, and the tank nose, which consisted of the retro-package being removed and replaced by a 16.5 inch full scale radius nose.

  12. Supersonic-Spray Cleaner

    NASA Technical Reports Server (NTRS)

    Caimi, Raoul E. B.; Lin, Feng-Nan; Thaxton, Eric A.

    1995-01-01

    Spraying system for cleaning mechanical components uses less liquid and operates at pressures significantly lower. Liquid currently used is water. Designed to replace chlorofluorocarbon (CFC) solvent-based cleaning and cleanliness verification methods. Consists of spray head containing supersonic converging/diverging nozzles, source of gas at regulated pressure, pressurized liquid tank, and various hoses, fittings, valves, and gauges. Parameters of nozzles set so any of large variety of liquids and gases combined in desired ratio and rate of flow. Size and number of nozzles varied so system built in configurations ranging from small hand-held spray heads to large multinozzle cleaners. Also used to verify part adequately cleaned. Runoff liquid from spray directed at part collected. Liquid analyzed for presence of contaminants, and part recleaned if necessary.

  13. Supersonic Pulsed Injection

    NASA Technical Reports Server (NTRS)

    Cutler, A. D.; Harding, G. C.; Diskin, G. S.

    2001-01-01

    An injector has been developed to provide high-speed high-frequency (order 10 kHz) pulsed a supersonic crossflow. The injector nozzle is formed between the fixed internal surface of the nozzle and a freely rotating three- or four-sided wheel embedded within the device. Flow-induced rotation of the wheel causes the nozzle throat to open and close at a frequency proportional to the speed of sound of the injected gas. Measurements of frequency and mass flow rate as a function of supply pressure are discussed for various injector designs. Preliminary results are presented for wall-normal injection of helium into a Mach-2 ducted airflow. The data include schlieren images in the injectant plume in a plane normal to the flow, downstream of injection.

  14. Supersonics--Airport Noise

    NASA Technical Reports Server (NTRS)

    Bridges, James

    2007-01-01

    At this, the first year-end meeting of the Fundamental Aeronautics Program, an overview of the Airport Noise discipline of the Supersonics Project leads the presentation of technical plans and achievements in this area of the Project. The overview starts by defining the Technical Challenges targeted by Airport Noise efforts, and the Approaches planned to meet these challenges. These are fleshed out in Elements, namely Prediction, Diagnostics, and Engineering, and broken down into Tasks. The Tasks level is where individual researchers' work is defined and from whence the technical presentations to follow this presentation come. This overview also presents the Milestones accomplished to date and to be completed in the next year. Finally, the NASA Research Announcement cooperative agreement activities are covered and tied to the Tasks and Milestones.

  15. Study of the supersonic propeller

    NASA Technical Reports Server (NTRS)

    Fabri, Jean; Siestrunck, Raymond

    1953-01-01

    In this paper a propeller having all sections operating at supersonic speeds is designated a supersonic propeller regardless of flight speed. Analyses assume subsonic flight speeds but very high rotational speeds. A very elementary analysis of the efficiency of a jet-propeller system is presented. A propeller analysis based on conventional vortex blade element theory is presented and reduced to a single point method which leads to an expression for optimum advance ratio in terms of hub-tip diameter ratio and airfoil fineness ratio. An expression for propeller efficiency in terms of advance ratio, hub-tip diameter ratio, and airfoil thickness ratio is also presented. Use is made of theoretical airfoil characteristics at supersonic speeds. A study of blade section interference, blade shock and expansion fields, at supersonic section speeds is presented. An example taken indicates that an efficiency of seventy percent can be obtained with a propeller having a tip Mach number of 2.3.

  16. Supersonic laminar flow control research

    NASA Technical Reports Server (NTRS)

    Lo, Ching F.; Wiberg, Clark G.

    1995-01-01

    The objective is to understand supersonic laminar flow stability, transition, and active control. Some prediction techniques will be developed or modified to analyze laminar flow stability. The effects of distributed heating and cooling as an active boundary layer control technique will be studied. The primary tasks of the research apply to the NASA/Ames Proof of Concept (PoC) and Laminar Flow Supersonic Wind Tunnel's (LFSWT's) nozzle design with laminar flow control and are listed as follows: (1) predictions of supersonic laminar boundary layer stability and transition, (2) effects of wall heating and cooling on supersonic laminar flow control, (3) performance evaluation of the PoC and LFSWT nozzle designs with wall heating and cooling applied at different locations and various lengths, and (4) effects of a conducted versus pulse wall temperature distribution for the LFSWT.

  17. Supersonic reacting internal flow fields

    NASA Technical Reports Server (NTRS)

    Drummond, J. Philip

    1989-01-01

    The national program to develop a trans-atmospheric vehicle has kindled a renewed interest in the modeling of supersonic reacting flows. A supersonic combustion ramjet, or scramjet, has been proposed to provide the propulsion system for this vehicle. The development of computational techniques for modeling supersonic reacting flow fields, and the application of these techniques to an increasingly difficult set of combustion problems are studied. Since the scramjet problem has been largely responsible for motivating this computational work, a brief history is given of hypersonic vehicles and their propulsion systems. A discussion is also given of some early modeling efforts applied to high speed reacting flows. Current activities to develop accurate and efficient algorithms and improved physical models for modeling supersonic combustion is then discussed. Some new problems where computer codes based on these algorithms and models are being applied are described.

  18. Advanced supersonic cruise aircraft technology

    NASA Technical Reports Server (NTRS)

    Baber, H. T., Jr.; Driver, C.

    1977-01-01

    A multidiscipline approach is taken to the application of the latest technology to supersonic cruise aircraft concept definition, and current problem areas are identified. Particular attention is given to the performance of the AST-100 advanced supersonic cruise vehicle with emphasis on aerodynamic characteristics, noise and chemical emission, and mission analysis. A recently developed aircraft sizing and performance computer program was used to determine allowable wing loading and takeoff gross weight sensitivity to structural weight reduction.

  19. Investigations for Supersonic Transports at Transonic and Supersonic Conditions

    NASA Technical Reports Server (NTRS)

    Rivers, S. Melissa B.; Owens, Lewis R.; Wahls, Richard A.

    2007-01-01

    Several computational studies were conducted as part of NASA s High Speed Research Program. Results of turbulence model comparisons from two studies on supersonic transport configurations performed during the NASA High-Speed Research program are given. The effects of grid topology and the representation of the actual wind tunnel model geometry are also investigated. Results are presented for both transonic conditions at Mach 0.90 and supersonic conditions at Mach 2.48. A feature of these two studies was the availability of higher Reynolds number wind tunnel data with which to compare the computational results. The transonic wind tunnel data was obtained in the National Transonic Facility at NASA Langley, and the supersonic data was obtained in the Boeing Polysonic Wind Tunnel. The computational data was acquired using a state of the art Navier-Stokes flow solver with a wide range of turbulence models implemented. The results show that the computed forces compare reasonably well with the experimental data, with the Baldwin-Lomax with Degani-Schiff modifications and the Baldwin-Barth models showing the best agreement for the transonic conditions and the Spalart-Allmaras model showing the best agreement for the supersonic conditions. The transonic results were more sensitive to the choice of turbulence model than were the supersonic results.

  20. Tesseract supersonic business transport

    NASA Technical Reports Server (NTRS)

    Reshotko, Eli; Garbinski, Gary; Fellenstein, James; Botting, Mary; Hooper, Joan; Ryan, Michael; Struk, Peter; Taggart, Ben; Taillon, Maggie; Warzynski, Gary

    1992-01-01

    This year, the senior level Aerospace Design class at Case Western Reserve University developed a conceptual design of a supersonic business transport. Due to the growing trade between Asia and the United States, a transpacific range was chosen for the aircraft. A Mach number of 2.2 was chosen, too, because it provides reasonable block times and allows the use of a large range of materials without a need for active cooling. A payload of 2,500 lbs. was assumed corresponding to a complement of nine passengers and crew, plus some light cargo. With these general requirements set, the class was broken down into three groups. The aerodynamics of the aircraft were the responsibility of the first group. The second developed the propulsion system. The efforts of both the aerodynamics and propulsion groups were monitored and reviewed for weight considerations and structural feasibility by the third group. Integration of the design required considerable interaction between the groups in the final stages. The fuselage length of the final conceptual design was 107.0 ft, while the diameter of the fuselage was 7.6 ft. The delta wing design consisted of an aspect ratio of 1.9 with a wing span of 47.75 ft and mid-chord length of 61.0 ft. A SNECMA MCV 99 variable-cycle engine design was chosen for this aircraft.

  1. Tesseract: Supersonic business transport

    NASA Technical Reports Server (NTRS)

    Reshotko, Eli; Garbinski, Gary

    1992-01-01

    This year, the senior level Aerospace Design class at Case Western Reserve University developed a conceptual design of a supersonic business transport. Due to the growing trade between Asia and the United States, a transpacific range has been chosen for the aircraft. A Mach number of 2.2 was chosen too because it provides reasonable block times and allows the use of a large range of materials without a need for active cooling. A payload of 2500 lbs. has been assumed corresponding to a complement of nine (passengers and crew) plus some light cargo. With these general requirements set, the class was broken down into three groups. The aerodynamics of the aircraft were the responsibility of the first group. The second developed the propulsion system. The efforts of both the aerodynamics and propulsion groups were monitored and reviewed for weight considerations and structural feasibility by the third group. Integration of the design required considerable interaction between the groups in the final stages. The fuselage length of the final conceptual design was 107.0 ft. while the diameter of the fuselage was 7.6 ft. The delta wing design consisted of an aspect ratio of 1.9 with a wing span of 47.75 ft and midcord length of 61.0 ft. A SNEMCA MCV 99 variable-cycle engine design was chosen for this aircraft.

  2. Diagnostic experiments on supersonic jet noise

    NASA Technical Reports Server (NTRS)

    Mclaughlin, D. K.

    1981-01-01

    Experiments and computations on the flowfield and radiated noise of supersonic model jets are discussed. The shock associated noise produced by large scale instabilities in underexpanded supersonic jets, the nonlinear propagation distortion phenomenon in the noise radiated by supersonic model jets, and computations of instability evolution and radiated noise using the LSNOIS computer code are addressed.

  3. Supersonic coal water slurry fuel atomizer

    DOEpatents

    Becker, Frederick E.; Smolensky, Leo A.; Balsavich, John

    1991-01-01

    A supersonic coal water slurry atomizer utilizing supersonic gas velocities to atomize coal water slurry is provided wherein atomization occurs externally of the atomizer. The atomizer has a central tube defining a coal water slurry passageway surrounded by an annular sleeve defining an annular passageway for gas. A converging/diverging section is provided for accelerating gas in the annular passageway to supersonic velocities.

  4. Supersonic Laminar Flow Control Research

    NASA Technical Reports Server (NTRS)

    Lo, Ching F.; Wiberg, Clark G.

    1996-01-01

    The objective of this research is to understand supersonic laminar flow stability, transition and active control. Some prediction techniques will be developed or modified to analyze laminar flow stability. The effects of distributed heating and cooling as an active boundary layer control technique will be studied. The primary tasks of the research apply to the NASA/Ames PoC and LFSWT's nozzle design with laminar flow control and are listed as follows: Predictions of supersonic laminar boundary layer stability and transition; Effects of wall heating and cooling on supersonic laminar flow control on a flat plate; Performance evaluation of the PoC and LFSWT nozzle designs with wall heating and cooling applied at different locations and various lengths; Effects of a conducted-vs-pulse wall temperature distribution for the LFSWT; and Application of wall heating and/or cooling to laminar boundary layer and flow separation control of airfoils and investigation of related active control techniques.

  5. Supersonic laminar flow control research

    NASA Technical Reports Server (NTRS)

    Lo, Ching F.; Wiberg, Clark G.

    1995-01-01

    The objective of this research is to understand supersonic laminar flow stability, transition and active control. Some prediction techniques will be developed or modified to analyze laminar flow stability. The effects of distributed heating and cooling as an active boundary layer control technique will be studied. The primary tasks of the research apply to the NASA/Ames PoC and LFSWT's nozzle design with laminar flow control and are listed as follows: (1) predictions of supersonic laminar boundary layer stability and transition; (2) effects of wall heating and cooling on supersonic laminar flow control; (3) performance evaluation of the PoC and LFSWT nozzle designs with wall heating and cooling applied at different locations and various lengths; and (4) effects of a conducted -vs- pulse wall temperature distribution for the LFSWT.

  6. Supersonic LFC: Challenges and opportunities

    NASA Technical Reports Server (NTRS)

    Powell, Arthur G.

    1992-01-01

    The discussion and viewgraphs on supersonic laminar control are provided. The high fuel fractions required for long range supersonic airplanes give significant leverage to technologies for cruise drag reduction such as laminar flow control (LFC). Fuel burn benefits are further enhanced when sizing effects are considered. These effects may even be powerful enough to reduce airplane production cost over a turbulent baseline. This is an important goal for LFC technology development. The results of aerodynamics studies on the application of LFC technology to the highly swept wings of supersonic airplanes are presented. Important questions of applicability, realistic benefit, and critical application issues, addressed in a NASA-sponsored study conducted by McDonnell Douglas Corporation in 1987-88 are reviewed. Efforts aimed at establishing the feasibility of demonstrating extensive laminarization on the F-16XL-2 airplane are summarized.

  7. Supersonic Dislocation Bursts in Silicon

    DOE PAGES

    Hahn, E. N.; Zhao, S.; Bringa, E. M.; ...

    2016-06-06

    Dislocations are the primary agents of permanent deformation in crystalline solids. Since the theoretical prediction of supersonic dislocations over half a century ago, there is a dearth of experimental evidence supporting their existence. Here we use non-equilibrium molecular dynamics simulations of shocked silicon to reveal transient supersonic partial dislocation motion at approximately 15 km/s, faster than any previous in-silico observation. Homogeneous dislocation nucleation occurs near the shock front and supersonic dislocation motion lasts just fractions of picoseconds before the dislocations catch the shock front and decelerate back to the elastic wave speed. Applying a modified analytical equation for dislocation evolutionmore » we successfully predict a dislocation density of 1.5 x 10(12) cm(-2) within the shocked volume, in agreement with the present simulations and realistic in regards to prior and on-going recovery experiments in silicon.« less

  8. Supersonic Dislocation Bursts in Silicon

    SciTech Connect

    Hahn, E. N.; Zhao, S.; Bringa, E. M.; Meyers, M. A.

    2016-06-06

    Dislocations are the primary agents of permanent deformation in crystalline solids. Since the theoretical prediction of supersonic dislocations over half a century ago, there is a dearth of experimental evidence supporting their existence. Here we use non-equilibrium molecular dynamics simulations of shocked silicon to reveal transient supersonic partial dislocation motion at approximately 15 km/s, faster than any previous in-silico observation. Homogeneous dislocation nucleation occurs near the shock front and supersonic dislocation motion lasts just fractions of picoseconds before the dislocations catch the shock front and decelerate back to the elastic wave speed. Applying a modified analytical equation for dislocation evolution we successfully predict a dislocation density of 1.5 x 10(12) cm(-2) within the shocked volume, in agreement with the present simulations and realistic in regards to prior and on-going recovery experiments in silicon.

  9. Properties of Supersonic Evershed Downflows

    NASA Astrophysics Data System (ADS)

    Pozuelo, S. Esteban; Bellot Rubio, L. R.; de la Cruz Rodríguez, J.

    2016-12-01

    We study supersonic Evershed downflows in a sunspot penumbra by means of high spatial resolution spectropolarimetric data acquired in the Fe i 617.3 nm line with the CRISP instrument at the Swedish 1 m Solar Telescope. Physical observables, such as Dopplergrams calculated from line bisectors and Stokes V zero-crossing wavelengths, and Stokes V maps in the far red-wing, are used to find regions where supersonic Evershed downflows may exist. We retrieve the line-of-sight velocity and the magnetic field vector in these regions using two-component inversions of the observed Stokes profiles with the help of the SIR code. We follow these regions during their lifetime to study their temporal behavior. Finally, we carry out a statistical analysis of the detected supersonic downflows to characterize their physical properties. Supersonic downflows are contained in compact patches moving outward, which are located in the mid- and outer penumbra. They are observed as bright, roundish structures at the outer end of penumbral filaments that resemble penumbral grains. The patches may undergo fragmentations and mergings during their lifetime; some of them are recurrent. Supersonic downflows are associated with strong and rather vertical magnetic fields with a reversed polarity compared to that of the sunspot. Our results suggest that downflows returning back to the solar surface with supersonic velocities are abruptly stopped in dense deep layers and produce a shock. Consequently, this shock enhances the temperature and is detected as a bright grain in the continuum filtergrams, which could explain the existence of outward-moving grains in the mid- and outer penumbra.

  10. Supersonic laminar-flow control

    NASA Technical Reports Server (NTRS)

    Bushnell, Dennis M.; Malik, Mujeeb R.

    1987-01-01

    Detailed, up to date systems studies of the application of laminar flow control (LFC) to various supersonic missions and/or vehicles, both civilian and military, are not yet available. However, various first order looks at the benefits are summarized. The bottom line is that laminar flow control may allow development of a viable second generation SST. This follows from a combination of reduced fuel, structure, and insulation weight permitting operation at higher altitudes, thereby lowering sonic boom along with improving performance. The long stage lengths associated with the emerging economic importance of the Pacific Basin are creating a serious and renewed requirement for such a vehicle. Supersonic LFC techniques are discussed.

  11. Supersonic market and economic analyses

    NASA Technical Reports Server (NTRS)

    Rochte, L. S.

    1980-01-01

    Advanced supersonic transport markets of the free world were projected for the period 1985 to 2004. Passenger traffic volume and airplane range and seat capacity requirements were estimated for Mach 2.2 service by international regional area market areas and by city pairs within and between these areas. Market factors and traffic factors examined include variable loads, growth rates, supersonic transport market shares, and schedule frequencies considering the different makeup of passenger traffic and individual city pairs. Direct, indirect, and total operating costs and yield levels were economically analyzed for first class and full fare economy class traffic.

  12. The aeroacoustics of supersonic jets

    NASA Technical Reports Server (NTRS)

    Morris, Philip J.; McLaughlin, Dennis K.

    1995-01-01

    This research project was a joint experimental/computational study of noise in supersonic jets. The experiments were performed in a low to moderate Reynolds number anechoic supersonic jet facility. Computations have focused on the modeling of the effect of an external shroud on the generation and radiation of jet noise. This report summarizes the results of the research program in the form of the Masters and Doctoral theses of those students who obtained their degrees with the assistance of this research grant. In addition, the presentations and publications made by the principal investigators and the research students is appended.

  13. Program Solves Euler Equations Of Supersonic Flow

    NASA Technical Reports Server (NTRS)

    Szema, Kuo-Yen; Chakravarthy, Sukumar; Shankar, Vijaya

    1992-01-01

    EMTAC (Euler Marching Technique For Accurate Computation) computer program employs Euler marching algorithm for computing supersonic flows. Enables accurate nonlinear analysis of nonlinear dynamics of supersonic flow over wide range of supersonic mach numbers and angles of attack. Based on exact equations of dynamics of inviscid gas and valid for computations of flows at low and high supersonic mach numbers exhibiting strong shocks and rotational effects. Program also used for inviscid three-dimensional flows fully subsonic or transonic (subsonic with supersonic pockets). Written in FORTRAN V.

  14. Supersonic Laminar Flow Control Research

    NASA Technical Reports Server (NTRS)

    Lo, C. F.; Wiberg, Clark G.

    1996-01-01

    The objective of this research is to understand supersonic laminar flow stability, transition and active control. Some prediction techniques are developed or modified to analyze laminar flow stability. The effects of distributed heating and cooling as an active boundary layer control technique are studied. The primary tasks of the research apply to the NASA/Ames Proof-of-Concept (PoC) and the Laminar Flow Supersonic Wind Tunnel's (LFSWT's) nozzle design with laminar flow control and are listed as follows: (1) Predictions of supersonic laminar boundary layer stability and transition; (2) Effects of wall heating and cooling on supersonic laminar flow control on a flat plate; (3) Performance evaluation of the PoC and LFSWT nozzle designs with wall heating and cooling applied at different locations and various lengths; (4) Effects of a conducted -vs- pulse wall temperature distribution for the LFSWT; and (5) Application of wall heating and/or cooling to laminar boundary layer and flow separation control of airfoils and investigation of related active control techniques.

  15. Supersonic Wave Interference Affecting Stability

    NASA Technical Reports Server (NTRS)

    Love, Eugene S.

    1958-01-01

    Some of the significant interference fields that may affect stability of aircraft at supersonic speeds are briefly summarized. Illustrations and calculations are presented to indicate the importance of interference fields created by wings, bodies, wing-body combinations, jets, and nacelles.

  16. Supersonic jet shock noise reduction

    NASA Technical Reports Server (NTRS)

    Stone, J. R.

    1984-01-01

    Shock-cell noise is identified to be a potentially significant problem for advanced supersonic aircraft at takeoff. Therefore NASA conducted fundamental studies of the phenomena involved and model-scale experiments aimed at developing means of noise reduction. The results of a series of studies conducted to determine means by which supersonic jet shock noise can be reduced to acceptable levels for advanced supersonic cruise aircraft are reviewed. Theoretical studies were conducted on the shock associated noise of supersonic jets from convergent-divergent (C-D) nozzles. Laboratory studies were conducted on the influence of narrowband shock screech on broadband noise and on means of screech reduction. The usefulness of C-D nozzle passages was investigated at model scale for single-stream and dual-stream nozzles. The effect of off-design pressure ratio was determined under static and simulated flight conditions for jet temperatures up to 960 K. Annular and coannular flow passages with center plugs and multi-element suppressor nozzles were evaluated, and the effect of plug tip geometry was established. In addition to the far-field acoustic data, mean and turbulent velocity distributions were measured with a laser velocimeter, and shadowgraph images of the flow field were obtained.

  17. Liquid atomization in supersonic flows

    NASA Astrophysics Data System (ADS)

    Missoum, Azzedine

    An experimental investigation of the atomization of a round liquid jet by coaxial, costream injection into a supersonic, Mach 1.5 air flow is reported. Extensive flow visualization was conducted using schlieren/shadowgraph, flash photography, and short duration (ns) laser imaging. The finer details of the jet were revealed when viewed under high magnification with the help of a microscope. The liquid and air pressures were varied individually. Photographic evidence indicates the presence of three regions within the liquid jet: a primary region enclosed by the first shock cell where the primary breakup occurs, a secondary region in which the jet is totally broken because of its interaction with the supersonic wave structure, and a third, subsonic region further downstream. It was found that the breakup mechanism of liquid jets in supersonic airstreams is quite complex. The breakup seems to be initiated by the growth of the turbulent structure on the liquid surface and the subsequent detachment of the three-dimensional structure as fine droplets by the intense shear at the liquid-gas interface. This seems to confirm the boundary layer stripping mechanism. The liquid jet expands into a bubble like formation as it interacts with the first set of waves. Higher liquid injection pressures resulted in higher initial spray angles. The liquid jet displayed a geometry strongly dependent on the pressure distribution resulting from the wave structure present in the supersonic jet. Droplet size and velocity distributions were measured by the P/DPA (Phase/Doppler Particle Analyzer) system. The Sauter Mean Diameter (SMD) was measured at several axial and radial locations at various liquid and air pressures. The SMD shows a decrease with increase in both the air-to-liquid mass flow ratio and the Weber number. The drop size decreased towards the outer edges of the jet. The results lead one to conclude that the coaxial, coflowing configuration is very attractive for atomizing

  18. Supersonic aerodynamics of delta wings

    NASA Technical Reports Server (NTRS)

    Wood, Richard M.

    1988-01-01

    Through the empirical correlation of experimental data and theoretical analysis, a set of graphs has been developed which summarize the inviscid aerodynamics of delta wings at supersonic speeds. The various graphs which detail the aerodynamic performance of delta wings at both zero-lift and lifting conditions were then employed to define a preliminary wing design approach in which both the low-lift and high-lift design criteria were combined to define a feasible design space.

  19. Nonaxisymmetric Body Supersonic, Aerodynamic Prediction

    DTIC Science & Technology

    1987-08-01

    wing - tail configuration are compared in Figure 27. CN comparisons are good. C. is a sensitive computation for xcp close to x’. 7.2...Analytical and Experimental Supersonic Aerodynamic Characteristics of a Forward Control Missile , AIAA Paper No. 81-0398, AIAA 19th Aerospace Sciences...body diameter. The next computational example is for a body- wing - tail configuration from Reference 32 A body-alone comparison has been made earlier in

  20. Aerodynamics of Supersonic Lifting Bodies

    DTIC Science & Technology

    1981-02-01

    verso of front cover. 19 Y WOROS (Continue on rt.’,;erso side i recessary and identily by block number) Theoretical Aerodynamics Lifting Bodies Wind ...waverider solution, developed from the supersonic wedge flow solution, is then i Fused to fashion vertLcal stabilizer-likh control surfaces. Wind ...served as Project Engineers ror thE wind tunnel work. Important contributions were also made bv: Mr. iis±ung Miin; Lee, -M. Beom-Soo Kim, Mtr. Martin Weeks

  1. Supersonic Chordwise Bending Flutter in Cascades

    DTIC Science & Technology

    1975-05-31

    such a flutter boundary can be made by utilizing the trend lines predicted from a supersonic analysis based on supersonic cascade theory (Appendix I...bonding agent was injected via hypodermic needles after the blade tabs were properly inserted, The integrity and repeatability of the mounting of the indi...in conjunction with NASTRAN predictions and supersonic cascade aerodynamic computa- tions. Comparisons between theory and experiment are discussed. DD

  2. Supersonic Retropropulsion Flight Test Concepts

    NASA Technical Reports Server (NTRS)

    Post, Ethan A.; Dupzyk, Ian C.; Korzun, Ashley M.; Dyakonov, Artem A.; Tanimoto, Rebekah L.; Edquist, Karl T.

    2011-01-01

    NASA's Exploration Technology Development and Demonstration Program has proposed plans for a series of three sub-scale flight tests at Earth for supersonic retropropulsion, a candidate decelerator technology for future, high-mass Mars missions. The first flight test in this series is intended to be a proof-of-concept test, demonstrating successful initiation and operation of supersonic retropropulsion at conditions that replicate the relevant physics of the aerodynamic-propulsive interactions expected in flight. Five sub-scale flight test article concepts, each designed for launch on sounding rockets, have been developed in consideration of this proof-of-concept flight test. Commercial, off-the-shelf components are utilized as much as possible in each concept. The design merits of the concepts are compared along with their predicted performance for a baseline trajectory. The results of a packaging study and performance-based trade studies indicate that a sounding rocket is a viable launch platform for this proof-of-concept test of supersonic retropropulsion.

  3. Supersonic coaxial jet noise predictions

    NASA Technical Reports Server (NTRS)

    Dahl, Milo D.; Morris, Philip J.

    1995-01-01

    Predictions are made for the noise radiation from supersonic, coaxial jets. These predictions are based on the assumption that the noise radiation in the downstream direction of supersonic jets is dominated by sound generated by instability waves with supersonic phase velocities relative to ambient. Since the analysis requires a known mean flow and the coaxial jet mean flow is not described easily in terms of analytical functions, a numerical prediction is made for its development. The compressible, Reynolds averaged, boundary layer equations are solved with a modified mixing length turbulence model. The model has been calibrated to account for compressibility and temperature effects on the rate of mixing. Both normal and inverted velocity profile jets are considered. Predictions are made for the differences between the noise radiated by coaxial jets with difference operating conditions and a single reference jet with the same thrust, mass flow, and exit area. The effects of area ratio changes and simulated enhanced mixing on noise radiation are also considered.

  4. Is ozone doomed by supersonic plume?

    NASA Astrophysics Data System (ADS)

    Briebele, Elaine

    The simple desire to save time is reviving interest in supersonic flight. Travelers' demands for faster travel have compelled Congress to fund NASA's High Speed Research Program, which aims to overcome the obstacles that sent full development of the supersonic passenger plane to its demise over 25 years ago.Removing those obstacles presents substantial challenges. The current $6,000 to $9,000 fare for a supersonic trans-Atlantic trip on the Concorde lends credence to the view of supersonic aircraft as “chariots of the rich.”

  5. Supersonic Cruise Research 1979, part 1

    NASA Technical Reports Server (NTRS)

    1980-01-01

    Aerodynamics, stability and control, propulsion, and environmental factors of the supersonic cruise aircraft are discussed. Other topics include airframe structures and materials, systems integration, and economics.

  6. The supersonic fan engine: An advanced concept in supersonic cruise propulsion

    NASA Technical Reports Server (NTRS)

    Franciscus, L. C.

    1981-01-01

    Engine performance and mission studies were carried out for turbofan engines equipped with supersonic through-flow fans. The mission was for a commercial supersonic transport with a Mach 2.32 capability. The advantages of the supersonic fan engines are discussed in terms of mission range comparisons with other engine types. The effects of fan efficiency, inlet losses, and engine weight on engine performance and mission range are shown. The range of a supersonic transport with supersonic fan engines could be 10 to 20 percent better than with other types having the same technology core.

  7. Validation of OVERFLOW for Supersonic Retropropulsion

    NASA Technical Reports Server (NTRS)

    Schauerhamer, Guy

    2012-01-01

    The goal is to softly land high mass vehicles (10s of metric tons) on Mars. Supersonic Retropropulsion (SRP) is a potential method of deceleration. Current method of supersonic parachutes does not scale past 1 metric ton. CFD is of increasing importance since flight and experimental data at these conditions is difficult to obtain. CFD must first be validated at these conditions.

  8. Supersonic axial-flow fan flutter

    NASA Technical Reports Server (NTRS)

    Ramsey, John K.

    1988-01-01

    Lane's (1957) analytical formulation of the unsteady pressure distribution on an oscillating two-dimensional flat plate cascade in supersonic axial flow has been developed into a computer code. This unsteady aerodynamic code has shown good agreement with other published data. This code has also been incorporated into an existing aeroelastic code to analyze the NASA Lewis supersonic through-flow fan design.

  9. Supersonic through-flow fan assessment

    NASA Technical Reports Server (NTRS)

    Kepler, C. E.; Champagne, G. A.

    1988-01-01

    A study was conducted to assess the performance potential of a supersonic through-flow fan engine for supersonic cruise aircraft. It included a mean-line analysis of fans designed to operate with in-flow velocities ranging from subsonic to high supersonic speeds. The fan performance generated was used to estimate the performance of supersonic fan engines designed for four applications: a Mach 2.3 supersonic transport, a Mach 2.5 fighter, a Mach 3.5 cruise missile, and a Mach 5.0 cruise vehicle. For each application an engine was conceptualized, fan performance and engine performance calculated, weight estimates made, engine installed in a hypothetical vehicle, and mission analysis was conducted.

  10. Multifidelity Analysis and Optimization for Supersonic Design

    NASA Technical Reports Server (NTRS)

    Kroo, Ilan; Willcox, Karen; March, Andrew; Haas, Alex; Rajnarayan, Dev; Kays, Cory

    2010-01-01

    Supersonic aircraft design is a computationally expensive optimization problem and multifidelity approaches over a significant opportunity to reduce design time and computational cost. This report presents tools developed to improve supersonic aircraft design capabilities including: aerodynamic tools for supersonic aircraft configurations; a systematic way to manage model uncertainty; and multifidelity model management concepts that incorporate uncertainty. The aerodynamic analysis tools developed are appropriate for use in a multifidelity optimization framework, and include four analysis routines to estimate the lift and drag of a supersonic airfoil, a multifidelity supersonic drag code that estimates the drag of aircraft configurations with three different methods: an area rule method, a panel method, and an Euler solver. In addition, five multifidelity optimization methods are developed, which include local and global methods as well as gradient-based and gradient-free techniques.

  11. The rebirth of supersonic transport

    NASA Technical Reports Server (NTRS)

    Rosen, Robert; Williams, Louis J.

    1993-01-01

    A development status evaluation is presented for the aerodynamic and propulsion system technologies that will be required to make a second-generation SST, or High Speed Civil Transport (HSCT) sufficiently (20 dB) quieter near airports and in supersonic cruise (sonic boom), less NO(x)-emitting, and longer-range than Concorde. NASA's Ames, Lewis, and Langley facilities are intensively concerned with the development of such an environmentally benign and economically feasible Mach 2.4 HSCT; efforts have concentrated on the design of a powerplant with the requisite specific fuel consumption, low emissions, and low noise.

  12. Space shuttle launch vehicle (13 P-OTS) strut support interference effects study in the Rockwell International 7- by 7-foot trisonic wind tunnel (IA68)

    NASA Technical Reports Server (NTRS)

    Rogge, R. L.

    1974-01-01

    Strut support interference investigations were conducted on an 0.004-(-) scale representation of the space shuttle launch vehicle in order to determine transonic and supersonic model support interference effects for use in a future exhaust plume effects study. Strut configurations were also tested. Orbiter, external tank, and solid rocket booster pressures were recorded at Mach numbers 0.9, 1.2, 1.5, and 2.0. Angle of attack and angle of sideslip were varied between plus or minus 4 degrees in 2 degree increments. Parametric variations consisted only of the strut configurations.

  13. Vortex Flows at Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Wood, Richard M.; Wilcox, Floyd J., Jr.; Bauer, Steven X. S.; Allen, Jerry M.

    2003-01-01

    A review of research conducted at the National Aeronautics and Space Administration (NASA) Langley Research Center (LaRC) into high-speed vortex flows during the 1970s, 1980s, and 1990s is presented. The data are for flat plates, cavities, bodies, missiles, wings, and aircraft with Mach numbers of 1.5 to 4.6. Data are presented to show the types of vortex structures that occur at supersonic speeds and the impact of these flow structures on vehicle performance and control. The data show the presence of both small- and large-scale vortex structures for a variety of vehicles, from missiles to transports. For cavities, the data show very complex multiple vortex structures exist at all combinations of cavity depth to length ratios and Mach number. The data for missiles show the existence of very strong interference effects between body and/or fin vortices. Data are shown that highlight the effect of leading-edge sweep, leading-edge bluntness, wing thickness, location of maximum thickness, and camber on the aerodynamics of and flow over delta wings. Finally, a discussion of a design approach for wings that use vortex flows for improved aerodynamic performance at supersonic speeds is presented.

  14. Economic benefits of supersonic overland operation

    NASA Technical Reports Server (NTRS)

    Metwally, Munir

    1992-01-01

    Environmental concerns are likely to impose some restrictions on the next generation of supersonic commercial transport. There is a global concern over the effects of engine emissions on the ozone layer which protects life on Earth from ultraviolet radiation. There is also some concern over community noise. The High Speed Civil Transport (HSCT) must meet at least the current subsonic noise certification standards to be compatible with the future subsonic fleet. Concerns over sonic boom represent another environmental and marketing challenge to the HSCT program. The most attractive feature of the supersonic transport is speed, which offers the traveling public significant time-savings on long range routes. The sonic boom issue represents a major environmental and economic challenge as well. Supersonic operation overland produces the most desirable economic results. However, unacceptable overland sonic boom raise levels may force HSCT to use subsonic speeds overland. These environmental and economic challenges are likely to impose some restrictions on supersonic operation, thus introducing major changes to existing route structures and future supersonic network composition. The current subsonic route structure may have to be altered for supersonic transports to avoid sensitive areas in the stratosphere or to minimize overland flight tracks. It is important to examine the alternative route structure and the impact of these restrictions on the economic viability of the overall supersonic operation. Future market potential for HSCT fleets must be large enough to enable engine and airframe manufacturers to build the plane at a cost that provides them with an attractive return on investment and to sell it at a price that allows the airlines to operate with a reasonable margin of profit. Subsonic overland operation of a supersonic aircraft hinders its economic viability. Ways to increase the market potential of supersonic operation are described.

  15. Theoretical aspects of supersonic jet noise

    NASA Technical Reports Server (NTRS)

    Tam, Christopher K. W.

    1992-01-01

    The topics covered include the following: the three components of supersonic jet noise; shock cell structure of imperfectly expanded jets; large turbulence structures/instability waves; supersonic jet noise theory; generation of turbulent mixing noise; comparisons between predicted peak noise frequency and direction of radiation with measurements; Strouhal number of maximum SPL of hot supersonic jets; near field sound pressure level contours; generation of broadband shock associated noise; calculated and measured far field shock noise spectra; generation of screech tones; and calculated and measured Strouhal number of screech tones.

  16. Method and apparatus for starting supersonic compressors

    DOEpatents

    Lawlor, Shawn P [Bellevue, WA

    2012-04-10

    A supersonic gas compressor. The compressor includes aerodynamic duct(s) situated on a rotor journaled in a casing. The aerodynamic duct(s) generate a plurality of oblique shock waves for efficiently compressing a gas at supersonic conditions. The convergent inlet is adjacent to a bleed air collector, and during acceleration of the rotor, bypass gas is removed from the convergent inlet via a collector to enable supersonic shock stabilization. Once the oblique shocks are stabilized at a selected inlet relative Mach number and pressure ratio, the bleed of bypass gas from the convergent inlet via the bypass gas collectors is eliminated.

  17. Nonlinear Stability of Supersonic Jets

    NASA Technical Reports Server (NTRS)

    Bhat, T. R. S.; Seiner, J. M.; Tiwari, S. N.

    1995-01-01

    This paper presents stability calculations made for a shock-free supersonic jet using the model based on parabolized stability equations. In this analysis the large-scale structures, which play a dominant role in the mixing as well as the noise radiated, are modeled as instability waves. This model takes into consideration non-parallel flow effects and also nonlinear interaction of the instability waves. The stability calculations have been performed for different frequencies and mode numbers over a range of jet operating temperatures. Comparisons are made, where appropriate, with the solutions to Rayleigh's equation (linear, inviscid analysis with the assumption of parallel flow). The comparison of the solutions obtained using the two approaches show very good agreement.

  18. Modeling Combustion in Supersonic Flows

    NASA Technical Reports Server (NTRS)

    Drummond, J. Philip; Danehy, Paul M.; Bivolaru, Daniel; Gaffney, Richard L.; Tedder, Sarah A.; Cutler, Andrew D.

    2007-01-01

    This paper discusses the progress of work to model high-speed supersonic reacting flow. The purpose of the work is to improve the state of the art of CFD capabilities for predicting the flow in high-speed propulsion systems, particularly combustor flow-paths. The program has several components including the development of advanced algorithms and models for simulating engine flowpaths as well as a fundamental experimental and diagnostic development effort to support the formulation and validation of the mathematical models. The paper will provide details of current work on experiments that will provide data for the modeling efforts along with with the associated nonintrusive diagnostics used to collect the data from the experimental flowfield. Simulation of a recent experiment to partially validate the accuracy of a combustion code is also described.

  19. Supersonic Combustion Research at NASA

    NASA Technical Reports Server (NTRS)

    Drummond, J. P.; Danehy, Paul M.; Gaffney, Richard L., Jr.; Tedder, Sarah A.; Cutler, Andrew D.; Bivolaru, Daniel

    2007-01-01

    This paper discusses the progress of work to model high-speed supersonic reacting flow. The purpose of the work is to improve the state of the art of CFD capabilities for predicting the flow in high-speed propulsion systems, particularly combustor flowpaths. The program has several components including the development of advanced algorithms and models for simulating engine flowpaths as well as a fundamental experimental and diagnostic development effort to support the formulation and validation of the mathematical models. The paper will provide details of current work on experiments that will provide data for the modeling efforts along with the associated nonintrusive diagnostics used to collect the data from the experimental flowfield. Simulation of a recent experiment to partially validate the accuracy of a combustion code is also described.

  20. 9- by 15-Foot Low Speed Wind Tunnel Acoustic Improvements Expanded Overview

    NASA Technical Reports Server (NTRS)

    Stephens, David

    2016-01-01

    The 9- by 15-Foot Low Speed Wind Tunnel (9x15 LSWT) at NASA Glenn Research Center was built in 1969 in the return leg of the 8- by 6-Foot Supersonic Wind Tunnel (8x6 SWT). The 8x6 SWT was completed in 1949 and acoustically treated to mitigate community noise issues in 1950. This treatment included the addition of a large muffler downstream of the 8x6 SWT test section and diffuser. The 9x15 LSWT was designed for performance testing of V/STOL aircraft models, but with the addition of the current acoustic treatment in 1986 the tunnel been used principally for acoustic and performance testing of aircraft propulsion systems. The present document describes an anticipated acoustic upgrade to be completed in 2017.

  1. NASA Flight Tests Explore Supersonic Laminar Flow

    NASA Video Gallery

    In partnership with Aerion Corporation of Reno, Nevada, NASA's Dryden Flight Research Center’s tested supersonic airflow over a small experimental airfoil design on its F-15B Test Bed aircraft du...

  2. Streamline curvature in supersonic shear layers

    NASA Technical Reports Server (NTRS)

    Kibens, V.

    1992-01-01

    Results of an experimental investigation in which a curved shear layer was generated between supersonic flow from a rectangular converging/diverging nozzle and the freestream in a series of open channels with varying radii of curvature are reported. The shear layers exhibit unsteady large-scale activity at supersonic pressure ratios, indicating increased mixing efficiency. This effect contrasts with supersonic flow in a straight channel, for which no large-scale vortical structure development occurs. Curvature must exceed a minimum level before it begins to affect the dynamics of the supersonic shear layer appreciably. The curved channel flows are compared with reference flows consisting of a free jet, a straight channel, and wall jets without sidewalls on a flat and a curved plate.

  3. Supersonic fan engines for military aircraft

    NASA Technical Reports Server (NTRS)

    Franciscus, L. C.

    1983-01-01

    Engine performance and mission studies were performed for turbofan engines with supersonic through-flow fans. A Mach 2.4 CTOL aircraft was used in the study. Two missions were considered: a long range penetrator mission and a long range intercept mission. The supersonic fan engine is compared with an augmented mixed flow turbofan in terms of mission radius for a fixed takeoff gross weight of 75,000 lbm. The mission radius of aircraft powered by supersonic fan engines could be 15 percent longer than aircraft powered with conventional turbofan engines at moderate thrust to gross weight ratios. The climb and acceleration performance of the supersonic fan engines is better than that of the conventional turbofan engines.

  4. Low-Density Supersonic Decelerator Test

    NASA Video Gallery

    NASA recently performed a trial run on a rocket sled test fixture, powered by rockets, to replicate the forces a supersonic spacecraft would experience prior to landing. The sled tests will allow t...

  5. Measurements of Supersonic Wing Tip Vortices

    NASA Technical Reports Server (NTRS)

    Smart, Michael K.; Kalkhoran, Iraj M.; Benston, James

    1994-01-01

    An experimental survey of supersonic wing tip vortices has been conducted at Mach 2.5 using small performed 2.25 chords down-stream of a semi-span rectangular wing at angle of attack of 5 and 10 degrees. The main objective of the experiments was to determine the Mach number, flow angularity and total pressure distribution in the core region of supersonic wing tip vortices. A secondary aim was to demonstrate the feasibility of using cone probes calibrated with a numerical flow solver to measure flow characteristics at supersonic speeds. Results showed that the numerically generated calibration curves can be used for 4-hole cone probes, but were not sufficiently accurate for conventional 5-hole probes due to nose bluntness effects. Combination of 4-hole cone probe measurements with independent pitot pressure measurements indicated a significant Mach number and total pressure deficit in the core regions of supersonic wing tip vortices, combined with an asymmetric 'Burger like' swirl distribution.

  6. Supersonic cruise aircraft research: An annotated bibliography

    NASA Technical Reports Server (NTRS)

    Tuttle, M. H.

    1980-01-01

    This bibliography, with abstracts, consists of 69 publications arranged in chronological order. The material may be useful to those interested in supersonic cruise fighter/penetrator/interceptor airplanes. Two pertinent conferences on military supercruise aircraft are considered as single items; one contains 37 papers and the other 29 papers. In addition, several related bibliographies are included which cover supersonic civil aircraft and military aircraft studies at the Langley Research Center. There is also an author index.

  7. Conditions for supersonic bent Marshak waves

    SciTech Connect

    Xu, Qiang Ren, Xiao-dong; Li, Jing; Dan, Jia-kun; Wang, Kun-lun; Zhou, Shao-tong

    2015-03-15

    Supersonic radiation diffusion approximation is an useful method to study the radiation transportation. Considering the 2-d Marshak theory, and an invariable source temperature, conditions for supersonic radiation diffusion are proved to be coincident with that for radiant flux domination in the early time when √(ε)x{sub f}/L≪1. However, they are even tighter than conditions for radiant flux domination in the late time when √(ε)x{sub f}/L≫1, and can be expressed as M>4(1+ε/3)/3 and τ>1. A large Mach number requires the high temperature, while the large optical depth requires the low temperature. Only when the source temperature is in a proper region the supersonic diffusion conditions can be satisfied. Assuming a power-low (in temperature and density) opacity and internal energy, for a given density, the supersonic diffusion regions are given theoretically. The 2-d Marshak theory is proved to be able to bound the supersonic diffusion conditions in both high and low temperature regions, however, the 1-d theory only bounds it in low temperature region. Taking SiO{sub 2} and the Au, for example, these supersonic regions are shown numerically.

  8. Low Density Supersonic Decelerator Parachute Decelerator System

    NASA Technical Reports Server (NTRS)

    Gallon, John C.; Clark, Ian G.; Rivellini, Tommaso P.; Adams, Douglas S.; Witkowski, Allen

    2013-01-01

    The Low Density Supersonic Decelerator Project has undertaken the task of developing and testing a large supersonic ringsail parachute. The parachute under development is intended to provide mission planners more options for parachutes larger than the Mars Science Laboratory's 21.5m parachute. During its development, this new parachute will be taken through a series of tests in order to bring the parachute to a TRL-6 readiness level and make the technology available for future Mars missions. This effort is primarily focused on two tests, a subsonic structural verification test done at sea level atmospheric conditions and a supersonic flight behind a blunt body in low-density atmospheric conditions. The preferred method of deploying a parachute behind a decelerating blunt body robotic spacecraft in a supersonic flow-field is via mortar deployment. Due to the configuration constraints in the design of the test vehicle used in the supersonic testing it is not possible to perform a mortar deployment. As a result of this limitation an alternative deployment process using a ballute as a pilot is being developed. The intent in this alternate approach is to preserve the requisite features of a mortar deployment during canopy extraction in a supersonic flow. Doing so will allow future Mars missions to either choose to mortar deploy or pilot deploy the parachute that is being developed.

  9. Design project: LONGBOW supersonic interceptor

    NASA Technical Reports Server (NTRS)

    Stoney, Robert; Baker, Matt; Capstaff, Joseph G.; Dishman, Robert; Fick, Gregory; Frick, Stephen N.; Kelly, Mark

    1993-01-01

    A recent white paper entitled 'From the Sea' has spotlighted the need for Naval Aviation to provide overland support to joint operations. The base for this support, the Aircraft Carrier (CVN), will frequently be unable to operate within close range of the battleground because of littoral land-based air and subsurface threats. A high speed, long range, carrier capable aircraft would allow the CVN to provide timely support to distant battleground operations. Such an aircraft, operating as a Deck-Launched Interceptor (DLI), would also be an excellent counter to Next Generation Russian Naval Aviation (NGRNA) threats consisting of supersonic bombers, such as the Backfire, equipped with the next generation of high-speed, long-range missiles. Additionally, it would serve as an excellent high speed Reconnaissance airplane, capable of providing Battle Force commanders with timely, accurate pre-mission targeting information and post-mission Bomb Damage Assessment (BDA). Recent advances in computational hypersonic airflow modeling has produced a method of defining aircraft shapes that fit a conical shock flow model to maximize the efficiency of the vehicle. This 'Waverider' concept provides one means of achieving long ranges at high speeds. A Request for Proposal (RFP) was issued by Professor Conrad Newberry that contained design requirements for an aircraft to accomplish the above stated missions, utilizing Waverider technology.

  10. Toward Supersonic Retropropulsion CFD Validation

    NASA Technical Reports Server (NTRS)

    Kleb, Bil; Schauerhamer, D. Guy; Trumble, Kerry; Sozer, Emre; Barnhardt, Michael; Carlson, Jan-Renee; Edquist, Karl

    2011-01-01

    This paper begins the process of verifying and validating computational fluid dynamics (CFD) codes for supersonic retropropulsive flows. Four CFD codes (DPLR, FUN3D, OVERFLOW, and US3D) are used to perform various numerical and physical modeling studies toward the goal of comparing predictions with a wind tunnel experiment specifically designed to support CFD validation. Numerical studies run the gamut in rigor from code-to-code comparisons to observed order-of-accuracy tests. Results indicate that this complex flowfield, involving time-dependent shocks and vortex shedding, design order of accuracy is not clearly evident. Also explored is the extent of physical modeling necessary to predict the salient flowfield features found in high-speed Schlieren images and surface pressure measurements taken during the validation experiment. Physical modeling studies include geometric items such as wind tunnel wall and sting mount interference, as well as turbulence modeling that ranges from a RANS (Reynolds-Averaged Navier-Stokes) 2-equation model to DES (Detached Eddy Simulation) models. These studies indicate that tunnel wall interference is minimal for the cases investigated; model mounting hardware effects are confined to the aft end of the model; and sparse grid resolution and turbulence modeling can damp or entirely dissipate the unsteadiness of this self-excited flow.

  11. Quiet Supersonic Wind Tunnel Development

    NASA Technical Reports Server (NTRS)

    King, Lyndell S.; Kutler, Paul (Technical Monitor)

    1994-01-01

    The ability to control the extent of laminar flow on swept wings at supersonic speeds may be a critical element in developing the enabling technology for a High Speed Civil Transport (HSCT). Laminar boundary layers are less resistive to forward flight than their turbulent counterparts, thus the farther downstream that transition from laminar to turbulent flow in the wing boundary layer is extended can be of significant economic impact. Due to the complex processes involved experimental studies of boundary layer stability and transition are needed, and these are performed in "quiet" wind tunnels capable of simulating the low-disturbance environment of free flight. At Ames, a wind tunnel has been built to operate at flow conditions which match those of the HSCT laminar flow flight demonstration 'aircraft, the F-16XL, i.e. at a Mach number of 1.6 and a Reynolds number range of 1 to 3 million per foot. This will allow detailed studies of the attachment line and crossflow on the leading edge area of the highly swept wing. Also, use of suction as a means of control of transition due to crossflow and attachment line instabilities can be studied. Topics covered include: test operating conditions required; design requirements to efficiently make use of the existing infrastructure; development of an injector drive system using a small pilot facility; plenum chamber design; use of computational tools for tunnel and model design; and early operational results.

  12. 6. VIEW NORTH, INTERIOR VIEW OF BUILDING 11, SUPERSONIC WIND ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    6. VIEW NORTH, INTERIOR VIEW OF BUILDING 11, SUPERSONIC WIND TUNNEL - Naval Surface Warfare Center, Supersonic Wind Tunnel Building, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  13. Supersonic Parachute Aerodynamic Testing and Fluid Structure Interaction Simulation

    NASA Astrophysics Data System (ADS)

    Lingard, J. S.; Underwood, J. C.; Darley, M. G.; Marraffa, L.; Ferracina, L.

    2014-06-01

    The ESA Supersonic Parachute program expands the knowledge of parachute inflation and flying characteristics in supersonic flows using wind tunnel testing and fluid structure interaction to develop new inflation algorithms and aerodynamic databases.

  14. Time-Lapse Assembly of NASA Supersonic Model

    NASA Video Gallery

    Supersonic passenger airplanes are another step closer to reality as NASA and Lockheed Martin began the first high-speed wind tunnel tests for the Quiet Supersonic Technology, or QueSST, X-plane pr...

  15. Supersonics Project - Airport Noise Tech Challenge

    NASA Technical Reports Server (NTRS)

    Bridges, James

    2010-01-01

    The Airport Noise Tech Challenge research effort under the Supersonics Project is reviewed. While the goal of "Improved supersonic jet noise models validated on innovative nozzle concepts" remains the same, the success of the research effort has caused the thrust of the research to be modified going forward in time. The main activities from FY06-10 focused on development and validation of jet noise prediction codes. This required innovative diagnostic techniques to be developed and deployed, extensive jet noise and flow databases to be created, and computational tools to be developed and validated. Furthermore, in FY09-10 systems studies commissioned by the Supersonics Project showed that viable supersonic aircraft were within reach using variable cycle engine architectures if exhaust nozzle technology could provide 3-5dB of suppression. The Project then began to focus on integrating the technologies being developed in its Tech Challenge areas to bring about successful system designs. Consequently, the Airport Noise Tech Challenge area has shifted efforts from developing jet noise prediction codes to using them to develop low-noise nozzle concepts for integration into supersonic aircraft. The new plan of research is briefly presented by technology and timelines.

  16. SUPERSONIC TRANSPORT DEVELOPMENT AND PRODUCTION. VOLUME I. COST ANALYSIS PROGRAM.

    DTIC Science & Technology

    SUPERSONIC AIRCRAFT, *COSTS), (*AIRCRAFT INDUSTRY, INDUSTRIAL PRODUCTION ), MANAGEMENT ENGINEERING, AIRFRAMES, ECONOMICS, COMPUTER PROGRAMS, STATISTICAL ANALYSIS, MONEY, AIRCRAFT ENGINES, FEASIBILITY STUDIES

  17. Drag Reduction Tests on Supersonic Transport Design

    NASA Technical Reports Server (NTRS)

    1998-01-01

    Langley researchers recently completed supersonic tests in the Unitary Plan Wind Tunnel on a nonlinear design for a supersonic transport. Although the drag reduction measured during the tests was not as great as that predicted using computational methods, significant drag reductions were achieved. Future tests will be conducted at a higher Reynolds number, which will be more representative of flight conditions. These tests will be used to identify a supersonic transport configuration that provides maximum drag reduction. Reducing drag decreases operating cost by improving fuel consumption and lowering aircraft weight. As a result, this research has the potential to help make a future high-speed civil transport (HSCT) an affordable means of travel for the flying public.

  18. Transonic and supersonic ground effect aerodynamics

    NASA Astrophysics Data System (ADS)

    Doig, G.

    2014-08-01

    A review of recent and historical work in the field of transonic and supersonic ground effect aerodynamics has been conducted, focussing on applied research on wings and aircraft, present and future ground transportation, projectiles, rocket sleds and other related bodies which travel in close ground proximity in the compressible regime. Methods for ground testing are described and evaluated, noting that wind tunnel testing is best performed with a symmetry model in the absence of a moving ground; sled or rail testing is ultimately preferable, though considerably more expensive. Findings are reported on shock-related ground influence on aerodynamic forces and moments in and accelerating through the transonic regime - where force reversals and the early onset of local supersonic flow is prevalent - as well as more predictable behaviours in fully supersonic to hypersonic ground effect flows.

  19. Supersonic combustion engine testbed, heat lightning

    NASA Technical Reports Server (NTRS)

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

    1990-01-01

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

  20. Method and apparatus for starting supersonic compressors

    DOEpatents

    Lawlor, Shawn P

    2013-08-06

    A supersonic gas compressor with bleed gas collectors, and a method of starting the compressor. The compressor includes aerodynamic duct(s) situated for rotary movement in a casing. The aerodynamic duct(s) generate a plurality of oblique shock waves for efficiently compressing a gas at supersonic conditions. A convergent inlet is provided adjacent to a bleed gas collector, and during startup of the compressor, bypass gas is removed from the convergent inlet via the bleed gas collector, to enable supersonic shock stabilization. Once the oblique shocks are stabilized at a selected inlet relative Mach number and pressure ratio, the bleed of bypass gas from the convergent inlet via the bypass gas collectors is effectively eliminated.

  1. Advanced Supersonic Technology Study: Engine Program Summary. Supersonic Propulsion: 1971 to 1976

    NASA Technical Reports Server (NTRS)

    Krebs, J. N.

    1976-01-01

    Sustained supersonic cruise propulsion systems for military applications are studied. The J79-5 in the Mach 2 B-58; YJ93 in the Mach 3.0 B-70 and the current F101 in the B-1, are all examples of military propulsion systems and airplanes operated at sustained supersonic cruise speeds. The Mach 2.7 B2707 transport powered by GE4 turbojet engines was the only non-military, sustained supersonic cruise vehicle intended for commercial passenger service.

  2. Laser transit anemometer experiences in supersonic flow

    NASA Technical Reports Server (NTRS)

    Hunter, William W., Jr.; Humphreys, William M., Jr.

    1988-01-01

    The purpose of this paper is to present examples of velocity measurements obtained in supersonic flow fields with the laser transit anemometer system. Velocity measurements of a supersonic jet exhausting in a transonic flow field, a cone boundary survey in a Mach 4 flow field, and a determination of the periodic disturbance frequencies of a sonic nozzle flow field are presented. Each of the above three cases also serves to illustrate different modes of laser transit anemometer operation. A brief description of the laser transit anemometer system is also presented.

  3. Systems integration studies for supersonic cruise aircraft

    NASA Technical Reports Server (NTRS)

    Mascitti, V. R.

    1975-01-01

    Technical progress in each of the disciplinary research areas affecting the design of supersonic cruise aircraft is discussed. The NASA AST/SCAR Program supported the integration of these technical advances into supersonic cruise aircraft configuration concepts. While the baseline concepts reflect differing design philosophy, all reflect a level of economic performance considerably above the current foreign aircraft as well as the former U.S. SST. Range-payload characteristics of the study configurating show significant improvement, while meeting environmental goals such as takeoff and landing noise and upper atmospheric pollution.

  4. The supersonic fan engine - An advanced concept in supersonic cruise propulsion

    NASA Technical Reports Server (NTRS)

    Franciscus, L. C.

    1981-01-01

    Engine performance and mission studies were conducted for a novel turbofan engine concept incorporating a supersonic through-flow fan, and comparisons were made with two supersonic transport (SST) engine concepts of equivalent thrust and technological sophistication. It was found that in the case of an SST with a cruise speed of Mach 2.32, the through-flow fan engine may yield ranges 10 to 20% greater than the two alternatives considered. The engine has a conventional core, with the supersonic fan being driven by a concentric low-pressure turbine that is uncoupled with the single, high pressure turbine/compressor core spool. Among the topics discussed are the methods of analysis employed and perturbation studies concerning supersonic fan adiabatic efficiency, fan discharge characteristics and propulsion system weight.

  5. A supersonic fan equipped variable cycle engine for a Mach 2.7 supersonic transport

    NASA Technical Reports Server (NTRS)

    Tavares, T. S.

    1985-01-01

    The concept of a variable cycle turbofan engine with an axially supersonic fan stage as powerplant for a Mach 2.7 supersonic transport was evaluated. Quantitative cycle analysis was used to assess the effects of the fan inlet and blading efficiencies on engine performance. Thrust levels predicted by cycle analysis are shown to match the thrust requirements of a representative aircraft. Fan inlet geometry is discussed and it is shown that a fixed geometry conical spike will provide sufficient airflow throughout the operating regime. The supersonic fan considered consists of a single stage comprising a rotor and stator. The concept is similar in principle to a supersonic compressor, but differs by having a stator which removes swirl from the flow without producing a net rise in static pressure. Operating conditions peculiar to the axially supersonic fan are discussed. Geometry of rotor and stator cascades are presented which utilize a supersonic vortex flow distribution. Results of a 2-D CFD flow analysis of these cascades are presented. A simple estimate of passage losses was made using empirical methods.

  6. Supersonic flow with feeding of energy

    NASA Technical Reports Server (NTRS)

    Zaremba, W.

    1985-01-01

    The present work discusses the results of some experimental studies on the possibility of attenuating shock waves in a supersonic flow. The shock waves were formed by an external source of electrical energy. An electromechanical method is described that permits partial recovery of the expended energy.

  7. Supersonic Injection of Aerated Liquid Jet

    NASA Astrophysics Data System (ADS)

    Choudhari, Abhijit; Sallam, Khaled

    2016-11-01

    A computational study of the exit flow of an aerated two-dimensional jet from an under-expanded supersonic nozzle is presented. The liquid sheet is operating within the annular flow regime and the study is motivated by the application of supersonic nozzles in air-breathing propulsion systems, e.g. scramjet engines, ramjet engines and afterburners. The simulation was conducted using VOF model and SST k- ω turbulence model. The test conditions included: jet exit of 1 mm and mass flow rate of 1.8 kg/s. The results show that air reaches transonic condition at the injector exit due to the Fanno flow effects in the injector passage. The aerated liquid jet is alternately expanded by Prandtl-Meyer expansion fan and compressed by oblique shock waves due to the difference between the back (chamber) pressure and the flow pressure. The process then repeats itself and shock (Mach) diamonds are formed at downstream of injector exit similar to those typical of exhaust plumes of propulsion system. The present results, however, indicate that the flow field of supersonic aerated liquid jet is different from supersonic gas jets due to the effects of water evaporation from the liquid sheet. The contours of the Mach number, static pressure of both cases are compared to the theory of gas dynamics.

  8. Pulsed supersonic expansion of nonvolatile solids

    SciTech Connect

    Christen, Wolfgang; Geggier, Stephanie; Grigorenko, Svitlana; Rademann, Klaus

    2004-11-01

    A compact apparatus for transferring nonvolatile particles into the gas phase and depositing them on a solid surface has been built and tested successfully. As initial experiment, solid caffeine with a vanishingly low vapor pressure has been dissolved in supercritical carbon dioxide, expanded into vacuum using a pulsed, supersonic molecular beam, and detected using a simple residual gas analyzer.

  9. The Aeroacoustics of Supersonic Coaxial Jets

    NASA Technical Reports Server (NTRS)

    Dahl, Milo D.

    1994-01-01

    Instability waves have been established as the dominant source of mixing noise radiating into the downstream arc of a supersonic jet when the waves have phase velocities that are supersonic relative to ambient conditions. Recent theories for supersonic jet noise have used the concepts of growing and decaying linear instability waves for predicting radiated noise. This analysis is extended to the prediction of noise radiation from supersonic coaxial jets. Since the analysis requires a known mean flow and the coaxial jet mean flow is not described easily in terms of analytic functions, a numerical prediction is made for its development. The Reynolds averaged, compressible, boundary layer equations are solved using a mixing length turbulence model. Empirical correlations are developed for the effects of velocity and temperature ratios and Mach number. Both normal and inverted velocity profile coaxial jets are considered. Comparisons with measurements for both single and coaxial jets show good agreement. The results from mean flow and stability calculations are used to predict the noise radiation from coaxial jets with different operating conditions. Comparisons are made between different coaxial jets and a single equivalent jet with the same total thrust, mass flow, and exit area. Results indicate that normal velocity profile jets can have noise reductions compared to the single equivalent jet. No noise reductions are found for inverted velocity profile jets operated at the minimum noise condition compared to the single equivalent jet. However, it is inferred that changes in area ratio may provide noise reduction benefits for inverted velocity profile jets.

  10. Aerodynamic Design Opportunities for Future Supersonic Aircraft

    NASA Technical Reports Server (NTRS)

    Wood, Richard M.; Bauer, Steven X. S.; Flamm, Jeffrey D.

    2002-01-01

    A discussion of a diverse set of aerodynamic opportunities to improve the aerodynamic performance of future supersonic aircraft has been presented and discussed. These ideas are offered to the community in a hope that future supersonic vehicle development activities will not be hindered by past efforts. A number of nonlinear flow based drag reduction technologies are presented and discussed. The subject technologies are related to the areas of interference flows, vehicle concepts, vortex flows, wing design, advanced control effectors, and planform design. The authors also discussed the importance of improving the aerodynamic design environment to allow creativity and knowledge greater influence. A review of all of the data presented show that pressure drag reductions on the order of 50 to 60 counts are achievable, compared to a conventional supersonic cruise vehicle, with the application of several of the discussed technologies. These drag reductions would correlate to a 30 to 40% increase in cruise L/D (lift-to-drag ratio) for a commercial supersonic transport.

  11. Supersonic biplane design via adjoint method

    NASA Astrophysics Data System (ADS)

    Hu, Rui

    In developing the next generation supersonic transport airplane, two major challenges must be resolved. The fuel efficiency must be significantly improved, and the sonic boom propagating to the ground must be dramatically reduced. Both of these objectives can be achieved by reducing the shockwaves formed in supersonic flight. The Busemann biplane is famous for using favorable shockwave interaction to achieve nearly shock-free supersonic flight at its design Mach number. Its performance at off-design Mach numbers, however, can be very poor. This dissertation studies the performance of supersonic biplane airfoils at design and off-design conditions. The choked flow and flow-hysteresis phenomena of these biplanes are studied. These effects are due to finite thickness of the airfoils and non-uniqueness of the solution to the Euler equations, creating over an order of magnitude more wave drag than that predicted by supersonic thin airfoil theory. As a result, the off-design performance is the major barrier to the practical use of supersonic biplanes. The main contribution of this work is to drastically improve the off-design performance of supersonic biplanes by using an adjoint based aerodynamic optimization technique. The Busemann biplane is used as the baseline design, and its shape is altered to achieve optimal wave drags in series of Mach numbers ranging from 1.1 to 1.7, during both acceleration and deceleration conditions. The optimized biplane airfoils dramatically reduces the effects of the choked flow and flow-hysteresis phenomena, while maintaining a certain degree of favorable shockwave interaction effects at the design Mach number. Compared to a diamond shaped single airfoil of the same total thickness, the wave drag of our optimized biplane is lower at almost all Mach numbers, and is significantly lower at the design Mach number. In addition, by performing a Navier-Stokes solution for the optimized airfoil, it is verified that the optimized biplane improves

  12. High Altitude Supersonic Decelerator Test Vehicle

    NASA Technical Reports Server (NTRS)

    Cook, Brant T.; Blando, Guillermo; Kennett, Andrew; Von Der Heydt, Max; Wolff, John Luke; Yerdon, Mark

    2013-01-01

    The Low Density Supersonic Decelerator (LDSD) project is tasked by NASA's Office of the Chief Technologist (OCT) to advance the state of the art in Mars entry and descent technology in order to allow for larger payloads to be delivered to Mars at higher altitudes with better accuracy. The project will develop a 33.5 m Do Supersonic Ringsail (SSRS) parachute, 6m attached torus, robotic class Supersonic Inflatable Aerodynamic Decelerator (SIAD-R), and an 8 m attached isotensoid, exploration class Supersonic Inflatable Aerodynamic Decelerator (SIAD-E). The SSRS and SIAD-R should be brought to TRL-6, while the SIAD-E should be brought to TRL-5. As part of the qualification and development program, LDSD must perform a Mach-scaled Supersonic Flight Dynamics Test (SFDT) in order to demonstrate successful free flight dynamic deployments at Mars equivalent altitude, of all three technologies. In order to perform these tests, LDSD must design and build a test vehicle to deliver all technologies to approximately 180,000 ft and Mach 4, deploy a SIAD, free fly to approximately Mach 2, deploy the SSRS, record high-speed and high-resolution imagery of both deployments, as well as record data from an instrumentation suite capable of characterizing the technology induced vehicle dynamics. The vehicle must also be recoverable after splashdown into the ocean under a nominal flight, while guaranteeing forensic data protection in an off nominal catastrophic failure of a test article that could result in a terminal velocity, tumbling water impact.

  13. Supersonic shock wave/vortex interaction

    NASA Technical Reports Server (NTRS)

    Settles, G. S.; Cattafesta, L.

    1993-01-01

    Although shock wave/vortex interaction is a basic and important fluid dynamics problem, very little research has been conducted on this topic. Therefore, a detailed experimental study of the interaction between a supersonic streamwise turbulent vortex and a shock wave was carried out at the Penn State Gas Dynamics Laboratory. A vortex is produced by replaceable swirl vanes located upstream of the throat of various converging-diverging nozzles. The supersonic vortex is then injected into either a coflowing supersonic stream or ambient air. The structure of the isolated vortex is investigated in a supersonic wind tunnel using miniature, fast-response, five-hole and total temperature probes and in a free jet using laser Doppler velocimetry. The cases tested have unit Reynolds numbers in excess of 25 million per meter, axial Mach numbers ranging from 2.5 to 4.0, and peak tangential Mach numbers from 0 (i.e., a pure jet) to about 0.7. The results show that the typical supersonic wake-like vortex consists of a non-isentropic, rotational core, where the reduced circulation distribution is self similar, and an outer isentropic, irrotational region. The vortex core is also a region of significant turbulent fluctuations. Radial profiles of turbulent kinetic energy and axial-tangential Reynolds stress are presented. The interactions between the vortex and both oblique and normal shock waves are investigated using nonintrusive optical diagnostics (i.e. schlieren, planar laser scattering, and laser Doppler velocimetry). Of the various types, two Mach 2.5 overexpanded-nozzle Mach disc interactions are examined in detail. Below a certain vortex strength, a 'weak' interaction exists in which the normal shock is perturbed locally into an unsteady 'bubble' shock near the vortex axis, but vortex breakdown (i.e., a stagnation point) does not occur. For stronger vortices, a random unsteady 'strong' interaction results that causes vortex breakdown. The vortex core reforms downstream of

  14. Turbulence Model Comparisons for Supersonic Transports at Transonic and Supersonic Conditions

    NASA Technical Reports Server (NTRS)

    Rivers, S. M. B.; Wahls, R. A.

    2003-01-01

    Results of turbulence model comparisons from two studies on supersonic transport configurations performed during the NASA High-speed Research program are given. Results are presented for both transonic conditions at Mach 0.90 and supersonic conditions at Mach 2.48. A feature of these two studies was the availability of higher Reynolds number wind tunnel data with which to compare the computational results. The transonic wind tunnel data was obtained in the National Transonic Facility at NASA Langley, and the supersonic data was obtained in the Boeing Polysonic Wind Tunnel. The computational data was acquired using a state of the art Navier-Stokes flow solver with a wide range of turbulence models implemented. The results show that the computed forces compare reasonably well with the experimental data, with the Baldwin- Lomax with Degani-Schiff modifications and the Baldwin-Barth models showing the best agreement for the transonic conditions and the Spalart-Allmaras model showing the best agreement for the supersonic conditions. The transonic results were more sensitive to the choice of turbulence model than were the supersonic results.

  15. Supersonic through-flow fan engine and aircraft mission performance

    NASA Technical Reports Server (NTRS)

    Franciscus, Leo C.; Maldonado, Jaime J.

    1989-01-01

    A study was made to evaluate potential improvement to a commercial supersonic transport by powering it with supersonic through-flow fan turbofan engines. A Mach 3.2 mission was considered. The three supersonic fan engines considered were designed to operate at bypass ratios of 0.25, 0.5, and 0.75 at supersonic cruise. For comparison a turbine bypass turbojet was included in the study. The engines were evaluated on the basis of aircraft takeoff gross weight with a payload of 250 passengers for a fixed range of 5000 N.MI. The installed specific fuel consumption of the supersonic fan engines was 7 to 8 percent lower than that of the turbine bypass engine. The aircraft powered by the supersonic fan engines had takeoff gross weights 9 to 13 percent lower than aircraft powered by turbine bypass engines.

  16. An asymptotic theory of supersonic propeller noise

    NASA Astrophysics Data System (ADS)

    Envia, Edmane

    1992-05-01

    A theory for predicting the noise field of supersonic propellers with realistic blade geometries is presented. The theory, which utilizes a large-blade-count approximation, provides an efficient formula for predicting the radiation of sound from all three sources of propeller noise. Comparisons with a full numerical integration indicate that the levels predicted by this formula are quite accurate. Calculations also show that, for high speed propellers, the noise radiated by the Lighthill quadrupole source is rather substantial when compared with the noise radiated by the blade thickness and loading sources. Results from a preliminary application of the theory indicate that the peak noise level generated by a supersonic propeller initially increases with increasing tip helical Mach number, but is eventually reaches a plateau and does not increase further. The predicted trend shows qualitative agreement with the experimental observations.

  17. Study of active cooling for supersonic transports

    NASA Technical Reports Server (NTRS)

    Brewer, G. D.; Morris, R. E.

    1975-01-01

    The potential benefits of using the fuel heat sink of hydrogen fueled supersonic transports for cooling large portions of the aircraft wing and fuselage are examined. The heat transfer would be accomplished by using an intermediate fluid such as an ethylene glycol-water solution. Some of the advantages of the system are: (1) reduced costs by using aluminum in place of titanium, (2) reduced cabin heat loads, and (3) more favorable environmental conditions for the aircraft systems. A liquid hydrogen fueled, Mach 2.7 supersonic transport aircraft design was used for the reference uncooled vehicle. The cooled aircraft designs were analyzed to determine their heat sink capability, the extent and location of feasible cooled surfaces, and the coolant passage size and spacing.

  18. Aerodynamics of sounding rockets at supersonic speeds

    NASA Astrophysics Data System (ADS)

    Vira, N. R.

    This dissertation presents a practical and low cost method of computing the aerodynamic characteristics of vehicles such as sounding rockets, high speed bombs, projectiles and guided missiles in supersonic flight. The vehicle configuration consists of a slender axisymmetric body with a conical or ogive noise, cylinders, shoulders and boattails, if any, and have sets of two, three or four fins. Geometry of the fin cross section can be single wedge, double wedge, modified single wedge or modified double wedge. First the aerodynamics of the fins and the body are analyzed separately; then fin body and fore and aft fin interferences are accounted for when they are combined to form the total vehicle. Results and formulas documented in this work are the basis of the supersonic portion of the Theoretical Aerodynamic Derivatives (TAD) computer program operating at the NASA Goddard Space Flight Center.

  19. Supersonic burning in separated flow regions

    NASA Technical Reports Server (NTRS)

    Zumwalt, G. W.

    1982-01-01

    The trough vortex phenomena is used for combustion of hydrogen in a supersonic air stream. This was done in small sizes suitable for igniters in supersonic combustion ramjets so long as the boundary layer displacement thickness is less than 25% of the trough step height. A simple electric spark, properly positioned, ignites the hydrogen in the trough corner. The resulting flame is self sustaining and reignitable. Hydrogen can be injected at the base wall or immediately upstream of the trough. The hydrogen is introduced at low velocity to permit it to be drawn into the corner vortex system and thus experience a long residence time in the combustion region. The igniters can be placed on a skewed back step for angles at least up to 30 deg. without affecting the igniter performance significantly. Certain metals (platinum, copper) act catalytically to improve ignition.

  20. Turbulent jet mixing in a supersonic stream

    NASA Technical Reports Server (NTRS)

    Swanson, R. C.; Schetz, J. A.

    1971-01-01

    An experimental study of turbulent, subsonic, coaxial jet mixing of air in a supersonic air stream is presented. Data taken at five axial stations downstream of the exit of the jet supply tube, which was suspended through the nozzle throat of a supersonic wind tunnel, are given in the form of total pressure, Mach number, and velocity distributions. An investigation of the effect of swirl as a mixing aid was conducted. Swirl, produced by tangential injection of 50% of the total air mass flow leaving the jet supply tube, was examined through Schlieren photographs and total pressure surveys. From a comparison of nonswirl and swirl data, it is concluded that the swirl has no discernible effect on the mixing.

  1. Large Eddy Simulation of Supersonic Inlet Flows

    DTIC Science & Technology

    1998-04-01

    SIMULATION OF SUPERSONIC INLET FLOWS 6. AUTHOR(S) PROF. PARVIZ MOIN PROF. SANJIVA K. LELE 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) STANFORD... Parviz Moin and Sanjiva K. Lele Stanford University Mechanical Engineering, Flow Physics & Computation Division Stanford, CA 94305-3030 Prepared...monitor. I am thankful to Professor Sanjiva Lele and Profes- sor Parviz Moin, and Keith Lucas for useful discussions! I am grateful to Professor Peter

  2. Supersonic wing design concepts employing nonlinear flows

    NASA Technical Reports Server (NTRS)

    Miller, D. S.

    1984-01-01

    Three nonlinear flow concepts for the design of supersonic wings are reviewed. The specific concepts are: leading-edge thrust, supercritical crossflow, and leading-edge vortex flow. The major results of the experimental-theoretical studies supporting the development of these concepts are presented and discussed. Also, supporting aerodynamic prediction methods are described and example applications are given. Recommendations for further development of each concept are made.

  3. Supersonic Cruise Efficiency - Propulsion Tech Challenge Overview

    NASA Technical Reports Server (NTRS)

    DeBonis James R.

    2007-01-01

    This presentation provides a brief overview of the research underway in the Cruise Efficiency -- Propulsion technical challenge area of NASA's Fundamental Aeronautics Supersonics project. The research involves both computational and experimental efforts in the areas of Advanced Inlet Concepts, High Performance/Wide Operability Fan and Compressors, Advanced Nozzle Concepts and Intelligent Sensors/Actuators. The work consists of both internal NASA research and external efforts funded through the NASA Research Announcement process.

  4. CFD Prediction of Magnus Effect in Subsonic to Supersonic Flight

    DTIC Science & Technology

    2009-09-01

    CFD Prediction of Magnus Effect in Subsonic to Supersonic Flight by James DeSpirito ARL-TR-4929 September 2009...of Magnus Effect in Subsonic to Supersonic Flight James DeSpirito Weapons and Materials Research Directorate, ARL...TITLE AND SUBTITLE CFD Prediction of Magnus Effect in Subsonic to Supersonic Flight 5a. CONTRACT NUMBER 5b. GRANT NUMBER 5c. PROGRAM ELEMENT

  5. Elliptic Length Scales in Laminar, Two-Dimensional Supersonic Flows

    DTIC Science & Technology

    2015-06-01

    AFRL-RQ-WP-TP-2015-0109 ELLIPTIC LENGTH SCALES IN LAMINAR, TWO- DIMENSIONAL SUPERSONIC FLOWS James H. Miller Vehicle Technology Branch...SUBTITLE ELLIPTIC LENGTH SCALES IN LAMINAR, TWO-DIMENSIONAL SUPERSONIC FLOWS 5a. CONTRACT NUMBER In-house 5b. GRANT NUMBER 5c. PROGRAM ELEMENT...ANSI Std. Z39-18 1 Approved for public release; distribution unlimited. Elliptic Length Scales in Laminar, Two-Dimensional Supersonic Flows

  6. Overview of NASA's Supersonic Cruise Efficiency - Propulsion Research

    NASA Technical Reports Server (NTRS)

    DeBonis, James R.

    2009-01-01

    The research in Supersonic Cruise Efficiency Propulsion (SCE-P) Technical Challenge area of NASA's Supersonics project is discussed. The research in SCE-P is being performed to enable efficient supersonic flight over land. Research elements in this area include: Advance Inlet Concepts, High Performance/Wider Operability Fan and Compressor, Advanced Nozzle Concepts, and Intelligent Sensors/Actuators. The research under each of these elements is briefly discussed.

  7. Supersonic Gas-Liquid Cleaning System

    NASA Technical Reports Server (NTRS)

    Kinney, Frank

    1996-01-01

    The Supersonic Gas-Liquid Cleaning System Research Project consisted mainly of a feasibility study, including theoretical and engineering analysis, of a proof-of-concept prototype of this particular cleaning system developed by NASA-KSC. The cleaning system utilizes gas-liquid supersonic nozzles to generate high impingement velocities at the surface of the device to be cleaned. The cleaning fluid being accelerated to these high velocities may consist of any solvent or liquid, including water. Compressed air or any inert gas is used to provide the conveying medium for the liquid, as well as substantially reduce the total amount of liquid needed to perform adequate surface cleaning and cleanliness verification. This type of aqueous cleaning system is considered to be an excellent way of conducting cleaning and cleanliness verification operations as replacements for the use of CFC 113 which must be discontinued by 1995. To utilize this particular cleaning system in various cleaning applications for both the Space Program and the commercial market, it is essential that the cleaning system, especially the supersonic nozzle, be characterized for such applications. This characterization consisted of performing theoretical and engineering analysis, identifying desirable modifications/extensions to the basic concept, evaluating effects of variations in operating parameters, and optimizing hardware design for specific applications.

  8. 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.

  9. Simulating Supersonic Turbulence in Galaxy Formation

    NASA Astrophysics Data System (ADS)

    Scannapieco, E.; Brüggen, M.; Pan, L.

    2012-07-01

    We present a series of simulations that study the role of supersonic turbulence in galaxy formation. First, we describe a detailed numerical study to compare the mixing rate of metals to the thermalization rate of kinetic energy in a supersonic turbulent medium. Our simulations show that the scalar mixing timescale increases modestly with Mach number, as compressible modes are less efficient in enhancing mixing than solenoidal modes. However, since most of kinetic energy is always contained in solenoidal modes, the ratio of mixing and thermalization timescales varies by only a factor of ≍ 2 over a wide range of Mach numbers. Building on these results, we present adaptive mesh simulations of galaxy outflows that use a subgrid model to track turbulent velocities and length scales. Unlike previous approaches, this method allows us to deposit energy from supernovae directly into supersonic turbulence, and we are able to simulate a starbursting galaxy, with realistic radiative cooling throughout the simulation. Pockets of hot, diffuse gas around individual stellar associations sweep up shells of material that persist for long times due to the cooling instability, and this leads to a complex, distribution of bubbles, loops and filaments as observed in outflowing galaxies. The overlapping of hot, rarefied regions leads to a collective central outflow that escapes the galaxy by eating away at the exterior gas through turbulent mixing.

  10. Turbulent mixing noise from supersonic jets

    NASA Technical Reports Server (NTRS)

    Tam, Christopher K. W.; Chen, Ping

    1994-01-01

    There is now a substantial body of theoretical and experimental evidence that the dominant part of the turbulent noise of supersonic jets is generated directly by the large turbulence structures/instability waves of the jet flow. Earlier, Tam and Burton provided a description of the physical mechanism by which supersonically traveling instability waves can generate sound efficiently. They used the method of matched asymptotic expansions to construct an instability wave solution which is valid in the far field. The present work is an extension of the theory of Tam and Burton. It is argued that the instability wave spectrum of the jet may be regarded as generated by stochastic white noise excitation at the nozzle lip region. The reason why the excitation has white noise characteristics is that near the nozzle lip region the flow in the jet mixing layer has no intrinsic length and time scales. The present stochastic wave model theory of supersonic jet noise contains a single unknown multiplicative constant. Comparisons between the calculated noise directivities at selected Strouhal numbers and experimental measurements of a Mach 2 jet at different jet temperatures have been carried out. Favorable agreements are found.

  11. Supersonic Decelerator on 'Right Track' for Future Mars Missions

    NASA Video Gallery

    Project Manager, Mark Adler, and Principal Investigator, Ian Clark describe the innovative testing being conducted by the Low Density Supersonic Decelerator (LDSD) project. Combining very large sup...

  12. Supersonic Air-Breathing Stage For Commercial Launch Rocket

    NASA Technical Reports Server (NTRS)

    Martin, James A.

    1993-01-01

    Concept proposed to expand use of air-breathing, reusable stages to put more payload into orbit at less cost. Stage with supersonic air-breathing engines added to carry expendable stages from subsonic airplane to supersonic velocity. Carry payload to orbit. Expendable stages and payload placed in front of supersonic air-breathing stage. After releasing expendable stages, remotely piloted supersonic air-breathing stage returns to takeoff site and land for reuse. New concept extends use of low-cost reusable hardware and increases payload delivered from B-52.

  13. Analysis of supersonic combustion flow fields with embedded subsonic regions

    NASA Technical Reports Server (NTRS)

    Dash, S.; Delguidice, P.

    1972-01-01

    The viscous characteristic analysis for supersonic chemically reacting flows was extended to include provisions for analyzing embedded subsonic regions. The numerical method developed to analyze this mixed subsonic-supersonic flow fields is described. The boundary conditions are discussed related to the supersonic-subsonic and subsonic-supersonic transition, as well as a heuristic description of several other numerical schemes for analyzing this problem. An analysis of shock waves generated either by pressure mismatch between the injected fluid and surrounding flow or by chemical heat release is also described.

  14. Feasibility of supersonic diode pumped alkali lasers: Model calculations

    SciTech Connect

    Barmashenko, B. D.; Rosenwaks, S.

    2013-04-08

    The feasibility of supersonic operation of diode pumped alkali lasers (DPALs) is studied for Cs and K atoms applying model calculations, based on a semi-analytical model previously used for studying static and subsonic flow DPALs. The operation of supersonic lasers is compared with that measured and modeled in subsonic lasers. The maximum power of supersonic Cs and K lasers is found to be higher than that of subsonic lasers with the same resonator and alkali density at the laser inlet by 25% and 70%, respectively. These results indicate that for scaling-up the power of DPALs, supersonic expansion should be considered.

  15. Supersonic rotation in the Maryland Centrifugal Experiment

    NASA Astrophysics Data System (ADS)

    Messer, Sarah

    The Maryland Centrifugal Experiment (MCX) has been built to study the confinement of supersonically-rotating plasmas and velocity shear stabilization of MHD instabilities. Theory predicts improved stability and confinement when a strong radial electric field is introduced into a magnetic-mirror geometry. The resulting radial currents establish a stable highly sheared plasma rotating at supersonic velocities in the azimuthal direction under the influence of J x B forces. This arrangement leads to increased confinement because the supersonic rotation creates an artificial radial gravity which draws the plasma away from the mirrors, closing the mirror loss cone. The large vφ shear stabilizes the plasma and enforces laminar flow. Based on these concepts, we have designed and constructed a machine to produce supersonically rotating highly-ionized plasmas. It typically does this by introducing a radial voltage of 7 kV in a magnetic-mirror geometry, 2 kG at the midplane and 19 kG at each mirror. MCX has completed its main construction phase and is acquiring data, here analyzed primarily in terms of a circuit model which infers plasma characteristics from the radial voltage across the plasma and the total radial current. The theory and simulations supporting the MCX centrifugal confinement scheme are presented here with the data and analysis from its first nine months of operation, including a description of basic plasma characteristics and evidence for both stability and confinement. Theory simulation, and initial experimental data all indicate that this centrifugal confinement scheme provides good stability and confinement at the temperatures and densities under study, as well as at the larger temperatures, fields, and dimensions expected for a fusion reactor. In particular, spectroscopic and circuit-model data indicate rotational velocities in MCX of up to 100 km/s, ion temperatures of approximately 30 eV, and ion densities upwards of 1020m-3. These parameters give

  16. Low Density Supersonic Decelerator (LDSD) Supersonic Flight Dynamics Test (SFDT) Plume Induced Environment Modelling

    NASA Technical Reports Server (NTRS)

    Mobley, B. L.; Smith, S. D.; Van Norman, J. W.; Muppidi, S.; Clark, I

    2016-01-01

    Provide plume induced heating (radiation & convection) predictions in support of the LDSD thermal design (pre-flight SFDT-1) Predict plume induced aerodynamics in support of flight dynamics, to achieve targeted freestream conditions to test supersonic deceleration technologies (post-flight SFDT-1, pre-flight SFDT-2)

  17. 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.

  18. Supersonic transport vis-a-vis energy savings

    NASA Technical Reports Server (NTRS)

    Cormery, G.

    1979-01-01

    The energy and economic saving modifications in supersonic transportation are studied. Modifications in the propulsion systems and in the aerodynamic configurations of the Concorde aircraft to reduce noise generation and increase fuel efficiency are discussed. The conversion of supersonic aircraft from fuel oils to synthetic fuels is examined.

  19. 14 CFR 91.821 - Civil supersonic airplanes: Noise limits.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... 14 Aeronautics and Space 2 2012-01-01 2012-01-01 false Civil supersonic airplanes: Noise limits... Noise Limits § 91.821 Civil supersonic airplanes: Noise limits. Except for Concorde airplanes having... airplane that does not comply with Stage 2 noise limits of part 36 in effect on October 13, 1977,...

  20. 14 CFR 91.821 - Civil supersonic airplanes: Noise limits.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... 14 Aeronautics and Space 2 2013-01-01 2013-01-01 false Civil supersonic airplanes: Noise limits... Noise Limits § 91.821 Civil supersonic airplanes: Noise limits. Except for Concorde airplanes having... airplane that does not comply with Stage 2 noise limits of part 36 in effect on October 13, 1977,...

  1. 14 CFR 91.821 - Civil supersonic airplanes: Noise limits.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 14 Aeronautics and Space 2 2014-01-01 2014-01-01 false Civil supersonic airplanes: Noise limits... Noise Limits § 91.821 Civil supersonic airplanes: Noise limits. Except for Concorde airplanes having... airplane that does not comply with Stage 2 noise limits of part 36 in effect on October 13, 1977,...

  2. 14 CFR 91.821 - Civil supersonic airplanes: Noise limits.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 2 2011-01-01 2011-01-01 false Civil supersonic airplanes: Noise limits... Noise Limits § 91.821 Civil supersonic airplanes: Noise limits. Except for Concorde airplanes having... airplane that does not comply with Stage 2 noise limits of part 36 in effect on October 13, 1977,...

  3. 14 CFR 91.821 - Civil supersonic airplanes: Noise limits.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 2 2010-01-01 2010-01-01 false Civil supersonic airplanes: Noise limits... Noise Limits § 91.821 Civil supersonic airplanes: Noise limits. Except for Concorde airplanes having... airplane that does not comply with Stage 2 noise limits of part 36 in effect on October 13, 1977,...

  4. Supersonic Flight Dynamics Test 2: Trajectory, Atmosphere, and Aerodynamics Reconstruction

    NASA Technical Reports Server (NTRS)

    Karlgaard, Christopher D.; O'Farrell, Clara; Ginn, Jason M.; Van Norman, John W.

    2016-01-01

    The Supersonic Flight Dynamics Test is a full-scale flight test of aerodynamic decelerator technologies developed by the Low Density Supersonic Decelerator technology demonstration project. The purpose of the project is to develop and mature aerodynamic decelerator technologies for landing large-mass payloads on the surface of Mars. The technologies include a Supersonic Inflatable Aerodynamic Decelerator and supersonic parachutes. The first Supersonic Flight Dynamics Test occurred on June 28th, 2014 at the Pacific Missile Range Facility. The purpose of this test was to validate the test architecture for future tests. The flight was a success and, in addition, was able to acquire data on the aerodynamic performance of the supersonic inflatable decelerator. The Supersonic Disksail parachute developed a tear during deployment. The second flight test occurred on June 8th, 2015, and incorporated a Supersonic Ringsail parachute which was redesigned based on data from the first flight. Again, the inflatable decelerator functioned as predicted but the parachute was damaged during deployment. This paper describes the instrumentation, analysis techniques, and acquired flight test data utilized to reconstruct the vehicle trajectory, main motor thrust, atmosphere, and aerodynamics.

  5. General purpose computer program for interacting supersonic configurations: Programmer's manual

    NASA Technical Reports Server (NTRS)

    Crill, W.; Dale, B.

    1977-01-01

    The program ISCON (Interacting Supersonic Configuration) is described. The program is in support of the problem to generate a numerical procedure for determining the unsteady dynamic forces on interacting wings and tails in supersonic flow. Subroutines are presented along with the complete FORTRAN source listing.

  6. Airloads investigation of an 0.030 scale model of the space shuttle vehicle 140A/B launch configuration (model 47-OTS) in the ARC 9 by 7 foot unitary plan wind tunnel for Mach 1.55 and 2.2 (iA14B), volume 1

    NASA Technical Reports Server (NTRS)

    Gillins, R. L.

    1975-01-01

    Tests were conducted in September, 1973 on a launch vehicle configuration consisting of a space shuttle orbiter, an external tank, two solid rocket boosters, and associated intercomponent attach hardware. Model variables included elevon, rudder, and speed brake deflections. Aerodynamic loads were measured and surface pressure distributions were obtained simultaneously with six-component stability and control force data on the complete configuration. Angles of attack and sideslip were also investigated.

  7. Selected Examples of NACA/NASA Supersonic Flight Research

    NASA Technical Reports Server (NTRS)

    Saltzman, Edwin J.; Ayers, Theodore G.

    1995-01-01

    The present Dryden Flight Research Center, a part of the National Aeronautics and Space Administration, has a flight research history that extends back to the mid-1940's. The parent organization was a part of the National Advisory Committee for Aeronautics and was formed in 1946 as the Muroc Flight Test Unit. This document describes 13 selected examples of important supersonic flight research conducted from the Mojave Desert location of the Dryden Flight Research Center over a 4 decade period beginning in 1946. The research described herein was either obtained at supersonic speeds or enabled subsequent aircraft to penetrate or traverse the supersonic region. In some instances there accrued from these research efforts benefits which are also applicable at lower or higher speed regions. A major consideration in the selection of the various research topics was the lasting impact they have had, or will have, on subsequent supersonic flight vehicle design, efficiency, safety, and performance or upon improved supersonic research techniques.

  8. Line-vortex theory for calculation of supersonic downwash

    NASA Technical Reports Server (NTRS)

    Mirels, Harold; Haefeli, Rudolph C

    1950-01-01

    The perturbation field induced by a line vortex in a supersonic stream and the downwash behind a supersonic lifting surface are examined to establish approximate methods for determining the downwash behind supersonic wings. Lifting-lines methods are presented for calculating supersonic downwash. A bent lifting-line method is proposed for computing the downwash field behind swept wings. When applied to triangular wings with subsonic leading edges, this method gives results that, in general, are in good agreement with the exact linearized solution. An unbent lifting-line method (horseshoe-vortex system) is proposed for unswept wings. This method is applied to determine downwash behind rectangular wings with aspect ratios of 2 and 4. Excellent agreement with exact linearized theory is obtained for both aspect ratios by placing the lifting line at the 1/2-chord point. The use of lifting-lines therefore appears promising for obtaining estimates of the downwash behind supersonic wings.

  9. Coherent Raman spectroscopy for supersonic flow measurments

    NASA Technical Reports Server (NTRS)

    She, C. Y.

    1986-01-01

    In collaboration with NASA/Langley Research Center, a truly nonintrusive and nonseeding method for measuring supersonic molecular flow parameters was proposed and developed at Colorado State University. The feasibility of this Raman Doppler Velocimetry (RDV), currently operated in a scanning mode, was demonstrated not only in a laboratory environment at Colorado State University, but also in a major wind tunnel at NASA/Langley Research Center. The research progress of the RDV development is summarized. In addition, methods of coherent Rayleigh-Brillouin spectroscopy and single-pulse coherent Raman spectroscopy are investigated, respectively, for measurements of high-pressure and turbulent flows.

  10. IPCS implications for future supersonic transport aircraft

    NASA Technical Reports Server (NTRS)

    Billig, L. O.; Kniat, J.; Schmidt, R. D.

    1976-01-01

    The Integrated Propulsion Control System (IPCS) demonstrates control of an entire supersonic propulsion module - inlet, engine afterburner, and nozzle - with an HDC 601 digital computer. The program encompasses the design, build, qualification, and flight testing of control modes, software, and hardware. The flight test vehicle is an F-111E airplane. The L.H. inlet and engine will be operated under control of a digital computer mounted in the weapons bay. A general description and the current status of the IPCS program are given.

  11. Advanced supersonic propulsion study, phase 4

    NASA Technical Reports Server (NTRS)

    Howlett, R. A.

    1977-01-01

    Installation characteristics for a Variable Stream Control Engine (VSCE) were studied for three advanced supersonic airplane designs. Sensitivity of the VSCE concept to change in technology projections was evaluated in terms of impact on overall installed performance. Based on these sensitivity results, critical technology requirements were reviewed, resulting in the reaffirmation of the following requirements: low-noise nozzle system; a high performance, low emissions duct burner and main burner; hot section technology; variable geometry components; and propulsion integration features, including an integrated electronic control system.

  12. Boundary conditions for unsteady supersonic inlet analyses

    NASA Astrophysics Data System (ADS)

    Mayer, David W.; Paynter, Gerald C.

    1994-06-01

    New bleed and compressor face boundary conditions have been developed to improve the accuracy of unsteady supersonic inlet calculations. The new bleed boundary conditions relate changes in the bleed hole discharge coefficient to changes in the local flow conditions; the local bleed flow rate can more than double as a shock moves forward over a bleed band in response to inlet flow disturbances. The effects of inlet flow disturbances on the flow at the compressor face are represented more realistically with this new boundary condition than with traditional fixed static pressure or mass flow conditions.

  13. Design features of a low-disturbance supersonic wind tunnel for transition research at low supersonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Wolf, Stephen W. D.; Laub, James A.; King, Lyndell S.; Reda, Daniel C.

    1992-01-01

    A unique, low-disturbance supersonic wind tunnel is being developed at NASA-Ames to support supersonic laminar flow control research at cruise Mach numbers of the High Speed Civil Transport (HSCT). The distinctive design features of this new quiet tunnel are a low-disturbance settling chamber, laminar boundary layers along the nozzle/test section walls, and steady supersonic diffuser flow. This paper discusses these important aspects of our quiet tunnel design and the studies necessary to support this design. Experimental results from an 1/8th-scale pilot supersonic wind tunnel are presented and discussed in association with theoretical predictions. Natural laminar flow on the test section walls is demonstrated and both settling chamber and supersonic diffuser performance is examined. The full-scale wind tunnel should be commissioned by the end of 1993.

  14. Numerical Simulation of Supersonic Gap Flow

    PubMed Central

    Jing, Xu; Haiming, Huang; Guo, Huang; Song, Mo

    2015-01-01

    Various gaps in the surface of the supersonic aircraft have a significant effect on airflows. In order to predict the effects of attack angle, Mach number and width-to-depth ratio of gap on the local aerodynamic heating environment of supersonic flow, two-dimensional compressible Navier-Stokes equations are solved by the finite volume method, where convective flux of space term adopts the Roe format, and discretization of time term is achieved by 5-step Runge-Kutta algorithm. The numerical results reveal that the heat flux ratio is U-shaped distribution on the gap wall and maximum at the windward corner of the gap. The heat flux ratio decreases as the gap depth and Mach number increase, however, it increases as the attack angle increases. In addition, it is important to find that chamfer in the windward corner can effectively reduce gap effect coefficient. The study will be helpful for the design of the thermal protection system in reentry vehicles. PMID:25635395

  15. Pressure Modulated Sonic Jet in Supersonic Crossflow

    NASA Astrophysics Data System (ADS)

    Rossmann, Tobias

    2014-11-01

    Sonic transverse jets in supersonic crossflow are modulated using high-amplitude variations in jet stagnation pressure to enhance jet penetration and mixing. An injection/modulation apparatus combining a powered resonance tube and acoustic resonator is used to create low momentum ratio jets (J = 1 , 2) in a supersonic cross-stream (M = 3 . 5). The injector has the capability to modulate the jet supply pressure at sufficiently high frequency (> 15 kHz) and amplitude (up to 190 dB) to access relevant Strouhal numbers (St = 0 - 0 . 3) and amplitudes (up to 10% of the jet stagnation pressure) related to mixing enhancement. Planar laser Mie scattering in both side and end views allows for instantaneous imaging of the jet fluid to quantify jet trajectory, spread, and mixing behavior. For modulated J = 2 transverse jets, the recirculation zone directly downstream of the injection location is eliminated and significantly faster centerline signal decay rates are seen. For the J = 1 modulated jets, substantial increases in centerline penetration, jet spread, and centerline signal decay rate are shown. Additionally, PDF analysis of the instantaneous jet fluid signal values is performed to compare local mixing efficiencies between the modulated and un-modulated cases.

  16. Supersonic gas-liquid cleaning system

    NASA Astrophysics Data System (ADS)

    Caimi, Raoul E. B.; Thaxton, Eric A.

    1994-02-01

    A system to perform cleaning and cleanliness verification is being developed to replace solvent flush methods using CFC 113 for fluid system components. The system is designed for two purposes: internal and external cleaning and verification. External cleaning is performed with the nozzle mounted at the end of a wand similar to a conventional pressure washer. Internal cleaning is performed with a variety of fixtures designed for specific applications. Internal cleaning includes tubes, pipes, flex hoses, and active fluid components such as valves and regulators. The system uses gas-liquid supersonic nozzles to generate high impingement velocities at the surface of the object to be cleaned. Compressed air or any inert gas may be used to provide the conveying medium for the liquid. The converging-diverging nozzles accelerate the gas-liquid mixture to supersonic velocities. The liquid being accelerated may be any solvent including water. This system may be used commercially to replace CFC and other solvent cleaning methods widely used to remove dust, dirt, flux, and lubricants. In addition, cleanliness verification can be performed without the solvents which are typically involved. This paper will present the technical details of the system, the results achieved during testing at KSC, and future applications for this system.

  17. Dynamic characteristics of pulsed supersonic fuel sprays

    NASA Astrophysics Data System (ADS)

    Pianthong, K.; Matthujak, A.; Takayama, K.; Milton, B. E.; Behnia, M.

    2008-06-01

    This paper describes the dynamic characteristics of pulsed, supersonic liquid fuel sprays or jets injected into ambient air. Simple, single hole nozzles were employed with the nozzle sac geometries being varied. Different fuel types, diesel fuel, bio-diesel, kerosene, and gasoline were used to determine the effects of fuel properties on the spray characteristics. A vertical two-stage light gas gun was employed as a projectile launcher to provide a high velocity impact to produce the liquid jet. The injection pressure was around 0.88-1.24 GPa in all cases. The pulsed, supersonic fuel sprays were visualized by using a high-speed video camera and shadowgraph method. The spray tip penetration and velocity attenuation and other characteristics were examined and are described here. An instantaneous spray tip velocity of 1,542 m/s (Mach number 4.52) was obtained. However, this spray tip velocity can be sustained for only a very short period (a few microseconds). It then attenuates very quickly. The phenomenon of multiple high frequency spray pulses generated by a single shot impact and the changed in the angle of the shock structure during the spray flight, which had already been observed in previous studies, is again noted. Multiple shock waves from the conical nozzle spray were also clearly captured.

  18. Characterization of supersonic radiation diffusion waves

    NASA Astrophysics Data System (ADS)

    Moore, Alastair S.; Guymer, Thomas M.; Morton, John; Williams, Benjamin; Kline, John L.; Bazin, Nicholas; Bentley, Christopher; Allan, Shelly; Brent, Katie; Comley, Andrew J.; Flippo, Kirk; Cowan, Joseph; Taccetti, J. Martin; Mussack-Tamashiro, Katie; Schmidt, Derek W.; Hamilton, Christopher E.; Obrey, Kimberly; Lanier, Nicholas E.; Workman, Jonathan B.; Stevenson, R. Mark

    2015-07-01

    Supersonic and diffusive radiation flow is an important test problem for the radiative transfer models used in radiation-hydrodynamics computer codes owing to solutions being accessible via analytic and numeric methods. We present experimental results with which we compare these solutions by studying supersonic and diffusive flow in the laboratory. We present results of higher-accuracy experiments than previously possible studying radiation flow through up to 7 high-temperature mean free paths of low-density, chlorine-doped polystyrene foam and silicon dioxide aerogel contained by an Au tube. Measurements of the heat front position and absolute measurements of the x-ray emission arrival at the end of the tube are used to test numerical and analytical models. We find excellent absolute agreement with simulations provided that the opacity and the equation of state are adjusted within expected uncertainties; analytical models provide a good phenomenological match to measurements but are not in quantitative agreement due to their limited scope.

  19. The Aeroacoustics of Slowly Diverging Supersonic Jets

    NASA Technical Reports Server (NTRS)

    Goldstein, M. E.; Leib, S. J.

    2008-01-01

    This paper is concerned with utilizing the acoustic analogy approach to predict the sound from unheated supersonic jets. Previous attempts have been unsuccessful at making such predictions over the Mach number range of practical interest. The present paper, therefore, focuses on implementing the necessary refinements needed to accomplish this objective. The important effects influencing peak supersonic noise turn out to be source convection, mean flow refraction, mean flow amplification, and source non-compactness. It appears that the last two effects have not been adequately dealt with in the literature. The first of these because the usual parallel flow models produce most of the amplification in the so called critical layer where the solution becomes singular and, therefore, causes the predicted sound field to become infinite as well. We deal with this by introducing a new weakly non parallel flow analysis that eliminates the critical layer singularity. This has a strong effect on the shape of the peak noise spectrum. The last effect places severe demands on the source models at the higher Mach numbers because the retarded time variations significantly increase the sensitivity of the radiated sound to the source structure in this case. A highly refined (non-separable) source model is, therefore, introduced in this paper.

  20. Calibration of transonic and supersonic wind tunnels

    NASA Technical Reports Server (NTRS)

    Reed, T. D.; Pope, T. C.; Cooksey, J. M.

    1977-01-01

    State-of-the art instrumentation and procedures for calibrating transonic (0.6 less than M less than 1.4) and supersonic (M less than or equal to 3.5) wind tunnels were reviewed and evaluated. Major emphasis was given to transonic tunnels. Continuous, blowdown and intermittent tunnels were considered. The required measurements of pressure, temperature, flow angularity, noise and humidity were discussed, and the effects of measurement uncertainties were summarized. A comprehensive review of instrumentation currently used to calibrate empty tunnel flow conditions was included. The recent results of relevant research are noted and recommendations for achieving improved data accuracy are made where appropriate. It is concluded, for general testing purposes, that satisfactory calibration measurements can be achieved in both transonic and supersonic tunnels. The goal of calibrating transonic tunnels to within 0.001 in centerline Mach number appears to be feasible with existing instrumentation, provided correct calibration procedures are carefully followed. A comparable accuracy can be achieved off-centerline with carefully designed, conventional probes, except near Mach 1. In the range 0.95 less than M less than 1.05, the laser Doppler velocimeter appears to offer the most promise for improved calibration accuracy off-centerline.

  1. Analysis of Buzz in a Supersonic Inlet

    NASA Technical Reports Server (NTRS)

    Chima, Rodrick V.

    2012-01-01

    A dual-stream, low-boom supersonic inlet designed for use on a small, Mach 1.6 aircraft was tested experimentally in the 8- by 6-Foot Supersonic Wind Tunnel (SWT) at the NASA Glenn Research Center (GRC). The tests showed that the inlet had good recovery and stable operation over large mass flow range. The inlet went into buzz at mass flows well below that needed for engine operation, and the experiments generated a wealth of data during buzz. High frequency response pressure measurements and high-speed schlieren videos were recorded for many buzz events. The objective of the present work was to use computational fluid dynamics (CFD) to predict some of the experimental data taken during buzz, compare those predictions to the experimental data, and to use both datasets to explain the physics of the buzz cycle. The calculations were done with the Wind-US CFD code using a second-order time-accurate differencing scheme and the SST turbulence model. Computed Mach number contours were compared with schlieren images, and ensemble-averaged unsteady pressures were compared to data. The results showed that the buzz cycle consisted partly of spike buzz, an unsteady oscillation of the main shock at the spike tip while the inlet pressure dropped, and partly of choked flow while the inlet repressurized. Most of the results could be explained by theory proposed by Dailey in 1954, but did not support commonly used acoustic resonance explanations.

  2. Supersonic gas-liquid cleaning system

    NASA Technical Reports Server (NTRS)

    Caimi, Raoul E. B.; Thaxton, Eric A.

    1994-01-01

    A system to perform cleaning and cleanliness verification is being developed to replace solvent flush methods using CFC 113 for fluid system components. The system is designed for two purposes: internal and external cleaning and verification. External cleaning is performed with the nozzle mounted at the end of a wand similar to a conventional pressure washer. Internal cleaning is performed with a variety of fixtures designed for specific applications. Internal cleaning includes tubes, pipes, flex hoses, and active fluid components such as valves and regulators. The system uses gas-liquid supersonic nozzles to generate high impingement velocities at the surface of the object to be cleaned. Compressed air or any inert gas may be used to provide the conveying medium for the liquid. The converging-diverging nozzles accelerate the gas-liquid mixture to supersonic velocities. The liquid being accelerated may be any solvent including water. This system may be used commercially to replace CFC and other solvent cleaning methods widely used to remove dust, dirt, flux, and lubricants. In addition, cleanliness verification can be performed without the solvents which are typically involved. This paper will present the technical details of the system, the results achieved during testing at KSC, and future applications for this system.

  3. Economy of flight at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Jones, R. T.

    1976-01-01

    Prandtl's theory is used to determine the airflow over bodies and wings adapted to supersonic flight. By making use of these results, and by incorporating in them an allowance for the probable skin friction, some estimates of expected lift-drag ratios are made for various flight speeds with the best configuration. At each speed a slender body and wings having the best angle of sweepback are considered. For the range of supersonic speeds shown an airplane of normal density and loading would be required to operate at an altitude of the order of 60,000 feet. The limiting value of 1-1/2 times the speed of sound corresponds to a flight speed of 1000 miles per hour. At this speed about 1.5 miles per gallon of fuel are expected. It is interesting to note that this value corresponds to a value of more than 15 miles per gallon when the weight is reduced to correspond to that of an ordinary automobile.

  4. Supersonic Flight Dynamics Test: Trajectory, Atmosphere, and Aerodynamics Reconstruction

    NASA Technical Reports Server (NTRS)

    Kutty, Prasad; Karlgaard, Christopher D.; Blood, Eric M.; O'Farrell, Clara; Ginn, Jason M.; Shoenenberger, Mark; Dutta, Soumyo

    2015-01-01

    The Supersonic Flight Dynamics Test is a full-scale flight test of a Supersonic Inflatable Aerodynamic Decelerator, which is part of the Low Density Supersonic Decelerator technology development project. The purpose of the project is to develop and mature aerodynamic decelerator technologies for landing large mass payloads on the surface of Mars. The technologies include a Supersonic Inflatable Aerodynamic Decelerator and Supersonic Parachutes. The first Supersonic Flight Dynamics Test occurred on June 28th, 2014 at the Pacific Missile Range Facility. This test was used to validate the test architecture for future missions. The flight was a success and, in addition, was able to acquire data on the aerodynamic performance of the supersonic inflatable decelerator. This paper describes the instrumentation, analysis techniques, and acquired flight test data utilized to reconstruct the vehicle trajectory, atmosphere, and aerodynamics. The results of the reconstruction show significantly higher lofting of the trajectory, which can partially be explained by off-nominal booster motor performance. The reconstructed vehicle force and moment coefficients fall well within pre-flight predictions. A parameter identification analysis indicates that the vehicle displayed greater aerodynamic static stability than seen in pre-flight computational predictions and ballistic range tests.

  5. Preliminary Investigation of a New Type of Supersonic Inlet

    NASA Technical Reports Server (NTRS)

    Ferri, Antonio; Nucci, Louis M

    1952-01-01

    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. Experimental results confirming the theoretical analysis give pressure recoveries which vary from 95 percent for Mach number 1.33 to 86 percent for number 2.00. These results were originally presented in a classified document of the NACA in 1946.

  6. Supersonic Test of the 10-Inch Bifurcated Two-Stage Supersonic Inlet (BTSSI)

    NASA Technical Reports Server (NTRS)

    Carlin, C. M.; Frisch, I.; Kolden, J. J.; Koncsek, J. L.; Marrs, K. J.; Mayer, D. W.; McMahon, S. L.; Milligan, K. H.; Sommerfield, D. M.

    2003-01-01

    Experimental results from testing of a novel supersonic inlet model in NASA Glenn Research Center's 10- by 10-foot supersonic wind tunnel are presented. The patented inlet concept, called Two-Stage Supersonic Inlet (TSSI), incorporates a large cavity, or throat slot, in the supersonic diffuser intended to enhance the stability of the normal shock. The present embodiment of the concept is a bifurcated twin-duct) design. During the course of testing an unusual 'semi-started' mode of operation was encountered. The inlet was able to spill up to 30 percent of the captured airstream without fully expelling the normal shock. In this mode, the total pressure recovery dropped approximately 6 percent without increasing steady-state distortion. Dynamic instrumentation at the cowl lip station indicates the semi-start mode may be a series of unstart/restart cycles with frequency ranging from 0.2 to 20 Hz. Engine face total pressure measurements indicate a modest impact due to this event. However, since the current test article does not have a representative subsonic diffuser (and is in fact separated), it is unclear how this mode of operation would effect an engine. Further investigation of this phenomenon is required before it is fully understood. Prior testing of the TSSI concept allowed extension of fully started inlet operation to regions of significantly reduced supply flow without reducing recovery. The test article was a smaller scale than the present test and was a single duct design. In the present test, the expanded range of stable operation with high recovery was not realized.

  7. AOFA- THREE-DIMENSIONAL SUPERSONIC VISCOUS FLOW

    NASA Technical Reports Server (NTRS)

    Rakich, J. V.

    1994-01-01

    This program, which is called 'AOFA', determines the complete viscous and inviscid flow around a body of revolution at a given angle of attack and traveling at supersonic speeds. The viscous calculations from this program agree with experimental values for surface and pitot pressures and with surface heating rates. At high speeds, lee-side flows are important because the local heating is difficult to correlate and because the shed vortices can interact with vehicle components such as a canopy or a vertical tail. This program should find application in the design analysis of any high speed vehicle. Lee-side flows are difficult to calculate because thin-boundary-layer theory is not applicable and the concept of matching inviscid and viscous flow is questionable. This program uses the parabolic approximation to the compressible Navier-Stokes equations and solves for the complete inviscid and viscous regions of flow, including the pressure. The parabolic approximation results from the assumption that the stress derivatives in the streamwise direction are small in comparison with derivatives in the normal and circumferential directions. This assumption permits the equation to be solved by an implicit finite difference marching technique which proceeds downstream from the initial data point, provided the inviscid portion of flow is supersonic. The viscous cross-flow separation is also determined as part of the solution. To use this method it is necessary to first determine an initial data point in a region where the inviscid portion of the flow is supersonic. Input to this program consists of two parts. Problem description is conveyed to the program by namelist input. Initial data is acquired by the program as formatted data. Because of the large amount of run time this program can consume the program includes a restart capability. Output is in printed format and magnetic tape for further processing. This program is written in FORTRAN IV and has been implemented on a CDC

  8. An asymptotic theory of supersonic propeller noise

    NASA Astrophysics Data System (ADS)

    Envia, Edmane

    A theory for predicting the noise field of a propeller with a realistic blade geometry is presented. The theory, which utilizes a large blade count approximation, provides an efficient formula for predicting the radiation of sound from all three sources of propeller noise. Comparisons with full numerical integration indicate that the noise levels predicted by this formula are quite accurate. Calculations based on this method also show that the radiation from the Lighthill quadrupole source is rather substantial when compared with thickness and loading noise for high speed propellers. A preliminary application of the theory to the problem of the sensitivity of the peak noise levels generated by a supersonic propeller to the variations in its tip helical Mach number has produced a trend that is in qualitative agreement with the experimental observations.

  9. Historical development of worldwide supersonic aircraft

    NASA Technical Reports Server (NTRS)

    Spearman, M. L.

    1983-01-01

    Some major milestones in the progression of airplane speeds from subsonic to supersonic are traced. Historical background is included on work done prior to the Twentieth Century, but the major emphasis is on the Twentieth Century developments after the man carrying airplane became a practical reality. The techniques of increasing airplane speed revolve around means of increasing the propulsive force and means of reducing the airframe resistance (drag). With the changes in speed, the attendant changes in flow patterns due to the compressibility of air introduce some aerodynamic problems. In addition, geometric changes introduced to combat the effects of compressibility also promote aerodynamic problems. Some of the solutions to these problems are illustrated, and many design features that evolved are discussed.

  10. The resonance of twin supersonic jets

    NASA Technical Reports Server (NTRS)

    Morris, Philip J.

    1989-01-01

    This paper presents an analytical study of the resonant interaction between twin supersonic jets. An instability wave model is used to describe the large scale coherent structures in the jet mixing layers. A linearized shock cell model is also given for the jets when operating off design. The problem's geometry admits four types of normal modes associated with each azimuthal mode number in the single jet. The stability of these modes is examined for both a vortex sheet model of the jet and a jet flow represented by realistic profiles. The growth rates of each mode number and type are found to vary with jet separation and mixing layer thickness and Strouhal number. Contours of equal pressure level are obtained for each mode. The region close to the symmetry axis is found to have the greatest pressure fluctuation amplitude.

  11. Compressible spreading rates of supersonic coaxial jets

    NASA Technical Reports Server (NTRS)

    Schadow, K. C.; Gutmark, E.; Wilson, K. J.

    1990-01-01

    The compressible spreading rates of two supersonic coaxial jets were studied experimentally. The center jet had a fully expanded Mach number of 3 and the outer jet of M = 1.8. The geometries of the center jet were circular and rectangular with two configurations, both with a 3:1 aspect ratio. The convective Mach numbers Mc were varied in the range between 0.25 and 2.25. The spreading rate of the center circular jet decreased with increasing Mc until it reached a constant value of 0.2 to 0.3 of the incompresible spreading rate for Mc larger than 1.4. The rectangular jets exhibited a similar drop at the same range of Mc, but their spreading rate was higher relative to the circular jet over the entire Mc range.

  12. Design of supersonic Coanda jet nozzles

    NASA Technical Reports Server (NTRS)

    Bevilaqua, Paul M.; Lee, John D.

    1987-01-01

    The thrust vectoring of supersonic Coanda jets was improved by designing a nozzle to skew the initial jet velocity profile. A new nozzle design procedure, based on the method of characteristics, was developed to design a nozzle which produces a specified exit velocity profile. The thrust vectoring of a simple convergent nozzle, a convergent-divergent nozzle, and a nozzle which produces a skewed velocity profile matched to the curvature of the Coanda surface were expermentially compared over a range of pressure ratios from 1.5 to 3.5. Elimination of the expansion shocks with the C-D nozzle is shown to greatly improve the thrust vectoring; elimination of turning shocks with the skewed profile nozzle further improves the vectoring.

  13. Laser fusion cutting using supersonic nozzles

    NASA Astrophysics Data System (ADS)

    Duan, Jun

    This research takes a systematic approach to study the flow patterns of gas jets from the conventional conical (subsonic) nozzle and the newly designed supersonic nozzle under a high-pressure gas regime by a three-dimensional mathematical model. Computer simulation was then confirmed by experimental shadowgraphic technique. The relationships between the type of nozzles, nozzle dimensions, stand-off distance, inlet stagnation pressure and the flow field distribution, incident shock and normal shock have been established. Compared with that of the subsonic nozzle, the gas jet from a supersonic nozzle possesses the desired dynamic characteristics such as uniform distribution, maximum and even momentum thrust and parallel jet boundary under the condition of the designed pressures. Then, a model calculating the three-dimensional stationary geometric shape of the cutting front obtained in the high-pressure laser fusion cutting regime was developed using numerical solution of energy balance. In this model, the energy absorbed by the workpiece includes not only the energy from the laser beam but also the energy from the multiple reflections generated by the beam impinging at the cutting front. The effects of the laser power, cutting speed, focus position, multiple reflections and inert gas pressure on the geometric shape of the cutting front have been analyzed systemically. The geometric shape of the cutting front was then used as boundary conditions in subsequent calculation of the gas flow field distribution inside a laser cut kerf. Finally, a third mathematical model was developed to calculate the distribution of the gas flow field at the entrance of the cut kerf, inside the cut kerf and at the exit of a cut kerf under the condition of a gas jet from a supersonic nozzle and the inlet stagnation pressure ≥5 bar. A two-dimensional analytical method was adopted to locate approximately the position and shape of the detached shock above the cut kerf. A method of two

  14. The noise from supersonic elliptic jets

    NASA Technical Reports Server (NTRS)

    Morris, Philip J.; Bhat, Thonse R. S.

    1992-01-01

    This paper presents calculations of the noise radiated by a supersonic elliptic jet. The large scale structures in the jet, that are the predominant source of noise in the downstream direction, are modeled as instability waves. The evolution of the instability waves is determined by a local, linear, inviscid analysis. An expression is derived for the acoustic field outside the jet and the far field directivity associated with each instability wave. Calculations are performed for a Mach 1.5 elliptic jet with aspect ratio 2:1 and a Mach 2.0 elliptic jet with aspect ratio 2:1 and a Mach 2.0 elliptic jet with aspect ratio 3:1. The mean flow development is taken from experimental results. Comparisons are made with far field acoustic measurements.

  15. An asymptotic theory of supersonic propeller noise

    NASA Technical Reports Server (NTRS)

    Envia, Edmane

    1992-01-01

    A theory for predicting the noise field of a propeller with a realistic blade geometry is presented. The theory, which utilizes a large blade count approximation, provides an efficient formula for predicting the radiation of sound from all three sources of propeller noise. Comparisons with full numerical integration indicate that the noise levels predicted by this formula are quite accurate. Calculations based on this method also show that the radiation from the Lighthill quadrupole source is rather substantial when compared with thickness and loading noise for high speed propellers. A preliminary application of the theory to the problem of the sensitivity of the peak noise levels generated by a supersonic propeller to the variations in its tip helical Mach number has produced a trend that is in qualitative agreement with the experimental observations.

  16. Base pressure in laminar supersonic flow.

    NASA Technical Reports Server (NTRS)

    Messiter, A. F.; Hough, G. R.; Feo, A.

    1973-01-01

    An asymptotic description is proposed for supersonic laminar flow over a wedge or a backward-facing step, for large Reynolds number and for a base or step height which is small compared with the boundary-layer length. The analysis is carried out for adiabatic wall conditions and a viscosity coefficient proportional to temperature. In a particular limit corresponding to a very thick boundary layer, a similarity law is obtained for the base pressure. For a thinner boundary layer an asymptotic form for the base pressure is obtained which shows the dependence on the parameters explicitly and which permits good agreement with experiment. This latter result is based on an inviscid-flow approximation for the corner expansion and for reattachment with viscous forces important primarily in a thin sublayer about the dividing streamline. A prediction of the pressure distribution at reattachment is given and the result is compared with experimental pressure distributions.

  17. Feasibility Study of Hydrogen Fueled Supersonic Transport

    NASA Astrophysics Data System (ADS)

    Yuhara, Tatsunori; Rinoie, Kenichi

    Recently, environmental problems have been paid great attention in the world. According to the current report of ICAO, CO2 emissions of aircraft to the total emissions will increase at an average annual rate of 3˜4%. Furthermore, stable supply of oil is becoming difficult. Without exceptions, those are hitting aviation industries. Therefore, in the near future, the time to shift from kerosene to liquid hydrogen (LH2) as fuel of aircraft will come. An idea of supersonic transport (SST) fueled by LH2 as one of the next-generation SST has been proposed by JAXA. The purpose of this study is to compare the weight, aerodynamics and cost of hydrogen fueled SST with those of kerosene fueled SST and to discuss its feasibilities. A possibility of two-bubble fuselage configuration is also evaluated.

  18. Supersonic projectile models for asynchronous shooter localization

    NASA Astrophysics Data System (ADS)

    Kozick, Richard J.; Whipps, Gene T.; Ash, Joshua N.

    2011-06-01

    In this work we consider the localization of a gunshot using a distributed sensor network measuring time differences of arrival between a firearm's muzzle blast and the shockwave induced by a supersonic bullet. This so-called MB-SW approach is desirable because time synchronization is not required between the sensors, however it suffers from increased computational complexity and requires knowledge of the bullet's velocity at all points along its trajectory. While the actual velocity profile of a particular gunshot is unknown, one may use a parameterized model for the velocity profile and simultaneously fit the model and localize the shooter. In this paper we study efficient solutions for the localization problem and identify deceleration models that trade off localization accuracy and computational complexity. We also develop a statistical analysis that includes bias due to mismatch between the true and actual deceleration models and covariance due to additive noise.

  19. The gaseous jet in supersonic crossflow

    SciTech Connect

    Heister, S.D.; Karagozian, A.R.

    1989-01-01

    An analytical/numerical model for the deflection and mixing of a single gaseous jet in a supersonic crossflow is presented. The jet cross-section is described in terms of the compressible vortex pair resulting from viscous and impulsive forces acting at the jet periphery, and the vortex pair data are combined with data for the mass and momentum balance along the jet axis in order to model the trajectory and mixing of the injected fluid. A numerical technique is employed to solve for the inviscid outer flow and the position of the bow shock which envelopes the jet. The model is shown to be capable of predicting overall jet penetration (for perfectly or slightly underexpanded jets) to within 10 percent of experimental findings, while requiring only a few seconds of computer time. 24 refs.

  20. Separation of winged vehicles in supersonics

    NASA Astrophysics Data System (ADS)

    Bonnefond, T.; Kharitonov, A. M.; Brodetsky, M. D.; Vasenyov, L. G.; Adamov, N. P.; Derunov, E. K.

    The aim of this paper is to present the aerodynamic interferences occurring during the separation of two vehicles at supersonic speeds. After a description of the flowfield structure between the two stages, an experimental study prepared in cooperation between AEROSPATIALE and the Institute of Theoretical and Applied Mechanics (ITAM), Russian Academy of Sciences, Siberian Division and performed in the T-313 facility at ITAM is presented, including the justification of the models and test conditions. Results are then presented concerning both wall pressure measurements on the models and global aerodynamic coefficients of the two stages (drag, lift and pitching moment). Some comparisons with computations are also shown. Shortcomings of the present study are given in a last part, together with possible future complementary activities on this topic.

  1. Supersonic jet mixing enhancement by vortex generators

    NASA Technical Reports Server (NTRS)

    Samimy, M.; Reeder, M. F.; Zaman, K. B. M. Q.

    1991-01-01

    Experiments were conducted to explore the effects of vortex generators, in the form of tabs projecting normally into the jet, on the mixing and the far-field noise characteristics of a jet. A converging-diverging nozzle with a design Mach number of 1.36 was used in the experiments. The flow regimes from subsonic to highly underexpanded supersonic conditions were studied. One, two, and four tabs were used and some of the findings of previous investigators were examined and confirmed. The tabs eliminated screech noise from moderately overexpanded cases to highly underexpanded cases. Detailed flow visualizations and measurements showed that two tabs bifurcated the jet at all Mach numbers. While the effect of two tabs was persistent and the jet remained bifurcated, the distortions produced by one and four tabs disappeared by a streamwise distance of approximately 16 jet diameters. Two and four tabs significantly increased the entrainment of ambient air into the jet.

  2. Axisymmetric supersonic flow in rotating impellers

    NASA Technical Reports Server (NTRS)

    Goldstein, Arthur W

    1952-01-01

    General equations are developed for isentropic, frictionless, axisymmetric flow in rotating impellers with blade thickness taken into account and with blade forces eliminated in favor of the blade-surface function. It is shown that the total energy of the gas relative to the rotating coordinate system is dependent on the stream function only, and that if the flow upstream of the impeller is vortex-free, a velocity potential exists which is a function of only the radial and axial distances in the impeller. The characteristic equations for supersonic flow are developed and used to investigate flows in several configurations in order to ascertain the effect of variations of the boundary conditions on the internal flow and the work input. Conditions varied are prerotation of the gas, blade turning rate, gas velocity at the blade tips, blade thickness, and sweep of the leading edge.

  3. Gas turbine engine with supersonic compressor

    SciTech Connect

    Roberts, II, William Byron; Lawlor, Shawn P.

    2015-10-20

    A gas turbine engine having a compressor section using blades on a rotor to deliver a gas at supersonic conditions to a stator. The stator includes one or more of aerodynamic ducts that have converging and diverging portions for deceleration of the gas to subsonic conditions and to deliver a high pressure gas to combustors. The aerodynamic ducts include structures for changing the effective contraction ratio to enable starting even when designed for high pressure ratios, and structures for boundary layer control. In an embodiment, aerodynamic ducts are provided having an aspect ratio of two to one (2:1) or more, when viewed in cross-section orthogonal to flow direction at an entrance to the aerodynamic duct.

  4. Thin oblique airfoils at supersonic speed

    NASA Technical Reports Server (NTRS)

    Jone, Robert T

    1946-01-01

    The well-known methods of thin-airfoil theory have been extended to oblique or sweptback airfoils of finite aspect ratio moving at supersonic speeds. The cases considered thus far are symmetrical airfoils at zero lift having plan forms bounded by straight lines. Because of the conical form of the elementary flow fields, the results are comparable in simplicity to the results of the two-dimensional thin-airfoil theory for subsonic speeds. In the case of untapered airfoils swept back behind the Mach cone the pressure distribution at the center section is similar to that given by the Ackeret theory for a straight airfoil. With increasing distance from the center section the distribution approaches the form given by the subsonic-flow theory. The pressure drag is concentrated chiefly at the center section and for long wings a slight negative drag may appear on outboard sections. (author)

  5. Supersonic Wing Optimization Using SpaRibs

    NASA Technical Reports Server (NTRS)

    Locatelli, David; Mulani, Sameer B.; Liu, Qiang; Tamijani, Ali Y.; Kapania, Rakesh K.

    2014-01-01

    This research investigates the advantages of using curvilinear spars and ribs, termed SpaRibs, to design a supersonic aircraft wing-box in comparison to the use of classic design concepts that employ straight spars and ribs. The objective is to achieve a more efficient load-bearing mechanism and to passively control the deformation of the structure under the flight loads. Moreover, the use of SpaRibs broadens the design space and allows for natural frequencies and natural mode shape tailoring. The SpaRibs concept is implemented in a new optimization MATLAB-based framework referred to as EBF3SSWingOpt. This optimization scheme performs both the sizing and the shaping of the internal structural elements, connecting the optimizer with the analysis software. The shape of the SpaRibs is parametrically defined using the so called Linked Shape method. Each set of SpaRibs is placed in a one by one square domain of the natural space. The set of curves is subsequently transformed in the physical space for creating the wing structure geometry layout. The shape of each curve of each set is unique; however, mathematical relations link the curvature in an effort to reduce the number of design variables. The internal structure of a High Speed Commercial Transport aircraft concept developed by Boeing is optimized subjected to stress, subsonic flutter and supersonic flutter constraints. The results show that the use of the SpaRibs allows for the reduction of the aircraft's primary structure weight without violating the constraints. A weight reduction of about 15 percent is observed.

  6. F-16XL Supersonic Laminar Flow Test Flight

    NASA Video Gallery

    An F-16XL aircraft was used by the Dryden Flight Research Center, Edwards, California, in a NASA-wide program to improve laminar airflow on aircraft flying at sustained supersonic speeds. It was th...

  7. 2. VIEW SOUTH OF TRANSONIC WIND TUNNEL BUILDING AND SUPERSONIC ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    2. VIEW SOUTH OF TRANSONIC WIND TUNNEL BUILDING AND SUPERSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  8. 4. VIEW NORTHWEST OF SUPERSONIC WIND TUNNEL BUILDING TO TRANSONIC ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    4. VIEW NORTHWEST OF SUPERSONIC WIND TUNNEL BUILDING TO TRANSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  9. Affordable/Acceptable Supersonic Flight: Is It Near?

    NASA Technical Reports Server (NTRS)

    Darden, Christine M.

    2003-01-01

    The author takes a historical look at supersonic flight and humankind's first encounter with the sonic boom. A review is given from the 1950s to the present of the quest to understand the sonic boom, quantify its disturbance on humans and structures, and minimize its effect through aircraft design and operation. Finally, the author reminds readers that sonic boom is only one factor, though critical, in enabling an economically viable commercial supersonic aircraft.

  10. The challenges and opportunities of supersonic transport propulsion technology

    NASA Technical Reports Server (NTRS)

    Strack, William C.; Morris, Shelby J., Jr.

    1988-01-01

    The major challenges confronting the propulsion community for civil supersonic transport applications are identified: high propulsion system efficiency at both supersonic and subsonic cruise conditions, low-cost fuel with adequate thermal stability at high temperatures, low noise cycles and exhaust systems, low emission combustion systems, and low drag installations. Both past progress and future opportunities are discussed in relation to perceived technology shortfalls for an economically successful airplane that satisfies environmental constraints.

  11. A Multiple Zone Method for Supersonic Tactical Missiles

    DTIC Science & Technology

    1986-06-01

    to missile shapes in supersonic flight. The explicit MacCormack scheme is used to advance interior points while a characteristic ...can be treated using a simple mapping. A sample zone structure for a body- wing - tail configuration is shown in Figure 3. This avoids...surface edges which models the flow in the vincinity of the edge. When the flow normal to the wing edge is sufficiently supersonic , this analysis

  12. Supersonic Flying Qualities Experience Using the SR-71

    NASA Technical Reports Server (NTRS)

    Cox, Timothy H.; Jackson, Dante

    1997-01-01

    Approximately 25 years ago NASA Dryden Flight Research Center, Edwards, California, initiated the evaluation of supersonic handling qualities issues using the XB-70 and the YF-12. Comparison of pilot comments and ratings with some of the classical handling qualities criteria for transport aircraft provided information on the usefulness of these criteria and insight into supersonic flying qualities issues. A second research study has recently been completed which again addressed supersonic flying qualities issues through evaluations of the SR-71 in flight at Mach 3. Additional insight into supersonic flying qualities issues was obtained through pilot ratings and comments. These ratings were compared with existing military specifications and proposed criteria for the High Speed Civil Transport. This paper investigates the disparity between pilot comments and the Neal/Smith criteria through a modification of the technique using vertical speed at the pilot station. The paper specifically addresses the pilot ability to control flightpath and pitch attitude in supersonic flight and pilot displays typical of supersonic maneuvering.

  13. Low-speed aerodynamic test of an axisymmetric supersonic inlet with variable cowl slot

    NASA Technical Reports Server (NTRS)

    Powell, A. G.; Welge, H. R.; Trefny, C. J.

    1985-01-01

    The experimental low-speed aerodynamic characteristics of an axisymmetric mixed-compression supersonic inlet with variable cowl slot are described. The model consisted of the NASA P-inlet centerbody and redesigned cowl with variable cowl slot powered by the JT8D single-stage fan simulator and driven by an air turbine. The model was tested in the NASA Lewis Research Center 9- by 15-foot low-speed tunnel at Mach numbers of 0, 0.1, and 0.2 over a range of flows, cowl slot openings, centerbody positions, and angles of attack. The variable cowl slot was effective in minimizing lip separation at high velocity ratios, showed good steady-state and dynamic distortion characteristics, and had good angle-of-attack tolerance.

  14. A Numerical Comparison of Symmetric and Asymmetric Supersonic Wind Tunnels

    NASA Astrophysics Data System (ADS)

    Clark, Kylen D.

    Supersonic wind tunnels are a vital aspect to the aerospace industry. Both the design and testing processes of different aerospace components often include and depend upon utilization of supersonic test facilities. Engine inlets, wing shapes, and body aerodynamics, to name a few, are aspects of aircraft that are frequently subjected to supersonic conditions in use, and thus often require supersonic wind tunnel testing. There is a need for reliable and repeatable supersonic test facilities in order to help create these vital components. The option of building and using asymmetric supersonic converging-diverging nozzles may be appealing due in part to lower construction costs. There is a need, however, to investigate the differences, if any, in the flow characteristics and performance of asymmetric type supersonic wind tunnels in comparison to symmetric due to the fact that asymmetric configurations of CD nozzle are not as common. A computational fluid dynamics (CFD) study has been conducted on an existing University of Michigan (UM) asymmetric supersonic wind tunnel geometry in order to study the effects of asymmetry on supersonic wind tunnel performance. Simulations were made on both the existing asymmetrical tunnel geometry and two axisymmetric reflections (of differing aspect ratio) of that original tunnel geometry. The Reynolds Averaged Navier Stokes equations are solved via NASAs OVERFLOW code to model flow through these configurations. In this way, information has been gleaned on the effects of asymmetry on supersonic wind tunnel performance. Shock boundary layer interactions are paid particular attention since the test section integrity is greatly dependent upon these interactions. Boundary layer and overall flow characteristics are studied. The RANS study presented in this document shows that the UM asymmetric wind tunnel/nozzle configuration is not as well suited to producing uniform test section flow as that of a symmetric configuration, specifically one

  15. Supersonic Flows in Micron-Sized Geometries

    NASA Astrophysics Data System (ADS)

    Bayt, Robert; Breuer, Kenneth

    1998-11-01

    The results of experiments and numerical simulations of flows in micromachined converging-diverging nozzles are presented. The nozzles are fabricated using deep Reactive Ion Etching (DRIE) and are typically 20-30 microns at the throat with expansion ratios ranging from 5 to 20. The flow channels are 300 microns deep, resulting in a 10:1 or better aspect ratio at the throat. Experimental measurements of mass flow and thrust vs. pressure ratio are presented demonstrating the presence of choked and supersonic flow in the micron-scale gemoetries. Mass flow and thrust efficiencies are also presented and compared with results from two-dimensional Navier-Stokes simulations. It is found that, while the efficiencies are reasonably large (much better than one might expect, considering the small dimension of the nozzles), the boundary layers have a considerable effect, particularly on the thrust efficiency of the device, due to the relatively large displacement thickness which reduces the effective expansion ratio. The boundary layers at the top and bottom of the nozzles also affect the performance, particularly at low Reynolds numbers. Additional experimental and numerical results are also discussed.

  16. Simulating Galaxy Outflows Driven by Supersonic Turbulence

    NASA Astrophysics Data System (ADS)

    Scannapieco, Evan

    Feedback from supernova-driven galaxy outflows plays a crucial role in setting the properties and assembly history of low-mass galaxies. Yet, numerically modeling such outflows has been by plagued by the fact that the underlying supernovae are much too small to be directly resolved, but have cooling times that are much too short to be modeled as thermal input. Here I will present three-dimensional, adaptive mesh simulations that overcome this problem using a subgrid model to deposit supernovae energy directly into supersonic turbulence, which acts on scales much smaller than the simulation grid spacing, but much larger than the particle mean free path. In this way, we are able to simulate a starbursting galaxy modeled after NGC 1569, including realistic radiative cooling throughout the simulation. Unlike in previous approaches, pockets of supernova-driven gas sweep up thick shells of material that persist for long times due to the cooling instability, and the overlapping of high-pressure, rarefied regions leads to a collective central outflow as observed in NGC 1569 and other outflowing starbursts. Such models will be directly comparable to a host of new observational constraints.

  17. Design and Analysis Tools for Supersonic Inlets

    NASA Technical Reports Server (NTRS)

    Slater, John W.; Folk, Thomas C.

    2009-01-01

    Computational tools are being developed for the design and analysis of supersonic inlets. The objective is to update existing tools and provide design and low-order aerodynamic analysis capability for advanced inlet concepts. The Inlet Tools effort includes aspects of creating an electronic database of inlet design information, a document describing inlet design and analysis methods, a geometry model for describing the shape of inlets, and computer tools that implement the geometry model and methods. The geometry model has a set of basic inlet shapes that include pitot, two-dimensional, axisymmetric, and stream-traced inlet shapes. The inlet model divides the inlet flow field into parts that facilitate the design and analysis methods. The inlet geometry model constructs the inlet surfaces through the generation and transformation of planar entities based on key inlet design factors. Future efforts will focus on developing the inlet geometry model, the inlet design and analysis methods, a Fortran 95 code to implement the model and methods. Other computational platforms, such as Java, will also be explored.

  18. Supersonic jet and crossflow interaction: Computational modeling

    NASA Astrophysics Data System (ADS)

    Hassan, Ez; Boles, John; Aono, Hikaru; Davis, Douglas; Shyy, Wei

    2013-02-01

    The supersonic jet-in-crossflow problem which involves shocks, turbulent mixing, and large-scale vortical structures, requires special treatment for turbulence to obtain accurate solutions. Different turbulence modeling techniques are reviewed and compared in terms of their performance in predicting results consistent with the experimental data. Reynolds-averaged Navier-Stokes (RANS) models are limited in prediction of fuel structure due to their inability to accurately capture unsteadiness in the flow. Large eddy simulation (LES) is not yet practical due to prohibitively large grid requirement near the wall. Hybrid RANS/LES can offer reasonable compromise between accuracy and efficiency. The hybrid models are based on various approaches such as explicit blending of RANS and LES, detached eddy simulation (DES), and filter-based multi-scale models. In particular, they can be used to evaluate the turbulent Schmidt number modeling techniques used in jet-in-crossflow simulations. Specifically, an adaptive approach can be devised by utilizing the information obtained from the resolved field to help assign the value of turbulent Schmidt number in the sub-filter field. The adaptive approach combined with the multi-scale model improves the results especially when highly refined grids are needed to resolve small structures involved in the mixing process.

  19. Particle Streak Velocimetry of Supersonic Nozzle Flows

    NASA Technical Reports Server (NTRS)

    Willits, J. D.; Pourpoint, T. L.

    2016-01-01

    A novel velocimetry technique to probe the exhaust flow of a laboratory scale combustor is being developed. The technique combines the advantages of standard particle velocimetry techniques and the ultra-fast imaging capabilities of a streak camera to probe high speed flows near continuously with improved spatial and velocity resolution. This "Particle Streak Velocimetry" technique tracks laser illuminated seed particles at up to 236 picosecond temporal resolution allowing time-resolved measurement of one-dimensional flows exceeding 2000 m/s as are found in rocket nozzles and many other applications. Developmental tests with cold nitrogen have been performed to validate and troubleshoot the technique with supersonic flows of much lower velocity and without background noise due to combusting flow. Flow velocities on the order of 500 m/s have been probed with titanium dioxide particles and a continuous-wave laser diode. Single frame images containing multiple streaks are analyzed to find the average slope of all incident particles corresponding to the centerline axial flow velocity. Long term objectives for these tests are correlation of specific impulse to theoretical combustion predictions and direct comparisons between candidate green fuels and the industry standard, monomethylhydrazine, each tested under identical conditions.

  20. Drag Force Anemometer Used in Supersonic Flow

    NASA Technical Reports Server (NTRS)

    Fralick, Gustave C.

    1998-01-01

    To measure the drag on a flat cantilever beam exposed transversely to a flow field, the drag force anemometer (beam probe) uses strain gauges attached on opposite sides of the base of the beam. This is in contrast to the hot wire anemometer, which depends for its operation on the variation of the convective heat transfer coefficient with velocity. The beam probe retains the high-frequency response (up to 100 kHz) of the hot wire anemometer, but it is more rugged, uses simpler electronics, is relatively easy to calibrate, is inherently temperature compensated, and can be used in supersonic flow. The output of the probe is proportional to the velocity head of the flow, 1/2 rho u(exp 2) (where rho is the fluid density and u is the fluid velocity). By adding a static pressure tap and a thermocouple to measure total temperature, one can determine the Mach number, static temperature, density, and velocity of the flow.

  1. Parametric study of supersonic STOVL flight characteristics

    NASA Technical Reports Server (NTRS)

    Rapp, David C.

    1985-01-01

    A number of different control devices and techniques are evaluated to determine their suitability for increasing the short takeoff performance of a supersonic short-takeoff/vertical landing (STOVL) aircraft. Analysis was based on a rigid-body mathematical model of the General Dynamics E-7, a single engine configuration that utilizes ejectors and thrust deflection for propulsive lift. Alternatives investigated include increased static pitch, the addition of a close-coupled canard, use of boundary layer control to increase the takeoff lift coefficient, and the addition of a vectorable aft fan air nozzle. Other performance studies included the impact of individual E-7 features, the sensitivity to ejector performance, the effect of removing the afterburners, and a determination of optional takeoff and landing transition methods. The results pertain to both the E-7 and other configurations. Several alternatives were not as well suited to the E-7 characteristics as they would be to an alternative configuration, and vice versa. A large amount of supporting data for each analysis is included.

  2. Supersonic gas injector for plasma fueling

    SciTech Connect

    Soukhanovskii, V A; Kugel, H W; Kaita, R; Roquemore, A L; Bell, M; Blanchard, W; Bush, C; Gernhardt, R; Gettelfinger, G; Gray, T; Majeski, R; Menard, J; Provost, T; Sichta, P; Raman, R

    2005-09-30

    A supersonic gas injector (SGI) has been developed for fueling and diagnostic applications on the National Spherical Torus Experiment (NSTX). It is comprised of a graphite converging-diverging Laval nozzle and a commercial piezoelectric gas valve mounted on a movable probe at a low field side midplane port location. Also mounted on the probe is a diagnostic package: a Langmuir probe, two thermocouples and five pickup coils for measuring toroidal, radial, vertical magnetic field components and magnetic fluctuations at the location of the SGI tip. The SGI flow rate is up to 4 x 10{sup 21} particles/s, comparable to conventional NSTX gas injectors. The nozzle operates in a pulsed regime at room temperature and a reservoir gas pressure up to 0.33 MPa. The deuterium jet Mach number of about 4, and the divergence half-angle of 5{sup o}-25{sup o} have been measured in laboratory experiments simulating NSTX environment. In initial NSTX experiments reliable operation of the SGI and all mounted diagnostics at distances 1-20 cm from the plasma separatrix has been demonstrated. The SGI has been used for fueling of ohmic and 2-4 MW NBI heated L- and H-mode plasmas. Fueling efficiency in the range 0.1-0.3 has been obtained from the plasma electron inventory analysis.

  3. Outflows from Supersonically-Moving Pulsars

    NASA Astrophysics Data System (ADS)

    Klingler, Noel; Kargaltsev, Oleg; Rangelov, Blagoy; Pavlov, George

    2015-08-01

    Pulsar wind nebulae (PWNe) are sources of nonthermal X-ray emission and prominent sites of particle acceleration. Among other parameters, the PWN appearance depends on the pulsar velocity. If a pulsar moves with a supersonic speed, the ram pressure exceeds the ambient medium pressure, resulting in a bow shock PWN with a tail behind the pulsar. We report on Chandra observations of extended pulsar tails behind PSR J1509-5850 and J1747-2958 ("the Mouse"), and the discovery of a puzzling outflow (in the J1509-5850 PWN) strongly misaligned with the pulsar's direction of motion. We resolve the structures of the heads of the two PWNe and interpret them in light of pulsar wind models. We perform spatially resolved spectral measurements and find only marginal evidence of cooling in the long tail of PSR J1509-5850. The morphologies of the PWN heads and the extended tails are discussed and compared with those of other bow shock PWNe detected by Chandra. A possible unifying scheme will be discussed.

  4. External-Compression Supersonic Inlet Design Code

    NASA Technical Reports Server (NTRS)

    Slater, John W.

    2011-01-01

    A computer code named SUPIN has been developed to perform aerodynamic design and analysis of external-compression, supersonic inlets. The baseline set of inlets include axisymmetric pitot, two-dimensional single-duct, axisymmetric outward-turning, and two-dimensional bifurcated-duct inlets. The aerodynamic methods are based on low-fidelity analytical and numerical procedures. The geometric methods are based on planar geometry elements. SUPIN has three modes of operation: 1) generate the inlet geometry from a explicit set of geometry information, 2) size and design the inlet geometry and analyze the aerodynamic performance, and 3) compute the aerodynamic performance of a specified inlet geometry. The aerodynamic performance quantities includes inlet flow rates, total pressure recovery, and drag. The geometry output from SUPIN includes inlet dimensions, cross-sectional areas, coordinates of planar profiles, and surface grids suitable for input to grid generators for analysis by computational fluid dynamics (CFD) methods. The input data file for SUPIN and the output file from SUPIN are text (ASCII) files. The surface grid files are output as formatted Plot3D or stereolithography (STL) files. SUPIN executes in batch mode and is available as a Microsoft Windows executable and Fortran95 source code with a makefile for Linux.

  5. Coherent structures in a supersonic complex nozzle

    NASA Astrophysics Data System (ADS)

    Magstadt, Andrew; Berry, Matthew; Glauser, Mark

    2016-11-01

    The jet flow from a complex supersonic nozzle is studied through experimental measurements. The nozzle's geometry is motivated by future engine designs for high-performance civilian and military aircraft. This rectangular jet has a single plane of symmetry, an additional shear layer (referred to as a wall jet), and an aft deck representative of airframe integration. The core flow operates at a Mach number of Mj , c = 1 . 6 , and the wall jet is choked (Mj , w = 1 . 0). This high Reynolds number jet flow is comprised of intense turbulence levels, an intricate shock structure, shear and boundary layers, and powerful corner vortices. In the present study, stereo PIV measurements are simultaneously sampled with high-speed pressure measurements, which are embedded in the aft deck, and far-field acoustics in the anechoic chamber at Syracuse University. Time-resolved schlieren measurements have indicated the existence of strong flow events at high frequencies, at a Strouhal number of St = 3 . 4 . These appear to result from von Kàrmàn vortex shedding within the nozzle and pervade the entire flow and acoustic domain. Proper orthogonal decomposition is applied on the current data to identify coherent structures in the jet and study the influence of this vortex street. AFOSR Turbulence and Transition Program (Grant No. FA9550-15-1-0435) with program managers Dr. I. Leyva and Dr. R. Ponnappan.

  6. Boundary conditions for unsteady supersonic inlet analyses

    NASA Astrophysics Data System (ADS)

    Mayer, David W.; Paynter, Gerald C.

    1994-06-01

    New bleed and compresor face boundary conditions have been developed to improve the accuracy of unsteady supersonic inlet calculations. The new bleed boundary condition relates changes in the bleed hole discharge coefficient to change the local flow conditions; the local bleed flow rate can more than double as a shock moves forward over a bleed band in response to inlet flow disturbances. The stability margin of the inlet is strongly dependent on the throat bleed configuration since the locally rapid increase in bleed flow has a stong effect on the motion of the normal shock. The new compressor face boundary condition accounts for changes in the unsteady flow conditions at the compressor face by specifying the compressor face corrected mass flow or Mach number either as a constant or as a linear function of the stagnation conditions. The effects of inlet flow disturbances on the flow at the compressor face are represented more realistically with this new boundary condition than with traditional fixed static pressure or mass flow conditions. Euler calculations of the dynamic response of an inlet flow to a flow disturbance at the compressor face with 20- and 90-deg throat bleed hole angles are reported. These results indicate that an extra margin of stability for the inlet is obtained with 90-deg bleed holes because the increase in bleed flow rate as the shock moves forward over a bleed is much larger for 90-deg holes than for 20-deg holes.

  7. Future Directions of Supersonic Combustion Research: Air Force/NASA Workshop on Supersonic Combustion

    NASA Technical Reports Server (NTRS)

    Tishkoff, Julian M.; Drummond, J. Philip; Edwards, Tim; Nejad, Abdollah S.

    1997-01-01

    The Air Force Office of Scientific Research, the Air Force Wright Laboratory Aero Propulsion and Power Directorate, and the NASA Langley Research Center held a joint supersonic combustion workshop on 14-16 May 1996. The intent of this meeting was to: (1) examine the current state-of-the-art in hydrocarbon and/or hydrogen fueled scramjet research; (2) define the future direction and needs of basic research in support of scramjet technology; and (3) when appropriate, help transition basic research findings to solve the needs of developmental engineering programs in the area of supersonic combustion and fuels. A series of topical sessions were planned. Opening presentations were designed to focus and encourage group discussion and scientific exchange. The last half-day of the workshop was set aside for group discussion of the issues that were raised during the meeting for defining future research opportunities and directions. The following text attempts to summarize the discussions that took place at the workshop.

  8. Preliminary supersonic flight test evaluation of performance seeking control

    NASA Technical Reports Server (NTRS)

    Orme, John S.; Gilyard, Glenn B.

    1993-01-01

    Digital flight and engine control, powerful onboard computers, and sophisticated controls techniques may improve aircraft performance by maximizing fuel efficiency, maximizing thrust, and extending engine life. An adaptive performance seeking control system for optimizing the quasi-steady state performance of an F-15 aircraft was developed and flight tested. This system has three optimization modes: minimum fuel, maximum thrust, and minimum fan turbine inlet temperature. Tests of the minimum fuel and fan turbine inlet temperature modes were performed at a constant thrust. Supersonic single-engine flight tests of the three modes were conducted using varied after burning power settings. At supersonic conditions, the performance seeking control law optimizes the integrated airframe, inlet, and engine. At subsonic conditions, only the engine is optimized. Supersonic flight tests showed improvements in thrust of 9 percent, increases in fuel savings of 8 percent, and reductions of up to 85 deg R in turbine temperatures for all three modes. The supersonic performance seeking control structure is described and preliminary results of supersonic performance seeking control tests are given. These findings have implications for improving performance of civilian and military aircraft.

  9. 14 CFR 91.819 - Civil supersonic airplanes that do not comply with part 36.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 2 2011-01-01 2011-01-01 false Civil supersonic airplanes that do not... RULES Operating Noise Limits § 91.819 Civil supersonic airplanes that do not comply with part 36. (a) Applicability. This section applies to civil supersonic airplanes that have not been shown to comply with...

  10. The enhancement of the mixing and combustion processes in supersonic flow applied to scramjet engine

    SciTech Connect

    Kopchenov, V.I.; Lomkov, K.E. )

    1992-07-01

    The Reynolds averaged parabolized Navier-Stokes equations are employed for the numerical study of turbulent mixing and combustion of a supersonic hydrogen jet in a supersonic airflow. A one-equation differential turbulence model is utilized. The simplified flame sheet model is employed for the numerical simulation of the supersonic combustion. 24 refs.

  11. Experimental Study of Dual Injections with a Cavity in Supersonic Flow

    NASA Astrophysics Data System (ADS)

    Ukai, T.; Zare-Behtash, H.; Lo, K. H.; Kontis, K.; Obayashi, S.

    The design of supersonic injection systems is a key issue for the development of a supersonic combustion ramjet (scramjet) [1]. Since supersonic flow is present in the combustion chamber of the scramjet, it is difficult to satisfy features which indicate fuel-air mixing and flame stabilisation.

  12. Supersonic Wind Tunnels (Latest Citations from the Aerospace Database)

    NASA Technical Reports Server (NTRS)

    1996-01-01

    The bibliography contains citations concerning the design, construction, operation, performance, and use of supersonic wind tunnels. References cover the design of flow nozzles, diffusers, test sections, and ejectors for tunnels driven by compressed air, high-pressure gases, or cryogenic liquids. Methods for flow calibration, boundary layer control, local and freestream turbulence reduction, and force measurement are discussed. Instrusive and non-intrusive instrumentation, sources of measurement error, and measurement corrections are also covered. The citations also include the testing of inlets, nozzles, airfoils, and other components of aerospace vehicles that must operate supersonically. Comprehensive coverage of wind tunnel force balancing systems, and blowdown and supersonic wind tunnels are covered in separate bibliographies.(Contains 50-250 citations and includes a subject term index and title list.)

  13. Computational Analysis of a Low-Boom Supersonic Inlet

    NASA Technical Reports Server (NTRS)

    Chima, Rodrick V.

    2011-01-01

    A low-boom supersonic inlet was designed for use on a conceptual small supersonic aircraft that would cruise with an over-wing Mach number of 1.7. The inlet was designed to minimize external overpressures, and used a novel bypass duct to divert the highest shock losses around the engine. The Wind-US CFD code was used to predict the effects of capture ratio, struts, bypass design, and angles of attack on inlet performance. The inlet was tested in the 8-ft by 6-ft Supersonic Wind Tunnel at NASA Glenn Research Center. Test results showed that the inlet had excellent performance, with capture ratios near one, a peak core total pressure recovery of 96 percent, and a stable operating range much larger than that of an engine. Predictions generally compared very well with the experimental data, and were used to help interpret some of the experimental results.

  14. Laser target using continuous supersonic jet in vacuum

    NASA Astrophysics Data System (ADS)

    Ogata, Yujin; Takahashi, Kazumasa; Kuwabara, Hajime; Nakajima, Mitsuo; Horioka, Kazuhiko

    2013-11-01

    For an extreme ultraviolet (EUV) light source, Sn plasmas, which can produce a strong radiation with high conversion efficiency at 13.5 nm, attract attention. The purpose of this study is to establish technology to form and recover a tin vapor supersonic jet in vacuum and to make clear the interaction between the supersonic free jet and an incident laser. In order to make the tin vapor, a tantalum crucible is inductively heated up and the radiation loss is compared with the input energy. We observe the shock waves in argon jet emitted from a supersonic nozzle by a Schlieren image. In the future, we will increase the number of shields and heat the tantalum crucible to the boiling point of tin. We are planning to make a study on the interaction process between laser and the free jet.

  15. Supersonic jet engine and method of operating the same

    SciTech Connect

    Klees, G.W.

    1981-10-13

    A supersonic jet engine is disclosed that is capable not only of developing sufficient power to accelerate up to supersonic cruise and maintain efficient operation at supersonic cruise, but also arrange to cruise at subsonic velocities with a relatively low specific fuel consumption. The engine is provided with a variable bypass passageway downstream of the compressor. Flow into the bypass passageway is controlled so that during low power setting the bypass passageway is closed so that all the gaseous flow is directed through the turbine. During higher power settings, the bypass passageway is opened to the extent that a selected portion of the gaseous flow is directed through the bypass passageway to bypass the turbine, so that the corrected flow to the turbine remains substantially constant for both high and low power settings of the engine.

  16. Dual-Pump CARS Development and Application to Supersonic Combustion

    NASA Technical Reports Server (NTRS)

    Magnotti, Gaetano; Cutler, Andrew D.

    2012-01-01

    A dual-pump Coherent Anti-Stokes Raman Spectroscopy (CARS) instrument has been developed to obtain simultaneous measurements of temperature and absolute mole fractions of N2, O2 and H2 in supersonic combustion and generate databases for validation and development of CFD codes. 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. 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 N2 co-flow (mixing case). Results are presented and the effects of the compressibility and of the heat release are discussed.

  17. Results from computational analysis of a mixed compression supersonic inlet

    NASA Technical Reports Server (NTRS)

    Saunders, J. D.; Keith, T. G.

    1991-01-01

    A numerical study was performed to simulate the critical flow through a supersonic inlet. This flow field has many phenomena such as shock waves, strong viscous effects, turbulent boundary layer development, boundary layer separations, and mass flow suction through the walls, (bleed). The computational tools used were two full Navier-Stokes (FNS) codes. The supersonic inlet that was analyzed is the Variable Diameter Centerbody, (VDC), inlet. This inlet is a candidate concept for the next generation supersonic involved effort in generating an efficient grid geometry and specifying boundary conditions, particularly in the bleed region and at the outflow boundary. Results for a critical inlet operation compare favorably to Method of Characteristics predictions and experimental data.

  18. Gas turbine power plant with supersonic shock compression ramps

    DOEpatents

    Lawlor, Shawn P.; Novaresi, Mark A.; Cornelius, Charles C.

    2008-10-14

    A gas turbine engine. The engine is based on the use of a gas turbine driven rotor having a compression ramp traveling at a local supersonic inlet velocity (based on the combination of inlet gas velocity and tangential speed of the ramp) which compresses inlet gas against a stationary sidewall. The supersonic compressor efficiently achieves high compression ratios while utilizing a compact, stabilized gasdynamic flow path. Operated at supersonic speeds, the inlet stabilizes an oblique/normal shock system in the gasdynamic flow path formed between the rim of the rotor, the strakes, and a stationary external housing. Part load efficiency is enhanced by use of a lean pre-mix system, a pre-swirl compressor, and a bypass stream to bleed a portion of the gas after passing through the pre-swirl compressor to the combustion gas outlet. Use of a stationary low NOx combustor provides excellent emissions results.

  19. The effects of temperature on supersonic jet noise emission

    NASA Technical Reports Server (NTRS)

    Seiner, John M.; Ponton, Michael K.; Jansen, Bernard J.; Lagen, Nicholas T.

    1992-01-01

    This paper examines the generation of sound produced by high temperature supersonic jets. In particular, the question of the importance of supersonic instability waves to noise emission is considered relative to the role of Kelvin-Helmholtz (K-H) instability waves. Here, these waves are taken to be synonymous with the Mach emission process. Jet total temperatures from 313 to 1534 K are investigated using an axisymmetric water cooled supersonic nozzle designed for Mach 2. The aerodynamic and acoustic results of this study indicate that the dominant noise contributors are the K-H waves over the entire temperature range. Good agreement between measured and numerically predicted plume properties are obtained and an elliptic nozzle is used to demonstrate reduction of the K-H waves.

  20. Calculation of Supersonic Combustion Using Implicit Schemes

    NASA Technical Reports Server (NTRS)

    Yoon, Seokkwan; Kwak, Dochan (Technical Monitor)

    2003-01-01

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

  1. Supersonic Quadrupole Noise Theory for High-Speed Helicopter Rotors

    NASA Technical Reports Server (NTRS)

    Farassat, F.; Brentner, Kenneth S.

    1997-01-01

    High-speed helicopter rotor impulsive noise prediction is an important problem of aeroacoustics. The deterministic quadrupoles have been shown to contribute significantly to high-speed impulsive (HSI) noise of rotors, particularly when the phenomenon of delocalization occurs. At high rotor-tip speeds, some of the quadrupole sources lie outside the sonic circle and move at supersonic speed. Brentner has given a formulation suitable for efficient prediction of quadrupole noise inside the sonic circle. In this paper, we give a simple formulation based on the acoustic analogy that is valid for both subsonic and supersonic quadrupole noise prediction. Like the formulation of Brentner, the model is exact for an observer in the far field and in the rotor plane and is approximate elsewhere. We give the full analytic derivation of this formulation in the paper. We present the method of implementation on a computer for supersonic quadrupoles using marching cubes for constructing the influence surface (Sigma surface) of an observer space- time variable (x; t). We then present several examples of noise prediction for both subsonic and supersonic quadrupoles. It is shown that in the case of transonic flow over rotor blades, the inclusion of the supersonic quadrupoles improves the prediction of the acoustic pressure signature. We show the equivalence of the new formulation to that of Brentner for subsonic quadrupoles. It is shown that the regions of high quadrupole source strength are primarily produced by the shock surface and the flow over the leading edge of the rotor. The primary role of the supersonic quadrupoles is to increase the width of a strong acoustic signal.

  2. Continuing Validation of Computational Fluid Dynamics for Supersonic Retropropulsion

    NASA Technical Reports Server (NTRS)

    Schauerhamer, Daniel Guy; Trumble, Kerry A.; Kleb, Bil; Carlson, Jan-Renee; Edquist, Karl T.

    2011-01-01

    A large step in the validation of Computational Fluid Dynamics (CFD) for Supersonic Retropropulsion (SRP) is shown through the comparison of three Navier-Stokes solvers (DPLR, FUN3D, and OVERFLOW) and wind tunnel test results. The test was designed specifically for CFD validation and was conducted in the Langley supersonic 4 x4 Unitary Plan Wind Tunnel and includes variations in the number of nozzles, Mach and Reynolds numbers, thrust coefficient, and angles of orientation. Code-to-code and code-to-test comparisons are encouraging and possible error sources are discussed.

  3. Stratospheric aerosol modification by supersonic transport operations with climate implications

    NASA Technical Reports Server (NTRS)

    Toon, O. B.; Turco, R. P.; Pollack, J. B.; Whitten, R. C.; Poppoff, I. G.; Hamill, P.

    1980-01-01

    The potential effects on stratospheric aerosois of supersonic transport emissions of sulfur dioxide gas and submicron size soot granules are estimated. An interactive particle-gas model of the stratospheric aerosol is used to compute particle changes due to exhaust emissions, and an accurate radiation transport model is used to compute the attendant surface temperature changes. It is shown that a fleet of several hundred supersonic aircraft, operating daily at 20 km, could produce about a 20% increase in the concentration of large particles in the stratosphere. Aerosol increases of this magnitude would reduce the global surface temperature by less than 0.01 K.

  4. Subsonic and Supersonic Jet Noise Calculations Using PSE and DNS

    NASA Technical Reports Server (NTRS)

    Balakumar, P.; Owis, Farouk

    1999-01-01

    Noise radiated from a supersonic jet is computed using the Parabolized Stability Equations (PSE) method. The evolution of the instability waves inside the jet is computed using the PSE method and the noise radiated to the far field from these waves is calculated by solving the wave equation using the Fourier transform method. We performed the computations for a cold supersonic jet of Mach number 2.1 which is excited by disturbances with Strouhal numbers St=.2 and .4 and the azimuthal wavenumber m=l. Good agreement in the sound pressure level are observed between the computed and the measured (Troutt and McLaughlin 1980) results.

  5. CARS Temperature Measurements in Turbulent and Supersonic Facilities

    NASA Technical Reports Server (NTRS)

    Jarrett, O., Jr.; Antcliff, R. R.; Smith, M. W.; Cutler, A. D.; Diskin, G. S.; Northam, G. B.

    1991-01-01

    This paper documents the development of the National Aeronautics and Space Administration s (NASA) Langley Research Center ( LaRC) Coherent Antistokes Raman Spectroscopy (CARS) systems for measurements of temperature in a turbulent subsonic or supersonic reacting hydrogen-air environment. Spectra data provides temperature data when compared to a precalculated library of nitrogen CARS spectra. Library validity was confirmed by comparing CARS temperatures derived through the library with three different techniques for determination of the temperature in hydrogen-air combustion and an electrically heated furnace. The CARS system has been used to survey temperature profiles in the simulated flow of a supersonic combustion ramjet (scramjet) model. Measurement results will be discussed.

  6. Extreme shockwave systems in problems of external supersonic aerodynamics

    NASA Astrophysics Data System (ADS)

    Uskov, V. N.; Chernyshov, M. V.

    2014-01-01

    The stationary shockwave systems (the sequences of shocks, isentropic expansion and compression waves), which arise at a planar supersonic flow of perfect inviscid gas around the bodies are investigated theoretically. The domains of the existence of shockwave systems under consideration are found analytically and numerically for the model problems of supersonic aerodynamics (the flow around a single plate, the plate with the frontal shield, polygonal profiles), the parameters of systems are determined, which provide the extrema of the force and thermal loadings as well as of the aerodynamic coefficients of streamlined bodies.

  7. Advanced supersonic technology concept study: Hydrogen fueled configuration

    NASA Technical Reports Server (NTRS)

    Brewer, G. D.

    1974-01-01

    Conceptual designs of hydrogen fueled supersonic transport configurations for the 1990 time period were developed and compared with equivalent technology Jet A-1 fueled vehicles to determine the economic and performance potential of liquid hydrogen as an alternate fuel. Parametric evaluations of supersonic cruise vehicles with varying design and transport mission characteristics established the basis for selecting a preferred configuration which was then studied in greater detail. An assessment was made of the general viability of the selected concept including an evaluation of costs and environmental considerations, i.e., exhaust emissions and sonic boom characteristics. Technology development requirements and suggested implementation schedules are presented.

  8. Numerical simulation of steady supersonic flow. [spatial marching

    NASA Technical Reports Server (NTRS)

    Schiff, L. B.; Steger, J. L.

    1981-01-01

    A noniterative, implicit, space-marching, finite-difference algorithm was developed for the steady thin-layer Navier-Stokes equations in conservation-law form. The numerical algorithm is applicable to steady supersonic viscous flow over bodies of arbitrary shape. In addition, the same code can be used to compute supersonic inviscid flow or three-dimensional boundary layers. Computed results from two-dimensional and three-dimensional versions of the numerical algorithm are in good agreement with those obtained from more costly time-marching techniques.

  9. USM3D Predictions of Supersonic Nozzle Flow

    NASA Technical Reports Server (NTRS)

    Carter, Melissa B.; Elmiligui, Alaa A.; Campbell, Richard L.; Nayani, Sudheer N.

    2014-01-01

    This study focused on the NASA Tetrahedral Unstructured Software System CFD code (USM3D) capability to predict supersonic plume flow. Previous studies, published in 2004 and 2009, investigated USM3D's results versus historical experimental data. This current study continued that comparison however focusing on the use of the volume souring to capture the shear layers and internal shock structure of the plume. This study was conducted using two benchmark axisymmetric supersonic jet experimental data sets. The study showed that with the use of volume sourcing, USM3D was able to capture and model a jet plume's shear layer and internal shock structure.

  10. Advanced supersonic technology and its implications for the future

    NASA Technical Reports Server (NTRS)

    Driver, C.

    1979-01-01

    A brief overview of the NASA Supersonic Cruise Research (SCR) program is presented. The SCR program has identified significant improvements in the areas of aerodynamics, structures, propulsion, noise reduction, takeoff and landing procedures, and advanced configuration concepts. These improvements tend to overcome most of the problems which led to the cancellation of the National SST program. They offer the promise of an advanced SST family of aircraft which are environmentally acceptable, have flexible range-payload capability, and are economically viable. The areas of technology addressed by the SCR program have direct application to advanced military aircraft and to supersonic executive aircraft.

  11. Estimation of additive forces and moments for supersonic inlets

    NASA Technical Reports Server (NTRS)

    Perkins, Stanley C., Jr.; Dillenius, Marnix F. E.

    1991-01-01

    A technique for estimating the additive forces and moments associated with supersonic, external compression inlets as a function of mass flow ratio has been developed. The technique makes use of a low order supersonic paneling method for calculating minimum additive forces at maximum mass flow conditions. A linear relationship between the minimum additive forces and the maximum values for fully blocked flow is employed to obtain the additive forces at a specified mass flow ratio. The method is applicable to two-dimensional inlets at zero or nonzero angle of attack, and to axisymmetric inlets at zero angle of attack. Comparisons with limited available additive drag data indicate fair to good agreement.

  12. Erosion of graphite surface exposed to hot supersonic hydrogen gas

    NASA Technical Reports Server (NTRS)

    Sharma, O. P.

    1972-01-01

    A theoretical model based on laminar boundary layer flow equations is developed to predict the erosion rate of a graphite (AGCarb-101) surface exposed to a hot supersonic stream of hydrogen gas. The supersonic flow in the nozzle outside the boundary layer formed over the surface of the specimen is determined by assuming one-dimensional isentropic conditions. An overall surface reaction rate expression based on the experimental studies by Clarke and Fox is used to describe the interaction of hydrogen with graphite. A satisfactory agreement is found between the results of the computation, and the available experimental data. Some shortcomings of the model, and further possible improvements are discussed.

  13. Two Programs for Supersonic Wing Design and Analysis

    NASA Technical Reports Server (NTRS)

    Mason, W. H.; Rosen, B. S.; Grossman, B.

    1986-01-01

    COREL and W12SC3 useful in aerodynamic design and analysis of wings for supersonic speeds. COREL (Conical Relaxation) program solves nonlinear full potential equation for spanwise section of wing in crossflow plane, and option exists to correct result for nonconical geometry. W12SC3 applies lineartheory panel methods to compute solutions for wing/body configuration. Programs restricted to supersonic flows and useful for many design, analysis, and optimization applications. COREL and W12SC3 written in FORTRAN IV for batch execution.

  14. The impact of emerging technologies on an advanced supersonic transport

    NASA Technical Reports Server (NTRS)

    Driver, C.; Maglieri, D. J.

    1986-01-01

    The effects of advances in propulsion systems, structure and materials, aerodynamics, and systems on the design and development of supersonic transport aircraft are analyzed. Efficient propulsion systems with variable-cycle engines provide the basis for improved propulsion systems; the propulsion efficienies of supersonic and subsonic engines are compared. Material advances consist of long-life damage-tolerant structures, advanced material development, aeroelastic tailoring, and low-cost fabrication. Improvements in the areas of aerodynamics and systems are examined. The environmental problems caused by engine emissions, airport noise, and sonic boom are studied. The characteristics of the aircraft designed to include these technical advances are described.

  15. Computational Aeroelastic Analyses of a Low-Boom Supersonic Configuration

    NASA Technical Reports Server (NTRS)

    Silva, Walter A.; Sanetrik, Mark D.; Chwalowski, Pawel; Connolly, Joseph

    2015-01-01

    An overview of NASA's Commercial Supersonic Technology (CST) Aeroservoelasticity (ASE) element is provided with a focus on recent computational aeroelastic analyses of a low-boom supersonic configuration developed by Lockheed-Martin and referred to as the N+2 configuration. The overview includes details of the computational models developed to date including a linear finite element model (FEM), linear unsteady aerodynamic models, unstructured CFD grids, and CFD-based aeroelastic analyses. In addition, a summary of the work involving the development of aeroelastic reduced-order models (ROMs) and the development of an aero-propulso-servo-elastic (APSE) model is provided.

  16. Two-dimensional unsteady lift problems in supersonic flight

    NASA Technical Reports Server (NTRS)

    Heaslet, Max A; Lomax, Harvard

    1949-01-01

    The variation of pressure distribution is calculated for a two-dimensional supersonic airfoil either experiencing a sudden angle-of-attack change or entering a sharp-edge gust. From these pressure distributions the indicial lift functions applicable to unsteady lift problems are determined for two cases. Results are presented which permit the determination of maximum increment in lift coefficient attained by an unrestrained airfoil during its flight through a gust. As an application of these results, the minimum altitude for safe flight through a specific gust is calculated for a particular supersonic wing of given strength and wing loading.

  17. On the Comparison of the Long Penetration Mode (LPM) Supersonic Counterflowing Jet to the Supersonic Screech Jet

    NASA Technical Reports Server (NTRS)

    Farr, Rebecca A.; Chang, Chau-Lyan; Jones, Jess H.; Dougherty, N. Sam

    2015-01-01

    Classic tonal screech noise created by under-expanded supersonic jets; Long Penetration Mode (LPM) supersonic phenomenon -Under-expanded counter-flowing jet in supersonic free stream -Demonstrated in several wind tunnel tests -Modeled in several computational fluid dynamics (CFD) simulations; Discussion of LPM acoustics feedback and fluid interactions -Analogous to the aero-acoustics interactions seen in screech jets; Lessons Learned: Applying certain methodologies to LPM -Developed and successfully demonstrated in the study of screech jets -Discussion of mechanically induced excitation in fluid oscillators in general; Conclusions -Large body of work done on jet screech, other aero-acoustic phenomenacan have direct application to the study and applications of LPM cold flow jets

  18. Shock position sensor for supersonic inlets. [measuring pressure in the throat of a supersonic inlet

    NASA Technical Reports Server (NTRS)

    Dustin, M. O. (Inventor)

    1975-01-01

    Static pressure taps or ports are provided in the throat of a supersonic inlet, and signals indicative of the pressure at each of the ports is fed to respective comparators. Means are also provided for directing a signal indicative of the total throat pressure to the comparators. A periodic signal is superimposed on the total throat pressure so that the signal from the static pressure tabs is compared to a varying scan signal rather than to total throat pressure only. This type of comparison causes each comparator to provide a pulse width modulated output which may vary from 0% 'time on' to 100% 'time on'. The pulse width modulated outputs of the comparators are summed, filtered, and directed to a controller which operates a bypass valve such as a door whereby air is dumped from the inlet to prevent the shock wave from being expelled out the front.

  19. Final Report for the Advanced Concept Studies for Supersonic Commercial Transports Entering Service in the 2030 to 2035 Period, N+3 Supersonic Program

    NASA Technical Reports Server (NTRS)

    Morgenstern, John; Norstrud, Nicole; Stelmack, Marc; Skoch, Craig

    2010-01-01

    The N+3 Final Report documents the work and progress made by Lockheed Martin Aeronautics in response to the NASA sponsored program "N+3 NRA Advanced Concept Studies for Supersonic Commercial Transports Entering Service in the 2030 to 2035 Period." The key technical objective of this effort was to generate promising supersonic concepts for the 2030 to 2035 timeframe and to develop plans for maturing the technologies required to make those concepts a reality. The N+3 program is aligned with NASA's Supersonic Project and is focused on providing alternative system-level solutions capable of overcoming the efficiency, environmental, and performance barriers to practical supersonic flight

  20. Experimental study of mixing enhancement using pylon in supersonic flow

    NASA Astrophysics Data System (ADS)

    Vishwakarma, Manmohan; Vaidyanathan, Aravind

    2016-01-01

    The Supersonic Combustion Ramjet (SCRAMJET) engine has been recognized as one of the most promising air breathing propulsion system for the supersonic/hypersonic flight mission requirements. Mixing and combustion of fuel inside scramjet engine is one of the major challenging tasks. In the current study the main focus has been to increase the penetration and mixing of the secondary jet inside the test chamber at supersonic speeds. In view of this, experiments are conducted to evaluate the effect of pylon on the mixing of secondary jet injection into supersonic mainstream flow at Mach 1.65. Two different pylons are investigated and the results are compared with those obtained by normal injection from a flat plate. The mixing studies are performed by varying the height of the pylon while keeping all other parameters the same. The study mainly focused on analyzing the area of spread and penetration depth achieved by different injection schemes based on the respective parameters. The measurements involved Mie scattering visualization and the flow features are analyzed using Schlieren images. The penetration height and spread area are the two parameters that are used for analyzing and comparing the performance of the pylons. It is observed that the secondary jet injection carried out from behind the big pylon resulted in maximum penetration and spread area of the jet as compared to the small pylon geometry. Moreover it is also evident that for obtaining maximum spreading and penetration of the jet, the same needs to be achieved at the injection location.

  1. Two-Dimensional, Supersonic, Linearized Flow with Heat Addition

    NASA Technical Reports Server (NTRS)

    Lomax, Harvard

    1959-01-01

    Calculations are presented for the forces on a thin supersonic wing underneath which the air is heated. The analysis is limited principally to linearized theory but nonlinear effects are considered. It is shown that significant advantages to external heating would exist if the heat were added well below and ahead of the wing.

  2. Noise from Supersonic Coaxial Jets. Part 2; Normal Velocity Profile

    NASA Technical Reports Server (NTRS)

    Dahl, M. D.; Morris, P. J.

    1997-01-01

    Instability waves have been established as noise generators in supersonic jets. Recent analysis of these slowly diverging jets has shown that these instability waves radiate noise to the far field when the waves have components with phase velocities that are supersonic relative to the ambient speed of sound. This instability wave noise generation model has been applied to supersonic jets with a single shear layer and is now applied to supersonic coaxial jets with two initial shear layers. In this paper the case of coaxial jets with normal velocity profiles is considered, where the inner jet stream velocity is higher than the outer jet stream velocity. To provide mean flow profiles at all axial locations, a numerical scheme is used to calculate the mean flow properties. Calculations are made for the stability characteristics in the coaxial jet shear layers and the noise radiated from the instability waves for different operating conditions with the same total thrust, mass flow and exit area as a single reference jet. The effects of changes in the velocity ratio, the density ratio and the area ratio are each considered independently.

  3. Science and Technology Text Mining: Hypersonic and Supersonic Flow

    DTIC Science & Technology

    2006-05-31

    human analyst. DT was used to derive technical intelligence from a hypersonic/ supersonic flow ( HSF ) database derived from the Science Citation Index and...the Engineering Compendex. Phrase frequency analysis by the technical domain expert provided the pervasive technical themes of the HSF database, and...the phrase proximity analysis provided the relationships among the pervasive technical themes. Bibliometric analysis of the HSF literature

  4. Attachment-line Baseline Model In 20' Supersonic Wind Tunne

    NASA Technical Reports Server (NTRS)

    1995-01-01

    This is the baseline attachment-line model mounted in the 20' supersonic wind tunnel. The tunnel flow is used to approximate an aircraft fuselage with a turbulent boundary layer. Follow-on models will include fairings to improve the flow in the juncture region between the wing and fuselage.

  5. Supersonic flow calculations for a cone with an elliptic flare

    NASA Technical Reports Server (NTRS)

    Lehrhaupt, H.

    1970-01-01

    A three-dimensional supersonic flow program is presented for calculating the flow field about a cone at zero angle of attack with an elliptical flare. The program is irrotational, and results remain valid in the region ahead of the first relected characteristic from the points of shock where the shock is no longer axisymmetric.

  6. Supersonic turbulence and structure of interstellar molecular clouds.

    PubMed

    Boldyrev, Stanislav; Nordlund, Ake; Padoan, Paolo

    2002-07-15

    The interstellar medium provides a unique laboratory for highly supersonic, driven hydrodynamic turbulence. We propose a theory of such turbulence, test it by numerical simulations, and use the results to explain observational scaling properties of interstellar molecular clouds, the regions where stars are born.

  7. Simulations of Direct Current Glow Discharges in Supersonic Air Flow

    NASA Astrophysics Data System (ADS)

    Mahadevan, Shankar; Raja, Laxminarayan

    2008-10-01

    In recent years, there have been a significant number of computational and experimental studies investigating the application of plasma discharges as actuators for high speed flow control. The relative importance of the actuation mechanisms: volumetric heating and electrostatic forcing can be established by developing self-consistent models of the plasma and bulk supersonic flow. To simulate the plasma discharge in a supersonic air stream, a fluid model of the glow discharge is coupled with a compressible Navier-Stokes solver in a self-consistent manner. Source terms for the momentum and energy equations are calculated from the plasma model and input into the Navier-Stokes solver. In turn, the pressure, gas temperature and velocity fields from the Navier-Stokes solution are fed back into the plasma model. The results include plasma species number density contour maps in the absence and presence of Mach 3 supersonic flow, and the corresponding effect of the glow discharge on gas dynamic properties such as the gas pressure and temperature. We also examine the effect of increasing the discharge voltage on the structure of the discharge and its corresponding effect on the supersonic flow.

  8. Titanium honeycomb structure. [for supersonic aircraft wing structure

    NASA Technical Reports Server (NTRS)

    Davis, R. A.; Elrod, S. D.; Lovell, D. T.

    1972-01-01

    A brazed titanium honeycomb sandwich system for supersonic transport wing cover panels provides the most efficient structure spanwise, chordwise, and loadwise. Flutter testing shows that high wing stiffness is most efficient in a sandwich structure. This structure also provides good thermal insulation if liquid fuel is carried in direct contact with the wing structure in integral fuel tanks.

  9. Near-Field Noise Computation for a Supersonic Circular Jet

    NASA Technical Reports Server (NTRS)

    Loh, Ching Y.; Hultgren, Lennart S.

    2005-01-01

    A fully expanded, high-Reynolds-number, supersonic circular jet of Mach number 1.4 is simulated, using a 3-D finite-volume Navier-Stokes solver, with emphasis on the near field noise. The numerical results are generally in good agreement with existing experimental findings.

  10. 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.

  11. NASA F-16XL supersonic laminar flow control program overview

    NASA Technical Reports Server (NTRS)

    Fischer, Michael C.

    1992-01-01

    The viewgraphs and discussion of the NASA supersonic laminar flow control program are provided. Successful application of laminar flow control to a High Speed Civil Transport (HSCT) offers significant benefits in reductions of take-off gross weight, mission fuel burn, cruise drag, structural temperatures, engine size, emissions, and sonic boom. The ultimate economic success of the proposed HSCT may depend on the successful adaption of laminar flow control, which offers the single most significant potential improvements in lift drag ratio (L/D) of all the aerodynamic technologies under consideration. The F-16XL Supersonic Laminar Flow Control (SLFC) Experiment was conceived based on the encouraging results of in-house and NASA supported industry studies to determine if laminar flow control is feasible for the HSCT. The primary objective is to achieve extensive laminar flow (50-60 percent chord) on a highly swept supersonic wing. Data obtained from the flight test will be used to validate existing Euler and Navier Stokes aerodynamic codes and transition prediction boundary layer stability codes. These validated codes and developed design methodology will be delivered to industry for their use in designing supersonic laminar flow control wings. Results from this experiment will establish preliminary suction system design criteria enabling industry to better size the suction system and develop improved estimates of system weight, fuel volume loss due to wing ducting, turbocompressor power requirements, etc. so that benefits and penalties can be more accurately assessed.

  12. On the axisymmetric stability of heated supersonic round jets

    PubMed Central

    2016-01-01

    We perform an inviscid, spatial stability analysis of supersonic, heated round jets with the mean properties assumed uniform on either side of the jet shear layer, modelled here via a cylindrical vortex sheet. Apart from the hydrodynamic Kelvin–Helmholtz (K–H) wave, the spatial growth rates of the acoustically coupled supersonic and subsonic instability waves are computed for axisymmetric conditions (m=0) to analyse their role on the jet stability, under increased heating and compressibility. With the ambient stationary, supersonic instability waves may exist for any jet Mach number Mj≥2, whereas the subsonic instability waves, in addition, require the core-to-ambient flow temperature ratio Tj/To>1. We show, for moderately heated jets at Tj/To>2, the acoustically coupled instability modes, once cut on, to govern the overall jet stability with the K–H wave having disappeared into the cluster of acoustic modes. Sufficiently high heating makes the subsonic modes dominate the jet near-field dynamics, whereas the supersonic instability modes form the primary Mach radiation at far field. PMID:27274691

  13. Wing-Design Program for Subsonic or Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Carlson, H. W.; Walkley, K. B.

    1986-01-01

    Surface of mildest possible camber generated. WINGDES provides analysis, design capability and is applicable to both subsonic and supersonic flows. Optimization carried out for entire wing or for designated leading- and trailing-edge areas, for design of missionadaptive surfaces. WINGDES written in FORTRAN IV.

  14. Commercial Supersonics Technology Project - Status of Airport Noise

    NASA Technical Reports Server (NTRS)

    Bridges, James

    2016-01-01

    The Commercial Supersonic Technology Project has been developing databases, computational tools, and system models to prepare for a level 1 milestone, the Low Noise Propulsion Tech Challenge, to be delivered Sept 2016. Steps taken to prepare for the final validation test are given, including system analysis, code validation, and risk reduction testing.

  15. Modeling of supersonic flows near flying vehicle elements

    NASA Astrophysics Data System (ADS)

    Kovenya, V. M.; Slyunyaev, A. Yu.

    2009-03-01

    Supersonic flows near flying vehicle elements are calculated in the approximation of the full Navier-Stokes equations for a viscous compressible heat-conducting gas with different values of free-stream. Mach and Reynolds numbers and angles of attack. The main laws of the flow near the lifting surface. and in the inlet are obtained.

  16. Characterization of a cold flow non-axisymmetric supersonic ejector

    NASA Astrophysics Data System (ADS)

    Lineberry, David M.

    Experimental investigations of dual and single nozzle non-axisymmetric strut based supersonic ejectors were carried out. The strut nozzles transitioned from a round throat to a square exit with an expansion ratio of 4.6. The ejector system entrained secondary air from the lab and exhausted to the lab at atmospheric pressure. The ejectors were operated at equivalent mass flow rates at primary chamber pressure to back pressure ratios ranging from 6.8 to 61.2 for the single nozzle strut and 3.4 to 30.6 for the dual nozzle strut. Under these conditions both struts demonstrated operation in three distinct regimes: mixed, saturated supersonic and supersonic. Secondary flow choking was demonstrated for both struts at equivalent primary mass flow rates. The mixing length was determined by pressure recovery or equalization with the back pressure. This length remained relatively constant at approximately 20 nozzle hydraulic diameters for the primary mass flow rates in the mixed regime. At higher mass flow rates, the pressure recovery length increased and appeared to be strongly affected by the primary nozzle exit pressure. Surveys of duct exit stagnation pressure indicated poor mixing at high mass flow rates with a supersonic core existing through the mixing duct. Shadow graph images revealed a complex shock structure in the recovery region of the mixing duct. Classical analytical models for axisymmetric ejectors were used to investigate the effect of non-axisymmetric geometry. Preliminary CFD simulations were performed to investigate ejector mixing.

  17. Computation of multi-dimensional viscous supersonic jet flow

    NASA Technical Reports Server (NTRS)

    Kim, Y. N.; Buggeln, R. C.; Mcdonald, H.

    1986-01-01

    A new method has been developed for two- and three-dimensional computations of viscous supersonic flows with embedded subsonic regions adjacent to solid boundaries. The approach employs a reduced form of the Navier-Stokes equations which allows solution as an initial-boundary value problem in space, using an efficient noniterative forward marching algorithm. Numerical instability associated with forward marching algorithms for flows with embedded subsonic regions is avoided by approximation of the reduced form of the Navier-Stokes equations in the subsonic regions of the boundary layers. Supersonic and subsonic portions of the flow field are simultaneously calculated by a consistently split linearized block implicit computational algorithm. The results of computations for a series of test cases relevant to internal supersonic flow is presented and compared with data. Comparison between data and computation are in general excellent thus indicating that the computational technique has great promise as a tool for calculating supersonic flow with embedded subsonic regions. Finally, a User's Manual is presented for the computer code used to perform the calculations.

  18. Equivalent sources method for supersonic intensity of arbitrarily shaped geometries

    NASA Astrophysics Data System (ADS)

    Valdivia, Nicolas P.; Williams, Earl G.; Herdic, Peter C.

    2015-07-01

    Supersonic acoustic intensity is utilized to locate radiating regions on a complex vibrating structure. The supersonic intensity is obtained by a special process that removes the evanescent waves from the near-field acoustical holography measurement. The filtering process is well understood for separable geometries, but unfortunately, there are few results for arbitrarily shaped objects. This work proposes a methodology based on a stable invertible representation of the radiated power operator. The power operator is approximated numerically by the equivalent source formulation and the appropriate complete spectral basis is employed to form the stable invertible operator. The operator is formed with the most efficient radiation modes and these modes are utilized to obtain the supersonic solution for the near-field holographic problem. This concept is tested using numerically generated data in a spherical geometry and the results are validated with the spherical harmonic, supersonic filter. Finally, a vibrating ship-hull structure provides a physical example for application and validation of the proposed methodology in a more complex geometry.

  19. 76 FR 34124 - Civil Supersonic Aircraft Panel Discussion

    Federal Register 2010, 2011, 2012, 2013, 2014

    2011-06-10

    ... From the Federal Register Online via the Government Publishing Office DEPARTMENT OF TRANSPORTATION Federal Aviation Administration Civil Supersonic Aircraft Panel Discussion Correction In notice document 2011-12742 appearing on page 30231 in the issue of Tuesday, May 24, 2011, make the following...

  20. 76 FR 30231 - Civil Supersonic Aircraft Panel Discussion

    Federal Register 2010, 2011, 2012, 2013, 2014

    2011-05-24

    ... for attendees. The purpose of the meeting is to raise public awareness of the continuing technological advancements in supersonic aircraft technology aimed at reducing the intensity of sonic boom. DATES: The public... to all interested parties; however, for building security requirements, please register by June...

  1. Advanced supersonic propulsion study. [with emphasis on noise level reduction

    NASA Technical Reports Server (NTRS)

    Sabatella, J. A. (Editor)

    1974-01-01

    A study was conducted to determine the promising propulsion systems for advanced supersonic transport application, and to identify the critical propulsion technology requirements. It is shown that noise constraints have a major effect on the selection of the various engine types and cycle parameters. Several promising advanced propulsion systems were identified which show the potential of achieving lower levels of sideline jet noise than the first generation supersonic transport systems. The non-afterburning turbojet engine, utilizing a very high level of jet suppression, shows the potential to achieve FAR 36 noise level. The duct-heating turbofan with a low level of jet suppression is the most attractive engine for noise levels from FAR 36 to FAR 36 minus 5 EPNdb, and some series/parallel variable cycle engines show the potential of achieving noise levels down to FAR 36 minus 10 EPNdb with moderate additional penalty. The study also shows that an advanced supersonic commercial transport would benefit appreciably from advanced propulsion technology. The critical propulsion technology needed for a viable supersonic propulsion system, and the required specific propulsion technology programs are outlined.

  2. 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.

  3. Thermal Design and Analysis of the Supersonic Flight Dynamics Test Vehicle for the Low Density Supersonic Decelerator Project

    NASA Technical Reports Server (NTRS)

    Mastropietro, A. J.; Pauken, Michael; Sunada, Eric; Gray, Sandria

    2013-01-01

    The thermal design and analysis of the experimental Supersonic Flight Dynamics Test (SFDT) vehicle is presented. The SFDT vehicle is currently being designed as a platform to help demonstrate key technologies for NASA's Low Density Supersonic Decelerator (LDSD) project. The LDSD project is charged by NASA's Office of the Chief Technologist (OCT) with the task of advancing the state of the art in Mars Entry, Descent, and Landing (EDL) systems by developing and testing three new technologies required for landing heavier payloads on Mars. The enabling technologies under development consist of a large 33.5 meter diameter Supersonic Ringsail (SSRS) parachute and two different types of Supersonic Inflatable Aerodynamic Decelerator (SIAD) devices - a robotic class, SIAD-R, that inflates to a 6 meter diameter torus, and an exploration class, SIAD-E, that inflates to an 8 meter diameter isotensoid. As part of the technology development effort, the various elements of the new supersonic decelerator system must be tested in a Mars-like environment. This is currently planned to be accomplished by sending a series of SFDT vehicles into Earth's stratosphere. Each SFDT vehicle will be lifted to a stable float altitude by a large helium carrier balloon. Once at altitude, the SFDT vehicles will be released from their carrier balloon and spun up via spin motors to provide trajectory stability. An onboard third stage solid rocket motor will propel each test vehicle to supersonic flight in the upper atmosphere. After main engine burnout, each vehicle will be despun and testing of the deceleration system will begin: first an inflatable decelerator will be deployed around the aeroshell to increase the drag surface area, and then the large parachute will be deployed to continue the deceleration and return the vehicle back to the Earth's surface. The SFDT vehicle thermal system must passively protect the vehicle structure and its components from cold temperatures experienced during the

  4. System-Level Experimental Validations for Supersonic Commercial Transport Aircraft Entering Service in the 2018-2020 Time Period

    NASA Technical Reports Server (NTRS)

    Magee, Todd E.; Fugal, Spencer R.; Fink, Lawrence E.; Adamson, Eric E.; Shaw, Stephen G.

    2015-01-01

    completed as a precursor to the selection of the facilities used for validation testing. As facility schedules allowed, the propulsion testing was done at the NASA Glenn Research Center (GRC) 8 x 6-Foot wind tunnel, while boom and force testing was done at the NASA Ames Research Center (ARC) 9 x 7-Foot wind tunnel. During boom testing, a live balance was used for gathering force data. This report is broken down into nine sections. The first technical section (Section 2) covers the general scope of the Phase II activities, goals, a description of the design and testing efforts, and the project plan and schedule. Section 3 covers the details of the propulsion system concepts and design evolution. A series of short tests to evaluate the suitability of different wind tunnels for boom, propulsion, and force testing was also performed under the Phase 2 effort, with the results covered in Section 4. The propulsion integration testing is covered in Section 5 and the boom and force testing in Section 6. CFD comparisons and analyses are included in Section 7. Section 8 includes the conclusions and lessons learned.

  5. Large Scale Turbulent Structures in Supersonic Jets

    NASA Technical Reports Server (NTRS)

    Rao, Ram Mohan; Lundgren, Thomas S.

    1997-01-01

    , turbulent stresses, etc. which will aid in turbulence modeling. This report will be presented in two chapters. The first chapter describes some work on the linear stability of a supersonic round jet and the implications of this for the jet noise problem. The second chapter is an extensive discussion of numerical work using the spectral method which we use to solve the compressible Navier-Stokes equations to study turbulent jet flows. The method uses Fourier expansions in the azimuthal and streamwise direction and a 1-D B-spline basis representation in the radial direction. The B-spline basis is locally supported and this ensures block diagonal matrix equations which can be solved in O(N) steps. This is a modification of a boundary layer code developed by Robert Moser. A very accurate highly resolved Direct Numerical Simulation (DNS) of a turbulent jet flow is produced.

  6. Large Scale Turbulent Structures in Supersonic Jets

    NASA Technical Reports Server (NTRS)

    Rao, Ram Mohan; Lundgren, Thomas S.

    1997-01-01

    velocities, turbulent stresses, etc. which will aid in turbulence modeling. This report will be presented in two chapters. The first chapter describes some work on the linear stability of a supersonic round jet and the implications of this for the jet noise problem. The second chapter is an extensive discussion of numerical work using the spectral method which we use to solve the compressible Navier-Stokes equations to study turbulent jet flows. The method uses Fourier expansions in the azimuthal and streamwise direction and a 1-D B-spline basis representation in the radial direction. The B-spline basis is locally supported and this ensures block diagonal matrix equations which can be solved in O(N) steps. This is a modification of a boundary layer code developed by Robert Moser. A very accurate highly resolved DNS of a turbulent jet flow is produced.

  7. Development and Testing of a New Family of Supersonic Decelerators

    NASA Technical Reports Server (NTRS)

    Clark, Ian G.; Adler, Mark; Rivellini, Tommaso P.

    2013-01-01

    The state of the art in Entry, Descent, and Landing systems for Mars applications is largely based on technologies developed in the late 1960's and early 1970's for the Viking Lander program. Although the 2011 Mars Science Laboratory has made advances in EDL technology, these are predominantly in the areas of entry (new thermal protection systems and guided hypersonic flight) and landing (the sky crane architecture). Increases in entry mass, landed mass, and landed altitude beyond MSL capabilities will require advances predominantly in the field of supersonic decelerators. With this in mind, a multi-year program has been initiated to advance three new types of supersonic decelerators that would enable future large-robotic and human-precursor class missions to Mars.

  8. Finite area method for nonlinear supersonic conical flows

    NASA Technical Reports Server (NTRS)

    Sritharan, S. S.; Seebass, A. R.

    1983-01-01

    A fully conservative numerical method for the computation of steady inviscid supersonic flow about general conical bodies at incidence is described. The procedure utilizes the potential approximation and implements a body conforming mesh generator. The conical potential is assumed to have its best linear variation inside each mesh cell; a secondary interlocking cell system is used to establish the flux balance required to conserve mass. In the supersonic regions the scheme is symmetrized by adding artificial viscosity in conservation form. The algorithm is nearly an order of a magnitude faster than present Euler methods and predicts known results accurately and qualitative features such as nodal point lift off correctly. Results are compared with those of other investigators.

  9. A higher order panel method for linearized supersonic flow

    NASA Technical Reports Server (NTRS)

    Ehlers, F. E.; Epton, M. A.; Johnson, F. T.; Magnus, A. E.; Rubbert, P. E.

    1979-01-01

    The basic integral equations of linearized supersonic theory for an advanced supersonic panel method are derived. Methods using only linear varying source strength over each panel or only quadratic doublet strength over each panel gave good agreement with analytic solutions over cones and zero thickness cambered wings. For three dimensional bodies and wings of general shape, combined source and doublet panels with interior boundary conditions to eliminate the internal perturbations lead to a stable method providing good agreement experiment. A panel system with all edges contiguous resulted from dividing the basic four point non-planar panel into eight triangular subpanels, and the doublet strength was made continuous at all edges by a quadratic distribution over each subpanel. Superinclined panels were developed and tested on s simple nacelle and on an airplane model having engine inlets, with excellent results.

  10. Generalized indical forces on deforming rectangular wings in supersonic flight

    NASA Technical Reports Server (NTRS)

    Lomax, Harvard; Fuller, Franklyn B; Sluder, Loma

    1955-01-01

    A method is presented for determining the time-dependent flow over a rectangular wing moving with a supersonic forward speed and undergoing small vertical distortions expressible as polynomials involving spanwise and chordwise distances. The solution for the velocity potential is presented in a form analogous to that for steady supersonic flow having the familiar "reflected area" concept discovered by Evvard. Particular attention is paid to indicial-type motions and results are expressed in terms of generalized indicial forces. Numerical results for Mach numbers equal to 1.1 and 1.2 are given for polynomials of the first and fifth degree in the chordwise and spanwise directions, respectively, on a wing having an aspect ratio of 4.

  11. Nonlinear potential analysis techniques for supersonic aerodynamic design

    NASA Technical Reports Server (NTRS)

    Shankar, V.; Szema, K. Y.

    1985-01-01

    A numerical method based on the conservation form of the full potential equation has been applied to the problem of three-dimensional supersonic flows with embedded subsonic regions. The governing equation is cast in a nonorthogonal coordinate system, and the theory of characteristics is used to accurately monitor the type-dependent flow field. A conservative switching scheme is employed to transition from the supersonic marching procedure to a subsonic relaxation algorithm and vice versa. The newly developed computer program can handle arbitrary geometries with fuselage, canard, wing, flow through nacelle, vertical tail and wake components at combined angles of attack and sideslip. Results are obtained for a variety of configurations that include a Langley advanced fighter concept with fuselage centerline nacelle, Rockwell's Advanced Tactical Fighter (ATF) with wing mounted nacelles, and the Shuttle Orbiter configuration. Comparisons with available experiments were good.

  12. Application of Doppler Global Velocimetry to Supersonic Flows

    NASA Technical Reports Server (NTRS)

    Meyers, James F.

    1996-01-01

    The design and implementation of Doppler Global Velocimetry (DGV) for testing in the Langley Unitary Plan Wind Tunnel is presented. The discussion begins by outlining the characteristics of the tunnel and the test environment, with potential problem areas highlighted. Modifications to the optical system design to implement solutions for these problems are described. Since this tunnel entry was the first ever use of DGV in a supersonic wind tunnel, the test series was divided into three phases, each with its own goal. Phase I determined if condensation provided sufficient scattered light for DGV applications. Phase II studied particle lag by measuring the flow about an oblique shock above an inclined flat plate. Phase III investigated the supersonic vortical flow field above a 75-degree delta wing at 24-degrees angle of attack. Example results from these tests are presented.

  13. Aerodynamic detuning analysis of an unstalled supersonic turbofan cascade

    NASA Technical Reports Server (NTRS)

    Hoyniak, D.; Fleeter, S.

    1985-01-01

    An approach to passive flutter control is aerodynamic detuning, defined as designed passage-to-passage differences in the unsteady aerodynamic flow field of a rotor blade row. Thus, aerodynamic detuning directly affects the fundamental driving mechanism for flutter. A model to demonstrate the enhanced supersonic aeroelastic stability associated with aerodynamic detuning is developed. The stability of an aerodynamically detuned cascade operating in a supersonic inlet flow field with a subsonic leading edge locus is analyzed, with the aerodynamic detuning accomplished by means of nonuniform circumferential spacing of adjacent rotor blades. The unsteady aerodynamic forces and moments on the blading are defined in terms of influence coefficients in a manner that permits the stability of both a conventional uniformally spaced rotor configuration as well as the detuned nonuniform circumferentially spaced rotor to be determined. With Verdon's uniformly spaced Cascade B as a baseline, this analysis is then utilized to demonstrate the potential enhanced aeroelastic stability associated with this particular type of aerodynamic detuning.

  14. MULTISHOCKED,THREE-DIMENSIONAL SUPERSONIC FLOWFIELDS WITH REAL GAS EFFECTS

    NASA Technical Reports Server (NTRS)

    Kutler, P.

    1994-01-01

    This program determines the supersonic flowfield surrounding three-dimensional wing-body configurations of a delta wing. It was designed to provide the numerical computation of three dimensional inviscid, flowfields of either perfect or real gases about supersonic or hypersonic airplanes. The governing equations in conservation law form are solved by a finite difference method using a second order noncentered algorithm between the body and the outermost shock wave, which is treated as a sharp discontinuity. Secondary shocks which form between these boundaries are captured automatically. The flowfield between the body and outermost shock is treated in a shock capturing fashion and therefore allows for the correct formation of secondary internal shocks . The program operates in batch mode, is in CDC update format, has been implemented on the CDC 7600, and requires more than 140K (octal) word locations.

  15. Supersonic laser-induced jetting of aluminum micro-droplets

    NASA Astrophysics Data System (ADS)

    Zenou, M.; Sa'ar, A.; Kotler, Z.

    2015-05-01

    The droplet velocity and the incubation time of pure aluminum micro-droplets, printed using the method of sub-nanosecond laser induced forward transfer, have been measured indicating the formation of supersonic laser-induced jetting. The incubation time and the droplet velocity were extracted by measuring a transient electrical signal associated with droplet landing on the surface of the acceptor substrate. This technique has been exploited for studying small volume droplets, in the range of 10-100 femto-litters for which supersonic velocities were measured. The results suggest elastic propagation of the droplets across the donor-to-acceptor gap, a nonlinear deposition dynamics on the surface of the acceptor and overall efficient energy transfer from the laser beam to the droplets.

  16. Supersonic flow past pointed-nose thin airfoils

    NASA Astrophysics Data System (ADS)

    Hemdan, Hamdi T.; Selim, Abdulfatah A.

    1992-05-01

    Accurate results for steady inviscid supersonic flow past any pointed-nose thin airfoil at small angle of attack are obtained in this paper. Straightforward perturbation from some basic wedge flow is applied to a recently developed hypersonic theory with second-order terms in the longitudinal curvature included. The hypersonic equations are thus reduced to two systems of simple linear ordinary differential equations for which polynomial solutions of the first and second degree are obtained. Further, the hypersonic theory is extended to include second-order effects on the basic wedge flow. This resulted in great increase in the accuracy of the results for curved airfoils besides, the range of validity of the theory extended to cover moderate supersonic speeds with quite reasonable accuracy.

  17. Supersonic laser-induced jetting of aluminum micro-droplets

    SciTech Connect

    Zenou, M.; Sa'ar, A.; Kotler, Z.

    2015-05-04

    The droplet velocity and the incubation time of pure aluminum micro-droplets, printed using the method of sub-nanosecond laser induced forward transfer, have been measured indicating the formation of supersonic laser-induced jetting. The incubation time and the droplet velocity were extracted by measuring a transient electrical signal associated with droplet landing on the surface of the acceptor substrate. This technique has been exploited for studying small volume droplets, in the range of 10–100 femto-litters for which supersonic velocities were measured. The results suggest elastic propagation of the droplets across the donor-to-acceptor gap, a nonlinear deposition dynamics on the surface of the acceptor and overall efficient energy transfer from the laser beam to the droplets.

  18. High-resolution submillimeter-wave radiometry of supersonic flow

    NASA Technical Reports Server (NTRS)

    Dionne, G. F.; Weiss, J. A.; Fitzgerald, J. F.; Fetterman, H. R.; Litvak, M. M.

    1983-01-01

    The recent development of a high-resolution submillimeter-wave heterodyne radiometer has made possible the first measurements of H2O molecule rotational line excitation temperatures and detailed profiles in supersonic flow. Absorption signals were measured across the flow for the 2/11/ from 2//02/ (752 GHz) para-H2O rotational transition against a hot background. These signals decrease downstream owing to the volume expansion of the gas away from the sonic nozle exit in the high-vacuum chamber. Radiative transfer calculations based on the large-velocity-gradient approximation and multilevel statistical equilibrium agree with these results and with the measured spectral line shapes. The data reveal nearly isentropic gas expansion and cooling. These studies have shown that submillimeter-wave heterodyne radiometry can be useful for remote sensing of supersonic flow with low mass flux, provided the signal transmission is through a dry or thin atmosphere.

  19. Volume Dynamics Propulsion System Modeling for Supersonics Vehicle Research

    NASA Technical Reports Server (NTRS)

    Kopasakis, George; Connolly, Joseph W.; Paxson, Daniel E.; Ma, Peter

    2008-01-01

    Under the NASA Fundamental Aeronautics Program, the Supersonics Project is working to overcome the obstacles to supersonic commercial flight. The proposed vehicles are long slim body aircraft with pronounced aero-servo-elastic modes. These modes can potentially couple with propulsion system dynamics; leading to performance challenges such as aircraft ride quality and stability. Other disturbances upstream of the engine generated from atmospheric wind gusts, angle of attack, and yaw can have similar effects. In addition, for optimal propulsion system performance, normal inlet-engine operations are required to be closer to compressor stall and inlet unstart. To study these phenomena an integrated model is needed that includes both airframe structural dynamics as well as the propulsion system dynamics. This paper covers the propulsion system component volume dynamics modeling of a turbojet engine that will be used for an integrated vehicle Aero-Propulso-Servo-Elastic model and for propulsion efficiency studies.

  20. Flight Research and Validation Formerly Experimental Capabilities Supersonic Project

    NASA Technical Reports Server (NTRS)

    Banks, Daniel

    2009-01-01

    This slide presentation reviews the work of the Experimental Capabilities Supersonic project, that is being reorganized into Flight Research and Validation. The work of Experimental Capabilities Project in FY '09 is reviewed, and the specific centers that is assigned to do the work is given. The portfolio of the newly formed Flight Research and Validation (FRV) group is also reviewed. The various projects for FY '10 for the FRV are detailed. These projects include: Eagle Probe, Channeled Centerbody Inlet Experiment (CCIE), Supersonic Boundary layer Transition test (SBLT), Aero-elastic Test Wing-2 (ATW-2), G-V External Vision Systems (G5 XVS), Air-to-Air Schlieren (A2A), In Flight Background Oriented Schlieren (BOS), Dynamic Inertia Measurement Technique (DIM), and Advanced In-Flight IR Thermography (AIR-T).

  1. Development of a plasma injector for supersonic drag reduction

    NASA Astrophysics Data System (ADS)

    Macnab, Angus; Hashim, Akel; Graham, Robert; Ziemba, Timothy; Miller, Kenneth

    2013-10-01

    Eagle Harbor Technologies, Inc. is developing and optimizing a magnetohydrodynamic (MHD) plasma injector, designed to reduce viscous skin friction of supersonic aircraft. The broad goals are to 1. Computationally investigate and verify the dominant physical mechanisms for MHD plasma drag reduction; 2. Develop a proof of concept plasma injector demo, which performs within the power limitations of an onboard flight-relevant system; and 3. Use insights gained through computational investigations to optimize the performance of our MHD plasma injector for maximum efficiency. This investigation focuses on flight-relevant Reynolds and magnetic Reynolds numbers at low supersonic (1

  2. User's manual: Subsonic/supersonic advanced panel pilot code

    NASA Technical Reports Server (NTRS)

    Moran, J.; Tinoco, E. N.; Johnson, F. T.

    1978-01-01

    Sufficient instructions for running the subsonic/supersonic advanced panel pilot code were developed. This software was developed as a vehicle for numerical experimentation and it should not be construed to represent a finished production program. The pilot code is based on a higher order panel method using linearly varying source and quadratically varying doublet distributions for computing both linearized supersonic and subsonic flow over arbitrary wings and bodies. This user's manual contains complete input and output descriptions. A brief description of the method is given as well as practical instructions for proper configurations modeling. Computed results are also included to demonstrate some of the capabilities of the pilot code. The computer program is written in FORTRAN IV for the SCOPE 3.4.4 operations system of the Ames CDC 7600 computer. The program uses overlay structure and thirteen disk files, and it requires approximately 132000 (Octal) central memory words.

  3. Development of Supersonic Combustion Experiments for CFD Modeling

    NASA Technical Reports Server (NTRS)

    Baurle, Robert; Bivolaru, Daniel; Tedder, Sarah; Danehy, Paul M.; Cutler, Andrew D.; Magnotti, Gaetano

    2007-01-01

    This paper describes the development of an experiment to acquire data for developing and validating computational fluid dynamics (CFD) models for turbulence in supersonic combusting flows. The intent is that the flow field would be simple yet relevant to flows within hypersonic air-breathing engine combustors undergoing testing in vitiated-air ground-testing facilities. Specifically, it describes development of laboratory-scale hardware to produce a supersonic combusting coaxial jet, discusses design calculations, operability and types of flames observed. These flames are studied using the dual-pump coherent anti- Stokes Raman spectroscopy (CARS) - interferometric Rayleigh scattering (IRS) technique. This technique simultaneously and instantaneously measures temperature, composition, and velocity in the flow, from which many of the important turbulence statistics can be found. Some preliminary CARS data are presented.

  4. Minimum energy, liquid hydrogen supersonic cruise vehicle study

    NASA Technical Reports Server (NTRS)

    Brewer, G. D.; Morris, R. E.

    1975-01-01

    The potential was examined of hydrogen-fueled supersonic vehicles designed for cruise at Mach 2.7 and at Mach 2.2. The aerodynamic, weight, and propulsion characteristics of a previously established design of a LH2 fueled, Mach 2.7 supersonic cruise vehicle (SCV) were critically reviewed and updated. The design of a Mach 2.2 SCV was established on a corresponding basis. These baseline designs were then studied to determine the potential of minimizing energy expenditure in performing their design mission, and to explore the effect of fuel price and noise restriction on their design and operating performance. The baseline designs of LH2 fueled aircraft were than compared with equivalent designs of jet A (conventional hydrocarbon) fueled SCV's. Use of liquid hydrogen for fuel for the subject aircraft provides significant advantages in performance, cost, noise, pollution, sonic boom, and energy utilization.

  5. Supersonic burning in separated flow regions. final report

    SciTech Connect

    Zumwalt, G.W.

    1982-04-01

    The trough vortex phenomena is used for combustion of hydrogen in a supersonic air stream. This was done in small sizes suitable for igniters in supersonic combustion ramjets so long as the boundary layer displacement thickness is less than 25% of the trough step height. A simple electric spark, properly positioned, ignites the hydrogen in the trough corner. The resulting flame is self sustaining and reignitable. Hydrogen can be injected at the base wall or immediately upstream of the trough. The hydrogen is introduced at low velocity to permit it to be drawn into the corner vortex system and thus experience a long residence time in the combustion region. The igniters can be placed on a skewed back step for angles at least up to 30 deg. without affecting the igniter performance significantly. Certain metals (platinum, copper) act catalytically to improve ignition.

  6. Off-Design Reynolds Number Effects for a Supersonic Transport

    NASA Technical Reports Server (NTRS)

    Owens, Lewis R.; Wahls, Richard A.; Rivers, S. Melissa

    2005-01-01

    A high Reynolds number wind tunnel test was conducted to assess Reynolds number effects on the aerodynamic performance characteristics of a realistic, second-generation supersonic transport concept. The tests included longitudinal studies at transonic and low-speed, high-lift conditions across a range of chord Reynolds numbers (8 million to 120 million). Results presented focus on Reynolds number and static aeroelastic sensitivities at Mach 0.30 and 0.90 for a configuration without a tail. Static aeroelastic effects, which mask Reynolds number effects, were observed. Reynolds number effects were generally small and the drag data followed established trends of skin friction as a function of Reynolds number. A more nose-down pitching moment was produced as Reynolds number increased because of an outward movement of the inboard leading-edge separation at constant angles of attack. This study extends the existing Reynolds number database for supersonic transports operating at off-design conditions.

  7. The inviscid stability of supersonic flow past a sharp cone

    NASA Technical Reports Server (NTRS)

    Duck, Peter W.; Shaw, Stephen J.

    1990-01-01

    The effects of lateral curvature on the development of supersonic laminar inviscid boundary-layer flow on a sharp cone with adiabatic wall conditions are investigated analytically, with a focus on the linear temporal inviscid stability properties. The derivation of the governing equations and of a 'triply generalized' inflexion condition is outlined, and numerical results for freestream Mach number 3.8 are presented in extensive graphs and characterized in detail. A third instability mode related to the viscous mode observed by Duck and Hall (1990) using triple-deck theory is detected and shown to be more unstable and to have larger growth rates than the second mode in some cases. It is found that the 'sonic' neutral mode is affected by the lateral curvature and becomes a supersonic neutral mode.

  8. Prediction, Measurement, and Suppression of High Temperature Supersonic Jet Noise

    NASA Technical Reports Server (NTRS)

    Seiner, John M.; Bhat, T. R. S.; Jansen, Bernard J.

    1999-01-01

    The photograph in figure 1 displays a water cooled round convergent-divergent supersonic nozzle operating slightly overexpanded near 2460 F. The nozzle is designed to produce shock free flow near this temperature at Mach 2. The exit diameter of this nozzle is 3.5 inches. This nozzle is used in the present study to establish properties of the sound field associated with high temperature supersonic jets operating fully pressure balanced (i.e. shock free) and to evaluate capability of the compressible Rayleigh model to account for principle physical features of the observed sound emission. The experiment is conducted statically (i.e. M(sub f) = 0.) in the NASA/LaRC Jet Noise Laboratory. Both aerodynamic and acoustic measurements are obtained in this study along with numerical plume simulation and theoretical prediction of jet noise. Detailed results from this study are reported previously by Seiner, Ponton, Jansen, and Lagen.

  9. Prediction of jet flows from the axisymmetric supersonic nozzle

    NASA Astrophysics Data System (ADS)

    Liu, Y.; Kendall, M. A. F.; Costigan, G.; Bellhouse, B. J.

    This study is motivated by the authors' interest in developing a needle-free powdered vaccine delivery device, the Epidermal Powdered Injection system(EPI). The behaviour of a supersonic jet, which accelerates powdered vaccines in micro-form to a sufficient momentum to penetrate the outer layer of human skin or mucosal tissue, is therefore of great importance. In this paper, a well established Modified Implicit Flux Vector Splitting (MIFVS) solver for the Navier-Stokes equations is extended to numerically study the transient supersonic jet flows of interest. A low Reynolds number k-ɛ turbulence model, with the compressibility effect considered, is integrated into MIFVS solver to the prediction of the turbulent structures and interactions with inherent shock systems. The results for the NASA validation case NPARC, Contoured Shock Tube and Venturi of EPI system are discussed.

  10. Volume Dynamics Propulsion System Modeling for Supersonics Vehicle Research

    NASA Technical Reports Server (NTRS)

    Kopasakis, George; Connolly, Joseph W.; Paxson, Daniel E.; Ma, Peter

    2008-01-01

    Under the NASA Fundamental Aeronautics Program the Supersonics Project is working to overcome the obstacles to supersonic commercial flight. The proposed vehicles are long slim body aircraft with pronounced aero-servo-elastic modes. These modes can potentially couple with propulsion system dynamics; leading to performance challenges such as aircraft ride quality and stability. Other disturbances upstream of the engine generated from atmospheric wind gusts, angle of attack, and yaw can have similar effects. In addition, for optimal propulsion system performance, normal inlet-engine operations are required to be closer to compressor stall and inlet unstart. To study these phenomena an integrated model is needed that includes both airframe structural dynamics as well as the propulsion system dynamics. This paper covers the propulsion system component volume dynamics modeling of a turbojet engine that will be used for an integrated vehicle Aero- Propulso-Servo-Elastic model and for propulsion efficiency studies.

  11. Volume Dynamics Propulsion System Modeling for Supersonics Vehicle Research

    NASA Technical Reports Server (NTRS)

    Kopasakis, George; Connolly, Joseph W.; Paxson, Daniel E.; Ma, Peter

    2010-01-01

    Under the NASA Fundamental Aeronautics Program the Supersonics Project is working to overcome the obstacles to supersonic commercial flight. The proposed vehicles are long slim body aircraft with pronounced aero-servo-elastic modes. These modes can potentially couple with propulsion system dynamics; leading to performance challenges such as aircraft ride quality and stability. Other disturbances upstream of the engine generated from atmospheric wind gusts, angle of attack, and yaw can have similar effects. In addition, for optimal propulsion system performance, normal inlet-engine operations are required to be closer to compressor stall and inlet unstart. To study these phenomena an integrated model is needed that includes both airframe structural dynamics as well as the propulsion system dynamics. This paper covers the propulsion system component volume dynamics modeling of a turbojet engine that will be used for an integrated vehicle Aero-Propulso-Servo-Elastic model and for propulsion efficiency studies.

  12. Effect of Microjet Injection on Supersonic Jet Noise

    NASA Technical Reports Server (NTRS)

    Zaman, K. B. M. Q.; Podboy, G. G.

    2010-01-01

    The effect of microjet (jet) injection on the noise from supersonic jets is investigated. Three convergent-divergent (C-D) nozzles and one convergent nozzle, all having the same exit diameters, are used in the study. The jets are injected perpendicular to the primary jet close to the nozzle lip from six equally-spaced ports having a jet-to-primary-jet diameter ratio of 0.0054. Effects in the over-expanded, fully expanded as well as underexpanded flow regimes are explored. Relative to the effect on subsonic jets, larger reductions in the overall sound pressure level (OASPL) are achieved in most supersonic conditions. The largest reductions are typically associated with suppression of screech and transonic tones. For a shock-free, fully expanded case, the OASPL reductions achieved are comparable to that in the subsonic case; the same correlation, found for subsonic jet noise reduction at shallow observation angle, applies.

  13. Modeling of supersonic combustor flows using parallel computing

    NASA Technical Reports Server (NTRS)

    Riggins, D.; Underwood, M.; Mcmillin, B.; Reeves, L.; Lu, E. J.-L.

    1992-01-01

    While current 3D CFD codes and modeling techniques have been shown capable of furnishing engineering data for complex scramjet flowfields, the usefulness of such efforts is primarily limited by solutions' CPU time requirements, and secondarily by memory requirements. Attention is presently given to the use of parallel computing capabilities for engineering CFD tools for the analysis of supersonic reacting flows, and to an illustrative incompressible CFD problem using up to 16 iPSC/2 processors with single-domain decomposition.

  14. Ion dynamics in a supersonic jet: Experiments and simulations

    NASA Astrophysics Data System (ADS)

    Caldirola, S.; Roman, H. E.; Riccardi, C.

    2016-03-01

    The authors suggest a model to simulate the dynamics of ions in a supersonic plasma jet. The model relies on experimental argon ion, Ar+, energy distribution functions measured by a quadrupole mass spectrometer at different positions along the central axis of a supersonic argon plasma jet. The latter is generated by the pressure difference between two vacuum chambers connected through a converging nozzle: a high-pressure chamber (P ≃3.20 Pa), where an inductively coupled argon plasma discharge is maintained, and a lower-pressure one (P ≃0.11 Pa), where the plasma jet expands. The model is based on the integration of the equations of motion of a single Ar+, moving along the supersonic jet in a reference system in which neutral species are at rest. Ar+-Ar induced dipole interactions are treated using a 12-4 Lennard-Jones potential. The resulting collisions are considered to be purely elastic, and in addition to them, we allow for charge transfer processes. The energy and position of 1000 Ar+ were calculated, using an integration time step of 10 ps for ion trajectories ranging from 5 mm to 20 mm from the nozzle, well inside the spatial extension of the supersonic jet. The numerically obtained ion energy distribution functions agree remarkably well with the experimental measurements. From our calculations we can draw conclusions about the energy loss and the mean free paths along the jet. In particular, we can distinguish between processes with and without charge transfer, allowing us to determine the effect of charge exchange phenomena in which the ion changes its nature. The calculated mean free paths were used to evaluate the effective cross sections for momentum transfer and charge transfer collisions, valid for ion energies in the range (0.5-10) eV, in very good agreement with those reported in the literature.

  15. Aerodynamic characteristics of reentry vehicles at supersonic velocities

    NASA Astrophysics Data System (ADS)

    Adamov, N. P.; Kharitonov, A. M.; Chasovnikov, E. A.; Dyad'kin, A. A.; Kazakov, M. I.; Krylov, A. N.; Skorovarov, A. Yu.

    2015-09-01

    Models of promising reentry vehicles, experimental equipment, and test program are described. The method used to determine the total aerodynamic characteristics of these models on the AB-313 mechanical balance in the T-313 supersonic wind tunnel and the method used for simulations are presented. The aerodynamic coefficients of the examined objects in wide ranges of Mach numbers and angles of attack are obtained. The experimental data are compared with the results of simulations.

  16. Numerical study of supersonic turbulent flow over small protuberances

    NASA Technical Reports Server (NTRS)

    Polak, A.; Werle, M. J.

    1975-01-01

    Supersonic turbulent boundary layers over two-dimensional protuberances are investigated, using the numerical finite difference alternating direction implicit (ADI) method. The turbulence is modeled mathematically. The turbulence is represented here by the eddy viscosity approach. The turbulent boundary layer structure as well as an interest in thick boundary layers and much larger protuberance heights than in the laminar case lead to new difficulties. The problems encountered and the means to remove them are discussed.

  17. Ion dynamics in a supersonic jet: Experiments and simulations.

    PubMed

    Caldirola, S; Roman, H E; Riccardi, C

    2016-03-01

    The authors suggest a model to simulate the dynamics of ions in a supersonic plasma jet. The model relies on experimental argon ion, Ar(+), energy distribution functions measured by a quadrupole mass spectrometer at different positions along the central axis of a supersonic argon plasma jet. The latter is generated by the pressure difference between two vacuum chambers connected through a converging nozzle: a high-pressure chamber (P ≃ 3.20 Pa), where an inductively coupled argon plasma discharge is maintained, and a lower-pressure one (P ≃ 0.11 Pa), where the plasma jet expands. The model is based on the integration of the equations of motion of a single Ar(+), moving along the supersonic jet in a reference system in which neutral species are at rest. Ar(+)-Ar induced dipole interactions are treated using a 12-4 Lennard-Jones potential. The resulting collisions are considered to be purely elastic, and in addition to them, we allow for charge transfer processes. The energy and position of 1000 Ar(+) were calculated, using an integration time step of 10 ps for ion trajectories ranging from 5 mm to 20 mm from the nozzle, well inside the spatial extension of the supersonic jet. The numerically obtained ion energy distribution functions agree remarkably well with the experimental measurements. From our calculations we can draw conclusions about the energy loss and the mean free paths along the jet. In particular, we can distinguish between processes with and without charge transfer, allowing us to determine the effect of charge exchange phenomena in which the ion changes its nature. The calculated mean free paths were used to evaluate the effective cross sections for momentum transfer and charge transfer collisions, valid for ion energies in the range (0.5-10) eV, in very good agreement with those reported in the literature.

  18. Cooperative airframe/propulsion control for supersonic cruise aircraft

    NASA Technical Reports Server (NTRS)

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

    1974-01-01

    Interactions between propulsion systems and flight controls have emerged as a major control problem on supersonic cruise aircraft. This paper describes the nature and causes of these interactions and the approaches to predicting and solving the problem. Integration of propulsion and flight control systems appears to be the most promising solution if the interaction effects can be adequately predicted early in the vehicle design. Significant performance, stability, and control improvements may be realized from a cooperative control system.

  19. Flightworthy active control landing gear for a supersonic aircraft

    NASA Technical Reports Server (NTRS)

    Ross, I.

    1980-01-01

    A flightworthy active control landing gear system for a supersonic aircraft was designed to minimize aircraft loads during takeoff, impact, rollout, and taxi. The design consists of hydromechanical modifications to the existing gear and the development of a fail-safe electronic controller. analytical RESULTS INDICATE that for an aircraft sink rate of 0.914 m/sec (3 ft/sec) the system achieves a peak load reduction of 36% during landing impact.

  20. Reaction of the French population to the supersonic bang

    NASA Technical Reports Server (NTRS)

    Bremond, J.

    1980-01-01

    A discussion of a survey dealing with the supersonic bang is presented. Topics include the position the bang has in today's pollution, annoyance caused by the bang and its dependence on sociological and psychological variables, and whether or not the perception of the ban is objective. Other questions raised are whether the frequency of exposure to the bang has an influence on attitudes and does the sensitivity to or annoyance from the bang have a linear increase with the frequency.

  1. Calculation of three-dimensional, inviscid supersonic, steady flows

    NASA Technical Reports Server (NTRS)

    Moretti, G.

    1982-01-01

    A numerical technique is described for the calculation of three dimensional, inviscid, supersonic, steady flows over wing-body configurations. A high degree of accuracy without increasing the number of computational nodes is obtained by means of a powerful conformal mapping technique. Results are presented for some simple body configurations and for a more complex arrow wing airframe. The numerical results show good agreement with experimental measurements.

  2. Advanced surface paneling method for subsonic and supersonic flow

    NASA Technical Reports Server (NTRS)

    Erickson, L. L.; Johnson, F. T.; Ehlers, F. E.

    1976-01-01

    Numerical results illustrating the capabilities of an advanced aerodynamic surface paneling method are presented. The method is applicable to both subsonic and supersonic flow, as represented by linearized potential flow theory. The method is based on linearly varying sources and quadratically varying doublets which are distributed over flat or curved panels. These panels are applied to the true surface geometry of arbitrarily shaped three dimensional aerodynamic configurations.

  3. Handbook of Supersonic Aerodynamics. Section 8. Bodies of Revolution

    DTIC Science & Technology

    1961-10-01

    degree of accuracy the characteristics method requires much more time. Heaviside Operational Method (Ref. 78).--The Heaviside opera- tor is applied to...Supersonic Blunt-Body Flows. JPL/CIT Progress Report No. 20-372, February 1959. 88. Oliver , R. E. An Experimental Investigation of Flow over Simple Blunt...272 Newtonian-Prandtl-Meyer theory, 160 skin fric t ion, effect on, 233 normal force coefficient, 163 Heaviside operator, 62 angle-of-attack effect on

  4. Studies on nonequilibrium phenomena in supersonic chemically reacting flows

    NASA Technical Reports Server (NTRS)

    Tiwari, S. N.; Chandrasekhar, Rajnish

    1993-01-01

    This study deals with a systematic investigation of nonequilibrium processes in supersonic combustion. The two-dimensional, elliptic Navier-Stokes equations are used to investigate supersonic flows with nonequilibrium chemistry and thermodynamics, coupled with radiation, for hydrogen-air systems. The explicit, unsplit MacCormack finite-difference scheme is used to advance the governing equations in time, until convergence is achieved. For a basic understanding of the flow physics, premixed flows undergoing finite rate chemical reactions are investigated. Results obtained for specific conditions indicate that the radiative interactions vary substantially, depending on reactions involving HO2 and NO species, and that this can have a noticeable influence on the flowfield. The second part of this study deals with premixed reacting flows under thermal nonequilibrium conditions. Here, the critical problem is coupling of the vibrational relaxation process with the radiative heat transfer. The specific problem considered is a premixed expanding flow in a supersonic nozzle. Results indicate the presence of nonequilibrium conditions in the expansion region of the nozzle. This results in reduction of the radiative interactions in the flowfield. Next, the present study focuses on investigation of non-premixed flows under chemical nonequilibrium conditions. In this case, the main problem is the coupled turbulence-chemistry interaction. The resulting formulation is validated by comparison with experimental data on reacting supersonic coflowing jets. Results indicate that the effect of heat release is to lower the turbulent shear stress and the mean density. The last part of this study proposes a new theoretical formulation for the coupled turbulence-radiation interactions. Results obtained for the coflowing jets experiment indicate that the effect of turbulence is to enhance the radiative interactions.

  5. Wave-driven Rotation in Supersonically Rotating Mirrors

    SciTech Connect

    A. Fetterman and N.J. Fisch

    2010-02-15

    Supersonic rotation in mirrors may be produced by radio frequency waves. The waves produce coupled diffusion in ion kinetic and potential energy. A population inversion along the diffusion path then produces rotation. Waves may be designed to exploit a natural kinetic energy source or may provide the rotation energy on their own. Centrifugal traps for fusion and isotope separation may benefit from this wave-driven rotation.

  6. Quasi 1D Modeling of Mixed Compression Supersonic Inlets

    NASA Technical Reports Server (NTRS)

    Kopasakis, George; Connolly, Joseph W.; Paxson, Daniel E.; Woolwine, Kyle J.

    2012-01-01

    The AeroServoElasticity task under the NASA Supersonics Project is developing dynamic models of the propulsion system and the vehicle in order to conduct research for integrated vehicle dynamic performance. As part of this effort, a nonlinear quasi 1-dimensional model of the 2-dimensional bifurcated mixed compression supersonic inlet is being developed. The model utilizes computational fluid dynamics for both the supersonic and subsonic diffusers. The oblique shocks are modeled utilizing compressible flow equations. This model also implements variable geometry required to control the normal shock position. The model is flexible and can also be utilized to simulate other mixed compression supersonic inlet designs. The model was validated both in time and in the frequency domain against the legacy LArge Perturbation INlet code, which has been previously verified using test data. This legacy code written in FORTRAN is quite extensive and complex in terms of the amount of software and number of subroutines. Further, the legacy code is not suitable for closed loop feedback controls design, and the simulation environment is not amenable to systems integration. Therefore, a solution is to develop an innovative, more simplified, mixed compression inlet model with the same steady state and dynamic performance as the legacy code that also can be used for controls design. The new nonlinear dynamic model is implemented in MATLAB Simulink. This environment allows easier development of linear models for controls design for shock positioning. The new model is also well suited for integration with a propulsion system model to study inlet/propulsion system performance, and integration with an aero-servo-elastic system model to study integrated vehicle ride quality, vehicle stability, and efficiency.

  7. Some unique characteristics of supersonic cruise vehicles and their effect on airport community noise

    NASA Technical Reports Server (NTRS)

    Driver, C.; Maglieri, D. J.

    1980-01-01

    The paper examines the differences between the supersonic and subsonic commercial aircraft in terms of their configuration, aerodynamic characteristics, propulsion systems, and the manner of operation. The unique characteristics of supersonic cruise vehicles should provide improved airport-community noise exposures if the vehicle is permitted to operate at its most efficient and effective flight modes. It is concluded that noise exposure levels for supersonic cruise vehicles can be comparable to those of its equivalent subsonic counterpart of that time period.

  8. Effects of the Air Traffic Control System on the Supersonic Transport

    NASA Technical Reports Server (NTRS)

    Silsby, Norman S.; McLaughlin, Milton D.; Fischer, Michael C.

    1965-01-01

    A study of the problems anticipated with the introduction of the supersonic transport into the air traffic control system indicated that supersonic transport design allowances for time and fuel for maneuvering during climbouts may not be sufficient, that there is a greater communications-navigation work- load for the supersonic transport than for the subsonic jet transport during descent, and that use of a flight director to command pitch control guidance for the pilot would be helpful.

  9. The calculation of downwash behind supersonic wings with an application to triangular plan forms

    NASA Technical Reports Server (NTRS)

    Lomax, Harvard; Sluder, Loma; Heaslet, Max A

    1950-01-01

    A method is developed consistent with the assumptions of small perturbation theory which provides a means of determining the downwash behind a wing in supersonic flow for a known load distribution. The analysis is based upon the use of supersonic doublets which are distributed over the plan form and wake of the wing in a manner determined from the wing loading. The equivalence in subsonic and supersonic flow of the downwash at infinity corresponding to a given load distribution is proved.

  10. Development of the triplet singularity for the analysis of wings and bodies in supersonic flow

    NASA Technical Reports Server (NTRS)

    Woodward, F. A.

    1981-01-01

    A supersonic triplet singularity was developed which eliminates internal waves generated by panels having supersonic edges. The triplet is a linear combination of source and vortex distributions which gives directional properties to the perturbation flow field surrounding the panel. The theoretical development of the triplet singularity is described together with its application to the calculation of surface pressures on wings and bodies. Examples are presented comparing the results of the new method with other supersonic methods and with experimental data.

  11. Noise from a Supersonic Round Jet Discharging Into a Duct

    NASA Technical Reports Server (NTRS)

    Zaman, K. B. M. Q.; Fagan, A. F.

    2014-01-01

    In an effort to understand an 'unwanted noise' problem occasionally encountered in ground test facilities, the interaction of a jet flow with a duct is studied in a model scale experiment. While the interaction of subsonic jets was studied earlier, that of supersonic jets is considered in this paper. The effect of the presence of a cylindrical duct in the path of the jet is studied through sound pressure level spectral measurements as well as schlieren flow visualization. When the jet involves screech tones the placement of the duct is found to make only minor effects on the tones themselves as well as on the high frequency noise. However, there is increased energy at low frequencies. The increase in low frequency noise becomes clearer when screech is eliminated from the jet by two small tabs placed at the nozzle exit. It is shown that spectral peaks and increased sound pressure levels occur at frequencies corresponding to the axial acoustic resonance modes of the duct. These peaks persist into the supersonic regime, however, their amplitudes diminish relative to increasing spectral amplitudes at other frequencies with increasing jet Mach number. A wire-mesh screen attached to the end of the duct effectively suppresses such unwanted noise at subsonic as well as supersonic conditions.

  12. Free-field propagation of high intensity noise. [supersonic jets

    NASA Technical Reports Server (NTRS)

    Mcdaniel, O. H.; Roth, S. D.; Welz, J. P.

    1981-01-01

    Research on high intensity (finite amplitude) acoustic waves shows that nonlinear distortion effects generally result in a shift of energy to higher frequencies. The higher intensities associated with supersonic jets would therefore indicate that high frequency enhancement of the spectrum should occur, resulting in the differences observed between subsonic and supersonic jets. A 10,000 acoustic watt source installed in an anechoic chamber generates sound levels such that acoustic shocks are readily observable. Dual frequency excitation of the source produces a strong parametric effect with a difference frequency comparable in level to the primary frequency. The test set up and recording equipment being used to determine the finite amplitude noise representative of an actual supersonic jet are described as well as the development of a computer program based on Burger's equation. The spectra of 1/2 octave band, 1 kHz sine wave, and dual frequency input and output are presented in graphs along with waveforms at Z = .025, 0.1, and 1.0.

  13. Effect of cavity fore wall modifications in supersonic flow

    NASA Astrophysics Data System (ADS)

    Jeyakumar, S.; Yuvaraj, K. Arun; Jayaraman, K.; Cardona, F.; Sultan, M. T. H.

    2016-10-01

    Understanding the complex behaviour of the cavity flow is essential for the design of supersonic combustor. The characteristics of the axisymmetric aft ramp cavity with fore wall modification have been experimentally studied in a blow-down type supersonic flow facility. The facility consists of a conventional CD nozzle that issues a flow Mach number of 1.88 to a supersonic combustor of circular cross section, which is placed immediately downstream of the nozzle. The axisymmetric cavities are incorporated within the combustor. The cavities are of open type and their length is kept constant while their depth varies. The aft wall of the cavities is inclined with three ramp angles and the fore wall is provided with a constant fillet radius of 3 mm. The performance of the cavity is analysed based on wall static pressures, momentum flux distribution at the exit of the combustor and stagnation pressure loss across the combustor. The study reveals that fore wall fillet cavities provide less cavity drag and stagnation pressure loss for various aft ramp angles under identical operating conditions.

  14. Supersonic jet noise reduction by coaxial rectangular nozzles

    NASA Technical Reports Server (NTRS)

    Ahuja, K. K.; Manes, J. P.; Massey, K. C.

    1992-01-01

    A physical understanding of noise reduction mechanisms in supersonic, single, and coaxial rectangular jets is quantified and obtained, with emphasis on shock noise reduction. For all conditions, corresponding acoustic measurements for an equivalent round jet are also obtained so that the noise characteristics of the two types of jets can be compared directly to quantify the noise reductions. Comparisons are thus provided for a single rectangular nozzle vs a single equivalent round nozzle, and a coaxial rectangular nozzle vs an equivalent round nozzle. It is shown that different operating conditions and nozzle arrangements for the same thrust, total exit area, and mass flow rate can produce different noise levels. With at least one stream operated supersonically, the coaxial rectangular nozzle operated in the inverted-velocity profile is always quieter than in the normal velocity profile mode for the same thrust, exit area, and mass flow rate. In general, the coaxial rectangular nozzle is shown to be quieter than an equivalent circular nozzle only for those conditions for which both nozzles are operated supersonically.

  15. Stability derivatives of triangular wings at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Ribner, Herbert S; Malvestuto, Frank S , Jr

    1948-01-01

    The analysis of the stability derivatives of low-aspect-ratio triangular wings at subsonic and supersonic speeds, given in NACA TN no. 1423, is extended to apply to triangular wings having large vertex angles and traveling at supersonic speeds. The lift, rolling moment due to sideslip, and damping in roll and pitch for this more general case have been treated elsewhere on the basis of the theory of small disturbances. The surface potentials for angle of attack and rolling taken therefrom are used to obtain the several side-force and yawing-moment derivatives that depend on leading-edge suction, and a tentative value for the rolling moment due to yawing. The lift and moment due to downward acceleration are obtained on the basis of an unpublished unsteady-flow solution. All the known stability derivatives of the triangular wing at supersonic speeds, regardless of source, are summarized for convenience and presented with respect to both body axes and stability axes. The results are limited to Mach numbers for which the triangular wing is contained within the Mach cone for its vertex. The spanwise variation of Mach number in the case of yawing is neglected, although the effect must be of importance.

  16. Interaction of a swept shock wave and a supersonic wake

    NASA Astrophysics Data System (ADS)

    He, G.; Zhao, Y. X.; Zhou, J.

    2017-03-01

    The interaction of a swept shock wave and a supersonic wake has been studied. The swept shock wave is generated by a swept compression sidewall, and the supersonic wake is generated by a wake generator. The flow field is visualized with the nanoparticle-based planar laser scattering method, and a supplementary numerical simulation is conducted by solving the Reynolds-averaged Navier-Stokes equations. The results show that the pressure rise induced by the swept shock wave can propagate upstream in the wake, which makes the location where vortices are generated move upstream, thickens the laminar section of the wake, and enlarges the generated vortices. The wake is swept away from the swept compression sidewall by the pressure gradient of the swept shock wave. This pressure gradient is not aligned with the density gradient of the supersonic wake, so the baroclinic torque generates streamwise vorticity and changes the distribution of the spanwise vorticity. The wake shock is curved, so the flow downstream of it is non-uniform, leaving the swept shock wave being distorted. A three-dimensional Mach disk structure is generated when the wake shock interacts with the swept shock wave.

  17. Development and analysis of a STOL supersonic cruise fighter concept

    NASA Technical Reports Server (NTRS)

    Dollyhigh, S. M.; Foss, W. E., Jr.; Morris, S. J., Jr.; Walkley, K. B.; Swanson, E. E.; Robins, A. W.

    1984-01-01

    The application of advanced and emerging technologies to a fighter aircraft concept is described. The twin-boom fighter (TBF-1) relies on a two dimensional vectoring/reversing nozzle to provide STOL performance while also achieving efficient long range supersonic cruise. A key feature is that the propulsion package is placed so that the nozzle hinge line is near the aircraft center-of-gravity to allow large vector angles and, thus, provide large values of direct lift while minimizing the moments to be trimmed. The configurations name is derived from the long twin booms extending aft of the engine to the twin vertical tails which have a single horizontal tail mounted atop and between them. Technologies utilized were an advanced engine (1985 state-of-the-art), superplastic formed/diffusion bonded titanium structure, advanced controls/avionics/displays, supersonic wing design, and conformal weapons carriage. The integration of advanced technologies into this concept indicate that large gains in takeoff and landing performance, maneuver, acceleration, supersonic cruise speed, and range can be acieved relative to current fighter concepts.

  18. Study on the characteristics of supersonic Coanda jet

    NASA Astrophysics Data System (ADS)

    Matsuo, Shigeru; Setoguchi, Toshiaki; Kudo, Takemasa; Yu, Shen

    1998-09-01

    Techniques using Coanda effect have been applied to the fluid control devices. In this field, experimental studies were so far performed for the spiral jet obtained by the Coanda jet issuing from a conical cylinder with an annular slit, thrust vectoring of supersonic Coanda jets and so on. It is important from the viewpoints of effective applications to investigate the characteristics of the supersonic Coanda jet in detail. In the present study, the effects of pressure ratios and nozzle configurations on the characteristics of the supersonic Coanda jet have been investigated experimentally by a schlieren optical method and pressure measurements. Furthermore, Navier-Stokes equations were solved numerically using a 2nd-order TVD finite-volume scheme with a 3rd-order three stage Runge-Kutta method for time integration. k - ɛ model was used in the computations. The effects of initial conditions on Coanda flow were investigated numerically. As a result, the simulated flow fields were compared with experimental data in good agreement qualitatively.

  19. Supersonic solutions of the solar-wind equations

    NASA Technical Reports Server (NTRS)

    Chamberlain, Joseph W.

    1995-01-01

    We re-examine the inviscid solar-wind equations with heat conduction from below, and establish a fundamentally new approach for finding stellar solutions. Although the problem is fourth order, only two independent integration constants can be assigned, since two boundary conditions that are required to specify well-behaved supersonic solutions determine the values of the other two constants. The solar-wind-model of Noble and Scarf are essentially accurate for practical purposes, but in a fundamental sense they are not self-consistent. At the supersonic level, the ratio of thermal energy to gravitational potential, (kTr/GMm)(sub s), must lie in a narrow range, between 0.4375 and about 0.40, to permit well-behaved solutions for ionized hydrogen. For a supersonic solution this ratio uniquely determines the constants A and epsilon, which in turn govern not only the escape flux and the energy flux per particle but also T'(lambda approaches 0), the temperature gradient at infinity. It is very difficult to find solutions by trial and error, without employing T'(0) as a guide to the iteration of A and epsilon, because not only must (kTr/GMm)(sub s) be chosen within the acceptable range but then either A or epsilon must be guessed with phenomenal accuracy. Otherwise the numerical integration of temperature and expansion velocity do not yield acceptable values at infinity.

  20. Supersonic HLFC: Potential benefits and technology development requirements

    NASA Technical Reports Server (NTRS)

    Neumann, Frank

    1992-01-01

    The viewgraphs and discussion of the potential benefits and technology development requirements of supersonic Hybrid Laminar Flow Control (HLFC) are reviewed. For the last three years, Boeing has performed studies on the application of laminar flow control to High Speed Civil Transport (HSCT) configurations. Large potential net benefits were identified for laminar flow control, even after accounting for the significant implementation penalties. However, the technical risks are high at this time, and an early, aggressive technology development program is required if laminar flow control (LFC) is to be incorporated in a year 2005 HSCT program. The benefits and required development effort are addressed. Of all the aerodynamic advances that are being considered for the HSCT, LFC has the largest potential for improving the supersonic lift drag ratio of a given configuration. The work accomplished to date, sponsored by NASA Langley, is summarized. This work includes studies on HLFC application to HSCT, cruise HLFC/low speed boundary layer control (BLC) compatibility, and impact of M 0.9 HLFC and high lift BLC. Requirements for production committments are listed. The need for a HSCT-HLFC Risk Reduction Program is addressed. Supersonic HLFC development planning and future applications are addressed. Laminar flow/high-lift integration is addressed.

  1. An experimental study of planar heterogeneous supersonic confined jets

    NASA Astrophysics Data System (ADS)

    Tanis, Frederick J., Jr.

    1994-12-01

    The effects of varying the exit pressure of a supersonic helium jet exhausting coaxially with two parallel supersonic air streams into a constant area duct were investigated. The method used to evaluate the mass entrainment rate was to measure helium molar concentration profiles and mass flux across the duct using a binary gas probe then calculate the mass entrainment into the helium jet. In order to conduct this study a novel binary gas probe was developed which allowed helium concentration and mass flux data to be obtained during continuous traverses across the supersonic flowfield. High exit pressure ratio (EPR) led to improved overall mixing compared to the baseline case with an EPR near unity. The high EPR caused low mass entrainment along the jet shear layers due to high convective Mach numbers and velocity ratios, but the high EPR caused oblique shocks to form which reflected off the duct walls and intersected with the helium jet several times causing significant mass entrainment due to numerous shock-shear layer interactions (SSLI's). A correlation between the vorticity generated during a SSLI and the mass entrainment into the jet was developed.

  2. Overview of Boeing supersonic transport efforts, 1971 - 1979

    NASA Technical Reports Server (NTRS)

    Sigalla, A.

    1980-01-01

    Following cancellation of the United States Supersonic Transport program, the status of the technology was assessed carefully and emphasis was put on finding solutions for what were considered the major technical difficulties. In particular, work on the breakthroughs needed to advance the technology was emphasized. Currently, solutions to all major technical problems are identified. Depending on the subject, either the problem is no longer a concern or the steps needed to bring about a solution are mapped out clearly. Throughout the NASA SCR program, important strides were made in the identification of design advances which would greatly improve supersonic airplane fuel efficiency, noise, and other performance and cost affecting parameters. Furthermore, these efforts created an atmosphere in which it was possible for new ideas to flourish and positive inventions to take place such as the variable cycle engine and the blended fuselage. These technical gains show that, given availability of such technology, advanced supersonic transports could be developed that would be economically successful and environmentally acceptable.

  3. Calculations of Supersonic and Hypersonic Flows using Compressible Wall Functions

    NASA Technical Reports Server (NTRS)

    Huang, P. G.; Coakley, T. J.

    1993-01-01

    The present paper presents a numerical procedure to calculate supersonic and hypersonic flows using the compressible law of the wall. The turbulence models under consideration include the Launder-Reece-Rodi-Gibson Reynolds-stress model and the k-epsilon model. The models coupled with the proposed wall function technique have been tested in both separated and unseparated flows. The flows include (1) an insulated flat plate flow over a range of Mach numbers, (2) a Mach 5 flat plate flow with cold wall conditions, (3) a two dimensional supersonic compression corner flow, (4) a hypersonic flow over an axisymmetric flare, and (5) a hypersonic flow over a 2-D compression corner. Results indicate that the wall function technique gives improved predictions of skin friction and heat transfer in separated flows compared with models using wall dampers. Predictions of the extent of separation are not improved over the wall damper models except with the Reynolds-stress model for the supersonic compression corner flow case.

  4. Simulated high speed flight effects on supersonic jet noise

    NASA Technical Reports Server (NTRS)

    Norum, Thomas D.; Brown, Martha C.

    1993-01-01

    A free jet is utilized to investigate the changes in the noise received from supersonic jets in high speed subsonic flight. Flight Mach numbers to 0.9 are simulated for supersonic jets with fully expanded Mach numbers between 1 and 2. Plume pressure measurements show only minor changes in the shock structure of off-design jets up to a Mach number of 0.6. Correspondingly, far-field noise measurements indicate little change to the broadband shock noise emitted at right angles to the jet. However, measurements within the free jet show that convection effects on the noise are substantial, and that the point source convective amplification that is proportional to the fourth power of the Doppler factor may apply for broadband shock noise in flight. Measurements of jet mixing noise for an on-design supersonic jet show that the current predictions of mixing noise in flight can be extended to flight Mach numbers of at least 0.5.

  5. Experimental and analytical investigations of wave enhanced supersonic combustors

    NASA Technical Reports Server (NTRS)

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

    1989-01-01

    Supersonic combustion ramjet (scramjet) engines rely on rapid mixing and combustion of fuel for good performance. However, both of these processes are relatively slow compared to the residence times in supersonic combusters. Methods of improving the mixing process include subjecting the fuel streams to shock waves and the generation of vortices by spiriling, bifurcating struts. The combustion process can also be enhanced by shock or detonation waves. An oblique shock wave can act as a flameholder by increasing the pressure and temperature of the air-fuel mixture and thereby decreasing the ignition delay. If the oblique shock is sufficiently strong, then the combustion front and the shock wave can couple into a detonation wave. In this case, combustion occurs almost instantaneously in a thin zone behind the wave front. While the existence of standing oblique detonation waves has been proven in computer simulations, the experimental validation is still in progress. Currently, there is an experimental and analytical program at NASA-Ames Research Center to study detonations and other means of wave enhanced supersonic mixing and combustion.

  6. Supersonic jet engine and method of operating the same

    SciTech Connect

    Klees, G.

    1981-09-08

    A supersonic jet engine is disclosed capable not only of developing sufficient power to accelerate up to supersonic cruise and maintain efficient operation at supersonic cruise, but also arrange to cruise at subsonic velocities with a relatively low specific fuel consumption. The engine is provided with a variable bypass passageway downstream of the compressor. Flow into the bypass passageway is controlled so that during low power setting the bypass passageway is closed so that all the gaseous flow is directed through the turbine. During higher power settings, the bypass passageway is opened to the extent that a selected portion of the gaseous flow is directed through the bypass passageway to bypass the turbine, so that the corrected flow to the turbine remains substantially constant for both high and low power settings of the engine. The turbine has first and second stages, with the first stage receiving all of the gaseous flow that is not directed through the bypass passageway. The gaseous flow from the turbine first stage is split into first and second portions, with the first portion being directed through a coannular inner passageway to drive the turbine second stage, and the second portion of the gaseous discharge being directed toward a coannular outer passageway to be discharged in a coannular pattern radially outwardly of the first portion of gaseous flow.

  7. Shock Positioning Controls Designs for a Supersonic Inlet

    NASA Technical Reports Server (NTRS)

    Kopasakis, George; Connolly, Joseph W.

    2010-01-01

    Under the NASA Fundamental Aeronautics Program, the Supersonics Project is working to overcome the obstacles to supersonic commercial flight. The supersonic inlet design that is utilized to efficiently compress the incoming air and deliver it to the engine has many design challenges. Among those challenges is the shock positioning of internal compression inlets, which requires active control in order to maintain performance and to prevent inlet unstarts due to upstream (freestream) and downstream (engine) disturbances. In this paper a novel feedback control technique is presented, which emphasizes disturbance attenuation among other control performance criteria, while it ties the speed of the actuation system(s) to the design of the controller. In this design, the desired performance specifications for the overall control system are used to design the closed loop gain of the feedback controller and then, knowing the transfer function of the plant, the controller is calculated to achieve this performance. The innovation is that this design procedure is methodical and allows maximization of the performance of the designed control system with respect to actuator rates, while the stability of the calculated controller is guaranteed.

  8. Mass flow and its pulsation measurements in supersonic wing wake

    NASA Astrophysics Data System (ADS)

    Shmakov, A. S.; Shevchenko, A. M.; Yatskikh, A. A.; Yermolaev, Yu. G.

    2016-10-01

    The results of experimental study of the flow in the wing wake are presented. Experiments were carried out in supersonic wind tunnel T-325 of ITAM SB RAS. Rectangle half-wing with sharp edges with a chord length of 30 mm and semispan of 95 mm was used to generate vortex wake. Experimental data were obtained in the cross section located 6 chord length downstream of the trailing edge at Mach numbers of 2.5 and 4 and at wing angles of attack of 4 and 10 degrees. Constant temperature hot-wire anemometer was used to measure disturbances in supersonic flow. Hot-wire was made of a tungsten wire with a diameter of 10 μm and length of 1.5 mm. Shlieren flow visualization were performed. As a result, the position and size of the vortex core in the wake of a rectangular wing were determined. For the first time experimental data on the mass flow distribution and its pulsations in the supersonic longitudinal vortex were obtained.

  9. Data Quality Assurance for Supersonic Jet Noise Measurements

    NASA Technical Reports Server (NTRS)

    Brown, Clifford A.; Henderson, Brenda S.; Bridges, James E.

    2010-01-01

    The noise created by a supersonic aircraft is a primary concern in the design of future high-speed planes. The jet noise reduction technologies required on these aircraft will be developed using scale-models mounted to experimental jet rigs designed to simulate the exhaust gases from a full-scale jet engine. The jet noise data collected in these experiments must accurately predict the noise levels produced by the full-scale hardware in order to be a useful development tool. A methodology has been adopted at the NASA Glenn Research Center s Aero-Acoustic Propulsion Laboratory to insure the quality of the supersonic jet noise data acquired from the facility s High Flow Jet Exit Rig so that it can be used to develop future nozzle technologies that reduce supersonic jet noise. The methodology relies on mitigating extraneous noise sources, examining the impact of measurement location on the acoustic results, and investigating the facility independence of the measurements. The methodology is documented here as a basis for validating future improvements and its limitations are noted so that they do not affect the data analysis. Maintaining a high quality jet noise laboratory is an ongoing process. By carefully examining the data produced and continually following this methodology, data quality can be maintained and improved over time.

  10. Seeding for laser velocimetry in confined supersonic flows with shocks

    NASA Technical Reports Server (NTRS)

    Lepicovsky, J.; Bruckner, R. J.

    1996-01-01

    There is a lack of firm conclusions or recommendations in the open literature to guide laser velocimeter (LV) users in minimizing the uncertainty of LV data acquired in confined supersonic flows with steep velocity gradients. This fact led the NASA Lewis Research Center (LeRC) in Cleveland (Ohio, USA), and the Institute of Propulsion Technology of DLR in Cologne (Germany) to a joint research effort to improve reliability of LV measurements in supersonic flows. Over the years, NASA and DLR have developed different expertise in laser velocimetry, using different LV systems: Doppler and two-spot (L2F). The goal of the joint program is to improve the reliability of LV measurements by comparing results from experiments in confined supersonic flows performed under identical test conditions but using two different LV systems and several seed particle generators. Initial experiments conducted at the NASA LERC are reported in this paper. The experiments were performed in a narrow channel with Mach number 2.5 flow containing an oblique shock wave generated by an immersed 25-dg wedge.

  11. Comparison Between Theory and Experiment for Wings at Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Vincenti, Walter G

    1951-01-01

    This paper presents a critical comparison made between experimental and theoretical results for the aerodynamic characteristics of wings at supersonic flight speeds. As a preliminary, a brief, nonmathematical review is given of the basic assumptions and general findings of supersonic wing theory in two and three dimensions. Published data from two-dimensional pressure-distribution tests are then used to illustrate the effects of fluid viscosity and to assess the accuracy of linear theory as compared with the more exact theories which are available in the two-dimensional case. Finally, an account is presented of an NACA study of the over-all force characteristics of three-dimensional wings at supersonic speed. In this study, the lift, pitching moment, and drag characteristics of several families of wings of varying plan form and section were measured in the wind tunnel and compared with values predicted by the three-dimensional linear theory. The regions of agreement and disagreement between experiment and theory are noted and discussed.

  12. The aerodynamic design of the oblique flying wing supersonic transport

    NASA Technical Reports Server (NTRS)

    Vandervelden, Alexander J. M.; Kroo, Ilan

    1990-01-01

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

  13. Mixed exhaust flow supersonic jet engine and method

    SciTech Connect

    Klees, G.W.

    1993-06-08

    A method of operating a supersonic jet engine installation is described comprising (a) providing an engine having a variable area air inlet means and an outlet to discharge engine exhaust; (b) providing a secondary air passageway means; (c) receiving ambient air in the air inlet means and providing the ambient air as primary air to the engine inlet and secondary air to the secondary air passageway means; (d) providing a mixing section having an inlet portion and an exit portion, utilizing the mixing section in directing the exhaust from the engine to primary convergent/divergent exit passageway segments, where the exhaust is discharged at supersonic velocity as primary flow components, and directing secondary air flow from the secondary air passageway means to secondary exit passageway segments which are interspersed with the primary segments and from which the secondary air is discharged at subsonic velocity as secondary flow components; and (e) providing an exhaust section to receive the primary and secondary flow components in a mixing region and causing the primary and secondary flow components to mix to create a supersonic mixed flow, the exhaust section having a variable area final nozzle through which the mixed flow is discharged.

  14. Numerical Simulation of Spray Atomization in Supersonic Flows

    NASA Astrophysics Data System (ADS)

    Wang, Jiangfeng; Liu, Chen; Wu, Yizhao

    With the rapid development of the air-breathing hypersonic vehicle design, an accurate description of the combustion properties becomes more and more important, where one of the key techniques is the procedure of the liquid fuel mixing, atomizing and burning coupled with the supersonic crossflow in the combustion chamber. The movement and distribution of the liquid fuel droplets in the combustion chamber will influence greatly the combustion properties, as well as the propulsion performance of the ramjet/scramjet engine. In this paper, numerical simulation methods on unstructured hybrid meshes were carried out for liquid spray atomization in supersonic crossflows. The Kelvin-Helmholtz/Rayleigh-Taylor hybrid model was used to simulate the breakup process of the liquid spray in a supersonic crossflow with Mach number 1.94. Various spray properties, including spray penetration height, droplet size distribution, were quantitatively compared with experimental results. In addition, numerical results of the complex shock wave structure induced by the presence of liquid spray were illustrated and discussed.

  15. Experimental Study on Restart Control of Supersonic Air Breathing Engine

    NASA Astrophysics Data System (ADS)

    Kojima, Takayuki; Sato, Tetsuya; Sawai, Shujiro; Tanatsugu, Nobuhiro

    In order to study dynamic response and establish control logic of supersonic air breathing engine, restart control tests of subscale engine model, that consists of axisymmetric intake and turbojet engine are done at ISAS supersonic wind tunnel (Mach 3). Assuming the condition that the combustion flame is blown out by the unstart, restart control sequences are set as follows. First, after a wind tunnel is started, the core engine is ignited. Second, the intake is restarted while the core engine is controlled. Third, the intake spike position and the terminal shock position are controlled and intake total pressure recovery becomes the designed value (60%). Tests are successful and the engine thrust is recovered for approximately 30-40 seconds after the intake unstart. Sudden increase of combustor flame temperature and rotational speed after the intake unstart is shown experimentally. This phenomenon is inevitable for supersonic engines that apply turbojet cycle as a core engine. To reduce sudden increase of the flame temperature, new sequence to close a fuel control valve after detection of the intake unstart is done and an increase of the flame temperature is reduced. Furthermore, necessity of avoidance of the intake buzz is shown experimentally. To avoid the intake buzz, buzz margin control by the bypass door is proposed and succeeded.

  16. 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.

  17. Multiscale turbulence effects in supersonic jets exhausting into still air

    NASA Technical Reports Server (NTRS)

    Abdol-Hamid, Khaled S.; Wilmoth, Richard G.

    1987-01-01

    A modified version of the multiscale turbulence model of Hanjalic has been applied to the problem of supersonic jets exhausting into still air. In particular, the problem of shock-cell decay through turbulent interaction with the mixing layer has been studied for both mildly interacting and strongly resonant jet conditions. The modified Hanjalic model takes into account the nonequilibrium energy transfer between two different turbulent spectral scales. The turbulence model was incorporated into an existing shock-capturing, parabolized Navier-Stokes computational model in order to perform numerical experiments. The results show that the two-scale turbulence model provides significant improvement over one-scale models in the prediction of plume shock structure for underexpanded supersonic (Mach 2) and sonic (Mach 1) jets. For the supersonic jet, excellent agreement with experiment was obtained for the centerline shock-cell pressure decay up to 40 jet radii. For the sonic jet, the agreement with experiment was not so good, but the two-scale model still showed significant improvement over the one-scale model. It is shown that by relating some of the coefficients in the turbulent-transport equations to the relative time scale for transfer of energy between scales the two-scale model can provide predictions that bound the measured shock-cell decay rate for the sonic jet.

  18. On the feedback mechanism in supersonic cavity flows

    NASA Astrophysics Data System (ADS)

    Li, Weipeng; Nonomura, Taku; Fujii, Kozo

    2013-05-01

    Self-sustained oscillations in supersonic cavity flows are investigated by implicit large-eddy simulations of a supersonic flow (M∞ = 2.0, ReD = 105) past a three-dimensional rectangular cavity with length-to-depth ratio of 2. Both turbulent and laminar inflows are considered, and a variation of boundary-layer thickness in the turbulent inflow case is conducted. An additional simulation of turbulent free shear layer is also performed to illustrate the relationship between shedding vortices and acoustic excitations. Feedback mechanism is identified as the dominant mechanism driving the self-sustained oscillations in supersonic open cavity flows, regardless of the upstream turbulent state and the boundary-layer thickness. The generation of discrete vortices in the cavity shear layer is shown to be highly associated with acoustic excitations rather than natural instabilities of the cavity shear layer. Simulation results support that the primary noise source arises from the successive passage of large-scale vortices over the cavity trailing edge. The effects of upstream boundary layer on the shear-layer characteristics and acoustic fields will also be discussed.

  19. Fundamental Aeronautics Program: Supersonics Project. Channeled Center-Body Inlet Experiment Overview

    NASA Technical Reports Server (NTRS)

    SaintJohn, Clint; Ratnayake, Nalin; Frederick, Mike

    2012-01-01

    The presentation describes supersonic flight testing accomplished on a novel mixed compression axisymmetric inlet utilizing channels for off design flow matching rather than a translating centerbody concept.

  20. A review of several propulsion integration features applicable to supersonic-cruise fighter aircraft

    NASA Technical Reports Server (NTRS)

    Berrier, B. L.

    1976-01-01

    A brief review has been made of the propulsion integration features which may impact the design of a supersonic cruise fighter type aircraft. The data used for this study were obtained from several investigations conducted in the Langley 16-foot transonic and 4 by 4 foot supersonic pressure wind tunnels. Results of this study show: (1) that for conventional nozzle installations, contradictory design guidelines exist between subsonic and supersonic flight condition, (2) that substantial drag penalties can be incurred by use of dry power nozzles during supersonic cruise; and (3) that a new and unique concept, the nonaxisymmetric nozzle, offers the potential for solving many of the current propulsion installation problems.

  1. Development of a quiet supersonic wind tunnel with a cryogenic adaptive nozzle

    NASA Technical Reports Server (NTRS)

    Wolf, Stephen D.

    1991-01-01

    The main objectives of this work is to demonstrate the potential of a cryogenic adaptive nozzle to generate quiet (low disturbance) supersonic flow. A drive system was researched for the Fluid Mechanics Laboratory (FML) Laminar Flow Supersonic Wind Tunnel (LFSWT) using a pilot tunnel. A supportive effort for ongoing Proof of Concept (PoC) research leading to the design of critical components of the LFSWT was maintained. The state-of-the-art in quiet supersonic wind tunnel design was investigated. A supersonic research capability was developed within the FML.

  2. Experimental investigation of the structure of supersonic two-dimensional air microjets

    NASA Astrophysics Data System (ADS)

    Timofeev, Ivan; Aniskin, Vladimir; Mironov, Sergey

    2016-10-01

    We have experimentally studied the structure of supersonic underexpanded room-temperature air jets escaping from micronozzles with characteristic heights from 47 to 175 µm and widths within 2410-3900 µm in a range of Reynolds numbers of 1280-9460. The dimensions of the first shock cell are established. The supersonic core length of supersonic underexpanded air jets has been determined for the first time. A flow regime with a large supersonic core length has observed for air jets escaping from a 47µm high nozzle.

  3. An experimental study of the structure of supersonic flat underexpanded microjets

    NASA Astrophysics Data System (ADS)

    Aniskin, V. M.; Maslov, A. A.; Mironov, S. G.; Tsyryulnikov, I. S.; Timofeev, I. V.

    2015-05-01

    We have experimentally studied the structure of supersonic flat underexpanded room-temperature air jets escaping from micro nozzles with characteristic heights from 47 to 175 μm and widths within 2410-3900 μm in a range of Reynolds numbers of 1280-9460. The dimensions of the first shock cell are established. The supersonic core length of supersonic flat underexpanded air jets has been determined for the first time. A flow regime with a large supersonic core length has been observed for air jets escaping from a 47-μm-high nozzle.

  4. Pressure distribution and aerodynamic coefficients associated with heat addition to supersonic air stream adjacent to two-dimensional supersonic wing

    NASA Technical Reports Server (NTRS)

    Pinkel, I Irving; Serafini, John S; Gregg, John L

    1952-01-01

    The modifications in the pressure distributions and the aerodynamic coefficients associated with additions of heat to the two-dimensional supersonic in viscid flow field adjacetnt to the lower surface of of a 5-percent-thickness symmetrical circular-arc wing are presented in this report. The pressure distributions are obtained by the use of graphical method which gives the two-dimensional supersonic inviscid flow field obtained with moderate heat addition. The variation is given of the lift-drag ratio and of the aerodynamic coefficients of lift, drag, and moment with free stream Mach number, angle of attack, and parameters defining extent and amount of heat addition. The six graphical solutions used in this study included Mach numbers of 3.0 and 5.0 and angles of attack of 0 degrees and 2 degrees.

  5. Gain and temperature in a slit nozzle supersonic chemical oxygen-iodine laser with transonic and supersonic injection of iodine

    NASA Astrophysics Data System (ADS)

    Rosenwaks, Salman; Barmashenko, Boris D.; Bruins, Esther; Furman, Dov; Rybalkin, Victor; Katz, Arje

    2002-05-01

    Spatial distributions of the gain and temperament across the flow were studied for transonic and supersonic schemes of the iodine injection in a slit nozzle supersonic chemical oxygen-iodine laser as a function of the iodine and secondary nitrogen flow rate, jet penetration parameter and gas pumping rate. The mixing efficiency for supersonic injection of iodine is found to be much larger than for transonic injection, the maximum values of the gain being approximately 0.65 percent/cm for both injection schemes. Measurements of the gain distribution as a function of the iodine molar flow rate nI2 were carried out. For transonic injection the optimal value of nI2 at the flow centerline is smaller than that at the off axis location. The temperature is distributed homogeneously across the flow, increasing only in the narrow boundary layers near the walls. Opening a leak downstream of the cavity in order to decease the Mach number results in a decrease of the gain and increase of the temperature. The mixing efficiency in this case is much larger than for closed leak.

  6. On the Comparison of the Long Penetration Mode (LPM) Supersonic Counterflowing Jet to the Supersonic Screech Jet

    NASA Technical Reports Server (NTRS)

    Farr, Rebecca A.; Chang, Chau-Lyan.; Jones, Jess H.; Dougherty, N. Sam

    2015-01-01

    The authors provide a brief overview of the classic tonal screech noise problem created by underexpanded supersonic jets, briefly describing the fluid dynamic-acoustics feedback mechanism that has been long established as the basis for this well-known aeroacoustics problem. This is followed by a description of the Long Penetration Mode (LPM) supersonic underexpanded counterflowing jet phenomenon which has been demonstrated in several wind tunnel tests and modeled in several computational fluid dynamics (CFD) simulations. The authors provide evidence from test and CFD analysis of LPM that indicates that acoustics feedback and fluid interaction seen in LPM are analogous to the aeroacoustics interactions seen in screech jets. Finally, the authors propose applying certain methodologies to LPM which have been developed and successfully demonstrated in the study of screech jets and mechanically induced excitation in fluid oscillators for decades. The authors conclude that the large body of work done on jet screech, other aeroacoustic phenomena, and fluid oscillators can have direct application to the study and applications of LPM counterflowing supersonic cold flow jets.

  7. Study of optical techniques for the Ames unitary wind tunnels. Part 2: Light sheet and vapor screen

    NASA Technical Reports Server (NTRS)

    Lee, George

    1992-01-01

    Light sheet and vapor screen methods have been studied with particular emphasis on those systems that have been used in large transonic and supersonic wind tunnels. The various fluids and solids used as tracers or light scatters and the methods for tracing generation have been studied. Light sources from high intensity lamps and various lasers have been surveyed. Light sheet generation and projection methods were considered. Detectors and location of detectors were briefly studied. A vapor screen system and a technique for location injection of tracers for the NASA Ames 9 by 7 foot Supersonic Wind Tunnel were proposed.

  8. On the Problems of Chaplygin for Mixed Sub-and Supersonic Flows

    NASA Technical Reports Server (NTRS)

    Frankl, F.

    1947-01-01

    There are investigated the problems of the flow of a supersonic jet out of a vessel with plane side walls and the problem of the supersonic flow about a wedge when there is a zone of local subsonic velocities ahead of the wedge.

  9. On the calculation of flow about objects traveling at high supersonic speeds

    NASA Technical Reports Server (NTRS)

    Eggers, A J

    1952-01-01

    A procedure for calculating three-dimensional steady and nonsteady supersonic flows with the method of characteristics is developed and discussed. An approximate method is deduced from the characteristics method and shown to be of practical value at high supersonic speeds.

  10. Experimental Characterization of Acoustic Wave Propagation through a Supersonic Ducted Flow

    DTIC Science & Technology

    2006-01-01

    which basically eliminates the condensate by reducing the temperature to freezing levels. Before air is delivered to the inlet of the test rig, it was...such as supersonic wind-tunnels, inlets of scramjets, as well as supersonic ejectors .35 At high speeds (M>1), these “multiple shock waves” interact with

  11. The role of finite-difference methods in design and analysis for supersonic cruise

    NASA Technical Reports Server (NTRS)

    Townsend, J. C.

    1976-01-01

    Finite-difference methods for analysis of steady, inviscid supersonic flows are described, and their present state of development is assessed with particular attention to their applicability to vehicles designed for efficient cruise flight. Current work is described which will allow greater geometric latitude, improve treatment of embedded shock waves, and relax the requirement that the axial velocity must be supersonic.

  12. A system for aerodynamic design and analysis of supersonic aircraft. Part 4: Test cases

    NASA Technical Reports Server (NTRS)

    Middleton, W. D.; Lundry, J. L.

    1980-01-01

    An integrated system of computer programs was developed for the design and analysis of supersonic configurations. The system uses linearized theory methods for the calculation of surface pressures and supersonic area rule concepts in combination with linearized theory for calculation of aerodynamic force coefficients. Interactive graphics are optional at the user's request. Representative test cases and associated program output are presented.

  13. Bibliography of Supersonic Cruise Research (SCR) program from 1980 to 1983

    NASA Technical Reports Server (NTRS)

    Hoffman, S.

    1984-01-01

    A bibliography for the Supersonic Cruise Research (SCR) and Variable Cycle Engine (VCE) Programs is presented. An annotated bibliography for the last 123 formal reports and a listing of titles for 44 articles and presentations is included. The studies identifies technologies for producing efficient supersonic commercial jet transports for cruise Mach numbers from 2.0 to 2.7.

  14. Advanced Concept Studies for Supersonic Commercial Transports Entering Service in the 2018 to 2020 Period

    NASA Technical Reports Server (NTRS)

    Morgenstern, John; Norstrud, Nicole; Sokhey, Jack; Martens, Steve; Alonso, Juan J.

    2013-01-01

    Lockheed Martin Aeronautics Company (LM), working in conjunction with General Electric Global Research (GE GR), Rolls-Royce Liberty Works (RRLW), and Stanford University, herein presents results from the "N+2 Supersonic Validations" contract s initial 22 month phase, addressing the NASA solicitation "Advanced Concept Studies for Supersonic Commercial Transports Entering Service in the 2018 to 2020 Period." This report version adds documentation of an additional three month low boom test task. The key technical objective of this effort was to validate integrated airframe and propulsion technologies and design methodologies. These capabilities aspired to produce a viable supersonic vehicle design with environmental and performance characteristics. Supersonic testing of both airframe and propulsion technologies (including LM3: 97-023 low boom testing and April-June nozzle acoustic testing) verified LM s supersonic low-boom design methodologies and both GE and RRLW's nozzle technologies for future implementation. The N+2 program is aligned with NASA s Supersonic Project and is focused on providing system-level solutions capable of overcoming the environmental and performance/efficiency barriers to practical supersonic flight. NASA proposed "Initial Environmental Targets and Performance Goals for Future Supersonic Civil Aircraft". The LM N+2 studies are built upon LM s prior N+3 100 passenger design studies. The LM N+2 program addresses low boom design and methodology validations with wind tunnel testing, performance and efficiency goals with system level analysis, and low noise validations with two nozzle (GE and RRLW) acoustic tests.

  15. Analysis and control of asymmetric vortex flows and supersonic vortex breakdown

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.

    1991-01-01

    Topics relative to the analysis and control of asymmetric vortex flow and supersonic vortex breakdown are discussed. Specific topics include the computation of compressible, quasi-axisymmetric slender vortex flow and breakdown; supersonic quasi-axisymmetric vortex breakdown; and three-dimensional Navier-Stokes asymmetric solutions for cones and cone-cylinder configurations.

  16. 3 TUNNELS IN THE ENGINE RESEARCH BUILDING ERB - IN CELL CE-26 VARIABLE REYNOLDS NUMBER SUPERSONIC NO

    NASA Technical Reports Server (NTRS)

    1956-01-01

    3 TUNNELS IN THE ENGINE RESEARCH BUILDING ERB - IN CELL CE-26 VARIABLE REYNOLDS NUMBER SUPERSONIC NOZZLE - CELL CE-4 6X6 INCH MACH NUMBER 2.96 SUPERSONIC AIRPLANE - CELL 1-NW 1X1 FOOT MACH 3.12 SUPERSONIC TUNNEL

  17. Unsteady Flow in a Supersonic Turbine with Variable Specific Heats

    NASA Technical Reports Server (NTRS)

    Dorney, Daniel J.; Griffin, Lisa W.; Huber, Frank; Sondak, Douglas L.; Turner, James (Technical Monitor)

    2001-01-01

    Modern high-work turbines can be compact, transonic, supersonic, counter-rotating, or use a dense drive gas. The vast majority of modern rocket turbine designs fall into these Categories. These turbines usually have large temperature variations across a given stage, and are characterized by large amounts of flow unsteadiness. The flow unsteadiness can have a major impact on the turbine performance and durability. For example, the Space Transportation Main Engine (STME) fuel turbine, a high work, transonic design, was found to have an unsteady inter-row shock which reduced efficiency by 2 points and increased dynamic loading by 24 percent. The Revolutionary Reusable Technology Turbopump (RRTT), which uses full flow oxygen for its drive gas, was found to shed vortices with such energy as to raise serious blade durability concerns. In both cases, the sources of the problems were uncovered (before turbopump testing) with the application of validated, unsteady computational fluid dynamics (CFD) to the designs. In the case of the RRTT and the Alternate Turbopump Development (ATD) turbines, the unsteady CFD codes have been used not just to identify problems, but to guide designs which mitigate problems due to unsteadiness. Using unsteady flow analyses as a part of the design process has led to turbine designs with higher performance (which affects temperature and mass flow rate) and fewer dynamics problems. One of the many assumptions made during the design and analysis of supersonic turbine stages is that the values of the specific heats are constant. In some analyses the value is based on an average of the expected upstream and downstream temperatures. In stages where the temperature can vary by 300 to 500 K, however, the assumption of constant fluid properties may lead to erroneous performance and durability predictions. In this study the suitability of assuming constant specific heats has been investigated by performing three-dimensional unsteady Navier

  18. Preliminary Analysis of Axial-Flow Compressors Having Supersonic Velocity at the Entrance of the Stator

    NASA Technical Reports Server (NTRS)

    Ferri, Antonio

    1949-01-01

    A supersonic compressor design having supersonic velocity at the entrance of the stator is analyzed on the assumption of two-dimensional flow. The rotor and stator losses assumed in the analysis are based on the results of preliminary supersonic cascade tests. The results of the analysis show that compression ratios per stage of 6 to 10 can be obtained with adiabatic efficiency between 70 and 80 percent. Consideration is also given in the analysis to the starting, stability, and range of efficient performance of this type of compressor. The desirability of employing variable-geometry stators and adjustable inlet guide vanes is indicated. Although either supersonic or subsonic axial component of velocity at the stator entrance can be used, the cascade test results suggest that higher pressure recovery can be obtained if the axial component is supersonic.

  19. Experimental investigation of leading-edge thrust at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Wood, R. M.; Miller, D. S.

    1983-01-01

    Wings, designed for leading edge thrust at supersonic speeds, were investigated in the Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.80, 2.00, 2.16, and 2.36. Experimental data were obtained on a uncambered wing which had three interchangeable leading edges that varied from sharp to blunt. The leading edge thrust concept was evaluated. Results from the investigation showed that leading edge flow separation characteristics of all wings tested agree well with theoretical predictions. The experimental data showed that significant changes in wing leading edge bluntness did not affect the zero lift drag of the uncambered wings.

  20. Engine/airframe compatibility studies for supersonic cruise aircraft

    NASA Technical Reports Server (NTRS)

    1975-01-01

    Technology assessment studies were conducted to provide an updated technology base from which an advanced supersonic cruise aircraft can be produced with a high probability of success. An assessment of the gains available through the application of advanced technologies in aerodynamics, propulsion, acoustics, structures, materials, and active controls is developed. The potential market and range requirements as well as economic factors including payload, speed, airline operating costs, and airline profitability are analyzed. The conceptual design of the baseline aircraft to be used in assessing the technology requirements is described.

  1. Structure and propagation of supersonic singularities from helicoidal sources

    NASA Technical Reports Server (NTRS)

    Myers, M. K.; Farassat, F.

    1987-01-01

    An asymptotic analysis of the acoustic field radiated by a supersonic helicoidal line source distribution is given. The asymptotic results are valid in the vicinity of the Mach surfaces associated with the moving sources. Particular attention is paid to the singular nature of the field on the Mach surfaces, which the analysis describes exactly. In addition, it is found that the asymptotic approximation predicts numerical values of the pressure with considerable accuracy. Some details on the field of a single source are derived as a special case.

  2. An improved higher order panel method for linearized supersonic flow

    NASA Technical Reports Server (NTRS)

    Ehlers, F. E.; Epton, M. A.; Johnson, F. T.; Magnus, A. E.; Rubbert, P. E.

    1978-01-01

    An improved higher order panel method for linearized supersonic flow is described. Each panel, defined by four points on the surface, is divided into eight subpanels in such a way that all subpanel and panel edges are contiguous. By prescribing a quadratic distribution of the doublet on each subpanel, the doublet strength is made strictly continuous on the paneled surface. A linear source distribution is also used. Numerical results are smoother and in better agreement with experiment than the previous method with less strict continuity. A brief discussion of superinclined panels used to eliminate interior interference in nacelles is included.

  3. Supersonic Coaxial Jet Experiment for CFD Code Validation

    NASA Technical Reports Server (NTRS)

    Cutler, A. D.; Carty, A. A.; Doerner, S. E.; Diskin, G. S.; Drummond, J. P.

    1999-01-01

    A supersonic coaxial jet facility has been designed to provide experimental data suitable for the validation of CFD codes used to analyze high-speed propulsion flows. The center jet is of a light gas and the coflow jet is of air, and the mixing layer between them is compressible. Various methods have been employed in characterizing the jet flow field, including schlieren visualization, pitot, total temperature and gas sampling probe surveying, and RELIEF velocimetry. A Navier-Stokes code has been used to calculate the nozzle flow field and the results compared to the experiment.

  4. A study of second-order supersonic flow theory

    NASA Technical Reports Server (NTRS)

    Van Dyke, Milton D

    1952-01-01

    Second-order solutions of supersonic-flow problems are sought by iteration, using the linearized solution as the first step. For plane and axially symmetric flows, particular solutions of the iteration equation are discovered which reduce the second-order problem to an equivalent linearized problem. Comparison of second-order solutions with exact and numerical results shows great improvement over linearized theory. For full three-dimensional flow, only a partial particular solution is found. The inclined cone is solved, and the possibility of treating more general problems is considered.

  5. Accuracy Of Hot-Wire Anemometry In Supersonic Turbulence

    NASA Technical Reports Server (NTRS)

    Logan, Pamela; Mckenzie, Robert L.; Bershader, Daniel

    1989-01-01

    Sensitivity of hot-wire probe compared to laser-induced-florescence measurements. Report discusses factors affecting readings of hot-wire anemometer in turbulent supersonic boundary layer. Presents theoretical analysis of responses of hot-wire probe to changes in flow; also compares measurements by hot-wire probe with measurements of same flows by laser-induced fluorescence (LIF). Because LIF provides spatially and temporally resolved data on temperature, density, and pressure, provides independent means to determine responses of hot-wire anemometers to these quantities.

  6. Supersonic minimum length nozzle design for dense gases

    NASA Technical Reports Server (NTRS)

    Aldo, Andrew C.; Argrow, Brian M.

    1993-01-01

    Recently, dense gases have been investigated for many engineering applications such as for turbomachinery and wind tunnels. Supersonic nozzle design for these gases is complicated by their nonclassical behavior in the transonic flow regime. In this paper a method of characteristics (MOC) is developed for two-dimensional (planar) and, primarily, axisymmetric flow of a van der Waals gas. Using a straight aortic line assumption, a centered expansion is used to generate an inviscid wall contour of minimum length. The van der Waals results are compared to previous perfect gas results to show the real gas effects on the flow properties and inviscid wall contours.

  7. Supersonic gas jets for laser-plasma experiments.

    PubMed

    Schmid, K; Veisz, L

    2012-05-01

    We present an in-depth analysis of De Laval nozzles, which are ideal for gas jet generation in a wide variety of experiments. Scaling behavior of parameters especially relevant to laser-plasma experiments as jet collimation, sharpness of the jet edges and Mach number of the resulting jet is studied and several scaling laws are given. Special attention is paid to the problem of the generation of microscopic supersonic jets with diameters as small as 150 μm. In this regime, boundary layers dominate the flow formation and have to be included in the analysis.

  8. Three-dimensional aerodynamic shape optimization of supersonic delta wings

    NASA Technical Reports Server (NTRS)

    Burgreen, Greg W.; Baysal, Oktay

    1994-01-01

    A recently developed three-dimensional aerodynamic shape optimization procedure AeSOP(sub 3D) is described. This procedure incorporates some of the most promising concepts from the area of computational aerodynamic analysis and design, specifically, discrete sensitivity analysis, a fully implicit 3D Computational Fluid Dynamics (CFD) methodology, and 3D Bezier-Bernstein surface parameterizations. The new procedure is demonstrated in the preliminary design of supersonic delta wings. Starting from a symmetric clipped delta wing geometry, a Mach 1.62 asymmetric delta wing and two Mach 1. 5 cranked delta wings were designed subject to various aerodynamic and geometric constraints.

  9. Nonlinear potential analysis techniques for supersonic-hypersonic aerodynamic design

    NASA Technical Reports Server (NTRS)

    Shankar, V.; Clever, W. C.

    1984-01-01

    Approximate nonlinear inviscid theoretical techniques for predicting aerodynamic characteristics and surface pressures for relatively slender vehicles at supersonic and moderate hypersonic speeds were developed. Emphasis was placed on approaches that would be responsive to conceptual configuration design level of effort. Second order small disturbance and full potential theory was utilized to meet this objective. Numerical codes were developed for relatively general three dimensional geometries to evaluate the capability of the approximate equations of motion considered. Results from the computations indicate good agreement with experimental results for a variety of wing, body, and wing-body shapes.

  10. Supersonic/hypersonic aerodynamic methods for aircraft design and analysis

    NASA Technical Reports Server (NTRS)

    Torres, Abel O.

    1992-01-01

    A methodology employed in engineering codes to predict aerodynamic characteristics over arbitrary supersonic/hypersonic configurations is considered. Engineering codes use a combination of simplified methods, based on geometrical impact angle and freestream conditions, to compute pressure distribution over the vehicle's surface in an efficient and timely manner. These approximate methods are valid for both hypersonic (Mach greater than 4) and lower speeds (Mach down to 2). It is concluded that the proposed methodology enables the user to obtain reasonable estimates of vehicle performance and engineering methods are valuable in the design process of these type of vehicles.

  11. Drag Minimization for Wings and Bodies in Supersonic Flow

    NASA Technical Reports Server (NTRS)

    Heaslet, Max A; Fuller, Franklyn B

    1958-01-01

    The minimization of inviscid fluid drag is studied for aerodynamic shapes satisfying the conditions of linearized theory, and subject to imposed constraints on lift, pitching moment, base area, or volume. The problem is transformed to one of determining two-dimensional potential flows satisfying either Laplace's or Poisson's equations with boundary values fixed by the imposed conditions. A general method for determining integral relations between perturbation velocity components is developed. This analysis is not restricted in application to optimum cases; it may be used for any supersonic wing problem.

  12. Asymptotic/numerical analysis of supersonic propeller noise

    NASA Technical Reports Server (NTRS)

    Myers, M. K.; Wydeven, R.

    1989-01-01

    An asymptotic analysis based on the Mach surface structure of the field of a supersonic helical source distribution is applied to predict thickness and loading noise radiated by high speed propeller blades. The theory utilizes an integral representation of the Ffowcs-Williams Hawkings equation in a fully linearized form. The asymptotic results are used for chordwise strips of the blade, while required spanwise integrations are performed numerically. The form of the analysis enables predicted waveforms to be interpreted in terms of Mach surface propagation. A computer code developed to implement the theory is described and found to yield results in close agreement with more exact computations.

  13. Sonic boom measurements from accelerating supersonic tracked sleds

    NASA Technical Reports Server (NTRS)

    Reed, J. W.

    1974-01-01

    Supersonic sled tests on the Sandia 1524-m (5000-ft) track generate sonic booms of sufficient intensity to allow some airblast measurements at distance scales not obtained from wind tunnel or flight tests. During acceleration, an emitted curved boom wave propagates to a caustic, or focus. Detailed measurements around these caustics may help to clarify the overpressure magnification which can occur from real aircraft operations. Six fixed pressure gages have been operated to document the general noise field, and a mobile array of twelve gages.

  14. Unsteady numerical simulations over the BLOODHOUND supersonic car

    NASA Astrophysics Data System (ADS)

    Araya, Guillermo; Evans, B.; Hassan, O.; Morgan, K.

    2010-11-01

    One of the main purposes of the BLOODHOUND SSC project consists on achieving the first 1000 mph record on land. In this study, unsteady flow predictions over the BLOODHOUND supersonic car (http://www.bloodhoundssc.swan.ac.uk/), are shown and discussed with Mach numbers up to 1.4. The governing equations of the flow are solved by implementing a hybrid RANS/LES approach. Close to walls the flow is treated with the RANS-equations and the Menter SST model is considered. In zones with separated flows and significant unsteadiness a subgrid-scale stress model is implemented.

  15. An economic study of an advanced technology supersonic cruise vehicle

    NASA Technical Reports Server (NTRS)

    Smith, C. L.; Williams, L. J.

    1975-01-01

    A description is given of the methods used and the results of an economic study of an advanced technology supersonic cruise vehicle. This vehicle was designed for a maximum range of 4000 n.mi. at a cruise speed of Mach 2.7 and carrying 292 passengers. The economic study includes the estimation of aircraft unit cost, operating cost, and idealized cash flow and discounted cash flow return on investment. In addition, it includes a sensitivity study on the effects of unit cost, manufacturing cost, production quantity, average trip length, fuel cost, load factor, and fare on the aircraft's economic feasibility.

  16. An Experimental and CFD Study of a Supersonic Coaxial Jet

    NASA Technical Reports Server (NTRS)

    Cutler, A. D.; White, J. A.

    2001-01-01

    A supersonic coaxial jet facility is designed and experimental data are acquired suitable for the validation of CFD codes employed in the analysis of high-speed air-breathing engines. The center jet is of a light gas, the coflow jet is of air, and the mixing layer between them is compressible. The jet flow field is characterized using schlieren imaging, surveys with pitot, total temperature and gas sampling probes, and RELIEF velocimetry. VULCAN, a structured grid CFD code, is used to solve for the nozzle and jet flow, and the results are compared to the experiment for several variations of the kappa - omega turbulence model

  17. Instantaneous planar visualization of reacting supersonic flows using silane seeding

    NASA Technical Reports Server (NTRS)

    Smith, Michael W.; Northam, G. B.

    1991-01-01

    A new visualization technique for reacting flows has been developed. This technique, which is suitable for supersonic combustion flows, has been demonstrated on a scramjet combustor model. In this application, gaseous silane (SiH4) was added to the primary hydrogen fuel. When the fuel reacted, so did the (SiH4), producing silica (SiO2) particles in situ. The particles were illuminated with a laser sheet formed from a frequency-doubled Nd:YAG laser (532 nm) beam and the Mie scattering signal was imaged. These planar images of the silica Mie scattering provided instantaneous 'maps' of combustion progress within the turbulent reacting flowfield.

  18. The inviscid stability of supersonic flow past axisymmetric bodies

    NASA Technical Reports Server (NTRS)

    Duck, Peter W.

    1990-01-01

    The supersonic flow past a sharp cone is studied. The associated boundary layer flow (i.e., the velocity and temperature field) is computed. The inviscid linear temporal stability of axisymmetric boundary layers in general is considered, and in particular, a so-called 'triply generalized' inflection condition for 'subsonic' nonaxisymmetric neutral modes is presented. Preliminary numerical results for the stability of the cone boundary layer are presented for a freestream Mach number of 3.8. In particular, a new inviscid mode of instability is seen to occur in certain regimes, and this is shown to be related to a viscous mode found by Duck and Hall (1988).

  19. Flying qualities design criteria applicable to supersonic cruise aircraft

    NASA Technical Reports Server (NTRS)

    Chalk, C. R.

    1980-01-01

    A comprehensive set of flying qualities design criteria was prepared for use in the supersonic cruise research program. The framework for stating the design criteria is established and design criteria are included which address specific failures, approach to dangerous flight conditions, flight at high angle of attack, longitudinal and lateral directional stability and control, the primary flight control system, and secondary flight controls. Examples are given of lateral directional design criteria limiting lateral accelerations at the cockpit, time to roll through 30 deg of bank, and time delay in the pilot's command path. Flight test data from the Concorde certification program are used to substantiate a number of the proposed design criteria.

  20. Numerical Simulation of a Spatially Evolving Supersonic Turbulent Boundary Layer

    NASA Technical Reports Server (NTRS)

    Gatski, T. B.; Erlebacher, G.

    2002-01-01

    The results from direct numerical simulations of a spatially evolving, supersonic, flat-plate turbulent boundary-layer flow, with free-stream Mach number of 2.25 are presented. The simulated flow field extends from a transition region, initiated by wall suction and blowing near the inflow boundary, into the fully turbulent regime. Distributions of mean and turbulent flow quantities are obtained and an analysis of these quantities is performed at a downstream station corresponding to Re(sub x)= 5.548 x10(exp 6) based on distance from the leading edge.

  1. Supersonic second order analysis and optimization program user's manual

    NASA Technical Reports Server (NTRS)

    Clever, W. C.

    1984-01-01

    Approximate nonlinear inviscid theoretical techniques for predicting aerodynamic characteristics and surface pressures for relatively slender vehicles at supersonic and moderate hypersonic speeds were developed. Emphasis was placed on approaches that would be responsive to conceptual configuration design level of effort. Second order small disturbance theory was utilized to meet this objective. Numerical codes were developed for analysis and design of relatively general three dimensional geometries. Results from the computations indicate good agreement with experimental results for a variety of wing, body, and wing-body shapes. Case computational time of one minute on a CDC 176 are typical for practical aircraft arrangement.

  2. Application of NASTRAN to large deflection supersonic flutter of panels

    NASA Technical Reports Server (NTRS)

    Mei, C.; Rogers, J. L., Jr.

    1976-01-01

    Flat panel flutter at high supersonic Mach number is analyzed using NASTRAN Level 16.0 by means of modifications to the code. Two-dimensional plate theory and quasi-steady aerodynamic theory are employed. The finite element formulation and solution procedure are presented. Modifications to the NASTRAN code are discussed. Convergence characteristics of the iteration processes are also briefly discussed. Effects of aerodynamic damping, boundary support condition and applied in-plane loading are included. Comparison of nonlinear vibration and linear flutter results with analytical solutions demonstrate that excellent accuracy is obtained with NASTRAN.

  3. Engine design considerations for 2nd generation supersonic transports

    NASA Technical Reports Server (NTRS)

    Howlett, R. A.

    1975-01-01

    The environmental and economic goals projected for advanced supersonic transports will require revolutionary improvements in propulsion systems. Variable cycle engine concepts that incorporate unique components and advanced technologies show promise in meeting these goals. Pratt & Whitney Aircraft is conducting conceptual design studies of variable cycle engine concepts under NASA sponsorship. This paper reviews some of the design considerations for these engine concepts. Emphasis is placed on jet noise abatement, reduction of emissions, performance improvements, installation considerations, hot-section characteristics and control system requirements. Two representative variable cycle engine concepts that incorporate these basic design considerations are described.

  4. A study of altitude-constrained supersonic cruise transport concepts

    NASA Technical Reports Server (NTRS)

    Tice, David C.; Martin, Glenn L.

    1992-01-01

    The effect of restricting maximum cruise altitude on the mission performance of two supersonic transport concepts across a selection of cruise Mach numbers is studied. Results indicate that a trapezoidal wing concept can be competitive with an arrow wing depending on the altitude and Mach number constraints imposed. The higher wing loading of trapezoidal wing configurations gives them an appreciably lower average cruise altitude than the lower wing loading of the arrow wing configurations, and this advantage increases as the maximum allowable cruise altitude is reduced.

  5. High Order Accuracy Methods for Supersonic Reactive Flows

    DTIC Science & Technology

    2008-06-25

    bubble interactions [12], the supersonic cavity flows [11], and etc. The difficulty of implement - ing the spectral method to these complex fluid...enhance the convergence properties of the approximation via a filter function σ(η) [30] with the following properties σ(η)= σ(−η), σ(±1) = 0, σ(0)= 1, σ... implementation of the Hybrid method depends on the ability to obtain accurate information on the smoothness of a function . In this work, we employ the Multi

  6. Supersonic flow past oscillating airfoils including nonlinear thickness effects

    NASA Technical Reports Server (NTRS)

    Van Dyke, Milton D

    1954-01-01

    A solution to second order in thickness is derived for harmonically oscillating two-dimensional airfoils in supersonic flow. For slow oscillations of an arbitrary profile, the result is found as a series including the third power of frequency. For arbitrary frequencies, the method of solution for any specific profile is indicated, and the explicit solution derived for a single wedge. Nonlinear thickness effects are found generally to reduce the torsional damping, and so enlarge the range of Mach numbers within which torsional instability is possible.

  7. Investigation of Supersonic Boundary Layer Transition and Turbulent Structure

    DTIC Science & Technology

    1990-07-15

    measurements in turbulent supersonic flows. The instrumentation obtained under this grant have helped to apply these new tools to flows of fundamental fluid...Intrusive Instrumentation 1S. ABSTRACT ,Ceaueapje Do, Wviepg fkV"p Uend 6444p4itI~j bu ,w .. q*V’ This report, is the final report for AFOSR Grant 89...0091, awarded under the Defense iversity Research Instrumentation Program (Control No. 08450-0810). In it, we 2describe -the research progress made

  8. Improvements in Modeling 90 degree Bleed Holes for Supersonic Inlets

    NASA Technical Reports Server (NTRS)

    Slater, John W.

    2009-01-01

    The modeling of porous bleed regions as boundary conditions in computational fluid dynamics (CFD) simulations of supersonic inlet flows has been improved through a scaling of sonic flow coefficient data for 90deg bleed holes. The scaling removed the Mach number as a factor in computing the sonic flow coefficient and allowed the data to be fitted with a quadratic equation, with the only factor being the ratio of the plenum static pressure to the surface static pressure. The implementation of the bleed model into the Wind-US CFD flow solver was simplified by no longer requiring the evaluation of the flow properties at the boundary-layer edge. The quadratic equation can be extrapolated to allow the modeling of small amounts of blowing, which can exist when recirculation of the bleed flow occurs within the bleed region. The improved accuracy of the bleed model was demonstrated through CFD simulations of bleed regions on a flat plate in supersonic flow with and without an impinging oblique shock. The bleed model demonstrated good agreement with experimental data and three-dimensional CFD simulations of bleed holes.

  9. Instability of a supersonic shock free elliptic jet

    SciTech Connect

    Baty, R.S. ); Seiner, J.M.; Ponton, M.K. . Langley Research Center)

    1990-01-01

    This paper presents a comparison of the measured and the computed spatial stability properties of an aspect ratio 2 supersonic shock free elliptic jet. The shock free nature of the elliptic jet provides an ideal test of validity of modeling the large scale coherent structures in the initial mixing region of noncircular supersonic jets with linear hydrodynamic stability theory. Both aerodynamic and acoustic data were measured. The data are used to compute the mean velocity profiles and to provide a description of the spatial composition of pressure waves in the elliptic jet. A hybrid numerical scheme is applied to solve the Rayleigh problem governing the inviscid linear spatial stability of the jet. The measured mean velocity profiles are used to provide a qualitative model for the cross sectional geometry and the smooth velocity profiles used in the stability analysis. Computational results are presented for several modes of instability at two jet cross sections. The acoustic measurements show that a varicose instability is the jet's perferred mode of motion. The stability analysis predicts that the Strouhal number varies linearly as a function of axial distance in the jet's initial mixing region, which is in good qualitative agreement with previous measurements. 18 refs., 18 figs., 1 tab.

  10. Supersonic mixing enhancement by vorticity for high-speed propulsion

    NASA Technical Reports Server (NTRS)

    Settles, Gary S.

    1991-01-01

    The results of a three year study on vortex enhancement of supersonic mixing are discussed. Recent interests in compressible mixing has spurred research in the field of high speed shear layers. It was established that shear layer growth diminishes with increasing convective Mach number; this Mach number is the relative Mach number of the large scale structures in the shear layer with respect to the Mach numbers on either side of the shear layer. The purpose was to evaluate the effect of swirl on compressible mixing rates. Previously analytical and experimental results seem to indicate that swirling flow may significantly modify the shear layer, in some cases resulting in enhanced mixing. Previous studies of the effect of swirl on compressible mixing were incomplete since the amount of swirl in the flowfield was not quantified. This study was undertaken to conclusively determine the effect of swirl on supersonic mixing, including the quantification of the swirl. Preliminary results indicate that the swirl modestly enhances the mixing rates.

  11. Noise from Supersonic Coaxial Jets. Part 1; Mean Flow Predictions

    NASA Technical Reports Server (NTRS)

    Dahl, Milo D.; Morris, Philip J.

    1997-01-01

    Recent theories for supersonic jet noise have used an instability wave noise generation model to predict radiated noise. This model requires a known mean flow that has typically been described by simple analytic functions for single jet mean flows. The mean flow of supersonic coaxial jets is not described easily in terms of analytic functions. To provide these profiles at all axial locations, a numerical scheme is developed to calculate the mean flow properties of a coaxial jet. The Reynolds-averaged, compressible, parabolic boundary layer equations are solved using a mixing length turbulence model. Empirical correlations are developed to account for the effects of velocity and temperature ratios and Mach number on the shear layer spreading. Both normal velocity profile and inverted velocity profile coaxial jets are considered. The mixing length model is modified in each case to obtain reasonable results when the two stream jet merges into a single fully developed jet. The mean flow calculations show both good qualitative and quantitative agreement with measurements in single and coaxial jet flows.

  12. Jet Noise Modeling for Supersonic Business Jet Application

    NASA Technical Reports Server (NTRS)

    Stone, James R.; Krejsa, Eugene A.; Clark, Bruce J.

    2004-01-01

    This document describes the development of an improved predictive model for coannular jet noise, including noise suppression modifications applicable to small supersonic-cruise aircraft such as the Supersonic Business Jet (SBJ), for NASA Langley Research Center (LaRC). For such aircraft a wide range of propulsion and integration options are under consideration. Thus there is a need for very versatile design tools, including a noise prediction model. The approach used is similar to that used with great success by the Modern Technologies Corporation (MTC) in developing a noise prediction model for two-dimensional mixer ejector (2DME) nozzles under the High Speed Research Program and in developing a more recent model for coannular nozzles over a wide range of conditions. If highly suppressed configurations are ultimately required, the 2DME model is expected to provide reasonable prediction for these smaller scales, although this has not been demonstrated. It is considered likely that more modest suppression approaches, such as dual stream nozzles featuring chevron or chute suppressors, perhaps in conjunction with inverted velocity profiles (IVP), will be sufficient for the SBJ.

  13. Instability of a supersonic shock free elliptic jet

    NASA Technical Reports Server (NTRS)

    Baty, Roy S.; Seiner, John M.; Ponton, Michael K.

    1990-01-01

    This paper presents a comparison of the measured and the computed spatial stability properties of an aspect ratio 2 supersonic shock free elliptic jet. The shock free nature of the elliptic jet provides an ideal test of validity of modeling the large scale coherent structures in the initial mixing region of noncircular supersonic jets with linear hydrodynamic stability theory. Both aerodynamic and acoustic data were measured. The data are used to compute the mean velocity profiles and to provide a description of the spatial composition of pressure waves in the elliptic jet. A hybrid numerical scheme is applied to solve the Rayleigh problem governing the inviscid linear spatial stability of the jet. The measured mean velocity profiles are used to provide a qualitative model for the cross sectional geometry and the smooth velocity profiles used in the stability analysis. Computational results are presented for several modes of instability at two jet cross sections. The acoustic measurements show that a varicose instability is the jet's perferred mode of motion. The stability analysis predicts that the Strouhal number varies linearly as a function of axial distance in the jet's initial mixing region, which is in good qualitative agreement with previous measurements.

  14. Subsonic and Supersonic Effects in Bose-Einstein Condensate

    NASA Technical Reports Server (NTRS)

    Zak, Michail

    2003-01-01

    A paper presents a theoretical investigation of subsonic and supersonic effects in a Bose-Einstein condensate (BEC). The BEC is represented by a time-dependent, nonlinear Schroedinger equation that includes terms for an external confining potential term and a weak interatomic repulsive potential proportional to the number density of atoms. From this model are derived Madelung equations, which relate the quantum phase with the number density, and which are used to represent excitations propagating through the BEC. These equations are shown to be analogous to the classical equations of flow of an inviscid, compressible fluid characterized by a speed of sound (g/Po)1/2, where g is the coefficient of the repulsive potential and Po is the unperturbed mass density of the BEC. The equations are used to study the effects of a region of perturbation moving through the BEC. The excitations created by a perturbation moving at subsonic speed are found to be described by a Laplace equation and to propagate at infinite speed. For a supersonically moving perturbation, the excitations are found to be described by a wave equation and to propagate at finite speed inside a Mach cone.

  15. SUPERSONIC LINE BROADENING WITHIN YOUNG AND MASSIVE SUPER STAR CLUSTERS

    SciTech Connect

    Tenorio-Tagle, Guillermo; Silich, Sergiy; Wuensch, Richard; Munoz-Tunon, Casiana; Palous, Jan E-mail: richard@wunsch.c E-mail: cmt@ll.iac.e

    2010-01-10

    The origin of supersonic infrared and radio recombination nebular lines often detected in young and massive superstar clusters is discussed. We suggest that these arise from a collection of repressurizing shocks (RSs), acting effectively to re-establish pressure balance within the cluster volume and from the cluster wind which leads to an even broader although much weaker component. The supersonic lines here are shown to occur in clusters that undergo a bimodal hydrodynamic solution, that is within clusters that are above the threshold line in the mechanical luminosity or cluster mass versus the size of the cluster plane. A plethora of RSs is due to frequent and recurrent thermal instabilities that take place within the matter reinserted by stellar winds and supernovae. We show that the maximum speed of the RSs and of the cluster wind are both functions of the temperature reached at the stagnation radius. This temperature depends only on the cluster heating efficiency (eta). Based on our two-dimensional simulations we calculate the line profiles that result from several models and confirm our analytical predictions. From a comparison between the predicted and observed values of the half-width zero intensity of the two line components, we conclude that the thermalization efficiency in young super star clusters above the threshold line must be lower than 20%.

  16. Subsonic/supersonic aerodynamic characteristics for a tactical supercruiser

    NASA Technical Reports Server (NTRS)

    Capone, F. J.; Bare, E. A.; Hollenback, D.; Hutchison, R.

    1984-01-01

    A series of cooperative NASA-Langley/Boeing experimental investigations have been conducted to determine the aeropropulsive characteristics of an advanced tactical fighter designed for supersonic cruise. These investigations were conducted in the Langley 16-Foot Transonic and Lewis 10 x 10-Foot Supersonic Wind Tunnels at Mach numbers from 0.60 to 2.47. This fighter is a Mach 2.0, 49,000 pound class vehicle that features a close-coupled canard and underwing propulsion units that utilize multifunction two-dimensional exhaust nozzles. Tests were conducted to determine the basic aerodynamic characteristics of the configuration with flow-through nacelles in which the spillage effects of representative inlets were measured. The effects of thrust-induced forces on overall aerodynamic performance were evaluated with a series of multifunction nozzles installed on air-powered nacelles. An axisymmetric nozzle configuration was also tested to obtain comparative aeropropulsive performance. Trim aerodynamic characteristics for the flow-through and powered configurations and the effect of thrust vectoring at subsonic speeds are presented.

  17. The wedge hot-film anemometer in supersonic flow

    NASA Technical Reports Server (NTRS)

    Seiner, J. M.

    1983-01-01

    A commercial wedge hot-film probe is studied to determine its heat transfer response in transonic to low supersonic flows of high unit Reynolds number. The results of this study show that its response in this flow regime differs from the response of cylindrical type sensors. Whereas the cylindrical sensor has the same sensitivity to velocity as to density for free-stream Mach numbers exceeding 1.3, the wedge probe sensitivity to velocity is always greater than its sensitivity to density over the entire flow regime. This property requires determination of three fluctuation components due to density, velocity, and temperature, in a transonic or supersonic turbulent flow. Sensitivity equations are derived based on the observed behavior of the wedge probe. Both the durability and the frequency response of the probe are excellent, the square wave insertion test indicating frequency response near 130 kHz. The directional response of the probe at sonic speed is poor and requires further examination before Reynolds stress measurements are attempted with dual sensor probes.

  18. The flip-flop nozzle extended to supersonic flows

    NASA Technical Reports Server (NTRS)

    Raman, Ganesh; Hailye, Michael; Rice, Edward J.

    1992-01-01

    An experiment studying a fluidically oscillated rectangular jet flow was conducted. The Mach number was varied over a range from low subsonic to supersonic. Unsteady velocity and pressure measurements were made using hot wires and piezoresistive pressure transducers. In addition smoke flow visualization using high speed photography was used to document the oscillation of the jet. For the subsonic flip-flop jet it was found that the apparent time-mean widening of the jet was not accompanied by an increase in mass flux. It was found that it is possible to extend the operation of these devices to supersonic flows. Most of the measurements were made for a fixed nozzle geometry for which the oscillations ceased at a fully expanded Mach number of 1.58. By varying the nozzle geometry this limitation was overcome and operation was extended to Mach 1.8. The streamwise velocity perturbation levels produced by this device were much higher than the perturbation levels that could be produced using conventional excitation sources such as acoustic drivers. In view of this ability to produce high amplitudes, the potential for using small scale fluidically oscillated jet as an unsteady excitation source for the control of shear flows in full scale practical applications seems promising.

  19. A compressible multiphase framework for simulating supersonic atomization

    NASA Astrophysics Data System (ADS)

    Regele, Jonathan D.; Garrick, Daniel P.; Hosseinzadeh-Nik, Zahra; Aslani, Mohamad; Owkes, Mark

    2016-11-01

    The study of atomization in supersonic combustors is critical in designing efficient and high performance scramjets. Numerical methods incorporating surface tension effects have largely focused on the incompressible regime as most atomization applications occur at low Mach numbers. Simulating surface tension effects in high speed compressible flow requires robust numerical methods that can handle discontinuities caused by both material interfaces and shocks. A shock capturing/diffused interface method is developed to simulate high-speed compressible gas-liquid flows with surface tension effects using the five-equation model. This includes developments that account for the interfacial pressure jump that occurs in the presence of surface tension. A simple and efficient method for computing local interface curvature is developed and an acoustic non-dimensional scaling for the surface tension force is proposed. The method successfully captures a variety of droplet breakup modes over a range of Weber numbers and demonstrates the impact of surface tension in countering droplet deformation in both subsonic and supersonic cross flows.

  20. Thermonuclear dynamo inside ultracentrifuge with supersonic plasma flow stabilization

    NASA Astrophysics Data System (ADS)

    Winterberg, F.

    2016-01-01

    Einstein's general theory of relativity implies the existence of virtual negative masses in the rotational reference frame of an ultracentrifuge with the negative mass density of the same order of magnitude as the positive mass density of a neutron star. In an ultracentrifuge, the repulsive gravitational field of this negative mass can simulate the attractive positive mass of a mini-neutron star, and for this reason can radially confine a dense thermonuclear plasma placed inside the centrifuge, very much as the positive mass of a star confines its plasma by its own attractive gravitational field. If the centrifuge is placed in an externally magnetic field to act as the seed field of a magnetohydrodynamic generator, the configuration resembles a magnetar driven by the release of energy through nuclear fusion, accelerating the plasma to supersonic velocities, with the magnetic field produced by the thermomagnetic Nernst effect insulating the hot plasma from the cold wall of the centrifuge. Because of the supersonic flow and the high plasma density the configuration is stable.

  1. An Investigation of Cavity Vortex Generators in Supersonic Flow

    NASA Technical Reports Server (NTRS)

    Hazlewood, Richard

    1996-01-01

    The purpose of this report is to document the results of experiments performed at the University of Kansas and at the NASA Langley Research Center (LaRC) into the use of shaped cavities to generate vortices in supersonic flow, as well as the progress made in simulating the observed flow using the PAB3D flow solver. The investigation was performed on 18 different cavity configurations installed in a convergent-divergent nozzle at the Jet Exit Facility at the LaRC. Pressure sensitive paint, static-pressure ports, focusing Schliern, and water tunnel flow visualization techniques were used to study the nature of the flow created by these cavities. The results of these investigations revealed that a shaped cavity can generate a pair of counter-rotating streamwise vortices in supersonic flow by creating weak, compression Mach waves and weak shocks. The PAB3D computer program, developed at the LaRC, was used to attempt to reproduce the experimental results. Unfortunately, due to problems with matching the grid blocks, no converged results were obtained. However, intermediate results, as well as a complete definition of the grid matching problems and suggested courses of actions are presented.

  2. Gas-Liquid Supersonic Cleaning and Cleaning Verification Spray System

    NASA Technical Reports Server (NTRS)

    Parrish, Lewis M.

    2009-01-01

    NASA Kennedy Space Center (KSC) recently entered into a nonexclusive license agreement with Applied Cryogenic Solutions (ACS), Inc. (Galveston, TX) to commercialize its Gas-Liquid Supersonic Cleaning and Cleaning Verification Spray System technology. This technology, developed by KSC, is a critical component of processes being developed and commercialized by ACS to replace current mechanical and chemical cleaning and descaling methods used by numerous industries. Pilot trials on heat exchanger tubing components have shown that the ACS technology provides for: Superior cleaning in a much shorter period of time. Lower energy and labor requirements for cleaning and de-scaling uper.ninih. Significant reductions in waste volumes by not using water, acidic or basic solutions, organic solvents, or nonvolatile solid abrasives as components in the cleaning process. Improved energy efficiency in post-cleaning heat exchanger operations. The ACS process consists of a spray head containing supersonic converging/diverging nozzles, a source of liquid gas; a novel, proprietary pumping system that permits pumping liquid nitrogen, liquid air, or supercritical carbon dioxide to pressures in the range of 20,000 to 60,000 psi; and various hoses, fittings, valves, and gauges. The size and number of nozzles can be varied so the system can be built in configurations ranging from small hand-held spray heads to large multinozzle cleaners. The system also can be used to verify if a part has been adequately cleaned.

  3. Thermonuclear dynamo inside ultracentrifuge with supersonic plasma flow stabilization

    SciTech Connect

    Winterberg, F.

    2016-01-15

    Einstein's general theory of relativity implies the existence of virtual negative masses in the rotational reference frame of an ultracentrifuge with the negative mass density of the same order of magnitude as the positive mass density of a neutron star. In an ultracentrifuge, the repulsive gravitational field of this negative mass can simulate the attractive positive mass of a mini-neutron star, and for this reason can radially confine a dense thermonuclear plasma placed inside the centrifuge, very much as the positive mass of a star confines its plasma by its own attractive gravitational field. If the centrifuge is placed in an externally magnetic field to act as the seed field of a magnetohydrodynamic generator, the configuration resembles a magnetar driven by the release of energy through nuclear fusion, accelerating the plasma to supersonic velocities, with the magnetic field produced by the thermomagnetic Nernst effect insulating the hot plasma from the cold wall of the centrifuge. Because of the supersonic flow and the high plasma density the configuration is stable.

  4. Investigation of radiative interactions in supersonic internal flows

    NASA Technical Reports Server (NTRS)

    Tiwari, Surendra N.; Thomas, A. M.

    1991-01-01

    Analyses and numerical procedures are presented to study the radiative interactions of absorbing emitting species in chemically reacting supersonic flow in various ducts. The 2-D time dependent Navier-Stokes equations in conjunction with radiative flux equation are used to study supersonic flows undergoing finite rate chemical reaction in a hydrogen air system. The specific problem considered is the flow of premixed radiating gas between parallel plates. Specific attention was directed toward studying the radiative contribution of H2O, OH, and NO under realistic physical and flow conditions. Results are presented for the radiative flux obtained for different gases and for various combination of these gases. The problem of chemically reacting and radiating flows was solved for the flow of premixed hydrogen-air through a 10 deg compression ramp. Results demonstrate that the radiative interaction increases with an increase in pressure, temperature, amount of participating species, plate spacing, and Mach number. Most of the energy, however, is transferred by convection in the flow direction. In general the results indicate that radiation can have a significant effect on the entire flow field.

  5. Supersonic Aerodynamic Characteristics of Proposed Mars '07 Smart Lander Configurations

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Horvath, Thomas J.; Erickson, Gary E.; Green, Joseph M.

    2002-01-01

    Supersonic aerodynamic data were obtained for proposed Mars '07 Smart Lander configurations in NASA Langley Research Center's Unitary Plan Wind Tunnel. The primary objective of this test program was to assess the supersonic aerodynamic characteristics of the baseline Smart Lander configuration with and without fixed shelf/tab control surfaces. Data were obtained over a Mach number range of 2.3 to 4.5, at a free stream Reynolds Number of 1 x 10(exp 6) based on body diameter. All configurations were run at angles of attack from -5 to 20 degrees and angles of sideslip of -5 to 5 degrees. These results were complemented with computational fluid dynamic (CFD) predictions to enhance the understanding of experimentally observed aerodynamic trends. Inviscid and viscous full model CFD solutions compared well with experimental results for the baseline and 3 shelf/tab configurations. Over the range tested, Mach number effects were shown to be small on vehicle aerodynamic characteristics. Based on the results from 3 different shelf/tab configurations, a fixed control surface appears to be a feasible concept for meeting aerodynamic performance metrics necessary to satisfy mission requirements.

  6. Noise reduction in supersonic jets by nozzle fluidic inserts

    NASA Astrophysics Data System (ADS)

    Morris, Philip J.; McLaughlin, Dennis K.; Kuo, Ching-Wen

    2013-08-01

    Professor Philip Doak spent a very productive time as a consultant to the Lockheed-Georgia Company in the early 1970s. The focus of the overall research project was the prediction and reduction of noise from supersonic jets. Now, 40 years on, the present paper describes an innovative methodology and device for the reduction of supersonic jet noise. The goal is the development of a practical active noise reduction technique for low bypass ratio turbofan engines. This method introduces fluidic inserts installed in the divergent wall of a CD nozzle to replace hard-wall corrugation seals, which have been demonstrated to be effective by Seiner (2005) [1]. By altering the configuration and operating conditions of the fluidic inserts, active noise reduction for both mixing and shock noise has been obtained. Substantial noise reductions have been achieved for mixing noise in the maximum noise emission direction and in the forward arc for broadband shock-associated noise. To achieve these reductions (on the order of greater than 4 and 2 dB for the two main components respectively), practically achievable levels of injection mass flow rates have been used. The total injected mass flow rates are less than 4% of the core mass flow rate and the effective operating injection pressure ratio has been maintained at or below the same level as the nozzle pressure ratio of the core flow.

  7. Experimental Study of Plasma Control of an Unstarting Supersonic Flow

    NASA Astrophysics Data System (ADS)

    Im, Seong-Kyun; Do, Hyungrok; Cappelli, Mark A.

    2011-11-01

    Experimental studies of the control of unstarting supersonic model inlet flows using Dielectric Barrier Discharges (DBD) is demonstrated at Mach 4.7 flow conditions and a static temperature of ~60K and static pressure of ~1kPa. Planar Laser Rayleigh Scattering (PLRS) is used to visualize important flow features, such as boundary layers and shockwaves. Supersonic flow unstart is initiated by injecting mass into model inlet flows of either laminar or tripped turbulent boundary layer flow conditions. DBD discharge actuation of the tripped turbulent flow delays the unstart process, shifting the unstart dynamics closer to what is seen for the laminar boundary layer case. In all studies, a single DBD actuator pair is used, oriented parallel to the freestream flow, generating spanwise disturbances. It is proposed that strong suction flow which brings high momentum freestream flow near exposed electrode can be a mechanism of this actuation. PLRS reveals that this actuation is spatially confined to the regions close to the actuator electrodes, greatly limiting their performance. This material is based upon work supported by the Department of Energy under Award Number(s)DE-FC52-08NA28614.

  8. Experimental Testing of the T-NSTAP in Supersonic Flow

    NASA Astrophysics Data System (ADS)

    Kokmanian, Katherine; Hultmark, Marcus

    2016-11-01

    A fast response nanoscale temperature sensor (T-NSTAP) was developed at Princeton University. This novel sensor has been shown to increase both the spatial and the temporal resolutions compared to conventional cold-wire probes, due to its large aspect ratio yet small overall size (100 nm x 2 μm x 200 μm). The T-NSTAP has been tested in various subsonic facilities, however it has not yet been tested under supersonic conditions. Here we will present the first measurements from supersonic flows using the T-NSTAP in Princeton's Low Turbulence Variable Geometry Facility at Mach 3 and later in Princeton's Hypersonic Boundary Layer Facility (HBLF) at Mach 8 in order to enable unfiltered data of the temperature field in high speed flows. Since the HBLF can generate more challenging conditions than these probes have previously been tested in, our attention will be focused on ensuring that the T-NSTAP can withstand these conditions. Assuming that a shock will form at the front edge of the sensor, the total force on the T-NSTAP was calculated to be on the order of μN, which is less than when it was tested in subsonic pressurized conditions. Investigations will be undertaken to ensure that the structural and electrical properties of the sensors are maintained during the tests. Air Force Office of Scientific Research (AFOSR).

  9. Investigation of supersonic chemically reacting and radiating channel flow

    NASA Technical Reports Server (NTRS)

    Mani, Mortaza; Tiwari, Surendra N.

    1988-01-01

    The 2-D time-dependent Navier-Stokes equations are used to investigate supersonic flows undergoing finite rate chemical reaction and radiation interaction for a hydrogen-air system. The explicit multistage finite volume technique of Jameson is used to advance the governing equations in time until convergence is achieved. The chemistry source term in the species equation is treated implicitly to alleviate the stiffness associated with fast reactions. The multidimensional radiative transfer equations for a nongray model are provided for a general configuration and then reduced for a planar geometry. Both pseudo-gray and nongray models are used to represent the absorption-emission characteristics of the participating species. The supersonic inviscid and viscous, nonreacting flows are solved by employing the finite volume technique of Jameson and the unsplit finite difference scheme of MacCormack. The specified problem considered is of the flow in a channel with a 10 deg compression-expansion ramp. The calculated results are compared with those of an upwind scheme. The problem of chemically reacting and radiating flows are solved for the flow of premixed hydrogen-air through a channel with parallel boundaries, and a channel with a compression corner. Results obtained for specific conditions indicate that the radiative interaction can have a significant influence on the entire flow field.

  10. Direct simulation of isothermal-wall supersonic channel flow

    NASA Technical Reports Server (NTRS)

    Coleman, Gary N.

    1993-01-01

    The motivation for this work is the fact that in turbulent flows where compressibility effects are important, they are often poorly understood. A few examples of such flows are those associated with astrophysical phenomena and those found in combustion chambers, supersonic diffusers and nozzles, and over high-speed airfoils. For this project, we are primarily interested in compressibility effects near solid surfaces. Our main objective is an improved understanding of the fundamentals of compressible wall-bounded turbulence, which can in turn be used to cast light upon modeling concepts such as the Morkovin hypothesis and the Van Driest transformation. To this end, we have performed a direct numerical simulation (DNS) study of supersonic turbulent flow in a plane channel with constant-temperature walls. All of the relevant spatial and temporal scales are resolved so that no sub grid scale or turbulence model is necessary. The channel geometry was chosen so that finite Mach number effects can be isolated by comparing the present results to well established incompressible channel data. Here the fluid is assumed to be an ideal gas with constant specific heats, constant Prandtl number, and power-law temperature-dependent viscosity. Isothermal-wall boundary conditions are imposed so that a statistically stationary state may be obtained. The flow is driven by a uniform (in space) body force (rather than a mean pressure gradient) to preserve stream wise homogeneity, with the body force defined so that the total mass flux is constant.

  11. Flow Simulation of Supersonic Inlet with Bypass Annular Duct

    NASA Technical Reports Server (NTRS)

    Kim, HyoungJin; Kumano, Takayasu; Liou, Meng-Sing; Povinelli, Louis A.; Conners, Timothy R.

    2011-01-01

    A relaxed isentropic compression supersonic inlet is a new concept that produces smaller cowl drag than a conventional inlet, but incurs lower total pressure recovery and increased flow distortion in the (radially) outer flowpath. A supersonic inlet comprising a bypass annulus to the relaxed isentropic compression inlet dumps out airflow of low quality through the bypass duct. A reliable computational fluid dynamics solution can provide considerable useful information to ascertain quantitatively relative merits of the concept, and further provide a basis for optimizing the design. For a fast and reliable performance evaluation of the inlet performance, an equivalent axisymmetric model whose area changes accounts for geometric and physical (blockage) effects resulting from the original complex three-dimensional configuration is proposed. In addition, full three-dimensional calculations are conducted for studying flow phenomena and verifying the validity of the equivalent model. The inlet-engine coupling is carried out by embedding numerical propulsion system simulation engine data into the flow solver for interactive boundary conditions at the engine fan face and exhaust plane. It was found that the blockage resulting from complex three-dimensional geometries in the bypass duct causes significant degradation of inlet performance by pushing the terminal normal shock upstream.

  12. Supersonic flow and shock formation in turbine tip gaps

    NASA Technical Reports Server (NTRS)

    Moore, John

    1993-01-01

    Shock formation due to overexpansion of supersonic flow at the inlet to the tip clearance gap of a turbomachine has been studied. As the flow enters the tip gap, it accelerates around the blade pressure-side corner creating a region of minimum static pressure. The 'free streamline' separates from the wall at the corner; and, for Mach numbers greater than about 1.3, it curves back to intersect the blade tip. At this point, the freestream flow is abruptly turned parallel to the surface, giving rise to an oblique shock. The results are consistent with compressible sharp-edged orifice flow calculations found in the literature and with the theory of oblique shock wave formation in supersonic flow over a wedge. For freestream Mach numbers of 1.4 to 1.8, wave angles are 43 to 54 deg, and turning angles are 9 to 20 deg; as the Mach number increases, the angle of turn also increases. It appears that in a turbine, after separating from the inlet corner, the flow reattaches on the blade tip and an oblique shock is formed at 0.4-1.4 tip gap heights into the clearance gap. The resulting shock-boundary layer interaction may contribute to further enhancement of already high heat transfer to the blade tip in this region. This in turn could lead to higher blade temperatures and adversely affect blade life and turbine efficiency.

  13. The Kirchhoff Formula for a Supersonically Moving Surface

    NASA Technical Reports Server (NTRS)

    Farassat, F.; Myers, M. K.

    1996-01-01

    The Kirchhoff formula for radiation from stationary surfaces first appeared in 1882, and it has since found many applications in wave propagation theory. In 1930, Morgans extended the formula to apply to surfaces moving at speeds below the wave propagation speed; we refer to Morgans formula as the subsonic formulation. A modern derivation of Morgans result was published by Farassat and Myers in 1988, and it has now been used extensively in acoustics, particularly for high speed helicopter rotor noise prediction. Under some common conditions in this application, however, the appropriate Kirchhoff surface must be chosen such that portions of it travel at supersonic speed. The available Kirchhoff formula for moving surfaces is not suitable for this situation. In the current paper we derive the Kirchhoff formula applicable to a supersonically moving surface using some results from generalized function theory. The new formula requires knowledge of the same surface data as in the subsonic case. Complications that arise from apparent singularities in the new formulation are discussed briefly in the paper.

  14. Pressure Probe Designs for Dynamic Pressure Measurements in a Supersonic Flow Field. [conducted in the Glenn Supersonic Wind Tunnel (SWT)

    NASA Technical Reports Server (NTRS)

    Porro, A. Robert

    2001-01-01

    A series of dynamic flow field pressure probes were developed for use in large-scale supersonic wind tunnels at NASA Glenn Research Center. These flow field probes include pitot, static, and five-hole conical pressure probes that are capable of capturing fast acting flow field pressure transients that occur on a millisecond time scale. The pitot and static probes can be used to determine local Mach number time histories during a transient event. The five-hole conical pressure probes are used primarily to determine local flow angularity, but can also determine local Mach number. These probes were designed, developed, and tested at the NASA Glenn Research Center. They were also used in a NASA Glenn 10- by 10-Foot Supersonic Wind Tunnel (SWT) test program where they successfully acquired flow field pressure data in the vicinity of a propulsion system during an engine compressor stall and inlet unstart transient event. Details of the design, development, and subsequent use of these probes are discussed in this report.

  15. Flow and Acoustic Properties of Low Reynolds Number Underexpanded Supersonic Jets. Ph.D. Thesis

    NASA Technical Reports Server (NTRS)

    Hu, Tieh-Feng

    1981-01-01

    Jet noise on underexpanded supersonic jets are studied with emphasis on determining the role played by large scale organized flow fluctuations in the flow and acoustic processes. The experimental conditions of the study were chosen as low Reynolds number (Re=8,000) Mach 1.4 and 2.1, and moderate Reynolds number (Re=68,000) Mach 1.6 underexpanded supersonic jets exhausting from convergent nozzles. At these chosen conditions, detailed experimental measurements were performed to improve the understanding of the flow and acoustic properties of underexpanded supersonic jets.

  16. Constant-temperature hot-wire anemometer practice in supersonic flows. II - The inclined wire

    NASA Technical Reports Server (NTRS)

    Smits, A. J.; Muck, K. C.

    1983-01-01

    The performance of a constant-temperature inclined hot-wire in a supersonic flow is critically examined. It is shown that calibration techniques applicable to subsonic flow, such as the cosine cooling law cannot be used when the flow is supersonic. Calibration and measurement procedures appropriate to supersonic flow are suggested, together with the possible limits on their validity. Experimental results for different wires indicate that the sensitivities do not seem to depend on flow direction according to any simple correlation. When the sensitivity exhibits a strong dependence on flow direction, the wire should be discarded to avoid errors due to nonlinear effects.

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

    NASA Technical Reports Server (NTRS)

    Fletcher, D. G.; Mcdaniel, J. C.

    1989-01-01

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

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

    NASA Technical Reports Server (NTRS)

    1990-01-01

    A preliminary study of a supersonic short takeoff and vertical landing (STOVL) fighter is presented. Three configurations (a lift plus lift/cruise concept, a hybrid fan vectored thrust concept, and a mixed flow vectored thrust concept) were initially investigated with one configuration selected for further design analysis. The selected configuration, the lift plus lift/cruise concept, was successfully integrated to accommodate the powered lift short takeoff and vertical landing requirements as well as the demanding supersonic cruise and point performance requirements. A supersonic fighter aircraft with a short takeoff and vertical landing capability using the lift plus lift/cruise engine concept seems a viable option for the next generation fighter.

  19. A Qualitative Piloted Evaluation of the Tupolev Tu-144 Supersonic Transport

    NASA Technical Reports Server (NTRS)

    Rivers, Robert A.; Jackson, E. Bruce; Fullerton, C. Gordon; Cox, Timothy H.; Princen, Norman H.

    2000-01-01

    Two U.S. research pilots evaluated the Tupolev Tu-144 supersonic transport aircraft on three dedicated flights: one subsonic and two supersonic profiles. The flight profiles and maneuvers were developed jointly by Tupolev and U.S. engineers. The vehicle was found to have unique operational and flight characteristics that serve as lessons for designers of future supersonic transport aircraft. Vehicle subsystems and observed characteristics are described as are flight test planning and ground monitoring facilities. Maneuver descriptions and extended pilot narratives for each flight are included as appendices.

  20. Review of passive shear-flow control research for improved subsonic and supersonic combustion

    SciTech Connect

    Schadow, K.C.; Gutmark, E.

    1989-01-01

    Shear-flow investigations have been conducted in the high-Re, turbulent initial-condition combustion regime representative of flow configurations encountered in ramjets and in supersonic plumes. Large-scale vortical structures were identified and characterized in both nonreacting and combustion-reaction experimental conditions; attention was given to these structures' role in mixing, and their breakup into fine-scale turbulence. Shear-flow/combustion control was obtained by actively enlisting duct acoustics and passively employing noncircular flow cross-sections. The investigations were extended to supersonic shear flows, yielding improved mixing for supersonic combustion. 44 refs.

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

    NASA Technical Reports Server (NTRS)

    Appa, Kari; Smith, Michael J. C.

    1987-01-01

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

  2. Verification Assessment of Flow Boundary Conditions for CFD Analysis of Supersonic Inlet Flows

    NASA Technical Reports Server (NTRS)

    Slater, John W.

    2002-01-01

    Boundary conditions for subsonic inflow, bleed, and subsonic outflow as implemented into the WIND CFD code are assessed with respect to verification for steady and unsteady flows associated with supersonic inlets. Verification procedures include grid convergence studies and comparisons to analytical data. The objective is to examine errors, limitations, capabilities, and behavior of the boundary conditions. Computational studies were performed on configurations derived from a "parameterized" supersonic inlet. These include steady supersonic flows with normal and oblique shocks, steady subsonic flow in a diffuser, and unsteady flow with the propagation and reflection of an acoustic disturbance.

  3. Transonic wind tunnel test of a supersonic nozzle installation

    NASA Technical Reports Server (NTRS)

    Yetter, J. A.; Evelyn, G. B.; Mercer, C.

    1982-01-01

    The design of the propulsion system installation affects strongly the total drag and overall performance of an aircraft, and the concept, placement, and integration details of the exhaust nozzle are major considerations in the configuration definition. As part of the NASA Supersonic Cruise Research (SCR) program, a wind tunnel test program has been conducted to investigate exhaust nozzle-airframe interactions at transonic speeds. First phase testing is to establish guidelines for follow-on testing. A summary is provided of the results of first phase testing, taking into account the test approach, the effect of nozzle closure on aircraft aerodynamic characteristics, nozzle installation effects and nacelle interference drag, and an analytical study of the effects of nozzle closure on the aircraft.

  4. Mixing Enhancement by Tabs in Round Supersonic Jets

    NASA Technical Reports Server (NTRS)

    Seiner, John M.; Grosch, C. E.

    1998-01-01

    The objective of this study was to analyze jet plume mass flow entrainment rates associated with the introduction of counter-rotating streamwise vorticity by prism shaped devices (tabs) located at the lip of the nozzle. We have examined the resulting mixing process through coordinated experimental tests and numerical simulations of the supersonic flow from a model axisymmetric nozzle. In the numerical simulations, the total induced vorticity was held constant while varying the distribution of counter-rotating vorticity around the nozzle lip training edge. In the experiment, the number of tabs applied was varied while holding the total projected area constant. Evaluations were also conducted on initial vortex strength. The results of this work show that the initial growth rate of the jet shear layer is increasingly enhanced as more tabs are added, but that the lowest tab count results in the largest entrained mass flow. The numerical simulations confirm these results.

  5. Variable geometry for supersonic mixed-compression inlets

    NASA Technical Reports Server (NTRS)

    Sorensen, N. E.; Latham, E. A.; Smeltzer, D. B.

    1974-01-01

    Study of two-dimensional and axisymmetric supersonic mixed-compression inlet systems has shown that the geometry of both systems can be varied to provide adequate transonic airflow to satisfy the airflow demand of most jet engines. Collapsing geometry systems for both types of inlet systems provide a generous amount of transonic airflow for any design Mach number inlet system. However, the mechanical practicality of collapsing centerbodies for axisymmetric inlet systems is doubtful. Therefore, translating centerbody axisymmetric inlets with auxiliary airflow systems to augment the transonic airflow capability are an attractive alternative. Estimates show that the capture mass-flow ratio at Mach number 1.0 can be increased approximately 0.20 for a very short axisymmetric inlet system designed for Mach number 2.37. With this increase in mass-flow ratio, even variable-cycle engine transonic airflow demand can be matched without oversizing the inlet at the design Mach number.

  6. Investigation of chemically reacting and radiating supersonic internal flows

    NASA Technical Reports Server (NTRS)

    Mani, M.; Tiwari, S. N.

    1986-01-01

    The two-dimensional spatially elliptic Navier-Stokes equations are used to investigate the chemically reacting and radiating supersonic flow of the hydrogen-air system between two parallel plates and in a channel with a ten degree compression-expansion ramp at the lower boundary. The explicit unsplit finite-difference technique of MacCormack is used to advance the governing equations in time until convergence is achieved. The chemistry source term in the species equation is treated implicitly to alleviate the stiffness associated with fast reactions. The tangent slab approximation is employed in the radiative flux formation. Both pseudo-gray and nongray models are used to represent the absorption characteristics of the participating species. Results obtained for specific conditions indicate that the radiative interaction can have a significant influence on the flow field.

  7. A shock wave approach to the noise of supersonic propellers

    NASA Astrophysics Data System (ADS)

    Dittmar, J. H.; Rice, E. J.

    1981-12-01

    To model propeller noise expected for a turboprop aircraft, the pressure ratio across the shock at the propeller tip was calculated and compared with noise data from three propellers. At helical tip Mach numbers over 1.0, using only the tip shock wave, the model gave a fairly good prediction of the noise for a bladed propeller and for a propeller swept for aerodynamic purposes. However for another propeller, which was highly swept and designed to have noise cancellations from the inboard propeller sections, the shock strength from the tip over predicted the noise. In general the good agreement indicates that shock theory is a viable method for predicting the noise from these supersonic propellers but that the shock strengths from all of the blade sections need to be properly included.

  8. Effect of Swirl on Turbulent Structures in Supersonic Jets

    NASA Technical Reports Server (NTRS)

    Rao, Ram Mohan; Lundgren, Thomas S.

    1998-01-01

    Direct Numerical Simulation (DNS) is used to study the mechanism of generation and evolution of turbulence structures in a temporally evolving supersonic swirling round jet and also to examine the resulting acoustic radiations. Fourier spectral expansions are used in the streamwise and azimuthal directions and a 1-D b-spline Galerkin representation is used in the radial direction. Spectral-like accuracy is achieved using this numerical scheme. Direct numerical simulations, using the b-spline spectral method, are carried out starting from mean flow initial conditions which are perturbed by the most unstable linear stability eigenfunctions. It is observed that the initial helical instability waves evolve into helical vortices which eventually breakdown into smaller scales of turbulence. 'Rib' structures similar to those seen in incompressible mixing layer flow of Rogers and Moserl are observed. The jet core breakdown stage exhibits increased acoustic radiations.

  9. Effect of Swirl on Turbulent Structures in Supersonic Jets

    NASA Technical Reports Server (NTRS)

    Rao, Ram Mohan; Lundgren, Thomas S.

    1998-01-01

    Direct numerical simulation (DNS) is used to study the mechanism of generation and evolution of turbulence structures in a temporally evolving supersonic swirling round jet and also to examine the resulting acoustic radiations. Fourier spectral expansions are used in the streamwise and azimuthal directions and a 1-D b-spline Galerkin representation is used in the radial direction. Spectral-like accuracy is achieved using this numerical scheme. Direct numerical simulations, using the b-spline spectral method, are carried out starting from mean flow initial conditions which are perturbed by the most unstable linear stability eigenfunctions. It is observed that the initial.helical instability waves evolve into helical vortices which eventually breakdown into smaller scales of turbulence. 'Rib' structures similar to those seen in incompressible mixing layer flow of Rogers and Moser are observed. The jet core breakdown stage exhibits increased acoustic radiations.

  10. A finite element formulation for supersonic flows around complex configurations

    NASA Technical Reports Server (NTRS)

    Morino, L.

    1974-01-01

    The problem of small perturbation potential supersonic flow around complex configurations is considered. This problem requires the solution of an integral equation relating the values of the potential on the surface of the body to the values of the normal derivative, which is known from the small perturbation boundary conditions. The surface of the body is divided into small (hyperboloidal quadrilateral) surface elements which are described in terms of the Cartesian components of the four corner points. The values of the potential (and its normal derivative) within each element are assumed to be constant and equal to its value at the centroid of the element. This yields a set of linear algebraic equations whose coefficients are given by source and doublet integrals over the surface elements. Closed form evaluations of the integrals are presented.

  11. Dynamic pressure loads associated with twin supersonic plume resonance

    NASA Technical Reports Server (NTRS)

    Seiner, J. M.; Manning, J. C.; Ponton, M. K.

    1986-01-01

    The phenomenon of twin supersonic plume resonance is defined and studied as it pertains to high level dynamic loads in the inter-nozzle region of aircraft like the F-15 and B1-A. Using a 1/40th scale model twin jet nacelle with powered choked nozzles, it is found that intense internozzle dynamic pressures are associated with the synchrophased coupling of each plume's jet flapping mode. This condition is found most prevalent when each plume's jet flapping mode has constituent elements composed of the B-type helical instability. Suppression of these fatigue bearing loads was accomplished by simple geometric modifications to only one plume's nozzle. These modifications disrupt the natural selection of the B-type mode and thereby decouple the plumes.

  12. Fluorescence of dissociating fragments from supersonic jet-electron collisions

    NASA Astrophysics Data System (ADS)

    Blake, Thomas A.; Smilgys, Russell V.; Lobue, James M.; Schiffman, Aram P.; Novick, Stewart E.

    1985-05-01

    Supersonically cooled jets of nitrogen, methane, ethane, cyclopropane, and azomethane are crossed with collimated streams of electrons. The CH (B 2Σ - → X 2Π) spectra resulting from the electron-induced dissociation of CH 4, C 2H 6, and CH 2) 3 can be fit with rotation temperatures between 4000 and 6000 K for an electron energy of 100 eV. Flourescence spectra of N 2+ (B 2Σ w+ → X 2Π) from the dissociative ionization of azomethane yield a rotational temperature of =8×10 3 K; from ionization of molecular nitrogen the rotational temperature of B 2Σ w+ N 2+ is 45 K. Mechanisms for these various processes are discussed.

  13. Integrated Flight-propulsion Control Concepts for Supersonic Transport Airplanes

    NASA Technical Reports Server (NTRS)

    Burcham, Frank W., Jr.; Gilyard, Glenn B.; Gelhausen, Paul A.

    1990-01-01

    Integration of propulsion and flight control systems will provide significant performance improvements for supersonic transport airplanes. Increased engine thrust and reduced fuel consumption can be obtained by controlling engine stall margin as a function of flight and engine operating conditions. Improved inlet pressure recovery and decreased inlet drag can result from inlet control system integration. Using propulsion system forces and moments to augment the flight control system and airplane stability can reduce the flight control surface and tail size, weight, and drag. Special control modes may also be desirable for minimizing community noise and for emergency procedures. The overall impact of integrated controls on the takeoff gross weight for a generic high speed civil transport is presented.

  14. Acoustic results of supersonic tip speed fan blade modification

    NASA Technical Reports Server (NTRS)

    Jutras, R. R.; Kazin, S. B.

    1974-01-01

    A supersonic tip speed single stage fan was modified with the intent of reducing multiple pure tone (MPT) or buzz saw noise. There were three modifications to the blades from the original design. The modifications to the blade resulted in an increase in cascade throat area causing the shock to start at a lower corrected fan speed. The acoustic results without acoustically absorbing liners showed substantial reduction in multiple pure tone levels. However, an increase in the blade passing frequency noise at takeoff fan speed accompanied the MPT reduction. The net result however, was a reduction in the maximum 1000-foot (304.8 m) altitude level flyover PNL. For the case with acoustic treatment in the inlet outer wall, the takeoff noise increased relative to an acoustically treated baseline. This was largely due to the increased blade passing frequency noise which was not effectively reduced by the liner.

  15. Guiding supersonic projectiles using optically generated air density channels

    NASA Astrophysics Data System (ADS)

    Johnson, Luke A.; Sprangle, Phillip

    2015-09-01

    We investigate the feasibility of using optically generated channels of reduced air density to provide trajectory correction (guiding) for a supersonic projectile. It is shown that the projectile experiences a force perpendicular to its direction of motion as one side of the projectile passes through a channel of reduced air density. A single channel of reduced air density can be generated by the energy deposited from filamentation of an intense laser pulse. We propose changing the laser pulse energy from shot-to-shot to build longer effective channels. Current femtosecond laser systems with multi-millijoule pulses could provide trajectory correction of several meters on 5 km trajectories for sub-kilogram projectiles traveling at Mach 3.

  16. Experimental transonic flutter characteristics of supersonic cruise configurations

    NASA Technical Reports Server (NTRS)

    Durham, Michael H.; Cole, Stanley R.; Cazier, F. W., Jr.; Keller, Donald F.; Parker, Ellen C.; Wilkie, W. Keats

    1990-01-01

    The flutter characteristics of a generic arrow-wing supersonic transport configuration are studied. The wing configuration has a 3 percent biconvex airfoil and a leading-edge sweep of 73 deg out to a cranked tip with a 60 deg leading-edge sweep. The ground vibration tests and flutter test procedure are described. The effects of flutter on engine nacelles, fuel loading, wing-mounted vertical fin, wing angle-of-attack, and wing tip mass and stiffness distributions are analyzed. The data reveal that engine nacelles reduce the transonic flutter dynamic pressure by 25-30 percent; fuel loadings decrease dynamic pressures by 25 percent; 4-6 deg wing angles-of-attack cause steep transonic boundaries; and 5-10 percent changes in flutter dynamic pressures are the result of the wing-mounted vertical fin and wing-tip mass and stiffness distributions.

  17. Status of noise technology for advanced supersonic cruise aircraft

    NASA Technical Reports Server (NTRS)

    Stone, J. R.; Gutierrez, O. A.

    1980-01-01

    Developments in acoustic technology applicable to advanced supersonic cruise aircraft, particularly those which relate to jet noise and its suppression are reviewed. The noise reducing potential of high radius ratio, inverted velocity profile coannular jets is demonstrated by model scale results from a wide range of nozzle geometries, including some simulated flight cases. These results were verified statistically at large scale on a variable cycle engine (VCE) testbed. A preliminary assessment of potential VCE noise sources such as fan and core noise is made, based on the testbed data. Recent advances in the understanding of flight effects are reviewed. The status of component noise prediction methods is assessed on the basis of recent test data, and the remaining problem areas are outlined.

  18. Modeling supersonic combustion using a fully-implicit numerical method

    NASA Technical Reports Server (NTRS)

    Maccormack, Robert W.; Wilson, Gregory J.

    1990-01-01

    A fully-implicit finite-volume algorithm for two-dimensional axisymmetric flows has been coupled to a detailed hydrogen-air reaction mechanism (13 species and 33 reactions) so that supersonic combustion phenomena may be investigated. Numerical computations are compared with ballistic-range shadowgraphs of Lehr (1972) that exhibit two discontinuities caused by a blunt body as it passes through a premixed stoichiometric hydrogen-air mixture. The suitability of the numerical procedure for simulating these double-front flows is shown. The requirements for the physical formulation and the numerical modeling of these flowfields are discussed. Finally, the sensitivity of these external flowfields to changes in certain key reaction rate constants is examined.

  19. Flutter, Postflutter, and Control of a Supersonic Wing Section

    NASA Technical Reports Server (NTRS)

    Marzocca, Piergiovanni; Librescu, Liviu; Silva, Walter A.

    2002-01-01

    A number of issues related to the flutter and postflutter of two-dimensional supersonic lifting surfaces are addressed. Among them there are the 1) investigation of the implications of the nonlinear unsteady aerodynamics and structural nonlinearities on the stable/unstable character of the limit cycle and 2) study of the implications of the incorporation of a control capability on both the flutter boundary and the postflutter behavior. To this end, a powerful methodology based on the Lyapunov first quantity is implemented. Such a treatment of the problem enables one to get a better understanding of the various factors involved in the nonlinear aeroelastic problem, including the stable and unstable limit cycle. In addition, it constitutes a first step toward a more general investigation of nonlinear aeroelastic phenomena of three-dimensional lifting surfaces.

  20. On Supersonic-Inlet Boundary-Layer Bleed Flow

    NASA Technical Reports Server (NTRS)

    Harloff, Gary J.; Smith, Gregory E.

    1995-01-01

    Boundary-layer bleed in supersonic inlets is typically used to avoid separation from adverse shock-wave/boundary-layer interactions and subsequent total pressure losses in the subsonic diffuser and to improve normal shock stability. Methodologies used to determine bleed requirements are reviewed. Empirical sonic flow coefficients are currently used to determine the bleed hole pattern. These coefficients depend on local Mach number, pressure ratio, hole geometry, etc. A new analytical bleed method is presented to compute sonic flow coefficients for holes and narrow slots and predictions are compared with published data to illustrate the accuracy of the model. The model can be used by inlet designers and as a bleed boundary condition for computational fluid dynamic studies.

  1. Ultra-precise particle velocities in pulsed supersonic beams

    SciTech Connect

    Christen, Wolfgang

    2013-07-14

    We describe an improved experimental method for the generation of cold, directed particle bunches, and the highly accurate determination of their velocities in a pulsed supersonic beam, allowing for high-resolution experiments of atoms, molecules, and clusters. It is characterized by a pulsed high pressure jet source with high brilliance and optimum repeatability, a flight distance of few metres that can be varied with a tolerance of setting of 50 {mu}m, and a precision in the mean flight time of particles of better than 10{sup -4}. The technique achieves unmatched accuracies in particle velocities and kinetic energies and also permits the reliable determination of enthalpy changes with very high precision.

  2. Vorticity generation mechanisms in parallel injection schemes for supersonic mixing

    NASA Astrophysics Data System (ADS)

    Donohue, James M.; Haj-Hariri, Hossein; McDaniel, James C., Jr.

    1992-07-01

    A numerical study is reported of the three-dimensional nonreacting supersonic flow field produced by three parallel fuel injection schemes. Such injections are being considered as a means of enhancing fuel mixing in future Scramjet engine designs. The strength of vortex structures produced by ramp injectors and by rocket/jet interactions are quantified and their relative effectiveness in enhancing the fuel mixing process is addressed. An experimental setup for validation of the above numerical results has been constructed and preliminary results are presented. For the flow field parameters chosen, the ramp generated vorticity is found to be considerably larger than that generated by the shock. The unsteady recirculating flow field in the wake of the injector appears also to give rise to important additional vorticity and mixing.

  3. Supersonic separated turbulent boundary - layer over a wavy wall

    NASA Technical Reports Server (NTRS)

    Polak, A.; Werle, M. J.

    1977-01-01

    A prediction method is developed for calculating distributions of surface heating rates, pressure and skin friction over a wavy wall in a two-dimensional supersonic flow. Of particular interest is the flow of thick turbulent boundary layers. The surface geometry and the flow conditions considered are such that there exists a strong interaction between the viscous and inviscid flow. First, using the interacting turbulent boundary layer equations, the problem is formulated in physical coordinates and then a reformulation of the governing equations in terms of Levy-Lees variables is given. Next, a numerical scheme for solving interacting boundary layer equations is adapted. A number of modifications which led to the improvement of the numerical algorithm are discussed. Finally, results are presented for flow over a train of up to six waves at various flow conditions.

  4. Vacuum chamber with a supersonic-flow aerodynamic window

    DOEpatents

    Hanson, C.L.

    1980-10-14

    A supersonic flow aerodynamic window is disclosed whereby a steam ejector situated in a primary chamber at vacuum exhausts superheated steam toward an orifice to a region of higher pressure, creating a barrier to the gas in the region of higher pressure which attempts to enter through the orifice. In a mixing chamber outside and in fluid communication with the primary chamber, superheated steam and gas are combined into a mixture which then enters the primary chamber through the orifice. At the point of impact of the ejector/superheated steam and the incoming gas/superheated steam mixture, a barrier is created to the gas attempting to enter the ejector chamber. This barrier, coupled with suitable vacuum pumping means and cooling means, serves to keep the steam ejector and primary chamber at a negative pressure, even though the primary chamber has an orifice to a region of higher pressure.

  5. Vacuum chamber with a supersonic flow aerodynamic window

    DOEpatents

    Hanson, Clark L.

    1982-01-01

    A supersonic flow aerodynamic window, whereby a steam ejector situated in a primary chamber at vacuum exhausts superheated steam toward an orifice to a region of higher pressure, creating a barrier to the gas in the region of higher pressure which attempts to enter through the orifice. In a mixing chamber outside and in fluid communication with the primary chamber, superheated steam and gas are combined into a mixture which then enters the primary chamber through the orifice. At the point of impact of the ejector/superheated steam and the incoming gas/superheated steam mixture, a barrier is created to the gas attempting to enter the ejector chamber. This barrier, coupled with suitable vacuum pumping means and cooling means, serves to keep the steam ejector and primary chamber at a negative pressure, even though the primary chamber has an orifice to a region of higher pressure.

  6. Fabrication research for supersonic cruise aircraft. [YF-12 skin structures

    NASA Technical Reports Server (NTRS)

    Hoffman, E. L.; Bales, T. T.; Payne, L.

    1979-01-01

    Advanced fabrication and joining processes for titanium and composite materials are being investigated by NASA to develop technology for the Supersonic Cruise Research (SCR) Program. Full-scale structural panels are being designed and fabricated to meet the criteria of an existing integrally stiffened shear panel on the upper wing surface of the NASA YF-12 aircraft. The program consists of laboratory testing and Mach 3 flight service of full-scale structural panels and laboratory testing of representative structural element specimens. Borsic/aluminum honeycomb-core, titanium clad Borsic/aluminum skin-stringer, graphite/PMR-15 polyimide honeycomb-core, and titanium superplastically formed/diffusion bonded panels have been designed, fabricated, and tested. Graphite/LARC-160 polyimide skin-stringer panels have been designed, and fabrication methods are being developed.

  7. Note: A modular and robust continuous supersonic expansion discharge source

    NASA Astrophysics Data System (ADS)

    Crabtree, Kyle N.; Kauffman, Carrie A.; McCall, Benjamin J.

    2010-08-01

    A direct current discharge has been coupled with a continuous supersonic expansion to provide a source of rotationally cold molecular ions for gas phase spectroscopy. Constructed primarily of machinable ceramic and stainless steel, this source design is modular, customizable, and robust. Its performance has been assessed by recording transitions within the ν2 fundamental band of H3+ using cavity ringdown spectroscopy to determine the rotational temperature of ions produced in the free-jet expansion. Temperature and column density were recorded as a function of discharge current as the source was operated over a period of 200 h. Observed temperatures ranged between 50-110 K, and the ion column densities between 8×1010 and 2×1012 cm-2.

  8. Supersonic combustion engine and method of combustion initiation and distribution

    SciTech Connect

    Stickler, D.B.; Ballantyne, A.; Kyuman Jeong.

    1993-06-29

    A supersonic combustion ramjet engine having a combustor with a combustion zone intended to channel gas flow at relatively high speed therethrough, the engine comprising: means for substantially continuously supplying fuel into the combustion zone; and means for substantially instantaneously igniting a volume of fuel in the combustion zone for providing a spatially controlled combustion distribution, the igniting means having means for providing a diffuse discharge of energy into the volume, the volume extending across a substantially complete cross-sectional area of the combustion zone, the means for discharging energy being capable of generating free radicals within the volume of reactive fuel in the combustion zone such that fuel in the volume can initiate a controlled relatively rapid combustion of fuel in the combustion zone whereby combustion distribution in relatively high speed gas flows through the combustion zone can be initiated and controlled without dependence upon a flame holder or relatively high local static temperature in the combustion zone.

  9. Survey of supersonic combustion ramjet research at Langley

    NASA Technical Reports Server (NTRS)

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

    1986-01-01

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

  10. Turbulent-flow separation criteria for overexpanded supersonic nozzles

    NASA Technical Reports Server (NTRS)

    Morrisette, E. L.; Goldberg, T. J.

    1978-01-01

    A comprehensive compilation of available turbulent flow separation data for overexpanded supersonic nozzles is presented with a discussion of correlation techniques, and prediction methods. Data are grouped by nozzle types: conical, contoured, and two dimensional wedge. Correlation of conical nozzle separation is found to be independent of nozzle divergence half-angle above the 9 deg, whereas the contoured nozzle data follow a different correlation curve. Zero pressure gradient prediction techniques are shown to predict adequately the higher divergence angle conical separation data, and an empirical equation is given for the contoured nozzle data correlation. Flow conditions for which the correlations are invalid are discussed and bounded. A nozzle boundary layer transition criterion is presented which can be used to show that much of the noncorrelating data in the literature are concerned with nonturbulent separation and which explains the previously reported external flow effects on nozzle separation.

  11. High-speed epitaxy using supersonic molecular jets

    SciTech Connect

    Eres, D.

    1990-01-01

    This paper discusses the use of supersonic jets of gaseous source molecules in thin films growth. Molecular jets in free form with no skimmers or collimators in the nozzle-substrate path were used in the investigation of basic film growth processes and in practical film growth applications. The Ge growth rates were found to depend linearly on the digermane jet intensity. Furthermore, the film thickness distributions showed excellent agreement with the distribution of digermane molecules in the jet. High epitaxial Ge growth rates were achieved on GaAs (100) substrates by utilizing high-intensity pulsed jets. The practical advantages and limitations of this film growth technique are evaluated, based on the results of microstructural and electrical measurements of heteroepitaxial Ge films on GaAs (100) substrates. 8 refs., 4 figs.

  12. Vortex development on slender missiles at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Allen, J. M.; Dillenius, M. F. E.

    1979-01-01

    A theoretical and experimental effort has been made to develop a vortex-prediction capability on circular and noncircular missiles at supersonic speeds. Predicted vortex patterns are computed by two linear-theory computer codes. One calculates the strengths and initial locations of the vortices, and the other calculates their trajectories. A short color motion picture has been produced from the calculations to illustrate the predicted vortex patterns on a typical missile. Experimental vapor-screen photographs are presented to show the longitudinal development of the vortices on a fin-control missile. Comparisons are made between these data and the predicted vortices to assess the accuracy of the theory. The theory appears to be fairly accurate in predicting the number, locations, and relative strengths of individual vortices which develop over the missile, but cannot predict vortex sheets or diffuse vorticity whenever they occur.

  13. Acoustic source analysis of supersonic jets from complex nozzles

    NASA Astrophysics Data System (ADS)

    Nichols, Joseph W.; Kreitzman, Jordan

    2015-11-01

    We compute acoustic source terms corresponding to Goldstein's generalized acoustic analogy from an unstructured high-fidelity large eddy simulation of a supersonic jet issuing from a rectangular nozzle with chevrons. The simulation data are validated against experimental measurements of mean and turbulence flow statistics as well as far-field noise. We evaluate fourth-order correlations from the simulation data to assess assumptions of quasi-normality and statistical axisymmetry that underpin reduced-order acoustic source models originally developed for round jets. A spatial analysis of these correlations in relation to the complex geometry of the nozzle reveals locations where the validity of these assumptions begins to break down. Using two point two-time correlations of the simulation data, we also directly evaluate and compare the accuracy of four different acoustic source models, including the Gaussian, moving-frame, fixed-frame, and modified distance models. Computational resources were provided by the Argonne Leadership Computing Facility.

  14. Nonaxisymmetric viscous lower branch modes in axisymmetric supersonic flows

    NASA Technical Reports Server (NTRS)

    Duck, Peter W.; Hall, Philip

    1988-01-01

    In a previous paper, the weakly nonlinear interaction of a pair of axisymmetric lower branch Tollmien-Schlichting instabilities in cylindrical supersonic flows was considered. Here the possibility that nonaxisymmetric modes might also exist is investigated. In fact, it is found that such modes do exist and, on the basis of linear theory, it appears that these modes are the most important. The nonaxisymmetric modes are found to exist for flows around cylinders with nondimensional radius alpha less than some critical value alpha sub c. This critical value alpha sub c is found to increase monotonically with the azimuthal wavenumber nu of the disturbance and it is found that unstable modes always occur in pairs. It is also shown that, in general, instability in the form of lower branch Tollmien-Schlichting waves will occur first for nonaxisymmetric modes and that in the unstable regime the largest growth rates correspond to the latter modes.

  15. Reynolds Number Effects on a Supersonic Transport at Transonic Conditions

    NASA Technical Reports Server (NTRS)

    Wahls, R. N.; Owens, L. R.; Rivers, S. M. B.

    2001-01-01

    A High Speed Civil Transport configuration was tested in the National Transonic Facility at the NASA Langley Research Center as part of NASA's High Speed Research Program. The primary purposes of the tests were to assess Reynolds number scale effects and the high Reynolds number aerodynamic characteristics of a realistic, second generation supersonic transport while providing data for the assessment of computational methods. The tests included longitudinal and lateral/directional studies at low speed high-lift and transonic conditions across a range of Reynolds numbers from that available in conventional wind tunnels to near flight conditions. Results are presented which focus on both the Reynolds number and static aeroelastic sensitivities of longitudinal characteristics at Mach 0.90 for a configuration without an empennage.

  16. Supersonic civil airplane study and design: Performance and sonic boom

    NASA Technical Reports Server (NTRS)

    Cheung, Samson

    1995-01-01

    Since aircraft configuration plays an important role in aerodynamic performance and sonic boom shape, the configuration of the next generation supersonic civil transport has to be tailored to meet high aerodynamic performance and low sonic boom requirements. Computational fluid dynamics (CFD) can be used to design airplanes to meet these dual objectives. The work and results in this report are used to support NASA's High Speed Research Program (HSRP). CFD tools and techniques have been developed for general usages of sonic boom propagation study and aerodynamic design. Parallel to the research effort on sonic boom extrapolation, CFD flow solvers have been coupled with a numeric optimization tool to form a design package for aircraft configuration. This CFD optimization package has been applied to configuration design on a low-boom concept and an oblique all-wing concept. A nonlinear unconstrained optimizer for Parallel Virtual Machine has been developed for aerodynamic design and study.

  17. Overexpanded viscous supersonic jet interacting with a unilateral barrier

    NASA Astrophysics Data System (ADS)

    Dobrynin, B. M.; Maslennikov, V. G.; Sakharov, V. A.; Serova, E. V.

    1986-07-01

    The interaction of a two-dimensional supersonic jet with a unilateral barrier parallel to the flow symmetry plane was studied to account for effects due to gas viscosity and backgound-gas ejection from the region into which the jet expands. In the present experiments, the incident shock wave was reflected at the end of a shock tube equipped with a nozzle. The jet emerged into a pressure chamber 6 cu m in volume and the environmental pressure ratio of the flow in the quasi-stationary phase remained constant. The light source was an OGM-20 laser operating in the giant-pulse mode. Due to background-gas ejection, the gas density in the vicinity of the barrier is much less than on the unconfined side of the jet. The resulting flow is characterized by two distinct environmental pressure ratios: the flow is underexpanded near the barrier, while on the other side it is overexpanded.

  18. Nonlinear flutter analysis of stiffened composite panels in supersonic flow

    NASA Astrophysics Data System (ADS)

    Yuan, Kaihua; Qiu, Zhiping

    2010-02-01

    The flutter instability of stiffened composite panels subjected to aerodynamic forces in the supersonic flow is investigated. Based on Hamilton’s principle, the aeroelastic model of the composite panel is established by using the von Karman large deflection plate theory, piston theory aerodynamics and the quasi-steady thermal stress theory. Then, using the finite element method along with Bogner-Fox-Schmit elements and three-dimensional beam elements, the nonlinear equations of motion are derived. The effect of stiffening scheme on the flutter critical dynamic pressure is demonstrated through the numerical example, and the nonlinear flutter characteristics of stiffened composite panels are also analyzed in the time domain. This will lay the foundation for design of panel structures employed in aerospace vehicles.

  19. SUPERSONIC AND HYPERSONIC INTERFERENCE FLOW FIELDS AND HEATING

    NASA Technical Reports Server (NTRS)

    Morris, D. J.

    1994-01-01

    Small areas of high heat transfer and pressure can occur on a vehicle surface due to the influence of an impinging shock on the local flow. A method was needed to determine peak pressure and heating of these areas. This package is a system of computer programs designed to calculate two-dimensional shock interference patterns for six types of interference flows. Results also include properties of the inviscid flow field and the inviscid-viscous interaction at the surface along with peak pressure and peak heating at the impingement point. The six types of interference flow patterns considered are: 1) Type I interference patterns, occurring when two weak shocks of opposite families, BS (bow shock) and IS (impingment shock), intersect when the flow upstream of the impingement point is supersonic, or in the case of a blunt body, takes place well below the sonic point. 2) Type II interference pattern occurs when two shocks of opposite families (bow shock and impinging shock) intersect. Both shocks are weak as in type I, but are of such strength that in order to turn the flow, a Mach reflection must exist in the center of the flow field with an embedded subsonic region occurring between the intersection points (A & B) and the accompanying shear layers. Type II interference occurs on a blunt body when the impinging shock intersects the bow shock near the sonic point. 3) Type III shock interference pattern occurs when a weak impinging shock intersects a strong detached bow shock. On a blunt body the shock intersection occurs near or above the lower sonic point. 4) Type IV interference can occur when the impinging shock intersects a strong bow shock ahead of a subsonic flow region. On a blunt body this shock intersection is located between the lower sonic point and just above the body axis. The impinging shock causes a displacement of the bow shock and the formation of a supersonic jet that is embedded in the subsonic region. A jet bow shock is produced when the jet impinges

  20. Wakes formed by dust grains in supersonically flowing plasmas

    SciTech Connect

    Willis, C. T. N.; Coppins, M.; Bacharis, M.; Allen, J. E.

    2011-10-15

    Interesting wake effects are found in simulations of dust grains in supersonically flowing plasma. A Mach cone is formed at an angle to the flow determined by the ratio of flow to Bohm speed. The latter is well approximated by [k(T{sub e}+{gamma}T{sub i})/m{sub i}]{sup 1/2} with {gamma}=3. For ion temperatures significantly lower than the electron temperature, a second (inner) cone forms due to flow convergence. An ''ion vacuum'' and stagnation point occur downstream. These latter effects cannot be described by conventional (cold-ion) gas dynamics. Critically, none of the cones observed are shocks but are more akin to weak discontinuities.

  1. Skin Friction and Transition Location Measurement on Supersonic Transport Models

    NASA Technical Reports Server (NTRS)

    Kennelly, Robert A., Jr.; Goodsell, Aga M.; Olsen, Lawrence E. (Technical Monitor)

    2000-01-01

    Flow visualization techniques were used to obtain both qualitative and quantitative skin friction and transition location data in wind tunnel tests performed on two supersonic transport models at Mach 2.40. Oil-film interferometry was useful for verifying boundary layer transition, but careful monitoring of model surface temperatures and systematic examination of the effects of tunnel start-up and shutdown transients will be required to achieve high levels of accuracy for skin friction measurements. A more common technique, use of a subliming solid to reveal transition location, was employed to correct drag measurements to a standard condition of all-turbulent flow on the wing. These corrected data were then analyzed to determine the additional correction required to account for the effect of the boundary layer trip devices.

  2. A k-omega-multivariate beta PDF for supersonic combustion

    NASA Technical Reports Server (NTRS)

    Alexopoulos, G. A.; Baurle, R. A.; Hassan, H. A.

    1992-01-01

    In an attempt to study the interaction between combustion and turbulence in supersonic flows, an assumed PDF has been employed. This makes it possible to calculate the time average of the chemical source terms that appear in the species conservation equations. In order to determine the averages indicated in an equation, two transport equations, one for the temperature (enthalpy) variance and one for Q, are required. Model equations are formulated for such quantities. The turbulent time scale controls the evolution. An algebraic model similar to that used by Eklund et al was used in an attempt to predict the recent measurements of Cheng et al. Predictions were satisfactory before ignition but were less satisfactory after ignition. One of the reasons for this behavior is the inadequacy of the algebraic turbulence model employed. Because of this, the objective of this work is to develop a k-omega model to remedy the situation.

  3. CARS Temperature Measurements in a Combustion-Heated Supersonic Jet

    NASA Technical Reports Server (NTRS)

    Tedder, S. A.; Danehy, P. M.; Magnotti, G.; Cutler, A. D.

    2009-01-01

    Measurements were made in a combustion-heated supersonic axi-symmetric free jet from a nozzle with a diameter of 6.35 cm using dual-pump Coherent Anti-Stokes Raman Spectroscopy (CARS). The resulting mean and standard deviation temperature maps are presented. The temperature results show that the gas temperature on the centerline remains constant for approximately 5 nozzle diameters. As the heated gas mixes with the ambient air further downstream the mean temperature decreases. The standard deviation map shows evidence of the increase of turbulence in the shear layer as the jet proceeds downstream and mixes with the ambient air. The challenges of collecting data in a harsh environment are discussed along with influences to the data. The yield of the data collected is presented and possible improvements to the yield is presented are discussed.

  4. Computation of supersonic turbulent flow past a spinning cone

    NASA Technical Reports Server (NTRS)

    Agarwal, R. K.

    1982-01-01

    Computational results are presented for supersonic laminar and turbulent flow past a pointed cone at angle of attack obtained with a parabolic Navier-Stokes marching code. The code takes into account the asymmetries in the flowfield resulting from spinning motion and computes the asymmetric shock shape, crossflow and streamwise shear, heat transfer, crossflow separation, and vortex structure. The Magnus force and moments are also computed. Comparisons are made with other analyses based on boundary-layer equations. For certain laminar flow conditions, an anomaly is discovered in the displacement thickness contribution to the Magnus force when compared with boundary-layer results. For turbulent flow, at small angles of attack, good agreement is obtained with the experimental data and other theoretical results.

  5. Design and analysis of a supersonic penetration/maneuvering fighter

    NASA Technical Reports Server (NTRS)

    Child, R. D.

    1975-01-01

    The design of three candidate air combat fighters which would cruise effectively at freestream Mach numbers of 1.6, 2.0, and 2.5 while maintaining good transonic maneuvering capability, is considered. These fighters were designed to deliver aerodynamically controlled dogfight missiles at the design Mach numbers. Studies performed by Rockwell International in May 1974 and guidance from NASA determined the shape and size of these missiles. The principle objective of this study is the aerodynamic design of the vehicles; however, configurations are sized to have realistic structures, mass properties, and propulsion systems. The results of this study show that air combat fighters in the 15,000 to 23,000 pound class would cruise supersonically on dry power and still maintain good transonic maneuvering performance.

  6. Performance of Several High Order Numerical Methods for Supersonic Combustion

    NASA Technical Reports Server (NTRS)

    Sjoegreen, Bjoern; Yee, H. C.; Don, Wai Sun; Mansour, Nagi N. (Technical Monitor)

    2001-01-01

    The performance of two recently developed numerical methods by Yee et al. and Sjoegreen and Yee using postprocessing nonlinear filters is examined for a 2-D multiscale viscous supersonic react-live flow. These nonlinear filters can improve nonlinear instabilities and at the same time can capture shock/shear waves accurately. They do not, belong to the class of TVD, ENO or WENO schemes. Nevertheless, they combine stable behavior at discontinuities and detonation without smearing the smooth parts of the flow field. For the present study, we employ a fourth-order Runge-Kutta in time and a sixth-order non-dissipative spatial base scheme for the convection and viscous terms. We denote the resulting nonlinear filter schemes ACM466-RK4 and WAV66-RK4.

  7. Particle Acceleration via Reconnection Processes in the Supersonic Solar Wind

    NASA Astrophysics Data System (ADS)

    Zank, G. P.; le Roux, J. A.; Webb, G. M.; Dosch, A.; Khabarova, O.

    2014-12-01

    An emerging paradigm for the dissipation of magnetic turbulence in the supersonic solar wind is via localized small-scale reconnection processes, essentially between quasi-2D interacting magnetic islands. Charged particles trapped in merging magnetic islands can be accelerated by the electric field generated by magnetic island merging and the contraction of magnetic islands. We derive a gyrophase-averaged transport equation for particles experiencing pitch-angle scattering and energization in a super-Alfvénic flowing plasma experiencing multiple small-scale reconnection events. A simpler advection-diffusion transport equation for a nearly isotropic particle distribution is derived. The dominant charged particle energization processes are (1) the electric field induced by quasi-2D magnetic island merging and (2) magnetic island contraction. The magnetic island topology ensures that charged particles are trapped in regions where they experience repeated interactions with the induced electric field or contracting magnetic islands. Steady-state solutions of the isotropic transport equation with only the induced electric field and a fixed source yield a power-law spectrum for the accelerated particles with index α = -(3 + MA )/2, where MA is the Alfvén Mach number. Considering only magnetic island contraction yields power-law-like solutions with index -3(1 + τ c /(8τdiff)), where τ c /τdiff is the ratio of timescales between magnetic island contraction and charged particle diffusion. The general solution is a power-law-like solution with an index that depends on the Alfvén Mach number and the timescale ratio τdiff/τ c . Observed power-law distributions of energetic particles observed in the quiet supersonic solar wind at 1 AU may be a consequence of particle acceleration associated with dissipative small-scale reconnection processes in a turbulent plasma, including the widely reported c -5 (c particle speed) spectra observed by Fisk & Gloeckler and Mewaldt et

  8. Plasma-enhanced mixing and flameholding in supersonic flow

    PubMed Central

    Firsov, Alexander; Savelkin, Konstantin V.; Yarantsev, Dmitry A.; Leonov, Sergey B.

    2015-01-01

    The results of experimental study of plasma-based mixing, ignition and flameholding in a supersonic model combustor are presented in the paper. The model combustor has a length of 600 mm and cross section of 72 mm width and 60 mm height. The fuel is directly injected into supersonic airflow (Mach number M=2, static pressure Pst=160–250 Torr) through wall orifices. Two series of tests are focused on flameholding and mixing correspondingly. In the first series, the near-surface quasi-DC electrical discharge is generated by flush-mounted electrodes at electrical power deposition of Wpl=3–24 kW. The scope includes parametric study of ignition and flame front dynamics, and comparison of three schemes of plasma generation: the first and the second layouts examine the location of plasma generators upstream and downstream from the fuel injectors. The third pattern follows a novel approach of combined mixing/ignition technique, where the electrical discharge distributes along the fuel jet. The last pattern demonstrates a significant advantage in terms of flameholding limit. In the second series of tests, a long discharge of submicrosecond duration is generated across the flow and along the fuel jet. A gasdynamic instability of thermal cavity developed after a deposition of high-power density in a thin plasma filament promotes the air–fuel mixing. The technique studied in this work has weighty potential for high-speed combustion applications, including cold start/restart of scramjet engines and support of transition regime in dual-mode scramjet and at off-design operation. PMID:26170434

  9. Transverse jet injection into a supersonic turbulent cross-flow

    NASA Astrophysics Data System (ADS)

    Rana, Z. A.; Thornber, B.; Drikakis, D.

    2011-04-01

    Jet injection into a supersonic cross-flow is a challenging fluid dynamics problem in the field of aerospace engineering which has applications as part of a rocket thrust vector control system for noise control in cavities and fuel injection in scramjet combustion chambers. Several experimental and theoretical/numerical works have been conducted to explore this flow; however, there is a dearth of literature detailing the instantaneous flow which is vital to improve the efficiency of the mixing of fluids. In this paper, a sonic jet in a Mach 1.6 free-stream is studied using a finite volume Godunov type implicit large eddy simulations technique, which employs fifth-order accurate MUSCL (Monotone Upstream-centered Schemes for Conservation Laws) scheme with modified variable extrapolation and a three-stage second-order strong-stability-preserving Runge-Kutta scheme for temporal advancement. A digital filter based turbulent inflow data generation method is implemented in order to capture the physics of the supersonic turbulent boundary layer. This paper details the averaged and instantaneous flow features including vortex structures downstream of the jet injection, along with the jet penetration, jet mixing, pressure distributions, turbulent kinetic energy, and Reynolds stresses in the downstream flow. It demonstrates that Kelvin-Helmholtz type instabilities in the upper jet shear layer are primarily responsible for mixing of the two fluids. The results are compared to experimental data and recently performed classical large eddy simulations (LES) with the same initial conditions in order to demonstrate the accuracy of the numerical methods and utility of the inflow generation method. Results here show equivalent accuracy for 1/45th of the computational resources used in the classical LES study.

  10. Brightness and virtual source size of a supersonic deuterium beam

    NASA Astrophysics Data System (ADS)

    Reisinger, Thomas; Greve, Martin M.; Eder, Sabrina D.; Bracco, Gianangelo; Holst, Bodil

    2012-10-01

    Supersonic beams have numerous applications in research fields ranging from spectroscopy with nanodroplets to surface science and matter-wave microscopy. Thus, measurement and prediction of their properties is of considerable interest. In this paper we present measurements of the virtual-source size and its brightness, as well as the terminal speed and terminal speed ratio of a supersonic deuterium (D2) beam. The speed distribution data were measured with time-of-flight experiments and Fresnel zone-plate imaging was used to measure virtual source size. The point-spread function of the zone plate was simulated based on the measured wavelength distribution and used to extract the width of the virtual source and its brightness from the focus measurement. The experiments were carried out with a 10-μm-diameter nozzle and a source temperature of T0=310 K in the pressure range p0=3-171 bars and for T0=106 K in the pressure range p0=3-131 bars. We found that using deuterium as opposed to helium results in a virtual source that is about a factor 2 brighter under similar stagnation conditions. A comparison between the measured data and the predictions from a theoretical model based on the Boltzmann equation, which explicitly include the coupling between translational and rotational degrees of freedom as well as the real-gas properties of D2, resulted in good correspondence for the two different interaction potentials we tried. A careful comparison with the experimental results shows that the potential by Buck [J. Chem. Phys.JCPSA60021-960610.1063/1.445336 78, 4439 (1983)] is moderately better than the Lennard-Jones potential at describing the expansion dynamics.

  11. Supersonic combustion ramjet propulsion experiments in a shock tunnel

    NASA Technical Reports Server (NTRS)

    Paull, A.; Stalker, R. J.; Mee, D. J.

    1995-01-01

    Measurements have been made of the propulsive effect of supersonic combustion ramjets incorporated into a simple axisymmetric model in a free piston shock tunnel. The nominal Mach number was 6, and the stagnation enthalpy varied from 2.8 MJ kg(exp -1) to 8.5 MJ kg(exp -1). A mixture of 13 percent silane and 87 percent hydrogen was used as fuel, and experiments were conducted at equivalence ratios up to approximately 0.8. The measurements involved the axial force on the model, and were made using a stress wave force balance, which is a recently developed technique for measuring forces in shock tunnels. A net thrust was experienced up to a stagnation enthalpy of 3.7 MJ kg(exp -1), but as the stagnation enthalpy increased, an increasing net drag was recorded. pitot and static pressure measurements showed that the combustion was supersonic. The results were found to compare satisfactorily with predictions based on established theoretical models, used with some simplifying approximations. The rapid reduction of net thrust with increasing stagnation enthalpy was seen to arise from increasing precombustion temperature, showing the need to control this variable if thrust performance was to be maintained over a range of stagnation enthalpies. Both the inviscid and viscous drag were seen to be relatively insensitive to stagnation enthalpy, with the combustion chambers making a particularly significant contribution to drag. The maximum fuel specific impulse achieved in the experiments was only 175 sec., but the theory indicates that there is considerable scope for improvement on this through aerodynamic design.

  12. Rocket Sled Propelled Testing of a Supersonic Inflatable Aerodynamic Decelerator

    NASA Technical Reports Server (NTRS)

    Meacham, Michael B.; Kennett, Andrew; Townsend, Derik J.; Marti, Benjamin

    2013-01-01

    Decelerators (IADs) have traditionally been tested in wind tunnels. As the limitations of these test facilities are reached, other avenues must be pursued. The IAD being tested is a Supersonic IAD (SIAD), which attaches just aft of the heatshield around the perimeter of an entry body. This 'attached torus' SIAD is meant to improve the accuracy of landing for robotic class missions to Mars and allow for potentially increased payloads. The SIAD Design Verification (SDV) test aims to qualify the SIAD by applying a targeted aerodynamic load to the vehicle. While many test architectures were researched, a rocket sled track was ultimately chosen to be the most cost effective way to achieve the desired dynamic pressures. The Supersonic Naval Ordnance Research Track (SNORT) at the Naval Air Warfare Center Weapons Division (NAWCWD) China Lake is a four mile test track, traditionally used for warhead and ejection seat testing. Prior to SDV, inflatable drag bodies have been tested on this particular track. Teams at Jet Propulsion Laboratory (JPL) and NAWCWD collaborate together to design and fabricate one of the largest sleds ever built. The SDV sled is comprised of three individual sleds: a Pusher Sled which holds the solid booster rockets, an Item Sled which supports the test vehicle, and a Camera Sled that is pushed in front for in-situ footage and measurements. The JPL-designed Test Vehicle has a full-scale heatshield shape and contains all instrumentation and inflation systems necessary to inflate and test a SIAD. The first campaign that is run at SNORT tested all hardware and instrumentation before the SIAD was ready to be tested. For each of the three tests in this campaign, the number of rockets and top speed was increased and the data analyzed to ensure the hardware is safe at the necessary accelerations and aerodynamic loads.

  13. Plasma-enhanced mixing and flameholding in supersonic flow.

    PubMed

    Firsov, Alexander; Savelkin, Konstantin V; Yarantsev, Dmitry A; Leonov, Sergey B

    2015-08-13

    The results of experimental study of plasma-based mixing, ignition and flameholding in a supersonic model combustor are presented in the paper. The model combustor has a length of 600 mm and cross section of 72 mm width and 60 mm height. The fuel is directly injected into supersonic airflow (Mach number M=2, static pressure P(st)=160-250 Torr) through wall orifices. Two series of tests are focused on flameholding and mixing correspondingly. In the first series, the near-surface quasi-DC electrical discharge is generated by flush-mounted electrodes at electrical power deposition of W(pl)=3-24 kW. The scope includes parametric study of ignition and flame front dynamics, and comparison of three schemes of plasma generation: the first and the second layouts examine the location of plasma generators upstream and downstream from the fuel injectors. The third pattern follows a novel approach of combined mixing/ignition technique, where the electrical discharge distributes along the fuel jet. The last pattern demonstrates a significant advantage in terms of flameholding limit. In the second series of tests, a long discharge of submicrosecond duration is generated across the flow and along the fuel jet. A gasdynamic instability of thermal cavity developed after a deposition of high-power density in a thin plasma filament promotes the air-fuel mixing. The technique studied in this work has weighty potential for high-speed combustion applications, including cold start/restart of scramjet engines and support of transition regime in dual-mode scramjet and at off-design operation.

  14. Expanding the Natural Laminar Flow Boundary for Supersonic Transports

    NASA Technical Reports Server (NTRS)

    Lynde, Michelle N.; Campbell, Richard L.

    2016-01-01

    A computational design and analysis methodology is being developed to design a vehicle that can support significant regions of natural laminar flow (NLF) at supersonic flight conditions. The methodology is built in the CDISC design module to be used in this paper with the flow solvers Cart3D and USM3D, and the transition prediction modules BLSTA3D and LASTRAC. The NLF design technique prescribes a target pressure distribution for an existing geometry based on relationships between modal instability wave growth and pressure gradients. The modal instability wave growths (both on- and off-axes crossflow and Tollmien-Schlichting) are balanced to produce a pressure distribution that will have a theoretical maximum NLF region for a given streamwise wing station. An example application is presented showing the methodology on a generic supersonic transport wingbody configuration. The configuration has been successfully redesigned to support significant regions of NLF (approximately 40% of the wing upper surface by surface area). Computational analysis predicts NLF with transition Reynolds numbers (ReT) as high as 36 million with 72 degrees of leading-edge sweep (?LE), significantly expanding the current boundary of ReT - ?LE combinations for NLF. This NLF geometry provides a total drag savings of 4.3 counts compared to the baseline wing-body configuration (approximately 5% of total drag). Off-design evaluations at near-cruise and low-speed, high-lift conditions are discussed, as well as attachment line contamination/transition concerns. This computational NLF design effort is a part of an ongoing cooperative agreement between NASA and JAXA researchers.

  15. Flight Tests of a Supersonic Natural Laminar Flow Airfoil

    NASA Technical Reports Server (NTRS)

    Frederick, Mike; Banks, Dan; Garzon, Andres; Matisheck, Jason

    2014-01-01

    IR thermography was used to characterize the transition front on a S-NLF test article at chord Reynolds numbers in excess of 30 million Changes in transition due to Mach number, Reynolds number, and surface roughness were investigated - Regions of laminar flow in excess of 80% chord at chord Reynolds numbers greater than 14 million IR thermography clearly showed the transition front and other flow features such as shock waves impinging upon the surface A series of parallel oblique shocks, of yet unknown origin, were found to cause premature transition at higher Reynolds numbers. NASA has a current goal to eliminate barriers to the development of practical supersonic transport aircraft Drag reduction through the use of supersonic natural laminar flow (S-NLF) is currently being explored as a means of increasing aerodynamic efficiency - Tradeoffs work best for business jet class at M<2 Conventional high-speed designs minimize inviscid drag at the expense of viscous drag - Existence of strong spanwise pressure gradient leads to crossflow (CF) while adverse chordwise pressure gradients amplifies and Tollmien-Schlichting (TS) instabilities Aerion Corporation has patented a S-NLF wing design (US Patent No. 5322242) - Low sweep to control CF - dp/dx < 0 on both wing surfaces to stabilize TS - Thin wing with sharp leading edge to minimize wave drag increase due to reduction in sweep NASA and Aerion have partnered to study S-NLF since 1999 Series of S-NLF experiments flown on the NASA F-15B research test bed airplane Infrared (IR) thermography used to characterize transition - Non-intrusive, global, good spatial resolution - Captures significant flow features well

  16. Dielectric Barrier Discharge control on an unstarting supersonic flow

    NASA Astrophysics Data System (ADS)

    Im, Seong-Kyun; Do, Hyungrok; Bak, Moonsoo; Cappelli, Mark

    2011-10-01

    The control of unstarting supersonic model inlet flow using Dielectric Barrier Discharge (DBD) is experimentally demonstrated at Mach 4.7 flow condition. Planar Laser Rayleigh Scattering (PLRS) technique is utilized to visualize important flow features, such as boundary layers and shockwaves, at low static temperature (~60 K) and pressure (~1 kPa) freestream condition. The unstart which is initiated by jet injection in three different model inlet flow, a laminar boundary layer, tripped boundary layer without actuation, and tripped boundary layer with actuation, is demonstrated with PLRS technique. The delay of unstart process is observed through the DBD actuation of the tripped boundary layer when a single DBD actuator pair is oriented parallel to the freestream flow, generates spanwise disturbances. However, this actuation on unstart process is limited to a region of actuation. The control of unstarting supersonic model inlet flow using Dielectric Barrier Discharge (DBD) is experimentally demonstrated at Mach 4.7 flow condition. Planar Laser Rayleigh Scattering (PLRS) technique is utilized to visualize important flow features, such as boundary layers and shockwaves, at low static temperature (~60 K) and pressure (~1 kPa) freestream condition. The unstart which is initiated by jet injection in three different model inlet flow, a laminar boundary layer, tripped boundary layer without actuation, and tripped boundary layer with actuation, is demonstrated with PLRS technique. The delay of unstart process is observed through the DBD actuation of the tripped boundary layer when a single DBD actuator pair is oriented parallel to the freestream flow, generates spanwise disturbances. However, this actuation on unstart process is limited to a region of actuation. This material is based upon work supported by the Department of Energy under Award Number(s)DE-FC52-08NA28614.

  17. Doppler Global Velocimetry Measurements for Supersonic Flow Fields

    NASA Technical Reports Server (NTRS)

    Meyers, James F.

    2005-01-01

    The application of Doppler Global Velocimetry (DGV) to high-speed flows has its origins in the original development of the technology by Komine et al (1991). Komine used a small shop-air driven nozzle to generate a 200 m/s flow. This flow velocity was chosen since it produced a fairly large Doppler shift in the scattered light, resulting in a significant transmission loss as the light passed through the Iodine vapor. This proof-of-concept investigation showed that the technology was capable of measuring flow velocity within a measurement plane defined by a single-frequency laser light sheet. The effort also proved that velocity measurements could be made without resolving individual seed particles as required by other techniques such as Fringe- Type Laser Velocimetry and Particle Image Velocimetry. The promise of making planar velocity measurements with the possibility of using 0.1-micron condensation particles for seeding, Dibble et al (1989), resulted in the investigation of supersonic jet flow fields, Elliott et al (1993) and Smith and Northam (1995) - Mach 2.0 and 1.9 respectively. Meyers (1993) conducted a wind tunnel investigation above an inclined flat plate at Mach 2.5 and above a delta wing at Mach 2.8 and 4.6. Although these measurements were crude from an accuracy viewpoint, they did prove that the technology could be used to study supersonic flows using condensation as the scattering medium. Since then several research groups have studied the technology and developed solutions and methodologies to overcome most of the measurement accuracy limitations:

  18. Chirped Pulse Microwave Spectroscopy in Pulsed Uniform Supersonic Flows

    NASA Astrophysics Data System (ADS)

    Abeysekera, Chamara; Oldham, James; Prozument, Kirill; Joalland, Baptiste; Park, Barratt; Field, Robert W.; Sims, Ian; Suits, Arthur; Zack, Lindsay

    2014-06-01

    We present preliminary results describing the development of a new instrument that combines two powerful techniques: Chirped Pulse-Fourier Transform MicroWave (CP-FTMW) spectroscopy and pulsed uniform supersonic flows. It promises a nearly universal detection method that can deliver quantitative isomer, conformer, and vibrational level specific detection, characterization of unstable reaction products and intermediates and perform unique spectroscopic, kinetics and dynamics measurements. We have constructed a new high-power K_a-band, 26-40 GHz, chirped pulse spectrometer with sub-MHz resolution, analogous to the revolutionary CP-FTMW spectroscopic technique developed in the Pate group at University of Virginia. In order to study smaller molecules, the E-band, 60-90 GHz, CP capability was added to our spectrometer. A novel strategy for generating uniform supersonic flow through a Laval nozzle is introduced. High throughput pulsed piezo-valve is used to produce cold (30 K) uniform flow with large volumes of 150 cm^3 and densities of 1014 molecules/cm3 with modest pumping facilities. The uniform flow conditions for a variety of noble gases extend as far as 20 cm from the Laval nozzle and a single compound turbo-molecular pump maintains the operating pressure. Two competing design considerations are critical to the performance of the system: a low temperature flow is needed to maximize the population difference between rotational levels, and high gas number densities are needed to ensure rapid cooling to achieve the uniform flow conditions. At the same time, collision times shorter than the chirp duration will give inaccurate intensities and reduced signal levels due to collisional dephasing of free induction decay. Details of the instrument and future directions and challenges will be discussed.

  19. Development of a quiet supersonic wind tunnel with a cryogenic adaptive nozzle

    NASA Technical Reports Server (NTRS)

    Wolf, Stephen W. D.

    1992-01-01

    Aspects of the design and construction of the Laminar Flow Supersonic Wind Tunnel at the NASA-Ames Fluid Mechanics Laboratory are discussed. The wind tunnel is to be used as part of the NASA High Speed Research Program (HSRP).

  20. Method of characteristics for three-dimensional axially symmetrical supersonic flows.

    NASA Technical Reports Server (NTRS)

    Sauer, R

    1947-01-01

    An approximation method for three-dimensional axially symmetrical supersonic flows is developed; it is based on the characteristics theory (represented partly graphically, partly analytically). Thereafter this method is applied to the construction of rotationally symmetrical nozzles. (author)

  1. The Calculation of Compressible Flows with Local Regions of Supersonic Velocity

    NASA Technical Reports Server (NTRS)

    Goethert, B.; Kawalki, K. H.

    1947-01-01

    This report addresses a method for the approximate calculation of compressible flows about profiles with local regions of supersonic velocity. The flow around a slender profile is treated as an example.

  2. Fundamental Aeronautics Program: Supersonics Project - Channeled Center-Body Inlet Experiment Overview

    NASA Technical Reports Server (NTRS)

    St. John, Clint; Ratnayake, Nalin A.; Frederick, Mike

    2012-01-01

    The presentation describes supersonic flight testing accomplished on a novel mixed-compression axisymmetric inlet utilizing channels for off-design flow matching rather than a translating centerbody concept.

  3. Sting Supported Bell XS-2 in the 9 Inch Supersonic Tunnel

    NASA Technical Reports Server (NTRS)

    1947-01-01

    A sting supported model of the Bell XS-2 was tested in the 9 Inch Supersonic Tunnel. Photograph published in Engineer in Charge: A History of the Langley Aeronautical Laboratory, 1917-1958 by James R. Hansen. Page 316.

  4. Opportunities and limitations in low earth subsonic testing for qualification of extraterrestrial supersonic parachute designs

    NASA Technical Reports Server (NTRS)

    Steltzner, A.; Cruz, J.; Bruno, R.; Mitcheltree, R.

    2003-01-01

    Parachutes for Mars and other planetary missions often need to operate at supersonic speeds in very low density atmospheres. Flight testing of such parachutes at appropriate conditions in the Earth's atmosphere is possible at high altitudes.

  5. Burning of the Supersonic Propane-Air Mixture in the Aerodynamic Channel With the Stagnant Zone

    DTIC Science & Technology

    2007-11-02

    V.Chernikov, V.Shibkov, O.Surkont. Mechanisms of transversal electric discharge sustention in supersonic air and propane-air flows. -American Institute of Aeronautics and Astronautics, AIAA Paper, 2003, No.03-0872, p. 1 -6 .

  6. Digital program analyzes supersonic flow field within bell-shaped rocket nozzles

    NASA Technical Reports Server (NTRS)

    Elliott, J. J.; Stromsta, R. R.

    1970-01-01

    Digital computer program computes and analyzes supersonic flow field in axisymmetric rocket nozzle for specified gas properties, nozzle geometry, and input or starting line. Method of characteristics is used for solution of set of hyperbolic partial differential equations.

  7. DEVELOPMENT OF A SUPERSONIC TRANSPORT AIRCRAFT ENGINE - PHASE II-A.

    DTIC Science & Technology

    JET TRANSPORT PLANES, *SUPERSONIC AIRCRAFT ) (U) TURBOJET ENGINES , PERFORMANCE( ENGINEERING ), TURBOFAN ENGINES , AFTERBURNING, SPECIFICATIONS...COMPRESSORS, GEOMETRY, TURBOJET INLETS, COMBUSTION, TEST EQUIPMENT, TURBINE BLADES , HEAT TRANSFER, AIRFOILS , CASCADE STRUCTURES, EVAPOTRANSPIRATION, PLUG NOZZLES, ANECHOIC CHAMBERS, BEARINGS, SEALS, DESIGN, FATIGUE(MECHANICS)

  8. Experimental Investigation of Laser-sustained Plasma in Supersonic Argon Flow

    SciTech Connect

    Sperber, David; Eckel, Hans-Albert; Moessinger, Peter; Fasoulas, Stefanos

    2011-11-10

    Laser-induced energy deposition is widely discussed as a flow control technique in supersonic transportation. In case of thermal laser-plasma upstream of a blunt body, a substantial adaptation of shock wave geometry and magnitude of wave drag is predicted. Related to the research on laser supported detonation, the paper describes the implementation of laser-sustained plasma in a supersonic Argon jet. The stable plasma state is generated by the intersection of a Q-switched Nd:YAG-laser and a continuous wave CO{sub 2}-laser beams, for ignition and maintenance of the plasma respectively. A miniature supersonic Ludwieg tube test facility generates a supersonic jet at velocities of Mach 2.1. Modifications of the flow and plasma conditions are investigated and characterized by Schlieren flow visualisation, laser energy transmission and plasma radiation measurements. The results include the discussions of the flow field as well as the required laser and gas parameters.

  9. Feasibility and benefits of laminar flow control on supersonic cruise airplanes

    NASA Technical Reports Server (NTRS)

    Powell, A. G.; Agrawal, S.; Lacey, T. R.

    1989-01-01

    An evaluation was made of the applicability and benefits of laminar flow control (LFC) technology to supersonic cruise airplanes. Ancillary objectives were to identify the technical issues critical to supersonic LFC application, and to determine how those issues can be addressed through flight and wind-tunnel testing. Vehicle types studied include a Mach 2.2 supersonic transport configuration, a Mach 4.0 transport, and two Mach 2-class fighter concepts. Laminar flow control methodologies developed for subsonic and transonic wing laminarization were extended and applied. No intractible aerodynamic problems were found in applying LFC to airplanes of the Mach 2 class, even ones of large size. Improvements of 12 to 17 percent in lift-drag ratios were found. Several key technical issues, such as contamination avoidance and excresence criteria were identified. Recommendations are made for their resolution. A need for an inverse supersonic wing design methodology is indicated.

  10. Design aspects of long range supersonic LFC airplanes with highly swept wings. [laminar flow control

    NASA Technical Reports Server (NTRS)

    Pfenninger, W.; Vemuru, C. S.

    1990-01-01

    Studies on supersonic long-range LFC (laminar flow control) aircraft were performed with the aim of maximizing L/D and alleviating sonic boom during supersonic cruise. It is found that configurations with highly swept LFC wings of very high structural aspect ratio, with the sweep increasing toward the wing root and braced externally by wide chord laminarized struts, appear especially promising. In the supersonic cruise design condition the wing upper surface isobars are swept such that the flow in the direction normal to them is transonic with embedded supersonic zones and practically shock-free over most of the span, with M-perpendicular equal to the two-dimensional design values of advanced SC LFC airfoils, e.g., of the X-787 or X-6 type.

  11. Developing Supersonic Impactor and Aerodynamic Lens for Separation and Handling of Nano-Sized Particles

    SciTech Connect

    Goodarz Ahmadi

    2008-06-30

    A computational model for supersonic flows of compressible gases in an aerodynamic lens with several lenses and in a supersonic/hypersonic impactor was developed. Airflow conditions in the aerodynamic lens were analyzed and contour plots for variation of Mach number, velocity magnitude and pressure field in the lens were evaluated. The nano and micro-particle trajectories in the lens and their focusing and transmission efficiencies were evaluated. The computational model was then applied to design of a aerodynamic lens that could generate focus particle beams while operating under atmospheric conditions. The computational model was also applied to airflow condition in the supersonic/hypersonic impactor. Variations of airflow condition and particle trajectories in the impactor were evaluated. The simulation results could provide understanding of the performance of the supersonic and hypersonic impactors that would be helpful for the design of such systems.

  12. A Fundamental Study for Aerodynamic Characteristics of Supersonic Biplane Wing and Wing-Body Configurations

    NASA Astrophysics Data System (ADS)

    Odaka, Yusuke; Kusunose, Kazuhiro

    In order to develop a quiet supersonic transport, it is necessary to reduce shock waves around the transport. Shock waves, in general, are the cause of the airplane's sonic boom. Authors have been studying an aerodynamic feasibility of supersonic biplanes based on the concept of the Busemann biplane. In this paper, the three dimensional effect of wing geometries on their wave drags, including wing tip effects and the interference effects between the wing and a body (Wing-Body configurations) are investigated, using CFD code in Euler (inviscid) mode. As a result, we can conclude that the supersonic biplane wings at their design Mach number (M∞=1.7) are still capable of reducing wave drag significantly similar to that of the 2-D supersonic biplane.

  13. Aerodynamic design and analysis system for supersonic aircraft. Part 2: User's manual

    NASA Technical Reports Server (NTRS)

    Middleton, W. D.; Lundry, J. L.; Coleman, R. G.

    1975-01-01

    An integrated system of computer programs for supersonic configurations is described. An explanation of system usage, the input definitions, and example output are included. For Part 1, see N75-18185; for Part 3, see N75-18186.

  14. A Preliminary Evaluation of Supersonic Transport Category Vehicle Operations in the National Airspace System

    NASA Technical Reports Server (NTRS)

    Underwood, Matthew C.; Guminsky, Michael D.

    2015-01-01

    Several public sector businesses and government agencies, including the National Aeronautics and Space Administration are currently working on solving key technological barriers that must be overcome in order to realize the vision of low-boom supersonic flights conducted over land. However, once these challenges are met, the manner in which this class of aircraft is integrated in the National Airspace System may become a potential constraint due to the significant environmental, efficiency, and economic repercussions that their integration may cause. Background research was performed on historic supersonic operations in the National Airspace System, including both flight deck procedures and air traffic controller procedures. Using this information, an experiment was created to test some of these historic procedures in a current-day, emerging Next Generation Air Transportation System (NextGen) environment and observe the interactions between commercial supersonic transport aircraft and modern-day air traffic. Data was gathered through batch simulations of supersonic commercial transport category aircraft operating in present-day traffic scenarios as a base-lining study to identify the magnitude of the integration problems and begin the exploration of new air traffic management technologies and architectures which will be needed to seamlessly integrate subsonic and supersonic transport aircraft operations. The data gathered include information about encounters between subsonic and supersonic aircraft that may occur when supersonic commercial transport aircraft are integrated into the National Airspace System, as well as flight time data. This initial investigation is being used to inform the creation and refinement of a preliminary Concept of Operations and for the subsequent development of technologies that will enable overland supersonic flight.

  15. Unique, clean-air, continuous-flow, high-stagnation-temperature facility for supersonic combustion research

    NASA Technical Reports Server (NTRS)

    Krauss, R. H.; Mcdaniel, J. C., Jr.; Scott, J. E., Jr.; Whitehurst, R. B., III; Segal, C.

    1988-01-01

    Accurate, spatially-resolved measurements can be conducted of a model supersonic combustor in a clean air/continuous flow supersonic combustion facility whose long run times will allow not only the point-by-point mapping of flow field variables with laser diagnostics but facilitate the simulation of steady-state combustor conditions. The facility will provide a Mach 2 freestream with static pressures in the 1 to 1/6 atm range, and stagnation temperatures of up to 2000 K.

  16. Effects Of Suspension-Line Damping On LADT #3 And Supersonic BLDT Parachute Inflation Dynamics

    NASA Technical Reports Server (NTRS)

    Poole, Lamont R.

    1972-01-01

    A two-body computerized mathematical model is used to calculate planar dynamics of the LADT #3 and supersonic BLDT parachute inflations. Results indicate that the calculated loads and motions of the LADT #3 inflation are not affected appreciably by variation in the suspension-line damping coefficient. However, variation of the coefficient results in significant changes in the calculated loads and strain rates of the supersonic BLDT inflation.

  17. Aerodynamic design and analysis system for supersonic aircraft. Part 1: General description and theoretical development

    NASA Technical Reports Server (NTRS)

    Middleton, W. D.; Lundry, J. L.

    1975-01-01

    An integrated system of computer programs has been developed for the design and analysis of supersonic configurations. The system uses linearized theory methods for the calculation of surface pressures and supersonic area rule concepts in combination with linearized theory for calculation of aerodynamic force coefficients. Interactive graphics are optional at the user's request. This part presents a general description of the system and describes the theoretical methods used.

  18. Ignition and Flameholding in a Supersonic Combustor by an Electrical Discharge Combined with a Fuel Injector

    DTIC Science & Technology

    2014-01-01

    1 Ignition and Flameholding in a Supersonic Combustor by an Electrical Discharge Combined with a Fuel Injector K. V. Savelkin 1 , D. A...presents the results of an experimental study of supersonic combustor operation enhanced by an electrical discharge. A novel scheme of plasma assisted...experimental combustor with the cross section of 72 mm (width)  60 mm (height) and length of 600 mm operates at a Mach number of M=2, initial stagnation

  19. Development of a Parachute System for Deceleration of Flying Vehicles in Supersonic Regimes

    NASA Astrophysics Data System (ADS)

    Pilyugin, N. N.; Khlebnikov, V. S.

    2010-09-01

    Aerodynamic problems arising during design and development of braking systems for re-entry vehicles are analyzed. Aerodynamic phenomena and laws valid in a supersonic flow around a pair of bodies having different shapes are studied. Results of this research can be used in solving application problems (arrangement and optimization of experiments; design and development of various braking systems for re-entry vehicles moving with supersonic speeds in the atmosphere).

  20. Confinement of Plasma along Shaped Open Magnetic Fields from the Centrifugal Force of Supersonic Plasma Rotation

    SciTech Connect

    Teodorescu, C.; Young, W. C.; Swan, G. W. S.; Ellis, R. F.; Hassam, A. B.; Romero-Talamas, C. A.

    2010-08-20

    Interferometric density measurements in plasmas rotating in shaped, open magnetic fields demonstrate strong confinement of plasma parallel to the magnetic field, with density drops of more than a factor of 10. Taken together with spectroscopic measurements of supersonic ExB rotation of sonic Mach 2, these measurements are in agreement with ideal MHD theory which predicts large parallel pressure drops balanced by centrifugal forces in supersonically rotating plasmas.