Science.gov

Sample records for 7-foot supersonic wind

  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. 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.

  6. 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.

  7. 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

  8. 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.

  9. 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

  10. 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

  11. 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

  12. 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.

  13. 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

  14. 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.)

  15. 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.

  16. 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.

  17. 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.

  18. Supersonic Flow Choking in Engine Wind Tunnels

    NASA Astrophysics Data System (ADS)

    Mitani, Tohru; Miyajima, Hiroshi; Tani, Koichiro; Kouchi, Toshinori; Sakuranaka, Noboru; Watanabe, Syuichi

    Breakdown of diffuser flow was often observed in our scramjet engine tests. This facility operation may damage the engine wind-tunnel and should be prevented. An one-dimensional analysis was applied to the diffuser flow to identify the causes of the flow breakdown. All the losses and gains by engine and friction loss in the diffuser were represented by point-sources of mass, momentum and energy. The thermal choking condition was calculated by uses of a chemical equilibrium code. The fuel rates causing the flow-choking successfully reproduced the limit fuel rates observed in our tests. Inlet-unstart of engine lost the ejector-pumping effect in the diffuser system to trigger the flow choking. The choking was also promoted by the drag of the gas sampling rakes. The choking in diffuser flow and the engine unstart may couple each other to cause hysteresis in the diffuser breakdown, which was also experienced in our tests. A rocket-based, combined-cycle (RBCC) engine will be tested under the Mach 4 condition. The engine easily causes the choking of diffuser because of the large propellant supply rates and the relatively-low specific impulse. Operation of the wind-tunnel was discussed to control the flow choking in the tests.

  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. 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.

  1. 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.

  2. 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.

  3. 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

  4. Damping insert materials for settling chambers of supersonic wind tunnels

    NASA Astrophysics Data System (ADS)

    Wu, Jie; Radespiel, Rolf

    2017-03-01

    This study describes the application of a novel damping insert material for reducing the flow fluctuations in a tandem nozzle supersonic wind tunnel. This new damping material is composed of multi-layer stainless steel wired meshes. The influences of the multi-layer mesh, such as the quantity of the mesh layer and the installed location in the settling chamber, to the freestream quality have been investigated. A Pitot probe instrumented with a Kulite pressure sensor and a hot-wire probe are employed to monitor the flow fluctuation in the test section of the wind tunnel. Thereafter, a combined modal analysis is applied for the disturbance qualification. Additionally, the transient Mach number in the test section is measured. The disturbance qualification indicates that the multi-layer mesh performs well in providing reduction of vorticity reduction and acoustic fluctuations. Comparable flow quality of the freestream was also obtained using a combination of flexible damping materials. However, the life-span of the new damping materials is much longer. The time transient of the Mach number measured in the test section indicates that the mean flow is rather constant over run time. Furthermore, the time-averaged pressure along the settling chamber is recorded and it shows the distribution of pressure drop by settling chamber inserts.

  5. Particle acceleration via reconnection processes in the supersonic solar wind

    SciTech Connect

    Zank, G. P.; Le Roux, J. A.; Webb, G. M.; Dosch, A.; Khabarova, O.

    2014-12-10

    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 + M{sub A} )/2, where M{sub A} is the Alfvén Mach number. Considering only magnetic island contraction yields power-law-like solutions with index –3(1 + τ {sub c}/(8τ{sub diff})), where τ {sub c}/τ{sub diff} 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 τ{sub diff}/τ {sub 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 {sup –5} (c particle

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

    NASA Technical Reports Server (NTRS)

    Wolf, Stephen W. D.

    1995-01-01

    Low-disturbance or 'quiet' wind tunnels are now considered an essential part of meaningful boundary layer transition research. Advances in Supersonic Laminar Flow Control (SLFC) technology for swept wings depends on a better understanding of the receptivity of the transition phenomena to attachment-line contamination and cross-flows. This need has provided the impetus for building the Laminar Flow Supersonic Wind Tunnel (LFSWT) at NASA-Ames, as part of the NASA High Speed Research Program (HSRP). The LFSWT was designed to provide NASA with an unequaled capability for transition research at low supersonic Mach numbers (<2.5). The following are the objectives in support of the new Fluid Mechanic Laboratory (FML) quiet supersonic wind tunnel: (I) Develop a unique injector drive system using the existing FML indraft compressor; (2) Develop an FML instrumentation capability for quiet supersonic wind tunnel evaluation and transition studies at NASA-Ames; (3) Determine the State of the Art in quiet supersonic wind tunnel design; (4) Build and commission the LFSWT; (5) Make detailed flow quality measurements in the LFSWT; (6) Perform tests of swept wing models in the LFSWT in support of the NASA HSR program; and (7) Provide documentation of research progress.

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

    NASA Technical Reports Server (NTRS)

    Wolf, Stephen W. D.

    1993-01-01

    The main objective of this work is to develop an interim Quiet (low-disturbance) supersonic wind tunnel for the NASA-Ames Fluid Mechanics Laboratory (FML). The main emphasis is to bring on-line a full-scale Mach 1.6 tunnel as rapidly as possible to impact the NASA High Speed Research Program (HSRP). The development of a cryogenic adaptive nozzle and other sophisticated features of the tunnel will now happen later, after the full scale wind tunnel is in operation. The work under this contract for the period of this report can be summarized as follows: provide aerodynamic design requirements for the NASA-Ames Fluid Mechanics Laboratory (FML) Laminar Flow Supersonic Wind Tunnel (LFSWT); research design parameters for a unique Mach 1.6 drive system for the LFSWT using an 1/8th-scale Proof-of-Concept (PoC) supersonic wind tunnel; carry out boundary layer transition studies in PoC to aid the design of critical components of the LFSWT; appraise the State of the Art in quiet supersonic wind tunnel design; and help develop a supersonic research capability within the FML particularly in the areas of high speed transition measurements and schlieren techniques. The body of this annual report summarizes the work of the Principal Investigator.

  8. Supersonic wind tunnel nozzles: A selected, annotated bibliography to aid in the development of quiet wind tunnel technology

    NASA Technical Reports Server (NTRS)

    Wolf, Stephen W. D.

    1990-01-01

    This bibliography, with abstracts, consists of 298 citations arranged in chronological order. The citations were selected to be helpful to persons engaged in the design and development of quiet (low disturbance) nozzles for modern supersonic wind tunnels. Author, subject, and corporate source indexes are included to assist with the location of specific information.

  9. 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.

  10. Increased Mach Number Capability for the NASA Glenn 10x10 Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Slater, J. W.; Saunders, J. D.

    2015-01-01

    Computational simulations and wind tunnel testing were conducted to explore the operation of the Abe Silverstein Supersonic Wind Tunnel at the NASA Glenn Research Center at test section Mach numbers above the current limit of Mach 3.5. An increased Mach number would enhance the capability for testing of supersonic and hypersonic propulsion systems. The focus of the explorations was on understanding the flow within the second throat of the tunnel, which is downstream of the test section and is where the supersonic flow decelerates to subsonic flow. Methods of computational fluid dynamics (CFD) were applied to provide details of the shock boundary layer structure and to estimate losses in total pressure. The CFD simulations indicated that the tunnel could be operated up to Mach 4.0 if the minimum width of the second throat was made smaller than that used for previous operation of the tunnel. Wind tunnel testing was able to confirm such operation of the tunnel at Mach 3.6 and 3.7 before a hydraulic failure caused a stop to the testing. CFD simulations performed after the wind tunnel testing showed good agreement with test data consisting of static pressures along the ceiling of the second throat. The CFD analyses showed increased shockwave boundary layer interactions, which was also observed as increased unsteadiness of dynamic pressures collected in the wind tunnel testing.

  11. Increased Mach Number Capability for the NASA Glenn 10x10 Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Slater, John; Saunders, John

    2014-01-01

    Computational simulations and wind tunnel testing were conducted to explore the operation of the Abe Silverstein Supersonic Wind Tunnel at the NASA Glenn Research Center at test section Mach numbers above the current limit of Mach 3.5. An increased Mach number would enhance the capability for testing of supersonic and hypersonic propulsion systems. The focus of the explorations was on understanding the flow within the second throat of the tunnel, which is downstream of the test section and is where the supersonic flow decelerates to subsonic flow. Methods of computational fluid dynamics (CFD) were applied to provide details of the shock boundary layer structure and to estimate losses in total pressure. The CFD simulations indicated that the tunnel could be operated up to Mach 4.0 if the minimum width of the second throat was made smaller than that used for previous operation of the tunnel. Wind tunnel testing was able to confirm such operation of the tunnel at Mach 3.6 and 3.7 before a hydraulic failure caused a stop to the testing. CFD simulations performed after the wind tunnel testing showed good agreement with test data consisting of static pressures along the ceiling of the second throat. The CFD analyses showed increased shockwave boundary layer interactions, which was also observed as increased unsteadiness of dynamic pressures collected in the wind tunnel testing.

  12. Low-Disturbance Flow Characteristics of the NASA-Ames Laminar Flow Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Wolf, Stephen W. D.; Laub, James A.; Davis, Sanford S. (Technical Monitor)

    1994-01-01

    A unique, low-disturbance (quiet) supersonic wind tunnel has been commissioned at the NASA-Ames Fluid Mechanics Laboratory (FML) to support Supersonic Laminar Flow Control (SLFC) research. Known as the Laminar Flow Supersonic Wind Tunnel (LFSWT), this tunnel is designed to operate at potential cruise Mach numbers and unit Reynolds numbers (Re) of the High Speed Civil Transport (HSCT). The need to better understand the receptivity of the transition phenomena on swept (HSCT) wings to attachment-line contamination and cross-flows has provided the impetus for building the LFSWT. Low-disturbance or "quiet" wind tunnels are known to be an essential part of any meaningful boundary layer transition research. In particular, the receptivity of supersonic boundary layers to wind tunnel disturbances can significantly alter the transition phenomena under investigation on a test model. Consequently, considerable effort has gone into the design of the LFSWT to provide quiet flow. The paper describes efforts to quantify the low-disturbance flows in the LFSWT operating at Mach 1.6, as a precursor to transition research on wing models. The research includes: (1) Flow measurements in both the test section and settling chamber of the LFSWT, using a full range of measurement techniques; (2) Study of the state of the test section boundary layer so far by using a single hot-wire mounted above the floor centerline, with and without boundary layer trips fitted at the test section entrance; (3) The effect of flow quality of unsteady supersonic diffuser flow, joint steps and gaps, and wall vibration.

  13. Orion Capsule and Launch Abort System (LAS) installed in the NASA Glenn 8x6 Supersonic Wind Tunnel f

    NASA Technical Reports Server (NTRS)

    2007-01-01

    Orion Capsule and Launch Abort System (LAS) installed in the NASA Glenn 8x6 Supersonic Wind Tunnel for testing. This test is an Aero Acoustic test of the LAS. Pictured is the calibration of the model's angle of attack

  14. Axisymmetric Calculations of a Low-Boom Inlet in a Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Chima, Rodrick V.; Hirt, Stefanie M.; Reger, Robert

    2011-01-01

    This paper describes axisymmetric CFD predictions made of a supersonic low-boom inlet with a facility diffuser, cold pipe, and mass flow plug within wind tunnel walls, and compares the CFD calculations with the experimental data. The inlet was designed for use on a small supersonic aircraft that would cruise at Mach 1.6, with a Mach number over the wing of 1.7. The inlet was tested in the 8-ft by 6-ft Supersonic Wind Tunnel at NASA Glenn Research Center in the fall of 2010 to demonstrate the performance and stability of a practical flight design that included a novel bypass duct. The inlet design is discussed here briefly. Prior to the test, CFD calculations were made to predict the performance of the inlet and its associated wind tunnel hardware, and to estimate flow areas needed to throttle the inlet. The calculations were done with the Wind-US CFD code and are described in detail. After the test, comparisons were made between computed and measured shock patterns, total pressure recoveries, and centerline pressures. The results showed that the dual-stream inlet had excellent performance, with capture ratios near one, a peak core total pressure recovery of 96 percent, and a large stable operating range. Predicted core recovery agreed well with the experiment but predicted bypass recovery and maximum capture ratio were high. Calculations of offdesign performance of the inlet along a flight profile agreed well with measurements and previous calculations.

  15. An Experimental Device for Generating High Frequency Perturbations in Supersonic Wind Tunnels

    NASA Technical Reports Server (NTRS)

    Melcher, Kevin J.; Ibrahim, Mounir B.

    1996-01-01

    This paper describes the analytical study of a device that has been proposed as a mechanism for generating gust-like perturbations in supersonic wind tunnels. The device is envisioned as a means to experimentally validate dynamic models and control systems designed for high-speed inlets. The proposed gust generator is composed of two flat trapezoidal plates that modify the properties of the flow ingested by the inlet. One plate may be oscillated to generate small perturbations in the flow. The other plate is held stationary to maintain a constant angle-of-attack. Using an idealized approach, design equations and performance maps for the new device were developed from the compressible flow relations. A two-dimensional CFD code was used to confirm the correctness of these results. The idealized approach was then used to design and evaluate a new gust generator for a 3.05-meter by 3.05-meter (10-foot by 10-foot) supersonic wind tunnel.

  16. NASA Lewis 8- by 6-foot supersonic wind tunnel user manual

    NASA Technical Reports Server (NTRS)

    Soeder, Ronald H.

    1993-01-01

    The 8- by 6-Foot Supersonic Wind Tunnel (SWT) at Lewis Research Center is available for use by qualified researchers. This manual contains tunnel performance maps which show the range of total temperature, total pressure, static pressure, dynamic pressure, altitude, Reynolds number, and mass flow as a function of test section Mach number. These maps are applicable for both the aerodynamic and propulsion cycle. The 8- by 6-Foot Supersonic Wind Tunnel is an atmospheric facility with a test section Mach number range from 0.36 to 2.0. General support systems (air systems, hydraulic system, hydrogen system, infrared system, laser system, laser sheet system, and schlieren system are also described as are instrumentation and data processing and acquisition systems. Pretest meeting formats are outlined. Tunnel user responsibility and personal safety requirements are also stated.

  17. Parametric Inlet Tested in Glenn's 10- by 10-Foot Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Slater, John W.; Davis, David O.; Solano, Paul A.

    2005-01-01

    The Parametric Inlet is an innovative concept for the inlet of a gas-turbine propulsion system for supersonic aircraft. The concept approaches the performance of past inlet concepts, but with less mechanical complexity, lower weight, and greater aerodynamic stability and safety. Potential applications include supersonic cruise aircraft and missiles. The Parametric Inlet uses tailored surfaces to turn the incoming supersonic flow inward toward an axis of symmetry. The terminal shock spans the opening of the subsonic diffuser leading to the engine. The external cowl area is smaller, which reduces cowl drag. The use of only external supersonic compression avoids inlet unstart--an unsafe shock instability present in previous inlet designs that use internal supersonic compression. This eliminates the need for complex mechanical systems to control unstart, which reduces weight. The conceptual design was conceived by TechLand Research, Inc. (North Olmsted, OH), which received funding through NASA s Small-Business Innovation Research program. The Boeing Company (Seattle, WA) also participated in the conceptual design. The NASA Glenn Research Center became involved starting with the preliminary design of a model for testing in Glenn s 10- by 10-Foot Supersonic Wind Tunnel (10 10 SWT). The inlet was sized for a speed of Mach 2.35 while matching requirements of an existing cold pipe used in previous inlet tests. The parametric aspects of the model included interchangeable components for different cowl lip, throat slot, and sidewall leading-edge shapes and different vortex generator configurations. Glenn researchers used computational fluid dynamics (CFD) tools for three-dimensional, turbulent flow analysis to further refine the aerodynamic design.

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

    NASA Technical Reports Server (NTRS)

    Wolf, Stephen W. D.

    1992-01-01

    The main objective of this work was the development of an interim quiet (low-disturbance) supersonic wind tunnel for the NASA-Ames Fluid Mechanics Laboratory (FML). This is a result of the need to bring the full-scale tunnel on-line as rapidly as possible to impact the NASA High Speed Research Program (HSRP). The development of a cryogenic adaptive nozzle and other sophisticated features of the tunnel will now happen later, after the full scale wind tunnel is in operation.

  19. Natural laminar flow wing for supersonic conditions: Wind tunnel experiments, flight test and stability computations

    NASA Astrophysics Data System (ADS)

    Vermeersch, Olivier; Yoshida, Kenji; Ueda, Yoshine; Arnal, Daniel

    2015-11-01

    In the framework of next supersonic transport airplane generation, the Japan Aerospace eXploration Agency (JAXA) has developed a new natural laminar flow highly swept wing. The design has been experimentally validated firstly in a supersonic wind tunnel and secondly accomplishing flight test. These experimental data were then analyzed and completed by numerical stability analyses in a joint research program between Onera and JAXA. At the design condition, for a Mach number M=2 at an altitude of h=18 km, results have confirmed the laminar design of the wing due to a strong attenuation of cross-flow instabilities ensuring an extended laminar zone. As the amplification of disturbances inside the boundary layer and transition process is very sensitive to external parameters, the impact of wall roughness of the models and the influence of Reynolds number on transition process have been carefully analyzed.

  20. Low-speed wind-tunnel test of a STOL supersonic-cruise fighter concept

    NASA Technical Reports Server (NTRS)

    Coe, Paul L., Jr.; Riley, Donald R.

    1988-01-01

    A wind-tunnel investigation was conducted to examine the low-speed static stability and control characteristics of a 0.10 scale model of a STOL supersonic cruise fighter concept. The concept, referred to as a twin boom fighter, was designed as a STOL aircraft capable of efficient long range supersonic cruise. The configuration name is derived from the long twin booms extending aft of the engine to the twin vertical tails which support a high center horizontal tail. The propulsion system features a two dimensional thrust vectoring exhaust nozzle which is located so that the nozzle hinge line is near the aircraft center of gravity. This arrangement is intended to allow large thrust vector angles to be used to obtain significant values of powered lift, while minimizing pitching moment trim changes. Low speed stability and control information was obtained over an angle of attack range including the stall. A study of jet induced power effects was included.

  1. 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.

  2. Effectiveness of a Wedge Probe to Measure Sonic Boom Signatures in a Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Wilcox, Floyd J., Jr.; Elmiligui, Alaa A.

    2013-01-01

    A wind tunnel investigation was conducted in the Langley Unitary Plan Wind Tunnel (UPWT) to determine the effectiveness of a wedge probe to measure sonic boom pressure signatures compared to a slender conical probe. A generic business jet model at a constant angle of attack and at a single model to probe separation distance was used to generate a sonic boom signature. Pressure signature data were acquired with both the wedge probe and a slender conical probe for comparison. The test was conducted at a Mach number of 2.0 and a free-stream unit Reynolds number of 2 million per foot. The results showed that the wedge probe was not effective in measuring the sonic boom pressure signature of the aircraft model in the supersonic wind tunnel. Data plots and a discussion of the results are presented. No tabulated data or flow visualization photographs are included.

  3. Uncertainty Analysis of NASA Glenn's 8- by 6-Foot Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Stephens, Julia E.; Hubbard, Erin P.; Walter, Joel A.; McElroy, Tyler

    2016-01-01

    An analysis was performed to determine the measurement uncertainty of the Mach Number of the 8- by 6-foot Supersonic Wind Tunnel at the NASA Glenn Research Center. This paper details the analysis process used, including methods for handling limited data and complicated data correlations. Due to the complexity of the equations used, a Monte Carlo Method was utilized for this uncertainty analysis. A summary of the findings are presented as pertains to understanding what the uncertainties are, how they impact various research tests in the facility, and methods of reducing the uncertainties in the future.

  4. High Response Dew Point Measurement System for a Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Blumenthal, Philip Z.

    1996-01-01

    A new high response on-line measurement system has been developed to continuously display and record the air stream dew point in the NASA Lewis 10 x 10 supersonic wind tunnel. Previous instruments suffered from such problems as very slow response, erratic readings, and high susceptibility to contamination. The system operates over the entire pressure level range of the 10 x 10 SWT, from less than 2 psia to 45 psia, without the need for a vacuum pump to provide sample flow. The system speeds up tunnel testing, provides large savings in tunnel power costs and provides the dew point input for the data-reduction subroutines which calculate test section conditions.

  5. Supersonic dynamic stability characteristics of a space shuttle orbiter. [wind tunnel tests of scale models

    NASA Technical Reports Server (NTRS)

    Freeman, D. C., Jr.; Boyden, R. P.; Davenport, E. E.

    1976-01-01

    Supersonic forced-oscillation tests of a 0.0165-scale model of a modified 089B Rockwell International shuttle orbiter were conducted in a wind tunnel for several configurations over a Mach range from 1.6 to 4.63. The tests covered angles of attack up to 30 deg. The period and damping of the basic unaugmented vehicle were calculated along the entry trajectory using the measured damping results. Some parameter analysis was made with the measured dynamic derivatives. Photographs of the test configurations and test equipment are shown.

  6. Simulation of pressure and temperature responses for the 20 Inch Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Motter, Mark A.

    1990-01-01

    A simulation of the pressure and temperature responses of the 20 inch Supersonic Wind Tunnel (SWT) is developed. The simulation models the tunnel system as a set of lumped parameter volumes connected by flow regulating elements such as valves and nozzles. Simulated transient responses of temperature and pressure for the five boundary points of the 20 inch SWT operating map are produced from their respective initial conditions, tunnel operating conditions, heater input power, and valve positions. Upon reaching steady state, a linearized model for each operating point is determined. Both simulated and actual tunnel responses are presented for comparison.

  7. Overview of Supersonic Aerodynamics Measurement Techniques in the NASA Langley Unitary Plan Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    2007-01-01

    An overview is given of selected measurement techniques used in the NASA Langley Research Center (NASA LaRC) Unitary Plan Wind Tunnel (UPWT) to determine the aerodynamic characteristics of aerospace vehicles operating at supersonic speeds. A broad definition of a measurement technique is adopted in this paper and is any qualitative or quantitative experimental approach that provides information leading to the improved understanding of the supersonic aerodynamic characteristics. On-surface and off-surface measurement techniques used to obtain discrete (point) and global (field) measurements and planar and global flow visualizations are described, and examples of all methods are included. The discussion is limited to recent experiences in the UPWT and is, therefore, not an exhaustive review of existing experimental techniques. The diversity and high quality of the measurement techniques and the resultant data illustrate the capabilities of a ground-based experimental facility and the key role that it plays in the advancement of our understanding, prediction, and control of supersonic aerodynamics.

  8. Entry, Descent, and Landing with Propulsive Deceleration: Supersonic Retropropulsion Wind Tunnel Testing and Shock Phenomena

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan

    2013-01-01

    The future exploration of the Solar System will require innovations in transportation and the use of entry, descent, and landing (EDL) systems at many planetary landing sites. The cost of space missions has always been prohibitive, and using the natural planetary and planet's moon atmospheres for entry, and descent can reduce the cost, mass, and complexity of these missions. This paper will describe some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be attractive for future Solar System missions. Future EDL systems may include an inflatable decelerator for the initial atmospheric entry and an additional supersonic retro-propulsion (SRP) rocket system for the final soft landing. A three engine retro-propulsion configuration with a 2.5 inch diameter sphere-cone aeroshell model was tested in the NASA Glenn 1x1 Supersonic Wind Tunnel (SWT). The testing was conducted to identify potential blockage issues in the tunnel, and visualize the rocket flow and shock interactions during supersonic and hypersonic entry conditions. Earlier experimental testing of a 70 degree Viking-like (sphere-cone) aeroshell was conducted as a baseline for testing of a supersonic retro-propulsion system. This baseline testing defined the flow field around the aeroshell and from this comparative baseline data, retro-propulsion options will be assessed. Images and analyses from the SWT testing with 300- and 500-psia rocket engine chamber pressures are presented here. In addition, special topics of electromagnetic interference with retro-propulsion induced shock waves and retro-propulsion for Earth launched booster recovery are also addressed.

  9. Entry, Descent, and Landing With Propulsive Deceleration: Supersonic Retropropulsion Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan

    2012-01-01

    The future exploration of the Solar System will require innovations in transportation and the use of entry, descent, and landing (EDL) systems at many planetary landing sites. The cost of space missions has always been prohibitive, and using the natural planetary and planet s moons atmosphere for entry, descent, and landing can reduce the cost, mass, and complexity of these missions. This paper will describe some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be attractive for future Solar System missions. Future EDL systems may include an inflatable decelerator for the initial atmospheric entry and an additional supersonic retro-propulsion (SRP) rocket system for the final soft landing. As part of those efforts, NASA began to conduct experiments to gather the experimental data to make informed decisions on the "best" EDL options. A model of a three engine retro-propulsion configuration with a 2.5 in. diameter sphere-cone aeroshell model was tested in the NASA Glenn 1- by 1-Foot Supersonic Wind Tunnel (SWT). The testing was conducted to identify potential blockage issues in the tunnel, and visualize the rocket flow and shock interactions during supersonic and hypersonic entry conditions. Earlier experimental testing of a 70 Viking-like (sphere-cone) aeroshell was conducted as a baseline for testing of a supersonic retro-propulsion system. This baseline testing defined the flow field around the aeroshell and from this comparative baseline data, retro-propulsion options will be assessed. Images and analyses from the SWT testing with 300- and 500-psia rocket engine chamber pressures are presented here. The rocket engine flow was simulated with a non-combusting flow of air.

  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. Tone Noise of Three Supersonic Helical Tip Speed Propellers in a Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Dittmar, J. H.; Jeracki, R. J.; Blaha, B. J.

    1979-01-01

    Three supersonic helical tip speed propellers were tested in the NASA Lewis 8- by 6-foot wind tunnel. This is a perforated-wall wind tunnel but it does not have acoustic damping material on its walls. The propellers were tested at tunnel through flow Mach numbers of 0.6, 0.7, 0.75, 0.8, and 0.85 with different rotational speeds and blade setting angles. The three propellers, which had approximately the same aerodynamic performance, incorporated different plan forms and different amounts of sweep and yielded different near field noise levels. The acoustically designed propeller had 45 deg of tip sweep and was significantly quieter at M = 0.8 cruise than the straight bladed propeller. The intermediate 30 deg tip sweep propeller, which was swept for aerodynamic purposes, exhibited noise that was between the other two propellers. Noise trends with varying helical tip Mach number and blade loading were also observed.

  12. Supersonic Retropropulsion CFD Validation with Ames Unitary Plan Wind Tunnel Test Data

    NASA Technical Reports Server (NTRS)

    Schauerhamer, Daniel G.; Zarchi, Kerry A.; Kleb, William L.; Edquist, Karl T.

    2013-01-01

    A validation study of Computational Fluid Dynamics (CFD) for Supersonic Retropropulsion (SRP) was conducted using three Navier-Stokes flow solvers (DPLR, FUN3D, and OVERFLOW). The study compared results from the CFD codes to each other and also to wind tunnel test data obtained in the NASA Ames Research Center 90 70 Unitary PlanWind Tunnel. Comparisons include surface pressure coefficient as well as unsteady plume effects, and cover a range of Mach numbers, levels of thrust, and angles of orientation. The comparisons show promising capability of CFD to simulate SRP, and best agreement with the tunnel data exists for the steadier cases of the 1-nozzle and high thrust 3-nozzle configurations.

  13. Development of Dynamic Flow Field Pressure Probes Suitable for Use in Large Scale Supersonic Wind Tunnels

    NASA Technical Reports Server (NTRS)

    Porro, A. Robert

    2000-01-01

    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 staff and inlet unstart transient event. Details of the design, development, and subsequent use of these probes are discussed in this report.

  14. Development of the NASA-Ames low disturbance supersonic wind tunnel for transition research up to Mach 2.5

    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 aerodynamic features of this new quiet tunnel will be a low-disturbance settling chamber, laminar boundary layers on the nozzle walls and steady supersonic diffuser flow. Furthermore, this new wind tunnel will operate continuously at uniquely low compression ratios (less than unity). This feature allows an existing non-specialist compressor to be used as a major part of the drive system. In this paper, we highlight activities associated with drive system development, the establishment of natural laminar flow on the test section walls, and instrumentation development for transition detection. Experimental results from an 1/8th-scale model of the supersonic wind tunnel are presented and discussed in association with theoretical predictions. Plans are progressing to build the full-scale wind tunnel by the end of 1993.

  15. Titan's Interaction With the Supersonic Solar Wind During T96: A Hybrid Simulation Study

    NASA Astrophysics Data System (ADS)

    Simon, S.; Feyerabend, M.; Neubauer, F. M.; Motschmann, U. M.; Bertucci, C.; Edberg, N. J. T.; Hospodarsky, G. B.; Kurth, W. S.

    2015-12-01

    We apply a hybrid simulation code (kinetic ions, fluid electrons) to study the structure of Titan's plasma environment during the T96 flyby on 01 December 2013. The T96 encounter constitutes the only event of the Cassini era during which Titan was exposed to the supersonic solar wind upstream of Saturn's bow shock. Cassini plasma and magnetic field observations revealed a Venus/Mars-like plasma interaction, including the formation of a bow shock and a magnetic barrier upstream of Titan. However, it has also been suggested that the ambient solar wind conditions were not stationary at the time of the T96 encounter: around closest approach Cassini probably passed through a region of fossilized magnetic fields, arising from Titan's interaction with a previously encountered, high-pressure solar wind regime. To substantiate this hypothesis, we carry out simulation runs for different ram pressures of the incident solar wind. The model reveals the interaction region to be dominated by kinetic effects, such as a splitting of the bow shock due to large ion gyroradii (shocklets). The simulation results are compared against Cassini magnetic field data from T96.

  16. Entry, Descent, and Landing with Propulsive Deceleration: Supersonic Retropropulsion Wind Tunnel Testing and Shock Phenomena

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan

    2014-01-01

    The future exploration of the Solar System will require innovations in transportation and the use of entry, descent, and landing (EDL) systems at many planetary landing sites. The cost of space missions has always been prohibitive, and using the natural planetary and planet's moon atmospheres for entry, and descent can reduce the cost, mass, and complexity of these missions. This paper will describe some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be attractive for future Solar System missions. Future EDL systems may include an inflatable decelerator for the initial atmospheric entry and an additional supersonic retropropulsion (SRP) rocket system for the final soft landing. A three engine retropropulsion configuration with a 2.5 in. diameter sphere-cone aeroshell model was tested in the NASA Glenn Research Center's 1- by 1-ft (1×1) Supersonic Wind Tunnel (SWT). The testing was conducted to identify potential blockage issues in the tunnel, and visualize the rocket flow and shock interactions during supersonic and hypersonic entry conditions. Earlier experimental testing of a 70deg Viking-like (sphere-cone) aeroshell was conducted as a baseline for testing of a SRP system. This baseline testing defined the flow field around the aeroshell and from this comparative baseline data, retropropulsion options will be assessed. Images and analyses from the SWT testing with 300- and 500-psia rocket engine chamber pressures are presented here. In addition, special topics of electromagnetic interference with retropropulsion induced shock waves and retropropulsion for Earth launched booster recovery are also addressed.

  17. Development of quiet-flow supersonic wind tunnels for laminar-turbulent transition research

    NASA Technical Reports Server (NTRS)

    Schneider, Steven P.

    1994-01-01

    This grant supported research into quiet-flow supersonic wind-tunnels, between May 1990 and December 1994. Quiet-flow nozzles operate with laminar nozzle-wall boundary layers, in order to provide low-disturbance flow for studies of laminar-turbulent transition under conditions comparable to flight. Major accomplishments include: (1) the design, fabrication, and performance-evaluation of a new kind of quiet tunnel, a quiet-flow Ludweig tube; (2) the integration of preexisting codes for nozzle design, 2D boundary-layer computation, and transition-estimation into a single user-friendly package for quiet-nozzle design; and (3) the design and preliminary evaluation of supersonic nozzles with square cross-section, as an alternative to conventional quiet-flow nozzles. After a brief summary of (1), a description of (2) is presented. Published work describing (3) is then summarized. The report concludes with a description of recent results for the Tollmien-Schlichting and Gortler instability in one of the square nozzles previously analyzed.

  18. Check Calibration of the NASA Glenn 10- by 10-Foot Supersonic Wind Tunnel (2014 Test Entry)

    NASA Technical Reports Server (NTRS)

    Johnson, Aaron; Pastor-Barsi, Christine; Arrington, E. Allen

    2016-01-01

    A check calibration of the 10- by 10-Foot Supersonic Wind Tunnel (SWT) was conducted in May/June 2014 using an array of five supersonic wedge probes to verify the 1999 Calibration. This check calibration was necessary following a control systems upgrade and an integrated systems test (IST). This check calibration was required to verify the tunnel flow quality was unchanged by the control systems upgrade prior to the next test customer beginning their test entry. The previous check calibration of the tunnel occurred in 2007, prior to the Mars Science Laboratory test program. Secondary objectives of this test entry included the validation of the new Cobra data acquisition system (DAS) against the current Escort DAS and the creation of statistical process control (SPC) charts through the collection of series of repeated test points at certain predetermined tunnel parameters. The SPC charts secondary objective was not completed due to schedule constraints. It is hoped that this effort will be readdressed and completed in the near future.

  19. Supersonic Retropropulsion Experimental Results from the NASA Langley Unitary Plan Wind Tunnel

    NASA Technical Reports Server (NTRS)

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

    2011-01-01

    A new supersonic retropropulsion experimental effort, intended to provide code validation data, was recently completed in the Langley Research Center Unitary Plan Wind Tunnel Test Section 2 over the Mach number range from 2.4 to 4.6. The experimental model was designed using insights gained from pre-test computations, which were instrumental for sizing and refining the model to minimize tunnel wall interference and internal flow separation concerns. A 5-in diameter 70-deg sphere-cone forebody with a roughly 10-in long cylindrical aftbody was the baseline configuration selected for this study. The forebody was designed to accommodate up to four 4:1 area ratio supersonic nozzles. Primary measurements for this model were a large number of surface pressures on the forebody and aftbody. Supplemental data included high-speed Schlieren video and internal pressures and temperatures. The run matrix was developed to allow for the quantification of various sources of experimental uncertainty, such as random errors due to run-to-run variations and bias errors due to flow field or model misalignments. Preliminary results and observations from the test are presented, while detailed data and uncertainty analyses are ongoing.

  20. Instantaneous flow measurements in a supersonic wind tunnel using spectrally resolved Rayleigh scattering

    NASA Technical Reports Server (NTRS)

    Seasholtz, Richard G.; Buggele, Alvin E.; Reeder, Mark F.

    1995-01-01

    Results of a feasibility study to apply laser Rayleigh scattering to non-intrusively measure flow properties in a small supersonic wind tunnel are presented. The technique uses an injection seeded, frequency doubled Nd:YAG laser tuned to an absorption band of iodine. The molecular Rayleigh scattered light is filtered with an iodine cell to block light at the laser frequency. The Doppler-shifted Rayleigh scattered light that passes through the iodine cell is analyzed with a planar mirror Fabry-Perot interferometer used in a static imaging mode. An intensified CCD camera is used to record the images. The images are analyzed at several subregions, where the flow velocity is determined. Each image is obtained with a single laser pulse, giving instantaneous measurements.

  1. New Model Exhaust System Supports Testing in NASA Lewis' 10- by 10-Foot Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Roeder, James W., Jr.

    1998-01-01

    In early 1996, the ability to run NASA Lewis Research Center's Abe Silverstein 10- by 10- Foot Supersonic Wind Tunnel (10x10) at subsonic test section speeds was reestablished. Taking advantage of this new speed range, a subsonic research test program was scheduled for the 10x10 in the fall of 1996. However, many subsonic aircraft test models require an exhaust source to simulate main engine flow, engine bleed flows, and other phenomena. This was also true of the proposed test model, but at the time the 10x10 did not have a model exhaust capability. So, through an in-house effort over a period of only 5 months, a new model exhaust system was designed, installed, checked out, and made ready in time to support the scheduled test program.

  2. NASA Glenn 1-by 1-Foot Supersonic Wind Tunnel User Manual

    NASA Technical Reports Server (NTRS)

    Seablom, Kirk D.; Soeder, Ronald H.; Stark, David E.; Leone, John F. X.; Henry, Michael W.

    1999-01-01

    This manual describes the NASA Glenn Research Center's 1 - by 1 -Foot Supersonic Wind Tunnel and provides information for customers who wish to conduct experiments in this facility. Tunnel performance envelopes of total pressure, total temperature, and dynamic pressure as a function of test section Mach number are presented. For each Mach number, maps are presented of Reynolds number per foot as a function of the total air temperature at the test section inlet for constant total air pressure at the inlet. General support systems-such as the service air, combustion air, altitude exhaust system, auxiliary bleed system, model hydraulic system, schlieren system, model pressure-sensitive paint, and laser sheet system are discussed. In addition, instrumentation and data processing, acquisition systems are described, pretest meeting formats and schedules are outlined, and customer responsibilities and personnel safety are addressed.

  3. User manual for NASA Lewis 10 by 10 foot supersonic wind tunnel. Revised

    NASA Technical Reports Server (NTRS)

    Soeder, Ronald H.

    1995-01-01

    This manual describes the 10- by 10-Foot Supersonic Wind Tunnel at the NASA Lewis Research Center and provides information for users who wish to conduct experiments in this facility. Tunnel performance operating envelopes of altitude, dynamic pressure, Reynolds number, total pressure, and total temperature as a function of test section Mach number are presented. Operating envelopes are shown for both the aerodynamic (closed) cycle and the propulsion (open) cycle. The tunnel test section Mach number range is 2.0 to 3.5. General support systems, such as air systems, hydraulic system, hydrogen system, fuel system, and Schlieren system, are described. Instrumentation and data processing and acquisition systems are also described. Pretest meeting formats and schedules are outlined. Tunnel user responsibility and personnel safety are also discussed.

  4. Uncertainty Analysis of the NASA Glenn 8x6 Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Stephens, Julia; Hubbard, Erin; Walter, Joel; McElroy, Tyler

    2016-01-01

    This paper presents methods and results of a detailed measurement uncertainty analysis that was performed for the 8- by 6-foot Supersonic Wind Tunnel located at the NASA Glenn Research Center. The statistical methods and engineering judgments used to estimate elemental uncertainties are described. The Monte Carlo method of propagating uncertainty was selected to determine the uncertainty of calculated variables of interest. A detailed description of the Monte Carlo method as applied for this analysis is provided. Detailed uncertainty results for the uncertainty in average free stream Mach number as well as other variables of interest are provided. All results are presented as random (variation in observed values about a true value), systematic (potential offset between observed and true value), and total (random and systematic combined) uncertainty. The largest sources contributing to uncertainty are determined and potential improvement opportunities for the facility are investigated.

  5. Two-Dimensional Bifurcated Inlet/Engine Tests Completed in 10- by 10-Foot Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Saunders, John D.

    1999-01-01

    A Two-Dimensional Bifurcated (2DB) Inlet was successfully tested in NASA Lewis Research Center s 10- by 10-Foot Supersonic Wind Tunnel. These tests were the culmination of a collaborative effort between the Boeing Company, General Electric, Pratt & Whitney, and Lewis. Extensive support in-house at Lewis contributed significantly to the progress and accomplishment of this test. The results, which met or exceeded many of the High-Speed Research (HSR) Program goals, were used to revise system studies within the HSR Program. The HSR Program is focused on developing low-noise, low-polluting, high-efficiency supersonic commercial aircraft. A supersonic inlet is an important component of an efficient, low-noise vehicle.

  6. 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.

  7. ACCRETION OF SUPERSONIC WINDS ONTO BLACK HOLES IN 3D: STABILITY OF THE SHOCK CONE

    SciTech Connect

    Gracia-Linares, M.; Guzmán, F. S.

    2015-10-10

    Using numerical simulations we present the accretion of supersonic winds onto a rotating black hole in three dimensions. We study five representative directions of the wind with respect to the axis of rotation of the black hole and focus on the evolution and stability of the high-density shock cone that is formed during the process. We explore both the regime in which the shock cone is expected to be stable in order to confirm previous results obtained with two-dimensional simulations, and the regime in which the shock cone is expected to show a flip–flop (FF) type of instability. The methods used to attempt a triggering of the instability were (i) the accumulation of numerical errors and (ii) the explicit application of a perturbation on the velocity field after the shock cone was formed. The result is negative, that is, we did not find the FF instability within the parameter space we explored, including cases that are expected to be unstable.

  8. Cone-Probe Rake Design and Calibration for Supersonic Wind Tunnel Models

    NASA Technical Reports Server (NTRS)

    Won, Mark J.

    1999-01-01

    A series of experimental investigations were conducted at the NASA Langley Unitary Plan Wind Tunnel (UPWT) to calibrate cone-probe rakes designed to measure the flow field on 1-2% scale, high-speed wind tunnel models from Mach 2.15 to 2.4. The rakes were developed from a previous design that exhibited unfavorable measurement characteristics caused by a high probe spatial density and flow blockage from the rake body. Calibration parameters included Mach number, total pressure recovery, and flow angularity. Reference conditions were determined from a localized UPWT test section flow survey using a 10deg supersonic wedge probe. Test section Mach number and total pressure were determined using a novel iterative technique that accounted for boundary layer effects on the wedge surface. Cone-probe measurements were correlated to the surveyed flow conditions using analytical functions and recursive algorithms that resolved Mach number, pressure recovery, and flow angle to within +/-0.01, +/-1% and +/-0.1deg , respectively, for angles of attack and sideslip between +/-8deg. Uncertainty estimates indicated the overall cone-probe calibration accuracy was strongly influenced by the propagation of measurement error into the calculated results.

  9. Computational Aeroelastic Analysis of the Semi-Span Super-Sonic Transport (S4T) Wind-Tunnel Model

    NASA Technical Reports Server (NTRS)

    Sanetrik, Mark D.; Silva, Walter A.; Hur, Jiyoung

    2012-01-01

    A summary of the computational aeroelastic analysis for the Semi-Span Super-Sonic Transport (S4T) wind-tunnel model is presented. A broad range of analysis techniques, including linear, nonlinear and Reduced Order Models (ROMs) were employed in support of a series of aeroelastic (AE) and aeroservoelastic (ASE) wind-tunnel tests conducted in the Transonic Dynamics Tunnel (TDT) at NASA Langley Research Center. This research was performed in support of the ASE element in the Supersonics Program, part of NASA's Fundamental Aeronautics Program. The analysis concentrated on open-loop flutter predictions, which were in good agreement with experimental results. This paper is one in a series that comprise a special S4T technical session, which summarizes the S4T project.

  10. Background Pressure Profiles for Sonic Boom Vehicle Testing in the NASA Glenn 8- by 6-Foot Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Castner, Raymond; Shaw, Stephen; Adamson, Eric; Simerly, Stephanie

    2013-01-01

    In an effort to identify test facilities that offer sonic boom measurement capabilities, an exploratory test program was initiated using wind tunnels at NASA research centers. The subject of this report is the sonic boom pressure rail data collected in the Glenn Research Center 8- by 6-Foot Supersonic Wind Tunnel. The purpose is to summarize the lessons learned based on the test activity, specifically relating to collecting sonic boom data which has a large amount of spatial pressure variation. The wind tunnel background pressure profiles are presented as well as data which demonstrated how both wind tunnel Mach number and model support-strut position affected the wind tunnel background pressure profile. Techniques were developed to mitigate these effects and are presented.

  11. Suprathermal Ion Acceleration in Multiple Contracting and Reconnecting Inertial-scale Flux Ropes in the Supersonic Solar Wind.

    NASA Astrophysics Data System (ADS)

    le Roux, J. A.; Zank, G. P.; Webb, G. M.

    2014-12-01

    3D and 2D MHD turbulence simulations with a strong large-scale magnetic field show that the turbulence is filled with quasi-2D inertial-scale flux ropes that intermittently reconnect, while test particle simulations stress how suprathermal particles can be efficiently accelerated to produce power law spectra (kappa distributions) when traversing multiple flux ropes. Solar wind observations indicate that the statistical properties of the turbulence agree well with the MHD turbulence simulation. In addition, recent observations show the presence of different size inertial-scale magnetic islands in the slow solar wind near the heliospheric current sheet, evidence of island merging, and of heating of ions and electrons in their vicinity. At the same time, observations in the supersonic solar wind suggest the existence of suprathermal ion spectra in the solar wind frame where the distribution function is a power law in momentum with a -5 exponent. We present a new statistical transport theory to model the acceleration of superthermal ions traversing multiple contracting and reconnecting inertial-scale quasi-2D flux ropes in the supersonic solar wind. Steady-state analytical solutions for the accelerated suprathermal particle spectrum in a radially expanding solar wind will be explored to show under what conditions one can reproduce the observed superthermal power-law slope.

  12. Facility Upgrade/Replacement Tasks ('planned') at the NASA Glenn Research Center 10x10 Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Giriunas, Julius A.

    2012-01-01

    Facility upgrades and large maintenance tasks needed at the NASA Glenn 10x10 Supersonic Wind Tunnel requires significant planning to make sure implementation proceeds in an efficiently and cost effective manner. Advanced planning to secure the funding, complete design efforts and schedule the installation needs to be thought out years in advance to avoid interference with wind tunnel testing. This presentation describes five facility tasks planned for implementation over the next few years. The main focus of the presentation highlights the efforts on possible replacement of the diesel generator and the rationale behind the effort.

  13. An Overview of the Semi-Span Super-Sonic Transport (S4T) Wind-Tunnel Model Program

    NASA Technical Reports Server (NTRS)

    Silva, Walter A.; Perry, Boyd, III; Florance, James R.; Sanetrik, Mark D.; Wieseman, Carol D.; Stevens, William L.; Funk, Christie J.; Christhilf, David M.; Coulson, David A.

    2012-01-01

    A summary of computational and experimental aeroelastic (AE) and aeroservoelastic (ASE) results for the Semi-Span Super-Sonic Transport (S4T) wind-tunnel model is presented. A broad range of analyses and multiple AE and ASE wind-tunnel tests of the S4T wind-tunnel model have been performed in support of the ASE element in the Supersonics Program, part of the NASA Fundamental Aeronautics Program. This paper is intended to be an overview of multiple papers that comprise a special S4T technical session. Along those lines, a brief description of the design and hardware of the S4T wind-tunnel model will be presented. Computational results presented include linear and nonlinear aeroelastic analyses, and rapid aeroelastic analyses using CFD-based reduced-order models (ROMs). A brief survey of some of the experimental results from two open-loop and two closed-loop wind-tunnel tests performed at the NASA Langley Transonic Dynamics Tunnel (TDT) will be presented as well.

  14. Tone noise of three supersonic helical tip speed propellers in a wind tunnel

    NASA Technical Reports Server (NTRS)

    Dittmar, J. H.; Jeracki, R. J.; Blaha, B. J.

    1979-01-01

    Three eight bladed supersonic helical tip speed propellers were tested in the NASA Lewis Wind Tunnel at through flow Mach numbers of 0.6, 0.7, 0.75, 0.8, and 0.85. Such propellers are being considered as propulsors for a future energy conservative aircraft. The three propellers of equal performance, incorporate different plan forms and different amounts of sweep in order to investigate their noise generation, which could cause a cabin environment problem. The acoustically designed propeller SR-3 had 45 deg of tip sweep and was significantly quieter at Mach 0.8 cruise than the straight bladed propeller, SR-2. An intermediate 30 deg tip sweep propeller, SR-1m exhibited a noise level between the other two. Enhanced pressure-time traces indicated that SR-2 and SR-1m exhibited shock-like pressure traces at Mach-0.8 cruise while the SR-3 did not. Noise trends with varying helical tip Mach number and blade loading were also observed.

  15. Analysis of Dynamic Data from Supersonic Retropropulsion Experiments in NASA Langley's Unitary Plan Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Codoni, Joshua R.; Berry, Scott A.

    2012-01-01

    Recent experimental supersonic retropropulsion tests were conducted at the NASA Langley Research Center Unitary Plan Wind Tunnel Test Section 2 for a range of Mach numbers from 2.4 to 4.6. A 5-inch 70-degree sphere-cone forebody model with a 10-inch cylindrical aftbody experimental model was used which is capable of multiple retrorocket configurations. These configurations include a single central nozzle on the center point of the forebody, three nozzles at the forebody half-radius, and a combination of the first two configurations with no jets being plugged. A series of measurements were achieved through various instrumentation including forebody and aftbody pressure, internal pressures and temperatures, and high speed Schlieren visualization. Specifically, several high speed pressure transducers on the forebody and in the plenum were implemented to look at unsteady flow effects. The following work focuses on analyzing frequency traits due to the unsteady flow for a range of thrust coefficients for single, tri, and quad-nozzle test cases at freestream Mach 4.6 and angle of attack ranging from -8 degrees to +20 degrees. This analysis uses Matlab s fast Fourier transform, Welch's method (modified average of a periodogram), to create a power spectral density and analyze any high speed pressure transducer frequency traits due to the unsteady flow.

  16. Calibration of the NASA Glenn 8- by 6-Foot Supersonic Wind Tunnel (1996 and 1997 Tests)

    NASA Technical Reports Server (NTRS)

    Arrington, E. Allen

    2012-01-01

    There were several physical and operational changes made to the NASA Glenn Research Center 8- by 6-Foot Supersonic Wind Tunnel during the period of 1992 through 1996. Following each of these changes, a facility calibration was conducted to provide the required information to support the research test programs. Due to several factors (facility research test schedule, facility downtime and continued facility upgrades), a full test section calibration was not conducted until 1996. This calibration test incorporated all test section configurations and covered the existing operating range of the facility. However, near the end of that test entry, two of the vortex generators mounted on the compressor exit tailcone failed causing minor damage to the honeycomb flow straightener. The vortex generators were removed from the facility and calibration testing was terminated. A follow-up test entry was conducted in 1997 in order to fully calibrate the facility without the effects of the vortex generators and to provide a complete calibration of the newly expanded low speed operating range. During the 1997 tunnel entry, all planned test points required for a complete test section calibration were obtained. This data set included detailed in-plane and axial flow field distributions for use in quantifying the test section flow quality.

  17. Blockage Testing in the NASA Glenn 225 Square Centimeter Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Sevier, Abigail; Davis, David; Schoenenberger, Mark

    2017-01-01

    A feasibility study is in progress at NASA Glenn Research Center to implement a magnetic suspension and balance system in the 225 sq cm Supersonic Wind Tunnel for the purpose of testing the dynamic stability of blunt bodies. An important area of investigation in this study was determining the optimum size of the model and the iron spherical core inside of it. In order to minimize the required magnetic field and thus the size of the magnetic suspension system, it was determined that the test model should be as large as possible. Blockage tests were conducted to determine the largest possible model that would allow for tunnel start at Mach 2, 2.5, and 3. Three different forebody model geometries were tested at different Mach numbers, axial locations in the tunnel, and in both a square and axisymmetric test section. Experimental results showed that different model geometries produced more varied results at higher Mach Numbers. It was also shown that testing closer to the nozzle allowed larger models to start compared with testing near the end of the test section. Finally, allowable model blockage was larger in the axisymmetric test section compared with the square test section at the same Mach number. This testing answered key questions posed by the feasibility study and will be used in the future to dictate model size and performance required from the magnetic suspension system.

  18. Supersonic Retropropulsion Test 1853 in NASA LaRC Unitary Plan Wind Tunnel Test Section 2

    NASA Technical Reports Server (NTRS)

    Berry, Scott A.; Rhode, Matthew N.

    2014-01-01

    A supersonic retropropulsion experiment was conducted in the Langley Research Center Unitary Plan Wind Tunnel Test Section 2 at Mach numbers of 2.4, 3.5, and 4.6. Intended as a code validation effort, this study used pretest computations to size and refine the model such that tunnel blockage and internal flow separations were minimized. A 5-in diameter 70 degree sphere-cone forebody, which can accommodate up to four 4:1 area ratio nozzles, followed by a 9.55 inches long cylindrical aft body was selected for this test after computational maturation. The primary measurements for this experiment were high spatial-density surface pressures. In addition, high speed schlieren video and internal pressures and temperatures were acquired. The test included parametric variations in the number of nozzles utilized, thrust coefficients (roughly 0 to 4), and angles of attack (-8 to 20 degrees). The run matrix was developed to also allow quantification of various sources of experimental uncertainty, such as random errors due to run-to-run variations and systematic errors due to flowfield or model misalignments. To accommodate the uncertainty assessment, many runs and replicates were conducted with the model at various locations within the tunnel and with model roll angles of 0, 60, 120, and 180 degrees. This test report provides operational details of the experiment, contains a review of trends, and provides all schlieren and pressure results within appendices.

  19. Estimation of Uncertainties for a Supersonic Retro-Propulsion Model Validation Experiment in a Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Rhode, Matthew N.; Oberkampf, William L.

    2012-01-01

    A high-quality model validation experiment was performed in the NASA Langley Research Center Unitary Plan Wind Tunnel to assess the predictive accuracy of computational fluid dynamics (CFD) models for a blunt-body supersonic retro-propulsion configuration at Mach numbers from 2.4 to 4.6. Static and fluctuating surface pressure data were acquired on a 5-inch-diameter test article with a forebody composed of a spherically-blunted, 70-degree half-angle cone and a cylindrical aft body. One non-powered configuration with a smooth outer mold line was tested as well as three different powered, forward-firing nozzle configurations: a centerline nozzle, three nozzles equally spaced around the forebody, and a combination with all four nozzles. A key objective of the experiment was the determination of experimental uncertainties from a range of sources such as random measurement error, flowfield non-uniformity, and model/instrumentation asymmetries. This paper discusses the design of the experiment towards capturing these uncertainties for the baseline non-powered configuration, the methodology utilized in quantifying the various sources of uncertainty, and examples of the uncertainties applied to non-powered and powered experimental results. The analysis showed that flowfield nonuniformity was the dominant contributor to the overall uncertainty a finding in agreement with other experiments that have quantified various sources of uncertainty.

  20. Supersonic Aftbody Closure Wind-Tunnel Testing, Data Analysis, and Computational Results

    NASA Technical Reports Server (NTRS)

    Allen, Jerry; Martin, Grant; Kubiatko, Paul

    1999-01-01

    This paper reports on the model, test, and results from the Langley Supersonic Aftbody Closure wind tunnel test. This project is an experimental evaluation of the 1.5% Technology Concept Aircraft (TCA) aftbody closure model (Model 23) in the Langley Unitary Plan Wind Tunnel. The baseline TCA design is the result of a multidisciplinary, multipoint optimization process and was developed using linear design and analysis methods, supplemented with Euler and Navier-Stokes numerical methods. After a thorough design review, it was decided to use an upswept blade attached to the forebody as the mounting system. Structural concerns dictated that a wingtip support system would not be feasible. Only the aftbody part of the model is metric. The metric break was chosen to be at the fuselage station where prior aft-sting supported models had been truncated. Model 23 is thus a modified version of Model 20. The wing strongback, flap parts, and nacelles from Model 20 were used, whereas new aftbodies, a common forebody, and some new tails were fabricated. In summary, significant differences in longitudinal and direction stability and control characteristics between the ABF and ABB aftbody geometries were measured. Correcting the experimental data obtained for the TCA configuration with the flared aftbody to the representative of the baseline TCA closed aftbody will result in a significant reduction in longitudinal stability, a moderate reduction in stabilizer effectiveness and directional stability, and a moderate to significant reduction in rudder effectiveness. These reductions in the stability and control effectiveness levels of the baseline TCA closed aftbody are attributed to the reduction in carry-over area.

  1. Investigation of Materials for Boundary Layer Control in a Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Braafladt, Alexander; Lucero, John M.; Hirt, Stefanie M.

    2013-01-01

    During operation of the NASA Glenn Research Center 15- by 15-Centimeter Supersonic Wind Tunnel (SWT), a significant, undesirable corner flow separation is created by the three-dimensional interaction of the wall and floor boundary layers in the tunnel corners following an oblique-shock/ boundary-layer interaction. A method to minimize this effect was conceived by connecting the wall and floor boundary layers with a radius of curvature in the corners. The results and observations of a trade study to determine the effectiveness of candidate materials for creating the radius of curvature in the SWT are presented. The experiments in the study focus on the formation of corner fillets of four different radii of curvature, 6.35 mm (0.25 in.), 9.525 mm (0.375 in.), 12.7 mm (0.5 in.), and 15.875 mm (0.625 in.), based on the observed boundary layer thickness of 11.43 mm (0.45 in.). Tests were performed on ten candidate materials to determine shrinkage, surface roughness, cure time, ease of application and removal, adhesion, eccentricity, formability, and repeatability. Of the ten materials, the four materials which exhibited characteristics most promising for effective use were the heavy body and regular type dental impression materials, the basic sculpting epoxy, and the polyurethane sealant. Of these, the particular material which was most effective, the heavy body dental impression material, was tested in the SWT in Mach 2 flow, and was observed to satisfy all requirements for use in creating the corner fillets in the upcoming experiments on shock-wave/boundary-layer interaction.

  2. Tone noise of three supersonic helical tip speed propellers in a wind tunnel at 0.8 Mach number

    NASA Technical Reports Server (NTRS)

    Dittmar, J. H.; Blaha, B. J.; Jeracki, R. J.

    1978-01-01

    Three supersonic helical tip speed propellers were tested in the NASA Lewis 8- by 6-foot wind tunnel. Noise data were obtained while these propellers were operating at a simulated cruise condition. The walls of this tunnel were not acoustically treated and therefore this was not an ideal location for taking noise data, but it was thought that the differences in noise among the three propellers would be meaningful. The straight bladed propeller which did not incorporate sweep was the noisiest with the aerodynamically swept propeller only slightly quieter. However, the acoustically swept propeller was significantly quieter than the straight propeller, thereby indicating the merit of this design technique.

  3. Effect of slot width on transition and noise attenuation of a sound shield panel for supersonic wind tunnels

    NASA Technical Reports Server (NTRS)

    Harvey, W. D.; Stainback, P. C.; Srokowski, A. J.

    1976-01-01

    The effect of slot width on transition and noise attenuation of a planar rod wall sound shield has been investigated at Mach 6. Boundary layer laminarization was obtained by induced suction through adjustable slots or gaps between the rods. Transition Reynolds numbers increased significantly with increasing gap width. The local mean flow field above the suction panel rod array was uniform. Important factors for design and performance of a supersonic wind-tunnel sound shield are rod diameter, gap-to-rod diameter ratio, surface finish of the rods, and leading-edge configuration.

  4. Plume and Shock Interaction Effects on Sonic Boom in the 1-foot by 1-foot Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Castner, Raymond; Elmiligui, Alaa; Cliff, Susan; Winski, Courtney

    2015-01-01

    The desire to reduce or eliminate the operational restrictions of supersonic aircraft over populated areas has led to extensive research at NASA. Restrictions are due to the disturbance of the sonic boom, caused by the coalescence of shock waves formed by the aircraft. A study has been performed focused on reducing the magnitude of the sonic boom N-wave generated by airplane components with a focus on shock waves caused by the exhaust nozzle plume. Testing was completed in the 1-foot by 1-foot supersonic wind tunnel to study the effects of an exhaust nozzle plume and shock wave interaction. The plume and shock interaction study was developed to collect data for computational fluid dynamics (CFD) validation of a nozzle plume passing through the shock generated from the wing or tail of a supersonic vehicle. The wing or tail was simulated with a wedgeshaped shock generator. This test entry was the first of two phases to collect schlieren images and off-body static pressure profiles. Three wedge configurations were tested consisting of strut-mounted wedges of 2.5- degrees and 5-degrees. Three propulsion configurations were tested simulating the propulsion pod and aft deck from a low boom vehicle concept, which also provided a trailing edge shock and plume interaction. Findings include how the interaction of the jet plume caused a thickening of the shock generated by the wedge (or aft deck) and demonstrate how the shock location moved with increasing nozzle pressure ratio.

  5. Propulsion and airframe aerodynamic interactions of supersonic V/STOL configurations. Volume 1: Wind tunnel test pressure data report

    NASA Technical Reports Server (NTRS)

    Zilz, D. E.; Devereaux, P. A.

    1985-01-01

    A wind tunnel model of a supersonic V/STOL fighter configuration has been tested to measure the aerodynamic interaction effects which can result from geometrically close-coupled propulsion system/airframe components. The approach was to configure the model to represent two different test techniques. One was a conventional test technique composed of two test modes. In the Flow-Through mode, absolute configuration aerodynamics are measured, including inlet/airframe interactions. In the Jet-Effects mode, incremental nozzle/airframe interactions are measured. The other test technique is a propulsion simulator approach, where a sub-scale, externally powered engine is mounted in the model. This allows proper measurement of inlet/airframe and nozzle/airframe interactions simultaneously. This is Volume 1 of 2: Wind Tunnel Test Pressure Data Report.

  6. FLPP IXV Re-Entry Vehicle, Supersonic Charectisation Based on DNW SST Wind Tunnel Tests and CFD

    NASA Astrophysics Data System (ADS)

    Kapteijn, C.; Maseland, H.; Chiarelli, C.; Mareschi, V.; Tribot, J.-P.; Binetti, P.; Walloscheck, T.

    2009-01-01

    measured for a number of discrete deflection settings in the Super-Sonic wind Tunnel (SST) of DNW. Enabling an improved understanding of the measured aerodynamic characteristics, complementary computations were performed by Thales Alenia Space. The complete set of data was analyzed and compared enabling a consolidation of the nominal aerodynamic and aerodynamic uncertainties as well. The paper presents the main objectives of the supersonic characterisation of IXV including WTTs, and the main outcomes of the current data comparisons.

  7. Wind tunnel investigation of the interaction and breakdown characteristics of slender wing vortices at subsonic, transonic, and supersonic speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    1991-01-01

    The vortex dominated aerodynamic characteristics of a generic 65 degree cropped delta wing model were studied in a wind tunnel at subsonic through supersonic speeds. The lee-side flow fields over the wing-alone configuration and the wing with leading edge extension (LEX) added were observed at M (infinity) equals 0.40 to 1.60 using a laser vapor screen technique. These results were correlated with surface streamline patterns, upper surface static pressure distributions, and six-component forces and moments. The wing-alone exhibited vortex breakdown and asymmetry of the breakdown location at the subsonic and transonic speeds. An earlier onset of vortex breakdown over the wing occurred at transonic speeds due to the interaction of the leading edge vortex with the normal shock wave. The development of a shock wave between the vortex and wing surface caused an early separation of the secondary boundary layer. With the LEX installed, wing vortex breakdown asymmetry did not occur up to the maximum angle of attack in the present test of 24 degrees. The favorable interaction of the LEX vortex with the wing flow field reduced the effects of shock waves on the wing primary and secondary vortical flows. The direct interaction of the wing and LEX vortex cores diminished with increasing Mach number. The maximum attainable vortex-induced pressure signatures were constrained by the vacuum pressure limit at the transonic and supersonic speeds.

  8. Pressure data for four analytically defined arrow wings in supersonic flow. [Langley Unitary Plan Wind Tunnel tests

    NASA Technical Reports Server (NTRS)

    Townsend, J. C.

    1980-01-01

    In order to provide experimental data for comparison with newly developed finite difference methods for computing supersonic flows over aircraft configurations, wind tunnel tests were conducted on four arrow wing models. The models were machined under numeric control to precisely duplicate analytically defined shapes. They were heavily instrumented with pressure orifices at several cross sections ahead of and in the region where there is a gap between the body and the wing trailing edge. The test Mach numbers were 2.36, 2.96, and 4.63. Tabulated pressure data for the complete test series are presented along with selected oil flow photographs. Comparisons of some preliminary numerical results at zero angle of attack show good to excellent agreement with the experimental pressure distributions.

  9. Some anomalies observed in wind-tunnel tests of a blunt body at transonic and supersonic speeds

    NASA Technical Reports Server (NTRS)

    Brooks, J. D.

    1976-01-01

    An investigation of anomalies observed in wind tunnel force tests of a blunt body configuration was conducted at Mach numbers from 0.20 to 1.35 in the Langley 8-foot transonic pressure tunnel and at Mach numbers of 1.50, 1,80, and 2.16 in the Langley Unitary Plan wind tunnel. At a Mach number of 1.35, large variations occurred in axial force coefficient at a given angle of attack. At transonic and low supersonic speeds, the total drag measured in the wind tunnel was much lower than that measured during earlier ballistic range tests. Accurate measurements of total drag for blunt bodies will require the use of models smaller than those tested thus far; however, it appears that accurate forebody drag results can be obtained by using relatively large models. Shock standoff distance is presented from experimental data over the Mach number range from 1.05 to 4.34. Theory accurately predicts the shock standoff distance at Mach numbers up to 1.75.

  10. An Overview of Preliminary Computational and Experimental Results for the Semi-Span Super-Sonic Transport (S4T) Wind-Tunnel Model

    NASA Technical Reports Server (NTRS)

    Silva, Walter A.; Perry, Boyd, III; Florance, James R.; Sanetrik, Mark D.; Wieseman, Carol D.; Stevens, William L.; Funk, Christie J.; Hur, Jiyoung; Christhilf, David M.; Coulson, David A.

    2011-01-01

    A summary of computational and experimental aeroelastic and aeroservoelastic (ASE) results for the Semi-Span Super-Sonic Transport (S4T) wind-tunnel model is presented. A broad range of analyses and multiple ASE wind-tunnel tests of the S4T have been performed in support of the ASE element in the Supersonics Program, part of NASA's Fundamental Aeronautics Program. The computational results to be presented include linear aeroelastic and ASE analyses, nonlinear aeroelastic analyses using an aeroelastic CFD code, and rapid aeroelastic analyses using CFD-based reduced-order models (ROMs). Experimental results from two closed-loop wind-tunnel tests performed at NASA Langley's Transonic Dynamics Tunnel (TDT) will be presented as well.

  11. Wind-tunnel/flight correlation study of aerodynamic characteristics of a large flexible supersonic cruise airplane (XB-701) 2: Extrapolation of wind-tunnel data to full-scale conditions

    NASA Technical Reports Server (NTRS)

    Peterson, J. B., Jr.; Mann, M. J.; Sorrells, R. B., III; Sawyer, W. C.; Fuller, D. E.

    1980-01-01

    The results of calculations necessary to extrapolate performance data on an XB-70-1 wind tunnel model to full scale at Mach numbers from 0.76 to 2.53 are presented. The extrapolation was part of a joint program to evaluate performance prediction techniques for large flexible supersonic airplanes similar to a supersonic transport. The extrapolation procedure included: interpolation of the wind tunnel data at the specific conditions of the flight test points; determination of the drag increments to be applied to the wind tunnel data, such as spillage drag, boundary layer trip drag, and skin friction increments; and estimates of the drag items not represented on the wind tunnel model, such as bypass doors, roughness, protuberances, and leakage drag. In addition, estimates of the effects of flexibility of the airplane were determined.

  12. Interaction of aerodynamic noise with laminar boundary layers in supersonic wind tunnels

    NASA Technical Reports Server (NTRS)

    Schopper, M. R.

    1984-01-01

    The interaction between incoming aerodynamic noise and the supersonic laminar boundary layer is studied. The noise field is modeled as a Mach wave radiation field consisting of discrete waves emanating from coherent turbulent entities moving downstream within the supersonic turbulent boundary layer. The individual disturbances are likened to miniature sonic booms and the laminar boundary layer is staffed by the waves as the sources move downstream. The mean, autocorrelation, and power spectral density of the field are expressed in terms of the wave shapes and their average arrival rates. Some consideration is given to the possible appreciable thickness of the weak shock fronts. The emphasis in the interaction analysis is on the behavior of the shocklets in the noise field. The shocklets are shown to be focused by the laminar boundary layer in its outer region. Borrowing wave propagation terminology, this region is termed the caustic region. Using scaling laws from sonic boom work, focus factors at the caustic are estimated to vary from 2 to 6 for incoming shocklet strengths of 1 to .01 percent of the free stream pressure level. The situation regarding experimental evidence of the caustic region is reviewed.

  13. Boundary Layer Measurements in a Supersonic Wind Tunnel Using Doppler Global Velocimetry

    NASA Technical Reports Server (NTRS)

    Meyers, James F.; Lee, Joseph W.; Cavone, Angelo A.

    2010-01-01

    A modified Doppler Global Velocimeter (DGV) was developed to measure the velocity within the boundary layer above a flat plate in a supersonic flow. Classic laser velocimetry (LV) approaches could not be used since the model surface was composed of a glass-ceramic insulator in support of heat-transfer measurements. Since surface flare limited the use of external LV techniques and windows placed in the model would change the heat transfer characteristics of the flat plate, a novel approach was developed. The input laser beam was divided into nine equal power beams and each transmitted through optical fibers to a small cavity within the model. The beams were then directed through 1.6-mm diameter orifices to form a series of orthogonal beams emitted from the model and aligned with the tunnel centerline to approximate a laser light sheet. Scattered light from 0.1-micron diameter water condensation ice crystals was collected by four 5-mm diameter lenses and transmitted by their respective optical fiber bundles to terminate at the image plane of a standard two-camera DGV receiver. Flow measurements were made over a range from 0.5-mm above the surface to the freestream at Mach 3.51 in steady state and heat pulse injected flows. This technique provides a unique option for measuring boundary layers in supersonic flows where seeding the flow is problematic or where the experimental apparatus does not provide the optical access required by other techniques.

  14. Characteristics of the NASA-Ames Laminar Flow Supersonic Wind Tunnel for Unique Mach 1.6 Transition Studies

    NASA Technical Reports Server (NTRS)

    Wolf, Stephen W. D.; Laub, James A.

    1997-01-01

    Flow quality measurements have been performed in the unique Laminar Flow Supersonic Wind Tunnel (LFSWT) to examine both mean and dynamic characteristics. The intent was to provide the necessary flow information about this ground test facility, to support meaningful transition research at Mach 1.6 and flight unit Reynolds numbers. This paper is intended to assist other experimentalists with similar goals of characterizing low-supersonic test environments. An array of instrumentation has been used to highlight the importance of proper selection of pressure instruments and data acquisition procedures. We conclude that the test section is low-disturbance (based on classical standards of pressure disturbances less than 0.1% with no specified data bandwidth), and has uniform flow. This is confirmation that the quiet design features of the LFSWT are effective. However, characterization of the test section flow over a 0.25k-5Ok bandwidth shows that the disturbance levels can be greater than classical standards particularly for stagnation pressures less than 9.5 psia (0.65 bar) with low stagnation temperatures. Variability of the flow disturbances in the settling chamber and test section is contained in a narrow frequency bandwidth below 5k Hz, which is associated with resonant frequencies from the pressure reduction system. So far, these disturbances have not impacted transition along the tunnel walls or a 10 degrees cone. However, continual vigilance is required to maintain a known low-disturbance environment for transition research in the LFSWT. Furthermore, the formation of standards for flow quality measurements is strongly recommended, so that transition research can be better isolated from tunnel disturbances.

  15. Demonstration of Clean Particle Seeding for Particle Image Velocimetry in a Closed Circuit Supersonic Wind Tunnel

    DTIC Science & Technology

    2007-03-01

    siphon drawing liquid CO2 through an injector. As the CO2 exits the injector, the 3 pressure and temperature reduction from the tank to the wind...blowdown system consists of two compressors with electronic driers that supply dry air pressurized to 145 psig into a 6,000 gallon tank , as well as a...6,000 gallon vacuum tank that is evacuated by a single vacuum pump. This vacuum assist reduces the pressure required to operate the wind tunnel and

  16. An efficient supersonic wind tunnel drive system for Mach 2.5 flows

    NASA Technical Reports Server (NTRS)

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

    1991-01-01

    A novel efficient drive system has been developed which provides for the continuous operation of a pitot Mach 2.5 wind tunnel at compression ratios down to 0.625:1. The drive system does not require an overpressure to start, and no hysteresis has been observed. The general design of the proof-of-concept wind tunnel using the new drive system and its modifications are described.

  17. Feasibility Study for Implementing Magnetic Suspension in the Glenn Research Center 225 cm2 Supersonic Wind Tunnel for Testing the Dynamic Stability of Blunt Bodies

    NASA Technical Reports Server (NTRS)

    Sevier, Abigail; Davis, David O.; Schoenenberger, Mark; Barnhart, Paul

    2016-01-01

    The implementation of a magnetic suspension system in the NASA Glenn Research Center (GRC) 225 cm2 Supersonic Wind Tunnel would be a powerful test technique that could accurately determine the dynamic stability of blunt body entry vehicles with no sting interference. This paper explores initial design challenges to be evaluated before implementation, including defining the lowest possible operating dynamic pressure and corresponding model size, developing a compatible video analysis technique, and incorporating a retractable initial support sting.

  18. 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.

  19. Characteristics of Control Laws Tested on the Semi-Span Super-Sonic Transport (S4T) Wind-Tunnel Model

    NASA Technical Reports Server (NTRS)

    Christhilf, David M.; Moulin, Boris; Ritz, Erich; Chen, P. C.; Roughen, Kevin M.; Perry, Boyd

    2012-01-01

    The Semi-Span Supersonic Transport (S4T) is an aeroelastically scaled wind-tunnel model built to test active controls concepts for large flexible supersonic aircraft in the transonic flight regime. It is one of several models constructed in the 1990's as part of the High Speed Research (HSR) Program. Control laws were developed for the S4T by M4 Engineering, Inc. and by Zona Technologies, Inc. under NASA Research Announcement (NRA) contracts. The model was tested in the NASA-Langley Transonic Dynamics Tunnel (TDT) four times from 2007 to 2010. The first two tests were primarily for plant identification. The third entry was used for testing control laws for Ride Quality Enhancement, Gust Load Alleviation, and Flutter Suppression. Whereas the third entry only tested FS subcritically, the fourth test demonstrated closed-loop operation above the open-loop flutter boundary. The results of the third entry are reported elsewhere. This paper reports on flutter suppression results from the fourth wind-tunnel test. Flutter suppression is seen as a way to provide stability margins while flying at transonic flight conditions without penalizing the primary supersonic cruise design condition. An account is given for how Controller Performance Evaluation (CPE) singular value plots were interpreted with regard to progressing open- or closed-loop to higher dynamic pressures during testing.

  20. 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.

  1. Design and preliminary test results at Mach 5 of an axisymmetric slotted sound shield. [for supersonic wind tunnels (noise reduction in wind tunnel nozzles)

    NASA Technical Reports Server (NTRS)

    Beckwith, I. E.; Spokowski, A. J.; Harvey, W. D.; Stainback, P. C.

    1975-01-01

    The basic theory and sound attenuation mechanisms, the design procedures, and preliminary experimental results are presented for a small axisymmetric sound shield for supersonic wind tunnels. The shield consists of an array of small diameter rods aligned nearly parallel to the entrance flow with small gaps between the rods for boundary layer suction. Results show that at the lowest test Reynolds number (based on rod diameter) of 52,000 the noise shield reduced the test section noise by about 60 percent ( or 8 db attenuation) but no attenuation was measured for the higher range of test reynolds numbers from 73,000 to 190,000. These results are below expectations based on data reported elsewhere on a flat sound shield model. The smaller attenuation from the present tests is attributed to insufficient suction at the gaps to prevent feedback of vacuum manifold noise into the shielded test flow and to insufficient suction to prevent transition of the rod boundary layers to turbulent flow at the higher Reynolds numbers. Schlieren photographs of the flow are shown.

  2. 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.

  3. CNSFV code development, virtual zone Navier-Stokes computations of oscillating control surfaces and computational support of the laminar flow supersonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Klopfer, Goetz H.

    1993-01-01

    The work performed during the past year on this cooperative agreement covered two major areas and two lesser ones. The two major items included further development and validation of the Compressible Navier-Stokes Finite Volume (CNSFV) code and providing computational support for the Laminar Flow Supersonic Wind Tunnel (LFSWT). The two lesser items involve a Navier-Stokes simulation of an oscillating control surface at transonic speeds and improving the basic algorithm used in the CNSFV code for faster convergence rates and more robustness. The work done in all four areas is in support of the High Speed Research Program at NASA Ames Research Center.

  4. Model aerodynamic test results for two variable cycle engine coannular exhaust systems at simulated takeoff and cruise conditions. [Lewis 8 by 6-foot supersonic wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Nelson, D. P.

    1980-01-01

    Wind tunnel tests were conducted to evaluate the aerodynamic performance of a coannular exhaust nozzle for a proposed variable stream control supersonic propulsion system. Tests were conducted with two simulated configurations differing primarily in the fan duct flowpaths: a short flap mechanism for fan stream control with an isentropic contoured flow splitter, and an iris fan nozzle with a conical flow splitter. Both designs feature a translating primary plug and an auxiliary inlet ejector. Tests were conducted at takeoff and simulated cruise conditions. Data were acquired at Mach numbers of 0, 0.36, 0.9, and 2.0 for a wide range of nozzle operating conditions. At simulated supersonic cruise, both configurations demonstrated good performance, comparable to levels assumed in earlier advanced supersonic propulsion studies. However, at subsonic cruise, both configurations exhibited performance that was 6 to 7.5 percent less than the study assumptions. At take off conditions, the iris configuration performance approached the assumed levels, while the short flap design was 4 to 6 percent less.

  5. Reduction of acoustic disturbances in the test section of supersonic wind tunnels by laminarizing their nozzle and test section wall boundary layers by means of suction

    NASA Technical Reports Server (NTRS)

    Pfenninger, W.; Syberg, J.

    1974-01-01

    The feasibility of quiet, suction laminarized, high Reynolds number (Re) supersonic wind tunnel nozzles was studied. According to nozzle wall boundary layer development and stability studies, relatively weak area suction can prevent amplified nozzle wall TS (Tollmien-Schlichting) boundary layer oscillations. Stronger suction is needed in and shortly upstream of the supersonic concave curvature nozzle area to avoid transition due to amplified TG (Taylor-Goertler) vortices. To control TG instability, moderately rapid and slow expansion nozzles require smaller total suction rates than rapid expansion nozzles, at the cost of larger nozzle length Re and increased TS disturbances. Test section mean flow irregularities can be minimized with suction through longitudinal or highly swept slots (swept behind local Mach cone) as well as finely perforated surfaces. Longitudinal slot suction is optimized when the suction-induced crossflow velocity increases linearly with surface distance from the slot attachment line toward the slot (through suitable slot geometry). Suction in supersonic blowdown tunnels may be operated by one or several individual vacuum spheres.

  6. NASA Ames Laminar Flow Supersonic Wind Tunnel (LFSWT) Tests of a 10 deg Cone at Mach 1.6

    NASA Technical Reports Server (NTRS)

    Wolf, Stephen W. D.; Laub, James A.

    1997-01-01

    This work is part of the ongoing qualification of the NASA Ames Laminar Flow Supersonic Wind Tunnel (LFSWT) as a low-disturbance (quiet) facility suitable for transition research. A 10 deg cone was tested over a range of unit Reynolds numbers (Re = 2.8 to 3.8 million per foot (9.2 to 12.5 million per meter)) and angles of incidence (O deg to 10 deg) at Mach 1.6. The location of boundary layer transition along the cone was measured primarily from surface temperature distributions, with oil flow interferometry and Schlieren flow visualization providing confirmation measurements. With the LFSWT in its normal quiet operating mode, no transition was detected on the cone in the test core, over the Reynolds number range tested at zero incidence and yaw. Increasing the pressure disturbance levels in the LFSWT test section by a factor of five caused transition onset on the cone within the test core, at zero incidence and yaw. When operating the LFSWT in its normal quiet mode, transition could only be detected in the test core when high angles of incidence (greater than 5 deg) for cones were set. Transition due to elevated pressure disturbances (Tollmien-Schlichting) and surface trips produced a skin temperature rise of order 4 F (2.2 C). Transition due to cross flows on the leeward side of the cone at incidence produced a smaller initial temperature rise of only order 2.5 F (1.4 C), which indicates a slower transition process. We can conclude that these cone tests add further proof that the LFSWT test core is normally low-disturbance (pressure fluctuations greater than 0.1%), as found by associated direct flow quality measurements discussed in this report. Furthermore, in a quiet test environment, the skin temperature rise is sensitive to the type of dominant instability causing transition. The testing of a cone in the LFSWT provides an excellent experiment for the development of advanced transition detection techniques.

  7. Propulsion and airframe aerodynamic interactions of supersonic V/STOL configurations. Volume 2: Wind tunnel test force and moment data report

    NASA Technical Reports Server (NTRS)

    Zilz, D. E.

    1985-01-01

    A wind tunnel model of a supersonic V/STOL fighter configuration has been tested to measure the aerodynamic interaction effects which can result from geometrically close-coupled propulsion system/airframe components. The approach was to configure the model to represent two different test techniques. One was a conventional test technique composed of two test modes. In the Flow-Through mode, absolute configuration aerodynamics are measured, including inlet/airframe interactions. In the Jet-Effects mode, incremental nozzle/airframe interactions are measured. The other test technique is a propulsion simulator approach, where a sub-scale, externally powered engine is mounted in the model. This allows proper measurement of inlet/airframe and nozzle/airframe interactions simultaneously. This is Volume 2 of 2: Wind Tunnel Test Force and Moment Data Report.

  8. Aerodynamic characteristics of a supersonic cruise airplane configuration at Mach numbers of 2.30, 2.96, and 3.30. [Langley Unitary Plan wind tunnel test

    NASA Technical Reports Server (NTRS)

    Shrout, B. L.; Fournier, R. H.

    1979-01-01

    An investigation was made in the Langley Unitary Plan wind tunnel at Mach numbers of 2.30, 2.96, and 3.30 to determine the static longitudinal and lateral aerodynamic characteristics of a model of a supersonic cruise airplane. The configuration, with a design Mach number of 3.0, has a highly swept arrow wing with tip panels of lesser sweep, a fuselage chine, outboard vertical tails, and outboard engines mounted in nacelles beneath the wings. For wind tunnel test conditions, a trimmed value above 6.0 of the maximum lift-drag ratio was obtained at the design Mach number. The configuration was statically stable, both longitudinally and laterally. Data are presented for variations of vertical-tail roll-out and toe-in and for various combinations of components. Some roll control data are shown as are data for the various sand grit sizes used in fixing the boundary layer transition location.

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

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

    A nonintrusive airdata-sensing system was calibrated in flight and wind-tunnel experiments to an angle of attack of 70 deg and to angles of sideslip of +/- 15 deg. Flight-calibration data have also been obtained to Mach 1.2. The sensor, known as the flush airdata sensor, was installed on the nosecap of an F-18 aircraft for flight tests and on a full-scale F-18 forebody for wind-tunnel tests. Flight tests occurred at the NASA Dryden Flight Research Facility, Edwards, California, using the F-18 High Alpha Research Vehicle. Wind-tunnel tests were conducted in the 30- by 60-ft wind tunnel at the NASA LaRC, Hampton, Virginia. The sensor consisted of 23 flush-mounted pressure ports arranged in concentric circles and located within 1.75 in. of the tip of the nosecap. An overdetermined mathematical model was used to relate the pressure measurements to the local airdata quantities. The mathematical model was based on potential flow over a sphere and was empirically adjusted based on flight and wind-tunnel data. For quasi-steady maneuvering, the mathematical model worked well throughout the subsonic, transonic, and low supersonic flight regimes. The model also worked well throughout the angles-of-attack and -sideslip regions studied.

  10. Aeroelastic Analyses of the SemiSpan SuperSonic Transport (S4T) Wind Tunnel Model at Mach 0.95

    NASA Technical Reports Server (NTRS)

    Hur, Jiyoung

    2014-01-01

    Detailed aeroelastic analyses of the SemiSpan SuperSonic Transport (S4T) wind tunnel model at Mach 0.95 with a 1.75deg fixed angle of attack are presented. First, a numerical procedure using the Computational Fluids Laboratory 3-Dimensional (CFL3D) Version 6.4 flow solver is investigated. The mesh update method for structured multi-block grids was successfully applied to the Navier-Stokes simulations. Second, the steady aerodynamic analyses with a rigid structure of the S4T wind tunnel model are reviewed in transonic flow. Third, the static analyses were performed for both the Euler and Navier-Stokes equations. Both the Euler and Navier-Stokes equations predicted a significant increase of lift forces, compared to the results from the rigid structure of the S4T wind-tunnel model, over various dynamic pressures. Finally, dynamic aeroelastic analyses were performed to investigate the flutter condition of the S4T wind tunnel model at the transonic Mach number. The condition of flutter was observed at a dynamic pressure of approximately 75.0-psf for the Navier-Stokes simulations. However, it was observed that the flutter condition occurred a dynamic pressure of approximately 47.27-psf for the Euler simulations. Also, the computational efficiency of the aeroelastic analyses for the S4T wind tunnel model has been assessed.

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

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

    A nonintrusive airdata-sensing system was calibrated in flight and wind-tunnel experiments to an angle of attack of 70 deg and to angles of sideslip of +/- 15 deg. Flight-calibration data have also been obtained to Mach 1.2. The sensor, known as the flush airdata sensor, was installed on the nosecap of an F-18 aircraft for flight tests and on a full-scale F-18 forebody for wind-tunnel tests. Flight tests occurred at the NASA Dryden Flight Research Facility, Edwards, California, using the F-18 High Alpha Research Vehicle. Wind-tunnel tests were conducted in the 30- by 60-ft wind tunnel at the NASA LaRC, Hampton, Virginia. The sensor consisted of 23 flush-mounted pressure ports arranged in concentric circles and located within 1.75 in. of the tip of the nosecap. An overdetermined mathematical model was used to relate the pressure measurements to the local airdata quantities. The mathematical model was based on potential flow over a sphere and was empirically adjusted based on flight and wind-tunnel data. For quasi-steady maneuvering, the mathematical model worked well throughout the subsonic, transonic, and low supersonic flight regimes. The model also worked well throughout the angle-of-attack and sideslip regions studied.

  12. 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

  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. Visualizing Flutter Mechanism as Traveling Wave Through Animation of Simulation Results for the Semi-Span Super-Sonic Transport Wind-Tunnel Model

    NASA Technical Reports Server (NTRS)

    Christhilf, David M.

    2014-01-01

    It has long been recognized that frequency and phasing of structural modes in the presence of airflow play a fundamental role in the occurrence of flutter. Animation of simulation results for the long, slender Semi-Span Super-Sonic Transport (S4T) wind-tunnel model demonstrates that, for the case of mass-ballasted nacelles, the flutter mode can be described as a traveling wave propagating downstream. Such a characterization provides certain insights, such as (1) describing the means by which energy is transferred from the airflow to the structure, (2) identifying airspeed as an upper limit for speed of wave propagation, (3) providing an interpretation for a companion mode that coalesces in frequency with the flutter mode but becomes very well damped, (4) providing an explanation for bursts of response to uniform turbulence, and (5) providing an explanation for loss of low frequency (lead) phase margin with increases in dynamic pressure (at constant Mach number) for feedback systems that use sensors located upstream from active control surfaces. Results from simulation animation, simplified modeling, and wind-tunnel testing are presented for comparison. The simulation animation was generated using double time-integration in Simulink of vertical accelerometer signals distributed over wing and fuselage, along with time histories for actuated control surfaces. Crossing points for a zero-elevation reference plane were tracked along a network of lines connecting the accelerometer locations. Accelerometer signals were used in preference to modal displacement state variables in anticipation that the technique could be used to animate motion of the actual wind-tunnel model using data acquired during testing. Double integration of wind-tunnel accelerometer signals introduced severe drift even with removal of both position and rate biases such that the technique does not currently work. Using wind-tunnel data to drive a Kalman filter based upon fitting coefficients to

  16. Multi-angular Flame Measurements and Analysis in a Supersonic Wind Tunnel Using Fiber-Based Endoscopes

    DTIC Science & Technology

    2016-09-14

    information for model validation and design optimization of propulsion systems. However, it is challenging to obtain such measurements due to various...1.4031306] Introduction Cavity-based flameholders are commonly used for hydrocarbon -fueled, supersonic, ramjet (scramjet) combustors [1]. This is due...operating continuously with peak stagnation conditions of 2860 kPa and 922 K at flow rates up to 15.4 kg/s [2,26]. For the current experiments, a generic

  17. An experimental wind-tunnel investigation of a ram-air-spoiler roll-control device on a forward-control missile at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Blair, A. B., Jr.

    1978-01-01

    A parametric experimental wind-tunnel investigation was made at supersonic Mach numbers to provide design data on a ram-air-spoiler roll-control device that is to be used on forward-control cruciform missile configurations. The results indicate that the ram-air-spoiler tail fin is an effective roll-control device and that roll control is generally constant with vehicle attitude and Mach number unless direct canard and/or forebody shock impingement occurs. The addition of the ram-air-spoiler tail fins resulted in only small changes in aerodynamic-center location. For the ram-air-spoiler configuration tested, there are large axial force coefficient effects associated with the increased fin thickness and ram-air momentum loss.

  18. Analytical and Experimental Evaluation of Digital Control Systems for the Semi-Span Super-Sonic Transport (S4T) Wind Tunnel Model

    NASA Technical Reports Server (NTRS)

    Wieseman, Carol D.; Christhilf, David; Perry, Boyd, III

    2012-01-01

    An important objective of the Semi-Span Super-Sonic Transport (S4T) wind tunnel model program was the demonstration of Flutter Suppression (FS), Gust Load Alleviation (GLA), and Ride Quality Enhancement (RQE). It was critical to evaluate the stability and robustness of these control laws analytically before testing them and experimentally while testing them to ensure safety of the model and the wind tunnel. MATLAB based software was applied to evaluate the performance of closed-loop systems in terms of stability and robustness. Existing software tools were extended to use analytical representations of the S4T and the control laws to analyze and evaluate the control laws prior to testing. Lessons were learned about the complex windtunnel model and experimental testing. The open-loop flutter boundary was determined from the closed-loop systems. A MATLAB/Simulink Simulation developed under the program is available for future work to improve the CPE process. This paper is one of a series of that comprise a special session, which summarizes the S4T wind-tunnel program.

  19. Repair of Corrosion in Air Supply Piping at the NASA Glenn Research Center's 1 by 1 Foot Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Henry, Michael

    2000-01-01

    During a test at the NASA Glenn Research Center's 1 x 1 Supersonic Wing Tunnel, it was discovered that particles entrained in the air flow were damaging the pressure sensitive paint on a test article. An investigation found the source of the entrained particles to be rust on the internal surfaces of the air supply piping. To remedy the situation, the air supply line components made from carbon steel were either refurbished or replaced with new stainless steel components. The refurbishment process included various combinations of chemical cleaning, bead blasting, painting and plating.

  20. 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.

  1. 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

  2. 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.

  3. 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.

  4. Effect of rod gap spacing on a suction panel for laminar flow and noise control in supersonic wind tunnels. M.S. Thesis - Old Dominion Univ.

    NASA Technical Reports Server (NTRS)

    Harvey, W. D.

    1975-01-01

    Results are presented of a coordinated experimental and theoretical study of a sound shield concept which aims to provide a means of noise reduction in the test section of supersonic wind tunnels at high Reynolds numbers. The model used consists of a planar array of circular rods aligned with the flow, with adjustable gaps between them for boundary layer removal by suction, i.e., laminar flow control. One of the basic requirements of the present sound shield concept is to achieve sonic cross flow through the gaps in order to prevent lee-side flow disturbances from penetrating back into the shielded region. Tests were conducted at Mach 6 over a local unit Reynolds number range from about 1.2 x 10 to the 6th power to 13.5 x 10 to the 6th power per foot. Measurements of heat transfer, static pressure, and sound levels were made to establish the transition characteristics of the boundary layer on the rod array and the sound shielding effectiveness.

  5. Modification of the Ames 40- by 80-foot wind tunnel for component acoustic testing for the second generation supersonic transport

    NASA Technical Reports Server (NTRS)

    Schmitz, F. H.; Allmen, J. R.; Soderman, P. T.

    1994-01-01

    The development of a large-scale anechoic test facility where large models of engine/airframe/high-lift systems can be tested for both improved noise reduction and minimum performance degradation is described. The facility development is part of the effort to investigate economically viable methods of reducing second generation high speed civil transport noise during takeoff and climb-out that is now under way in the United States. This new capability will be achieved through acoustic modifications of NASA's second largest subsonic wind tunnel: the 40-by 80-Foot Wind Tunnel at the NASA Ames Research Center. Three major items are addressed in the design of this large anechoic and quiet wind tunnel: a new deep (42 inch (107 cm)) test section liner, expansion of the wind tunnel drive operating envelope at low rpm to reduce background noise, and other promising methods of improving signal-to-noise levels of inflow microphones. Current testing plans supporting the U.S. high speed civil transport program are also outlined.

  6. Wind Tunnel Model Design for Sonic Boom Studies of Nozzle Jet Flows with Shock Interactions

    NASA Technical Reports Server (NTRS)

    Cliff, Susan E.; Denison, Marie; Moini-Yekta, Shayan; Morr, Donald E.; Durston, Donald A.

    2016-01-01

    NASA and the U.S. aerospace industry are performing studies of supersonic aircraft concepts with low sonic boom pressure signatures. The computational analyses of modern aircraft designs have matured to the point where there is confidence in the prediction of the pressure signature from the front of the vehicle, but uncertainty remains in the aft signatures due to boundary layer and nozzle exhaust jet effects. Wind tunnel testing without inlet and nozzle exhaust jet effects at lower Reynolds numbers than in-flight make it difficult to accurately assess the computational solutions of flight vehicles. A wind tunnel test in the NASA Ames 9- by 7-Foot Supersonic Wind Tunnel is planned for February 2016 to address the nozzle jet effects on sonic boom. The experiment will provide pressure signatures of test articles that replicate waveforms from aircraft wings, tails, and aft fuselage (deck) components after passing through cold nozzle jet plumes. The data will provide a variety of nozzle plume and shock interactions for comparison with computational results. A large number of high-fidelity numerical simulations of a variety of shock generators were evaluated to define a reduced collection of suitable test models. The computational results of the candidate wind tunnel test models as they evolved are summarized, and pre-test computations of the final designs are provided.

  7. Wind Tunnel Investigation of the Effects of Surface Porosity and Vertical Tail Placement on Slender Wing Vortex Flow Aerodynamics at Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    2007-01-01

    A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) Unitary Plan Wind Tunnel (UPWT) to determine the effects of passive surface porosity and vertical tail placement on vortex flow development and interactions about a general research fighter configuration at supersonic speeds. Optical flow measurement and flow visualization techniques were used that featured pressure sensitive paint (PSP), laser vapor screen (LVS), and schlieren, These techniques were combined with conventional electronically-scanned pressure (ESP) and six-component force and moment measurements to quantify and to visualize the effects of flow-through porosity applied to a wing leading edge extension (LEX) and the placement of centerline and twin vertical tails on the vortex-dominated flow field of a 65 cropped delta wing model. Test results were obtained at free-stream Mach numbers of 1.6, 1.8, and 2.1 and a Reynolds number per foot of 2.0 million. LEX porosity promoted a wing vortex-dominated flow field as a result of a diffusion and weakening of the LEX vortex. The redistribution of the vortex-induced suction pressures contributed to large nose-down pitching moment increments but did not significantly affect the vortex-induced lift. The trends associated with LEX porosity were unaffected by vertical tail placement. The centerline tail configuration generally provided more stable rolling moments and yawing moments compared to the twin wing-mounted vertical tails. The strength of a complex system of shock waves between the twin tails was reduced by LEX porosity.

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

    NASA Technical Reports Server (NTRS)

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

    2012-01-01

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

  9. A Wind-tunnel Test Technique for Measuring the Dynamic Rotary Stability Derivatives at Subsonic and Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Beam, Benjamin H

    1956-01-01

    A method is described for measuring the dynamic stability derivatives of a model airplane in a wind tunnel. The characteristic features of this system are that single-degree-of-freedom oscillations were used to obtain combinations of rolling, yawing and pitching motions; that the oscillations were excited and controlled by velocity feedback which permitted operation under conditions unfavorable for more conventional types of oscillatory testing; and that data processing was greatly simplified by using analog computer elements in the strain-gage circuitry. A small number of experimental data are included to illustrate the general scope of results obtainable with this system.

  10. Supersonic Wind Tunnel Tests of a Half-axisymmetric 12 Deg-spike Inlet to a Rocket-based Combined-cycle Propulsion System

    NASA Technical Reports Server (NTRS)

    DeBonis, J. R.; Trefny, C. J.

    2001-01-01

    Results of an isolated inlet test for NASA's GTX air-breathing launch vehicle concept are presented. The GTX is a Vertical Take-off/ Horizontal Landing reusable single-stage-to-orbit system powered by a rocket-based combined-cycle propulsion system. Tests were conducted in the NASA Glenn 1- by 1-Foot Supersonic Wind Tunnel during two entries in October 1998 and February 1999. Tests were run from Mach 2.8 to 6. Integrated performance parameters and static pressure distributions are reported. The maximum contraction ratios achieved in the tests were lower than predicted by axisymmetric Reynolds-averaged Navier-Stokes computational fluid dynamics (CFD). At Mach 6, the maximum contraction ratio was roughly one-half of the CFD value of 16. The addition of either boundary-layer trip strips or vortex generators had a negligible effect on the maximum contraction ratio. A shock boundary-layer interaction was also evident on the end-walls that terminate the annular flowpath cross section. Cut-back end-walls, designed to reduce the boundary-layer growth upstream of the shock and minimize the interaction, also had negligible effect on the maximum contraction ratio. Both the excessive turning of low-momentum comer flows and local over-contraction due to asymmetric end-walls were identified as possible reasons for the discrepancy between the CFD predictions and the experiment. It is recommended that the centerbody spike and throat angles be reduced in order to lessen the induced pressure rise. The addition of a step on the cowl surface, and planar end-walls more closely approximating a plane of symmetry are also recommended. Provisions for end-wall boundary-layer bleed should be incorporated.

  11. Transonic and Supersonic Wind-Tunnel Tests of Wing-Body Combinations Designed for High Efficiency at a Mach Number of 1.41

    NASA Technical Reports Server (NTRS)

    Grant, Frederick C.; Sevier, John R., Jr.

    1960-01-01

    Wind-tunnel force tests of a number of wing-body combinations designed for high lift-drag ratio at a Mach number of 1.41 are reported. Five wings and six bodies were used in making up the various wing-body combinations investigated. All the wings had the same highly swept dis- continuously tapered plan form with NACA 65A-series airfoil sections 4 percent thick at the root tapering linearly to 3 percent thick at the tip. The bodies were based on the area distribution of a Sears-Haack body of revolution for minimum drag with a given length and volume. These wings and bodies were used to determine the effects of wing twist., wing twist and camber, wing leading-edge droop, a change from circular to elliptical body cross-sectional shape, and body indentation by the area-rule and streamline methods. The supersonic test Mach numbers were 1.41 and 2.01. The transonic test Mach number range was from 0.6 to 1.2. For the transition-fixed condition and at a Reynolds number of 2.7 x 10(exp 6) based on the mean aerodynamic chord, the maximum value of lift- drag ratio at a Mach number of 1.41 was 9.6 for a combination with a twisted wing and an indented body of elliptical cross section. The tests indicated that the transonic rise in minimum drag was low and did not change appreciably up to the highest test Mach number of 2.01. The lower values of lift-drag ratio obtained at a Mach number of 2.01 can be attributed to the increase of drag due to lift with Mach number.

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

    NASA Technical Reports Server (NTRS)

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

    1980-01-01

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

  14. Wind tunnel test of the 0.015-scale Rockwell International space shuttle vehicle orbiter in the Ames 6 by 6 foot supersonic wind tunnel. [to determine longitudinal and lateral-directional characteristics

    NASA Technical Reports Server (NTRS)

    Milam, M. D.; Dziubala, T. J.

    1973-01-01

    Experimental investigations were performed in a 6- by 6-Foot Supersonic wind tunnel on a 0.015-scale model of the Rockwell International space shuttle vehicle (SSV) 2A orbiter. The purpose of the test was to investigate the longitudinal and lateral-directional characteristics of the vehicle. In addition, hinge moments were measured on the rudder and elevons. Buffet onset was investigated using wing trailing edge pressures and a strain gauge instrumented panel mounted in the wing. The model was tested through a Mach range from 0.6 to 2.0 at a constant unit Reynolds number of 2.5 million. Pitch runs were made at angles of attack from minus 2 deg to +26 deg with beta = 0 deg and 5 deg; yaw runs were made in the range from minus 5 deg to 10 deg of sideslip at angles of attack of 0 deg and 10 deg. Static pressures were measured at the fuselage base and the trailing edges of the wing and rudder. Boundary layer transition was fixed for some runs using distributed roughness strips.

  15. Wind Tunnel Model Design for Sonic Boom Studies of Nozzle Jet with Shock Interactions

    NASA Technical Reports Server (NTRS)

    Cliff, Susan E.; Denison, Marie; Sozer, Emre; Moini-Yekta, Shayan

    2016-01-01

    NASA and Industry are performing vehicle studies of configurations with low sonic boom pressure signatures. The computational analyses of modern configuration designs have matured to the point where there is confidence in the prediction of the pressure signature from the front of the vehicle, but uncertainty in the aft signatures with often greater boundary layer effects and nozzle jet pressures. Wind tunnel testing at significantly lower Reynolds numbers than in flight and without inlet and nozzle jet pressures make it difficult to accurately assess the computational solutions of flight vehicles. A wind tunnel test in the NASA Ames 9- by 7-Foot Supersonic Wind Tunnel from Mach 1.6 to 2.0 will be used to assess the effects of shocks from components passing through nozzle jet plumes on the sonic boom pressure signature and provide datasets for comparison with CFD codes. A large number of high-fidelity numerical simulations of wind tunnel test models with a variety of shock generators that simulate horizontal tails and aft decks have been studied to provide suitable models for sonic boom pressure measurements using a minimally intrusive pressure rail in the wind tunnel. The computational results are presented and the evolution of candidate wind tunnel models is summarized and discussed in this paper.

  16. A Wind-Tunnel Investigation of the Aerodynamic Characteristics of a Full-Scale Supersonic-Type Three-blade Propeller at Mach Numbers to 0.96

    NASA Technical Reports Server (NTRS)

    Evans, Albert J; Liner, George

    1958-01-01

    An investigation of the characteristics of a full-scale supersonic-type propeller has been made in the Langley 16-foot transonic tunnel with the 6000-horsepower propeller dynamometer. The tests covered a range of blade angles from 20.2 degrees to 60.2 degrees at forward Mach numbers up to 0.96. The results showed that envelope efficiency at an advance ratio of 2.8 decreased from 86 percent to 72 percent when the forward Mach number was increased from 0.70 to 0.96. A comparison of the experimental results with calculated results showed that maximum propeller efficiency can be calculated with good accuracy by using ordinary subsonic strip theory when the blade-section speeds are supersonic. The investigation also showed favorable power-absorption properties of the supersonic-type propeller at high speeds.

  17. Nonequilibrium Supersonic Magnetogasdynamic Wind Tunnel

    DTIC Science & Technology

    2007-12-31

    test section. 43 lrann,sionlin fo si alIransmission line for signal R:CaQ itive Feedl-throoph capacitor Probe Transmiission linefor pulsar Ir ansmission...line for pulsar Contact 49 Ou~ter shieldsufc nte wir C. 3Q9x 12 Figure 3. Schematic of a back current shunt. Figure 4. Schematic and an equivalent

  18. 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.

  19. Lessons in the Design and Characterization Testing of the Semi-Span Super-Sonic Transport (S4T) Wind-Tunnel Model

    NASA Technical Reports Server (NTRS)

    2012-01-01

    This paper focuses on some of the more challenging design processes and characterization tests of the Semi-Span Super-Sonic Transport (S4T)-Active Controls Testbed (ACT). The model was successfully tested in four entries in the National Aeronautics and Space Administration Langley Transonic Dynamics Tunnel to satisfy the goals and objectives of the Fundamental Aeronautics Program Supersonic Project Aero-Propulso-Servo-Elastic effort. Due to the complexity of the S4T-ACT, only a small sample of the technical challenges for designing and characterizing the model will be presented. Specifically, the challenges encountered in designing the model include scaling the Technology Concept Airplane to model scale, designing the model fuselage, aileron actuator, and engine pylons. Characterization tests included full model ground vibration tests, wing stiffness measurements, geometry measurements, proof load testing, and measurement of fuselage static and dynamic properties.

  20. Aerodynamic characteristics of a fixed arrow-wing supersonic cruise aircraft at Mach numbers of 2.30, 2.70, and 2.95. [Langley Unitary Plan wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Morris, O. A.; Fuller, D. E.; Watson, C. B.

    1978-01-01

    Tests were conducted in the Langley Unitary Plan wind tunnel at Mach numbers of 2.30. 2.70, and 2.95 to determine the performance, static stability, and control characteristics of a model of a fixed-wing supersonic cruise aircraft with a design Mach Number of 2.70 (SCAT 15-F-9898). The configuration had a 74 deg swept warped wing with a reflexed trailing edge and four engine nacelles mounted below the reflexed portion of the wing. A number of variations in the basic configuration were investigated; they included the effect of wing leading edge radius, the effect of various model components, and the effect of model control deflections.

  1. 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.

  2. A throat-bypass stability-bleed system using relief valves to increase the transient stability of a mixed-compression inlet. [YF-12 aircraft inlet tests in the Lewis 10 by 10 ft supersonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Neiner, G. H.; Dustin, M. O.; Cole, G. L.

    1979-01-01

    A stability-bleed system was installed in a YF-12 flight inlet that was subjected to internal and external airflow disturbances in the NASA Lewis 10 by 10 foot supersonic wind tunnel. The purpose of the system is to allow higher inlet performance while maintaining a substantial tolerance (without unstart) to internal and external disturbances. At Mach numbers of 2.47 and 2.76, the inlet tolerance to decreases in diffuser-exit corrected airflow was increased by approximately 10 percent of the operating-point airflow. The stability-bleed system complemented the terminal-shock-control system of the inlet and did not show interaction problems. For disturbances which caused a combined decrease in Mach number and increase in angle of attack, the system with valves operative kept the inlet started 4 to 28 times longer than with the valves inoperative. Hence, the stability system provides additional time for the inlet control system to react and prevent unstart. This was observed for initial Mach numbers of 2.55 and 2.68. For slow increase in angle of attack at Mach 2.47 and 2.76, the system kept the inlet started beyond the steady-state unstart angle. However, the maximum transient angles of attack without unstart could not be determined because wind-tunnel mechanical-stop limits for angle of attack were reached.

  3. Aerodynamic interactions from reaction controls for lateral control of the M2-F2 lifting-body entry configuration at transonic and supersonic and supersonic Mach numbers. [wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Bailey, R. O.; Brownson, J. J.

    1979-01-01

    Tests were conducted in the Ames 6 by 6 foot wind tunnel to determine the interaction of reaction jets for roll control on the M2-F2 lifting-body entry vehicle. Moment interactions are presented for a Mach number range of 0.6 to 1.7, a Reynolds number range of 1.2 x 10 to the 6th power to 1.6 x 10 to the 6th power (based on model reference length), an angle-of-attack range of -9 deg to 20 deg, and an angle-of-sideslip range of -6 deg to 6 deg at an angle of attack of 6 deg. The reaction jets produce roll control with small adverse yawing moment, which can be offset by horizontal thrust component of canted jets.

  4. 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.

  5. Low-speed aerodynamic characteristics from wind-tunnel tests of a large-scale advanced arrow-wing supersonic-cruise transport concept

    NASA Technical Reports Server (NTRS)

    Smith, P. M.

    1978-01-01

    Tests have been conducted to extend the existing low speed aerodynamic data base of advanced supersonic-cruise arrow wing configurations. Principle configuration variables included wing leading-edge flap deflection, wing trailing-edge flap deflection, horizontal tail effectiveness, and fuselage forebody strakes. A limited investigation was also conducted to determine the low speed aerodynamic effects due to slotted training-edge flaps. Results of this investigation demonstrate that deflecting the wing leading-edge flaps downward to suppress the wing apex vortices provides improved static longitudinal stability; however, it also results in significantly reduced static directional stability. The use of a selected fuselage forebody strakes is found to be effective in increasing the level of positive static directional stability. Drooping the fuselage nose, which is required for low-speed pilot vision, significantly improves the later-directional trim characteristics.

  6. 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.

  7. 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.

  8. 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.

  9. Wind tunnel pressure study and Euler code validation of a missile configuration with 77 deg swept delta wings at supersonic speeds. M.S. Thesis - George Washington Univ.

    NASA Technical Reports Server (NTRS)

    Fulton, Patsy S.

    1988-01-01

    A wind-tunnel pressure study was conducted on an axisymmetric missile configuration in the Unitary Plan Wind Tunnel at NASA Langley Research Center. The Mach numbers ranged from 1.70 to 2.86 and the angles of attack ranged from minus 4 degrees to plus 24 degrees. The computational accuracy for limited conditions of a space-marching Euler code was assessed.

  10. 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…

  11. 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.

  12. 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.

  13. Low-speed wind-tunnel tests of a 1/10-scale model of an advanced arrow-wing supersonic cruise configuration designed for cruise at Mach 2.2. [Langley Full Scale Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Yip, L. P.

    1979-01-01

    The low-speed longitudinal and lateral-directional characteristics of a scale model of an advanced arrow-wing supersonic cruise configuration were investigated in tests conducted at a Reynolds number of 4.19 x 10 to the 6th power based on the mean aerodynamic chord, with an angle of attack range from - 6 deg to 23 deg and sideslip angle range from -15 deg to 20 deg. The effects of segmented leading-edge flaps, slotted trailing-edge flaps, horizontal and vertical tails, and ailerons and spoilers were determined. Extensive pressure data and flow visualization pictures with non-intrusive fluorescent mini-tufts were obtained.

  14. Wind-Tunnel Investigation of Subsonic Longitudinal Aerodynamic Characteristics of a Tiltable-Wing Vertical-Take-Off-and-Landing Supersonic Bomber Configuration Including Turbojet Power Effects

    NASA Technical Reports Server (NTRS)

    Thompson, Robert F.; Vogler, Raymond D.; Moseley, William C., Jr.

    1959-01-01

    Jet-powered model tests were made to determine the low-speed longitudinal aerodynamic characteristics of a vertical-take-off and-landing supersonic bomber configuration. The configuration has an unique engine-wing arrangement wherein six large turbojet engines (three on each side of the fuselage) are buried in a low-aspect-ratio wing which is tilted into the vertical plane for take-off. An essentially two-dimensional variable inlet, spanning the leading edge of each wing semispan, provides air for the engines. Jet flow conditions were simulated for a range of military (nonafterburner) and afterburner turbojet-powered flight at subsonic speeds. Three horizontal tails were tested at a station down-stream of the jet exit and at three heights above the jet axes. A semi-span model was used and test parameters covered wing-fuselage incidence angles from 0 deg to 15 deg, wing angles of attack from -4 deg to 36 deg, a variable range of horizontal-tail incidence angles, and some variations in power simulation conditions. Results show that, with all horizontal tails tested, there were large variations in static stability throughout the lift range. When the wing and fuselage were alined, the model was statically stable throughout the test range only with the largest tail tested (tail span of 1.25 wing span) and only when the tail was located in the low test position which placed the tail nearest to the undeflected jet. For transition flight conditions, none of the tail configurations provided satisfactory longitudinal stability or trim throughout the lift range. Jet flow was destabilizing for most of the test conditions, and varying the jet-exit flow conditions at a constant thrust coefficient had little effect on the stability of this model. Wing leading-edge simulation had some important effects on the longitudinal aerodynamic characteristics.

  15. 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.

  16. 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.

  17. Flight effects on the aerodynamic and acoustic characteristics of inverted profile coannular nozzles, volume 1. [supersonic cruise aircraft research wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Kozlowski, H.; Packman, A. B.

    1978-01-01

    Jet noise spectra obtained at static conditions from an acoustic wind tunnel and an outdoor facility are compared. Data curves are presented for (1) the effect of relative velocity on OASPL directivity (all configurations); (2) the effect of relative velocity on noise spectra (all configurations); (3) the effect of velocity on PNL directivity (coannular nozzle configurations); (4) nozzle exhaust plume velocity profiles; and (5) the effect of relative velocity on aerodynamic performance.

  18. 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

  19. 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

  20. 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.

  1. Solar wind models

    NASA Technical Reports Server (NTRS)

    Leer, Egil; Sandbaek, Ornulf

    1991-01-01

    The understanding of the solar wind is based upon Parker's (1958) description of a thermally driven subsonic - supersonic outflow from a fully ionized electron-proton corona. The basic physical processes of thermally driven solar wind models are discussed. Also studied are the effect of alpha particles in the corona on the solar wind proton flux. The acceleration of the solar wind by Alfven waves is discussed.

  2. Wind-tunnel investigation at supersonic speeds of a canard-controlled missile with fixed and free-rolling tail fins

    NASA Technical Reports Server (NTRS)

    Blair, A. B., Jr.

    1978-01-01

    A wind tunnel investigation was made at free stream Mach numbers from 1.70 to 2.86 to determine the effects of fixed and free rolling tail fin afterbodies on the static longitudinal and lateral aerodynamic characteristics of a cruciform canard controlled missile model. The effect of small canard roll and yaw control deflections was also examined. The results indicate that the fixed and free rolling tail configurations have about the same lift curve slope and longitudinal stability level at low angles of attack. For the free rolling tail configuration, the canards provide conventional roll control with no roll control reversal at low angles of attack. The free rolling tail configuration reduced induced roll due to model roll angle and canard yaw control.

  3. 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.

  4. Wind-tunnel investigation at supersonic speeds of a remote-controlled canard missile with a free-rolling-tail brake torque system

    NASA Technical Reports Server (NTRS)

    Blair, A. B., Jr.

    1985-01-01

    Wind tunnel tests were conducted at Mach numbers 1.70, 2.16, and 2.86 to determine the static aerodynamic characteristics of a cruciform canard-controlled missile with fixed or free rolling tailfin afterbodies. Mechanical coupling effects of the free-rolling-tail afterbody were investigated by using an electronic electromagnetic brake system providing arbitrary tail-fin brake torques with continuous measurements of tail-to-mainframe torque and tail roll rate. Remote-controlled canards were deflected to provide pitch, yaw, and roll control. Results indicate that the induced rolling moment coefficients due to canard yaw control are reduced and linearized for the free-rolling-tail (free-tail) configuration. The canards of the latter provide conventional roll control for the entire angle-of-attack test range. For the free-tail configuration, the induced rolling moment coefficient due to canard yaw control increased and the canard roll control decreased with increases in brake torque, which simulated bearing friction torque. It appears that a compromise in regard to bearing friction, for example, low-cost bearings with some friction, may allow satisfactory free-tail aerodynamic characteristics that include reductions in adverse rolling-moment coefficients and lower tail roll rates.

  5. 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.

  6. High-speed Wind Tunnels

    NASA Technical Reports Server (NTRS)

    Ackeret, J

    1936-01-01

    Wind tunnel construction and design is discussed especially in relation to subsonic and supersonic speeds. Reynolds Numbers and the theory of compressible flows are also taken into consideration in designing new tunnels.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. Aerodynamic results of wind tunnel tests on a 0.010-scale model (32-QTS) space shuttle integrated vehicle in the AEDC VKF-40-inch supersonic wind tunnel (IA61)

    NASA Technical Reports Server (NTRS)

    Daileda, J. J.

    1976-01-01

    Plotted and tabulated aerodynamic coefficient data from a wind tunnel test of the integrated space shuttle vehicle are presented. The primary test objective was to determine proximity force and moment data for the orbiter/external tank and solid rocket booster (SRB) with and without separation rockets firing for both single and dual booster runs. Data were obtained at three points (t = 0, 1.25, and 2.0 seconds) on the nominal SRB separation trajectory.

  14. 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.

  15. 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.

  16. 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.

  17. Solar Wind Magnetic Fields

    NASA Technical Reports Server (NTRS)

    Smith, E. J.

    1995-01-01

    The magnetic fields originate as coronal fields that are converted into space by the supersonic, infinitely conducting, solar wind. On average, the sun's rotation causes the field to wind up and form an Archimedes Spiral. However, the field direction changes almost continuously on a variety of scales and the irregular nature of these changes is often interpreted as evidence that the solar wind flow is turbulent.

  18. 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.

  19. 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

  20. 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.

  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. 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

  3. 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).

  4. 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.

  5. Pressure recovery, drag, and subcritical stability characteristics of conical supersonic diffusers with boundary-layer removal

    NASA Technical Reports Server (NTRS)

    Obey, Leonard T; Englert, Gerald W; Nussdorfer, Theodore J , Jr

    1952-01-01

    A study of two 20 degrees half-angle, low mass-flow ratio conical supersonic inlets with cone boundary-layer bleed was made on a 16-inch ram-jet engine in the Lewis 8- by 6-foot supersonic wind tunnel. A greater stable subcritical range of operation was obtained with the bleed inlets than with the corresponding inlet without boundary-layer bleed. The drag added by the bleed system was small.

  6. 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.

  7. MSL Parachute Shock Oscillation

    NASA Video Gallery

    Ames engineers conducted tests of a two-percent scale model of the Mars Science Laboratory (MSL) entry vehicle in the NASA Ames 9-foot-by-7-foot Supersonic Wind Tunnel with a rigid representation o...

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  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 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. Low-speed wind-tunnel tests of a one-tenth-scale model of a blended-arrow advanced supersonic transport. [conducted in Langley full-scale tunnel

    NASA Technical Reports Server (NTRS)

    Lemore, H. C.; Parett, L. P.

    1975-01-01

    Tests were conducted in the Langley full scale tunnel to determine the low-speed aerodynamic characteristics of a 1/10 scale model of a blended-arrow advanced supersonic transport. Tests were made for the clean configuration and a high-lift configuration with several combinations of leading- and trailing-edge flaps deflected for providing improved lift and longitudinal stability in the landing and takeoff modes. The tests were conducted for a range of angles of attack from about -6 deg to 30 deg, sideslip angles from -5 deg to 10 deg, and for Reynolds numbers from 6.78 x 1,000,000 to 13.85 x 1,000,000 corresponding to test velocities of 41 knots to 85 knots, respectively.

  2. Wind-Tunnel Investigation of a 0.6-Scale Model of Hughes MX-904 Tail Surface at Supersonic Speeds Several Combinations of the Tail with Each of Two Foreshortened Body Segments

    NASA Technical Reports Server (NTRS)

    Guy, Lawrence D.; Conner, D. William

    1950-01-01

    An investigation has been made in the Langley 9- by 12-inch super-sonic blowdown tunnel at Mach numbers of 1.62 and 1.96 of a partial-span body with one tail surface, designed for use on the Hughes Falcon (MX-904) missile. The present paper extends the work reported in NACA-RM-SL50E10. Force and moment data including elevator hinge moment were obtained for the conditions of the tail in the presence of a small segment of the fore-shortened body, in the presence of a semi-span body and attached to a semi-span body, and for the condition of the foreshortened semi-span body alone.

  3. 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.

  4. Stellar Winds

    NASA Astrophysics Data System (ADS)

    Owocki, Stan

    A "stellar wind" is the continuous, supersonic outflow of matter from the surface layers of a star. Our sun has a solar wind, driven by the gas-pressure expansion of the hot (T > 106 K) solar corona. It can be studied through direct in situ measurement by interplanetary spacecraft; but analogous coronal winds in more distant solar-type stars are so tenuous and transparent that that they are difficult to detect directly. Many more luminous stars have winds that are dense enough to be opaque at certain wavelengths of the star's radiation, making it possible to study their wind outflows remotely through careful interpretation of the observed stellar spectra. Red giant stars show slow, dense winds that may be driven by the pressure from magnetohydrodyanmic waves. As stars with initial mass up to 8 M ⊙ evolve toward the Asymptotic Giant Branch (AGB), a combination of stellar pulsations and radiative scattering off dust can culminate in "superwinds" that strip away the entire stellar envelope, leaving behind a hot white dwarf stellar core with less than the Chandrasekhar mass of ˜ ​​ 1. 4M ⊙. The winds of hot, luminous, massive stars are driven by line-scattering of stellar radiation, but such massive stars can also exhibit superwind episodes, either as Red Supergiants or Luminous Blue Variable stars. The combined wind and superwind mass loss can strip the star's hydrogen envelope, leaving behind a Wolf-Rayet star composed of the products of earlier nuclear burning via the CNO cycle. In addition to such direct effects on a star's own evolution, stellar winds can be a substantial source of mass, momentum, and energy to the interstellar medium, blowing open large cavities or "bubbles" in this ISM, seeding it with nuclear processed material, and even helping trigger the formation of new stars, and influencing their eventual fate as white dwarves or core-collapse supernovae. This chapter reviews the properties of such stellar winds, with an emphasis on the various

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  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

    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.

  11. 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.

  12. 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.

  13. 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.

  14. 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

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. Ongoing Validation of Computational Fluid Dynamics for Supersonic Retro-Propulsion

    NASA Technical Reports Server (NTRS)

    Schauerhamer, Guy; Trumble, Kerry; Carlson, Jan-Renee; Edquist, Karl; Buning, Pieter; Sozer, Emre

    2011-01-01

    During the Entry, Decent, and Landing phase of planetary exploration, previous methods of deceleration do not scale with high mass spacecraft. Supersonic Retro-Propulsion(SRP)is a viable method to decelerate large spacecraft including those that will carry humans to Mars. Flow data at these conditions are difficult to obtain through flight or wind tunnel experiments

  1. 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.

  2. 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.

  3. F-16XL Wind Tunnel Models

    NASA Technical Reports Server (NTRS)

    1994-01-01

    This is a multiple exposure image of the F-16XL Supersonic Laminar Flow Control (SLFC) model in the Unitary Plan Wind Tunnel. This wind tunnel test was conducted to verify design pressure distributions for the SLFC flight experiment (see modified port wing) and to obtain simulator coefficients for stability and control investigations.

  4. Results of aeroheating DFI and ET design-data test on a 0.0175-scale model 60-OTS conducted in the Von Karman Gas Dynamics Facility (VKF) 40" supersonic and the 50" hypersonic wind tunnels A and C

    NASA Technical Reports Server (NTRS)

    Marroquin, J.; Leef, C. R.

    1985-01-01

    The 0.0175-scale thin-skin thermocouple Model 60-OTS was tested in the von Karman Gas Facility 40-inch continuous flow supersonic tunnel A and the VKF Tunnel C. Testing was conducted at Mach numbers 2.25 to 4.0, and Reynolds numbers from 0.4 x 10 to the 6th power/ft to 6.6 x 10 to the 6th power/ft. Angle of attack range was from -5.0 to +5.0 degrees, and angle of sideslip range was from -6 to +6 degrees. The primary objective of this test was to provide a valid base for the external tank (ET) and solid rocket booster (SRB) heating prediction methodology for ascent flight by taking heating data at Development Flight Instrumentation (DFI) locations for flight conditions simulating STS-1 through -4. A second objective was to obtain additional aeroheating data to support potential reduction of the thermal protection system (TPS) on the ET. The third phase of the test was funded and conducted by NASA/MSFC for the purpose of establishing confidence in the data base from the lower temperature tunnel A.

  5. Exhaust Nozzles for Propulsion Systems with Emphasis on Supersonic Cruise Aircraft

    NASA Technical Reports Server (NTRS)

    Stitt, Leonard E.

    1990-01-01

    This compendium summarizes the contributions of the NASA-Lewis and its contractors to supersonic exhaust nozzle research from 1963 to 1985. Two major research and technology efforts sponsored this nozzle research work; the U.S. Supersonic Transport (SST) Program and the follow-on Supersonic Cruise Research (SCR) Program. They account for two generations of nozzle technology: the first from 1963 to 1971, and the second from 1971 to 1985. First, the equations used to calculate nozzle thrust are introduced. Then the general types of nozzles are presented, followed by a discussion of those types proposed for supersonic aircraft. Next, the first-generation nozzles designed specifically for the Boeing SST and the second-generation nozzles designed under the SCR program are separately reviewed and then compared. A chapter on throttle-dependent afterbody drag is included, since drag has a major effect on the off-design performance of supersonic nozzles. A chapter on the performance of supersonic dash nozzles follows, since these nozzles have similar design problems, Finally, the nozzle test facilities used at NASA-Lewis during this nozzle research effort are identified and discussed. These facilities include static test stands, a transonic wind tunnel, and a flying testbed aircraft. A concluding section points to the future: a third generation of nozzles designed for a new era of high speed civil transports to produce even greater advances in performance, to meet new noise rules, and to ensure the continuity of over two decades of NASA research.

  6. 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.

  7. 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.

  8. 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%.

  9. 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.

  10. 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.

  11. 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.

  12. Supersonic control effectiveness for full and partial span elevon configurations on a 0.0165 scale model space shuttle orbiter tested in the LaRC unitary plan wind tunnel (LA49)

    NASA Technical Reports Server (NTRS)

    1977-01-01

    A wind tunnel test is reported on an early version of the space shuttle orbiter (designated 089B-139) 0.0165 scale model to systematically determine both longitudinal and lateral control effectiveness associated with various combinations of inboard, outboard, and full span wing trailing edge controls. The test Mach umbers were 2.5 and 4.63 over an angle of attack range from -4 deg to 42 deg at 0 deg sideslip.

  13. 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.

  14. 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.

  15. 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.

  16. Computational Analysis of the Large Scale Low-Boom Supersonic Inlet

    NASA Technical Reports Server (NTRS)

    Chima, ROdrick V.

    2011-01-01

    This presentation describes two computational fluid dynamic (CFD) analyses done in support of a supersonic inlet test performed at NASA Glenn Research Center in the fall of 2010. The large-scale-low-boom supersonic inlet was designed for a small supersonic aircraft that would cruise at a Mach number of 1.6. It uses an axisymmetric, external compression spike to reduce the Mach number to 0.65 at the fan face. The inlet was tested in the 8x6 supersonic wind tunnel at NASA GRC using conventional pressure probes, pressure sensitive paint, and high-speed schlieren. Two CFD analyses of the inlet were performed before the test, and compared to the experimental data afterwards. Both analyses used the WIND-US code. First, an axisymmetric analysis of the inlet, diffuser, cold pipe, and mass flow plug was performed to predict the performance of the entire system in the wind tunnel. Then a 3-D analysis of the inlet with all its interior struts was performed to predict details of the flow field and effects of angle of attack. Test results showed that the inlet had excellent performance, with a peak total pressure recovery of 96 percent, and a buzz point far outside the engine operating range. The computations agreed very well with the data, with predicted recoveries within 0.3 - 0.5 points of the measurements.

  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

    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.

  18. Supersonic flow around a cylinder with front gas-permeable insert which modeled by skeleton of porous material

    NASA Astrophysics Data System (ADS)

    Poplavskaya, T. V.; Kirilovskiy, S. V.; Mironov, S. G.

    2016-10-01

    Experimental data and results of numerical simulation of a supersonic flow around a streamwise aligned cylinder with a frontal gas-permeable insert made of a high-porosity cellular material are presented. The porous material structure is modeled by a system of staggered rings of different diameters (discrete model of a porous medium). The model skeleton of the material corresponds to the pore size (diameter 1mm) and porosity (0.95) of a real cellular porous material. The computed results are compared with the data of wind tunnel experiments performed in a T-327B supersonic continuous-flow wind tunnel at the flow Mach number M∞ = 4.85.

  19. 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

  20. 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.

  1. 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

  2. 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

  3. 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.

  4. 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.

  5. 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.”

  6. 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.

  7. Computational and Experimental Study of Supersonic Nozzle Flow and Shock Interactions

    NASA Technical Reports Server (NTRS)

    Carter, Melissa B.; Elmiligui, Alaa A.; Nayani, Sudheer N.; Castner, Ray; Bruce, Walter E., IV; Inskeep, Jacob

    2015-01-01

    This study focused on the capability of NASA Tetrahedral Unstructured Software System's CFD code USM3D capability to predict the interaction between a shock and supersonic plume flow. Previous studies, published in 2004, 2009 and 2013, investigated USM3D's supersonic plume flow results versus historical experimental data. This current study builds on that research by utilizing the best practices from the early papers for properly capturing the plume flow and then adding a wedge acting as a shock generator. This computational study is in conjunction with experimental tests conducted at the Glenn Research Center 1'x1' Supersonic Wind Tunnel. The comparison of the computational and experimental data shows good agreement for location and strength of the shocks although there are vertical shifts between the data sets that may be do to the measurement technique.

  8. Control of shock wave-boundary layer interactions by bleed in supersonic mixed compression inlets

    NASA Technical Reports Server (NTRS)

    Fukuda, M. K.; Hingst, W. G.; Reshotko, E.

    1975-01-01

    An experimental investigation was conducted to determine the effect of bleed on a shock wave-boundary layer interaction in an axisymmetric mixed-compression supersonic inlet. The inlet was designed for a free-stream Mach number of 2.50 with 60-percent supersonic internal area contraction. The experiment was conducted in the NASA Lewis Research Center 10-Foot Supersonic Wind Tunnel. The effects of bleed amount and bleed geometry on the boundary layer after a shock wave-boundary layer interaction were studied. The effect of bleed on the transformed form factor is such that the full realizable reduction is obtained by bleeding of a mass flow equal to about one-half of the incident boundary layer mass flow. More bleeding does not yield further reduction. Bleeding upstream or downstream of the shock-induced pressure rise is preferable to bleeding across the shock-induced pressure rise.

  9. Ejector nozzle test results at simulated flight conditions for an advanced supersonic transport propulsion system

    NASA Technical Reports Server (NTRS)

    Nelson, D. P.; Bresnahan, D. L.

    1983-01-01

    Results are presented of wind tunnel tests conducted to verify the performance improvements of a refined ejector nozzle design for advanced supersonic transport propulsion systems. The analysis of results obtained at simulated engine operating conditions is emphasized. Tests were conducted with models of approximately 1/10th scale which were configured to simulate nozzle operation at takeoff, subsonic cruise, transonic cruise, and supersonic cruise. Transonic cruise operation was not a consideration during the nozzle design phase, although an evaluation at this condition was later conducted. Test results, characterized by thrust and flow coefficients, are given for a range of nozzle pressure ratios, emphasizing the thrust performance at the engine operating conditions predicted for each flight Mach number. The results indicate that nozzle performance goals were met or closely approximated at takeoff and supersonic cruise, while subsonic cruise performance was within 2.3 percent of the goal with further improvement possible.

  10. 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.

  11. Aerodynamic results of a separation effects test on a 0.010-scale model (52-OTS) of the integrated SSV in the AEDC/VKF 40-by-40 inch supersonic wind tunnel A (IA111), volume 1

    NASA Technical Reports Server (NTRS)

    Chee, E.

    1976-01-01

    Graphical data obtained during experimental wind tunnel aerodynamic investigations of a 0.010 scale model (52-OTS) of the integrated space shuttle vehicle was presented. The purpose of this investigation was to obtain data with the solid rocket booster (SRB) in proximity to the orbiter/external tank (O/ET), over a large O/ET initial angle of attack and sideslip range, as well as data on the SRB alone (greatly separated from the O/ET). A captive trajectory system, which supported the SRB, was used with the tunnel primary sector (supporting the O/ET) to obtain grid type separation effects data. One symmetrical SRB model was used interchangeably to obtain right-hand and left-hand SRB data. The entire investigation was conducted at a free-stream Mach number of 4.5 at unit Reynolds number of 3.95 and 5.9 million per foot.

  12. Aerodynamic results of a separation effects test on a 0.010-scale model (52-OTS) of the integrated SSV in the AEDC/VKF 40-by-40 inch supersonic wind tunnel A (IA111), volume 2

    NASA Technical Reports Server (NTRS)

    Chee, E.

    1976-01-01

    Tabular data obtained during experimental wind tunnel aerodynamic investigations of a 0.010 scale model (52-OTS) of the integrated space shuttle vehicle was presented. The purpose of this investigation was to obtain data with the solid rocket booster (SRB) in proximity to the orbiter/external tank (O/ET), over a large O/ET initial angle of attack and sideslip range, as well as data on the SRB alone (greatly separated from the O/ET). A captive trajectory system, which supported the SRB, was used with the tunnel primary sector (supporting the O/ET) to obtain grid type separation effects data. One symmetrical SRB model was used interchangeably to obtain right-hand and left-hand SRB data. The entire investigation was conducted at free-stream Mach number of 4.5 at unit Reynolds number of 3.95 and 5.9 million per foot.

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

    NASA Technical Reports Server (NTRS)

    Foust, J. W.

    1979-01-01

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

  14. 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.

  15. 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.

  16. The influence of flow parameters on the transition to turbulence in supersonic boundary layer on swept wing

    NASA Astrophysics Data System (ADS)

    Semionov, N. V.; Yermolaev, Yu. G.; Kosinov, A. D.; Dryasov, A. D.; Semenov, A. N.; Yatskikh, A. A.

    2016-10-01

    The paper is devoted to an experimental study of laminar-turbulent transition in a three-dimensional supersonic boundary layer. The experiments were conducted at the low nose supersonic wind tunnel T-325 of ITAM at Mach numbers M=2 - 4. Model is a symmetrical wing with a 45° sweep angle, a 3 percent-thick circular-arc airfoil. The influence of flow parameters, such as the Mach number, unit Reynolds number, angle of attack, level of perturbations on the transitions to turbulence are on the consideration. Transition Reynolds numbers are obtained. Analysis of all obtained data allow to determine reliable value of Retr of swept wing supersonic boundary layer, that especially important at consideration of experiments fulfilled at different flow conditions in different wind tunnels.

  17. 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.

  18. 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.

  19. Wind-Tunnel Investigation at Subsonic and Supersonic Speeds of the Static and Dynamic Stability Derivatives of an Airplane Model with an Unswept Wing and a High Horizontal Tail

    NASA Technical Reports Server (NTRS)

    Lessing, Henry C.; Butler, James K.

    1959-01-01

    Results are presented of a wind-tunnel investigation to evaluate the static and dynamic stability derivatives of a model with a low-aspect-ratio unswept wing and a high horizontal tail. In addition to results for the complete model, results were also obtained of the body alone, body and wing, and body and tail. Data were obtained in the Mach number range from 0.65 to 2.2, at a Reynolds number of 2 million based on the wing mean aerodynamic chord. The angle-of-attack range for most of the data was -11.5 deg to 18 deg. A limited amount of data was obtained with fixed transition. A correspondence between the damping in pitch and the static stability, previously noted in other investigations, was also observed in the present results. The effect observed was that a decrease (or increase) in the static stability was accompanied by an increase (or decrease) in the damping in pitch. A similar correspondence was observed between the damping in yaw and the static-directional stability. Results from similar tests of the same model configuration in two other facilities over different speed ranges are presented for comparison. It was found that most of the results from the three investigations correlated reasonably well. Estimates of the rotary derivatives were made using available procedures. Comparison with the experimental results indicates the need for development of more precise estimation procedures.

  20. 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.

  1. Comparison of Theoretical and Experimental Unsteady Aerodynamics of Linear Oscillating Cascade With Supersonic Leading-Edge Locus. Video

    NASA Technical Reports Server (NTRS)

    Ramsey, John K.; Erwin, Dan

    2004-01-01

    The movie file contained in the DVD contains footage of the NASA/Ohio State Supersonic Oscillating Cascade Facility wind tunnel run with an oscillating mechanism. Links to view the movie can also be found in figures 4 and 8 in the online PDF version that this record references, CASI ID 20040082334.

  2. An NACA Vane-Type Angle-of-Attack Indicator for use at Subsonic and Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Mitchell, Jesse L.; Peck, Robert F.

    1949-01-01

    A vane-type angle-of-attack indicator suitable for measurements at both subsonic and supersonic speeds has been developed by the National Advisory Committee for Aeronautics. A brief history is given of the development, and a wind-tunnel calibration of the indicator is presented, together with a discussion of the corrections to be applied to the indicated readings.

  3. 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.

  4. Supersonic quiet-tunnel development for laminar-turbulent transition research

    NASA Technical Reports Server (NTRS)

    Schneider, Steven P.

    1995-01-01

    This grant supported research into quiet-flow supersonic wind-tunnels, between February 1994 and February 1995. Quiet-flow nozzles operate with laminar nozzle-wall boundary layers, in order to provide low-disturbance flow for studies of laminar-turbulent transition under conditions comparable to flight. Major accomplishments include: (1) development of the Purdue Quiet-Flow Ludwieg Tube, (2) computational evaluation of the square nozzle concept for quiet-flow nozzles, and (3) measurement of the presence of early transition on the flat sidewalls of the NASA LaRC Mach 3.5 supersonic low-disturbance tunnel. Since items (1) and (2) are described in the final report for companion grant NAG1-1133, only item (3) is described here. A thesis addressing the development of square nozzles for high-speed, low-disturbance wind tunnels is included as an appendix.

  5. Rudolf Hermann, wind tunnels and aerodynamics

    NASA Astrophysics Data System (ADS)

    Lundquist, Charles A.; Coleman, Anne M.

    2008-04-01

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

  6. 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.

  7. Calculation of vertical and ramp-assisted takeoffs for supersonic cruise fighters

    NASA Technical Reports Server (NTRS)

    Liu, G. C.

    1984-01-01

    A procedure that allows rapid preliminary evaluations of the vertical, short, and normal takeoff performance of supersonic cruise aircraft concepts was developed into a numerical computer program. The program is used to determine the effects on takeoff performance of various parameters, such as thrust-weight ratio, wing loading, thrust vector angle, and flap setting. Ramp-assisted takeoffs for overloaded configurations typical of a ground-attack mission are included. The effects of wind on the takeoff performance are also considered.

  8. A two-fluid model of the solar wind

    NASA Technical Reports Server (NTRS)

    Sandbaek, O.; Leer, E.; Holzer, T. E.

    1992-01-01

    A method is presented for the integration of the two-fluid solar-wind equations which is applicable to a wide variety of coronal base densities and temperatures. The method involves proton heat conduction, and may be applied to coronal base conditions for which subsonic-supersonic solar wind solutions exist.

  9. 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.

  10. 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.

  11. A review of ONERA aerodynamic research in support of a future supersonic transport aircraft

    NASA Astrophysics Data System (ADS)

    Thibert, J. J.; Arnal, D.

    2000-11-01

    The ONERA activities concerning the aerodynamics of the future supersonic transport aircraft are reviewed. Section 1 is devoted to the performance prediction and detailed comparisons between CFD and wind-tunnel data are presented and discussed. Section 2 addresses the problem of the drag prediction in cruise flight conditions from wind-tunnel data. Skin friction coefficients values measured in flight are compared to the results of boundary layer computations. Section 3 is devoted to wing designs with numerical optimisation techniques. Several examples are presented and discussed. Results concerning riblets and laminar flow control are given in Section 4 part which also will present experiments carried out for attachment line contamination investigation. Results from basic research on supersonic laminar flows are also be presented. Section 5 deals with activities on air intake aerodynamics. After a brief recall of supersonic air intakes operational modes and a description of the Concorde air intake, comparisons between CFD and wind tunnel data on a generic 2D intake are presented. Basic experiments on intake internal flow are described and the problem of the internal shock control is addressed.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  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. 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.

  18. 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...

  19. 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.

  20. 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.

  1. 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...

  2. 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.

  3. Research on a two-dimensional inlet for a supersonic V/STOL propulsion system. Appendix A

    NASA Technical Reports Server (NTRS)

    Mark, J. L.; Mcgarry, M. A.; Reagan, P. V.

    1984-01-01

    The inlet system performance requirements associated with supersonic V/STOL aircraft place extreme demands on the inlet designer. The present effort makes maximum use of flow improvement techniques, proven for high subsonic maneuvering flight and adapts them to the critical static and low speed/high angle-of-attack flight regime of the supersonic V/STOL aircraft. A description of the aerodynamic design, model characteristics, data analysis, discussion, and conclusions concerning the most promising inlet design approaches are contained. The appendix contains the reduced wind tunnel data plots and pressure distribution.

  4. The Experimental Measurement of Aerodynamic Heating About Complex Shapes at Supersonic Mach Numbers

    NASA Technical Reports Server (NTRS)

    Neumann, Richard D.; Freeman, Delma C.

    2011-01-01

    In 2008 a wind tunnel test program was implemented to update the experimental data available for predicting protuberance heating at supersonic Mach numbers. For this test the Langley Unitary Wind Tunnel was also used. The significant differences for this current test were the advances in the state-of-the-art in model design, fabrication techniques, instrumentation and data acquisition capabilities. This current paper provides a focused discussion of the results of an in depth analysis of unique measurements of recovery temperature obtained during the test.

  5. Effects of Passive Porosity on Interacting Vortex Flows At Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    2000-01-01

    A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) Unitary Plan Wind Tunnel (UPWT) to determine the effects of passive surface porosity on vortex flow interaction about a general research fighter configuration at supersonic speeds. Optical flow measurement and flow visualization techniques were used and included pressure-sensitive paint (PSP), schlieren, and laser vapor screen (LVS) These techniques were combined with force and moment and conventional electronically-scanned pressure (ESP) measurements to quantify and to visualize the effects of flow-through porosity applied to a wing leading-edge extension (LEX) mounted to a 65 deg cropped delta wing model.

  6. Supersonic aerodynamic characteristics of an advanced F-16 derivative aircraft configuration

    NASA Technical Reports Server (NTRS)

    Fox, Mike C.; Forrest, Dana K.

    1993-01-01

    A supersonic wind tunnel investigation was conducted in the NASA Langley Unitary Plan Wind Tunnel on an advanced derivative configuration of the United States Air Force F-16 fighter. Longitudinal and lateral directional force and moment data were obtained at Mach numbers of 1.60 to 2.16 to evaluate basic performance parameters and control effectiveness. The aerodynamic characteristics for the F-16 derivative model were compared with the data obtained for the F-16C model and also with a previously tested generic wing model that features an identical plan form shape and similar twist distribution.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. Results of the NASA/MSFC FA-23 plume technology test program performed in the NASA/Ames unitary wind tunnels

    NASA Technical Reports Server (NTRS)

    Hendershot, K. C.

    1977-01-01

    A 2.25% scale model of the space shuttle external tank and solid rocket boosters was tested in the NASA/Ames Unitary 11 x 11 foot transonic and 9 x 7 foot supersonic tunnels to obtain base pressure data with firing solid propellant exhaust plumes. Data system difficulties prevented the acquisition of any useful data in the 9 x 7 tunnel. However, 28 successful rocket test firings were made in the 11 x 11 tunnel, providing base pressure data at Mach numbers of 0.5, 0.9, 1.05, 1.2, and 1.3 and at plume pressure ratios ranging from 11 to 89.

  13. An Interactive Method of Characteristics Java Applet to Design and Analyze Supersonic Aircraft Nozzles

    NASA Technical Reports Server (NTRS)

    Benson, Thomas J.

    2014-01-01

    The Method of Characteristics (MOC) is a classic technique for designing supersonic nozzles. An interactive computer program using MOC has been developed to allow engineers to design and analyze supersonic nozzle flow fields. The program calculates the internal flow for many classic designs, such as a supersonic wind tunnel nozzle, an ideal 2D or axisymmetric nozzle, or a variety of plug nozzles. The program also calculates the plume flow produced by the nozzle and the external flow leading to the nozzle exit. The program can be used to assess the interactions between the internal, external and plume flows. By proper design and operation of the nozzle, it may be possible to lessen the strength of the sonic boom produced at the rear of supersonic aircraft. The program can also calculate non-ideal nozzles, such as simple cone flows, to determine flow divergence and nonuniformities at the exit, and its effect on the plume shape. The computer program is written in Java and is provided as free-ware from the NASA Glenn central software server.

  14. Thrust shock vector control of an axisymmetric conical supersonic nozzle via secondary transverse gas injection

    NASA Astrophysics Data System (ADS)

    Zmijanovic, V.; Lago, V.; Sellam, M.; Chpoun, A.

    2014-01-01

    Transverse secondary gas injection into the supersonic flow of an axisymmetric convergent-divergent nozzle is investigated to describe the effects of the fluidic thrust vectoring within the framework of a small satellite launcher. Cold-flow dry-air experiments are performed in a supersonic wind tunnel using two identical supersonic conical nozzles with the different transverse injection port positions. The complex three-dimensional flow field generated by the supersonic cross-flows in these test nozzles was examined. Valuable experimental data were confronted and compared with the results obtained from the numerical simulations. Different nozzle models are numerically simulated under experimental conditions and then further investigated to determine which parameters significantly affect thrust vectoring. Effects which characterize the nozzle and thrust vectoring performances are established. The results indicate that with moderate secondary to primary mass flow rate ratios, ranging around 5 %, it is possible to achieve pertinent vector side forces. It is also revealed that injector positioning and geometry have a strong effect on the shock vector control system and nozzle performances.

  15. Low-cost digital visualization and high-speed tracking of supersonic shockwaves

    NASA Astrophysics Data System (ADS)

    Bryanston-Cross, Peter J.; Skeen, Andrew J.; Timmerman, Brenda H.; Dunkley, P.; Paduano, James D.; Guenette, G. R., Jr.

    2003-11-01

    A low-cost and low-maintenance digital focused shadowgraph flow visualization system has been developed to provide fast diagnostics of rapidly changing phenomena in supersonic flows. The system is particularly designed for tracking shock positions in a supersonic inlet, enabling high-speed active shock control. It is based on a low-cost, high-intensity white LED light source, which can be flashed with microsecond pulses enabling freeze-frame imaging of constant illumination quality. The system features three modes of operation: (1) High-resolution digital still frames and sequences (1280 x 1024, 2fps), (2) High-resolution digital frames and sequences showing spatial-temporal variation in flow field (1280 x 1024, 12 fps), (3) Adjustable windowed digital frames at reduced resolution, but at high frame rates (980 fps at 1280 x 8 pixel viewing area). The three modes of operation allow high-speed tracking of flow features such as moving of shock waves (up to 980 Hz) as well as overall instantaneous views of the flow. Furthermore, it allows direct identification of areas where high-speed changes occur. The positional shock data can be transmitted directly to a shock-stabilizing control system. Results are presented of the unsteady flow generated by an aspirated cone-shaped nozzle in a supersonic flow in the supersonic wind tunnel of the MIT Gas Turbine Laboratory.

  16. Aerodynamic and performance characterization of supersonic retropropulsion for application to planetary entry and descent

    NASA Astrophysics Data System (ADS)

    Korzun, Ashley M.

    shock layer of a blunt body in supersonic flow. Although numerous wind tunnel tests of relevance to SRP have been conducted, the scope of the work is limited in the freestream conditions and composition, retropropulsion conditions and composition, and configurations and geometries explored. The SRP aerodynamic - propulsive interaction alters the aerodynamic characteristics of the vehicle, and models must be developed that accurately represent the impact of SRP on system mass and performance. Work within this thesis has defined and advanced the state of the art for supersonic retropropulsion. This has been achieved through the application of systems analysis, computational analysis, and analytical methods. The contributions of this thesis include a detailed performance analysis and exploration of the design space specific to supersonic retropropulsion, establishment of the relationship between vehicle performance and the aerodynamic - propulsive interaction, and an assessment of the required fidelity and computational cost in simulating supersonic retropropulsion flowfields, with emphasis on the effort required to develop aerodynamic databases for conceptual design.

  17. Experimental Study on Control of Supersonic Aerodynamic Characteristics of a Vertical Landing Rocket by Using Opposing Jet

    NASA Astrophysics Data System (ADS)

    Akita, Daisuke; Yamada, Kazuhiko; Suzuki, Kojiro

    In order to control a reentry trajectory of a vertical landing rocket, an opposing jet system is experimentally tested at supersonic speeds. Experiments are conducted at Mach number 4.0 in the supersonic wind tunnel of ISAS. Supersonic nozzles of the exit Mach number 2.4 are installed on a blunt nose of a rocket model. The most significant drag reduction due to the jet-spike effect is obtained when the jet is injected from the stagnation point of the body in the opposite direction to the freestream. Various types of the nozzle arrangement are investigated and the method to increase the L/D is discussed. It is found that the off-axis jet is effective both to increase the L/D and to enable a vehicle to be trimmed at an intended attack angle.

  18. Project SQUID: The Princeton Pilot Variable Density Supersonic Wind Tunnel

    DTIC Science & Technology

    1948-05-29

    adiabatic expansion of the air in tanks, neglecting the Joule- Thomson effect . • With the pressure drops experienced across the reducing-regulator...valves, the Joule- Thomson effect is quite large. It accounts for a drop öf approximately forty-five degrees Fahrenheit at the start of a run with a

  19. Wind tunnel tests of an 0.019-scale space shuttle integrated vehicle -2A configuration (model 14-OTS) in the NASA Ames 8 X 7 foot unitary wind tunnel, volume 3. [cold jet gas plumes and pressure distribution

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

    Tests were conducted to determine the effects of cold jet gas plumes on (1) the integrated vehicle longitudinal and lateral-directional force data, (2) exposed wing hinge moment, (3) wing pressure distributions, (4) orbiter MPS external pressure distributions, and (5) model base pressures. An investigation was undertaken to determine the similarity between solid and gaseous plumes; fluorescent oil flow visualization studies were also conducted. Tabulated data listings are included.

  20. Wind tunnel tests of an 0.019-scale space shuttle integrated vehicle -2A configuration (model 14-OTS) in the NASA Ames 8 X 7 foot unitary wind tunnel, volume 2. [cold jet gas plumes and pressure distribution

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

    The purpose of the test was to determine the effects of cold jet gas plumes on (1) the integrated vehicle longitudinal and lateral-directional force data, (2) exposed wing hinge moment, (3) wing pressure distributions, (4) orbiter MPS external pressure distributions, and (5) model base pressures. An investigation was undertaken to determine the similarity between solid and gaseous plumes; fluorescent oil flow visualization studies were also conducted. Plotted wing pressure data is tabulated.

  1. 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.

  2. 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.

  3. 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

  4. Findings from the Supersonic Qualification Program of the Mars Science Laboratory Parachute System

    NASA Technical Reports Server (NTRS)

    Sengupta, Anita; Steltzner, Adam; Witkowski, Allen; Candler, Graham; Pantano, Carlos

    2009-01-01

    supersonic wind tunnel testing of a rigid DGB parachute with entry-vehicle to validate two high fidelity computational fluid dynamics (CFD) tools. The computer codes utilized Large Eddy Simulation and Detached Eddy Simulation numerical approaches to accurately capture the turbulent wake of the entry vehicle and its coupling to the parachute bow-shock. The second phase was the development of fluid structure interaction (FSI) computational tools to predict parachute response to the supersonic flow field. The FSI development included the integration of the CFD from the first phase with a finite element structural model of the parachute membrane and cable elements. In this phase, a 4% of full-scale supersonic flexible parachute test program was conducted to provide validation data to the FSI code and an empirical dataset of the MSL parachute in a flight-like environment. The final phase is FSI simulations of the full-scale MSL parachute in a Mars type deployment. Findings from this program will be presented in terms of code development and validation, empirical findings from the supersonic testing, and drag performance during supersonic operation.

  5. A review and development of correlations for base pressure and base heating in supersonic flow

    SciTech Connect

    Lamb, J.P.; Oberkampf, W.L.

    1993-11-01

    A comprehensive review of experimental base pressure and base heating data related to supersonic and hypersonic flight vehicles has been completed. Particular attention was paid to free-flight data as well as wind tunnel data for models without rear sting support. Using theoretically based correlation parameters, a series of internally consistent, empirical prediction equations has been developed for planar and axisymmetric geometries (wedges, cones, and cylinders). These equations encompass the speed range from low supersonic to hypersonic flow and laminar and turbulent forebody boundary layers. A wide range of cone and wedge angles and cone bluntness ratios was included in the data base used to develop the correlations. The present investigation also included preliminary studies of the effect of angle of attack and specific-heat ratio of the gas.

  6. Low flight speed acoustic results for a supersonic inlet with auxiliary inlet doors

    NASA Technical Reports Server (NTRS)

    Woodward, R. P.; Glaser, F. W.; Lucas, J. G.

    1982-01-01

    A model supersonic inlet with auxiliary inlet doors and bounday layer bleeds was acoustically tested in simulated low speed flight up to Mach 0.2 in the NASA Lewis 9x15 Anechoic Wind Tunnel and statically in the NASA Lewis Anechoic Chamber. A JT8D refan model was used as the noise source. Data were also taken for a CTOL inlet and for an annular inlet with simulated centerbody support struts. Inlet operation with open auxiliary doors increased the blade passage tone by about 10 dB relative to the closed door configuration although noise radiation was primarily through the main inlet rather than the doors. Numerous strong spikes in the noise spectra were associated with the bleed system, and were strongly affected by the centerbody location. The supersonic inlet appeared to suppress multiple pure tone (MPT) generation at the fan source. Inlet length and the presence of support struts were shown not to cause this MPT suppression.

  7. Bibliography of Supersonic Cruise Research (SCR) program from 1977 to mid-1980

    NASA Technical Reports Server (NTRS)

    Hoffman, S.

    1980-01-01

    The supersonic cruise research (SCR) program, initiated in July 1972, includes system studies and the following disciplines: propulsion, stratospheric emission impact, structures and materials, aerodynamic performance, and stability and control. In a coordinated effort to provide a sound basis for any future consideration that may be given by the United States to the development of an acceptable commercial supersonic transport, integration of the technical disciplines was undertaken, analytical tools were developed, and wind tunnel, flight, and laboratory investigations were conducted. The present bibliography covers the time period from 1977 to mid-1980. It is arranged according to system studies and the above five SCR disciplines. There are 306 NASA reports and 135 articles, meeting papers, and company reports cited.

  8. Analytical and experimental performance of optimal controller designs for a supersonic inlet

    NASA Technical Reports Server (NTRS)

    Zeller, J. R.; Lehtinen, B.; Geyser, L. C.; Batterton, P. G.

    1973-01-01

    The techniques of modern optimal control theory were applied to the design of a control system for a supersonic inlet. The inlet control problem was approached as a linear stochastic optimal control problem using as the performance index the expected frequency of unstarts. The details of the formulation of the stochastic inlet control problem are presented. The computational procedures required to obtain optimal controller designs are discussed, and the analytically predicted performance of controllers designed for several different inlet conditions is tabulated. The experimental implementation of the optimal control laws is described, and the experimental results obtained in a supersonic wind tunnel are presented. The control laws were implemented with analog and digital computers. Comparisons are made between the experimental and analytically predicted performance results. Comparisons are also made between the results obtained with continuous analog computer controllers and discrete digital computer versions.

  9. Wall pressure fluctuations in the reattachment region of a supersonic free shear layer

    NASA Technical Reports Server (NTRS)

    Shen, Zhen-Hua; Smith, Douglas R.; Smith, Alexander J.

    1990-01-01

    Results are presented of a study designed to examine the unsteady shock wave interaction formed in the reattachment zone of a separated supersonic flow in which the separation occurred in the absence of a separation shock. Wall pressure fluctuation measurements were conducted in the Princeton University 203 mm-203 mm high-Reynolds number blowdown supersonic wind tunnel at a freestream Mach number of 2.92 and a unit Re number of 6.7 x 10 to the 7th/m. Reattachment accurred on a 20-deg ramp. It was found that, contrary to previous observations of such flow, the reattachment region was highly unsteady and the pressure fluctuations were significant. The overall behavior of the wall pressure loading was similar in scale and magnitude to the unsteadiness of the wall pressure field in compression ramp flows at the same Mach number.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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:

  17. An Aileron Flutter Experiment and Analysis Using Semi-Span Model for the Small Supersonic Experimental Aircraft

    NASA Astrophysics Data System (ADS)

    Saitoh, Kenichi; Tamayama, Masato; Kikuchi, Takao; Machida, Shigeru; Nakamichi, Jiro

    This paper reports a wind-tunnel experiment and analysis that have been conducted under the National Experimental Airplane for Supersonic Transports (NEXST-1) project of JAXA. In order to perform the flight experiment, the design of the vehicle was examined from the stand point of aeroelasticity. The aileron buzz as well as flutter was of much concern for its aileron system on the main wing. Therefore, both wind-tunnel test and analysis were carried out by using a semi-span model with fuselage. Although the buzz was not observed in the test, damping responses of the aileron rotation mode were obtained. Critical damping was observed in supersonic flow, that meant a buzz could occur in ``region C'' of Lambourne's classification. Linear unsteady aerodynamic analysis is applicable to this type of buzz and the characteristics of the buzz of the model is discussed.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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

  5. 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.

  6. Results for the hybrid laminar flow control experiment conducted in the NASA Langley 8-foot transonic pressure tunnel on a 7-foot chord model

    NASA Technical Reports Server (NTRS)

    Bobbitt, Percy J.; Ferris, James C.; Harvey, William D.; Goradia, Suresh H.

    1992-01-01

    A description is given of the development of, and results from, the hybrid laminar flow control (HLFC) experiment conducted in the NASA LaRC 8 ft Transonic Pressure Tunnel on a 7 ft chord, 23 deg swept model. The methods/codes used to obtain the contours of the HLFC model surface and to define the suction requirements are outlined followed by a discussion of the model construction, suction system, instrumentation, and some example results from the wind tunnel tests. Included in the latter are the effects of Mach number, suction level, and the extent of suction. An assessment is also given of the effect of the wind tunnel environment on the suction requirements. The data show that, at or near the design Mach number, large extents of laminar flow can be achieved with suction mass flows over the first 25 percent, or less, of the chord. Top surface drag coefficients with suction extending from the near leading edge to 20 percent of the chord were approximately 40 percent lower than those obtained with no suction. The results indicate that HLFC can be designed for transonic speeds with lift and drag coefficients approaching those of LFC designs but with much smaller extents and levels of suction.

  7. Development of a computational model for predicting solar wind flows past nonmagnetic terrestrial planets

    NASA Technical Reports Server (NTRS)

    Stahara, S. S.; Spreiter, J. R.

    1983-01-01

    A computational model for the determination of the detailed plasma and magnetic field properties of the global interaction of the solar wind with nonmagnetic terrestrial planetary obstacles is described. The theoretical method is based on an established single fluid, steady, dissipationless, magnetohydrodynamic continuum model, and is appropriate for the calculation of supersonic, super-Alfvenic solar wind flow past terrestrial ionospheres.

  8. Analysis of a stability valve system for extending the dynamic range of a supersonic inlet

    NASA Technical Reports Server (NTRS)

    Webb, J. A., Jr.; Dustin, M. O.

    1975-01-01

    A stability valve system designed for a full-scale, flight, supersonic, mixed-compression inlet was modeled dynamically by using analog computer techniques. The system uses poppet valves mounted in the inlet cowl to bypass airflow and augments the inlet shock position control system by preventing unstarts caused by high-frequency perturbations. The model was used as a design aid to investigate the effects of varying both the physical configurations of the valve and the flight and wind tunnel conditions. Results of the analysis indicate that the stability valve will provide a bandpass operation of 1 hertz to 17 hertz.

  9. A Status Review of the Commercial Supersonic Technology (CST) Aeroservoelasticity (ASE) Project

    NASA Technical Reports Server (NTRS)

    Silva, Walter A.; Sanetrik, Mark D.; Chwalowski, Pawel; Funk, Christy; Keller, Donald F.; Ringertz, Ulf

    2016-01-01

    An overview of recent progress regarding the computational aeroelastic and aeroservoelastic (ASE) analyses of a low-boom supersonic configuration is presented. The overview includes details of the computational models developed to date with a focus on unstructured CFD grids, computational aeroelastic analyses, sonic boom propagation studies that include static aeroelastic effects, and gust loads analyses. In addition, flutter boundaries using aeroelastic Reduced-Order Models (ROMs) are presented at various Mach numbers of interest. Details regarding a collaboration with the Royal Institute of Technology (KTH, Stockholm, Sweden) to design, fabricate, and test a full-span aeroelastic wind-tunnel model are also presented.

  10. An experimental and analytical study of flow through a supersonic open channel with contoured floor

    NASA Technical Reports Server (NTRS)

    Saheli, F. P.; Dunn, B.; Marrs, K.; Kumar, A.; Peery, K. M.

    1984-01-01

    A wind tunnel experiment was performed to study the characteristics of supersonic airflow (M(infinity) = 2.5-3.86) through an open channel with a contoured floor. The measured static pressures along the centerline of the channel floor exhibited an unexpected rise at the end of the channel. Complex three-dimensional interactions of compression and expansion waves within the channel coupled with external flow perturbations caused by model/tunnel wall interference were the suspected sources of this flow behavior. Three-dimensional inviscid flow analysis procedures were used to investigate and explain this phenomenon. The results of the computations and the experiment are presented and discussed.

  11. Experimental Study on the Control of the Supersonic Axisymmetric Intake under the Acceleration/Deceleration Conditions

    NASA Astrophysics Data System (ADS)

    Kojima, Takayuki; Sato, Tetsuya; Tanatsugu, Nobuhiro; Enomoto, Yoshinari

    A control system of variable geometry mixed compression axisymmetric intake is experimentally studied at ONERA S3 supersonic wind tunnel. The acceleration/deceleration of the space plane is simulated by changing the free stream velocity. The intake is successfully controlled with 90% of the maximum total pressure recovery and mass capture ratio. In this experiment, two subjects about control of axisymmetric intake are also cleared. First, the effect of the trapping of the terminal shock by bleed holes causes the disturbances in the terminal shock control system. Second, a special compression form change operation is necessary when the intake compression form change from all external compression to mixed compression.

  12. Pressure Measurement in Supersonic Air Flow by Differential Absorptive Laser-Induced Thermal Acoustics

    NASA Technical Reports Server (NTRS)

    Hart, Roger C.; Herring, Gregory C.; Balla, Robert J.

    2007-01-01

    Nonintrusive, off-body flow barometry in Mach-2 airflow has been demonstrated in a large-scale supersonic wind tunnel using seedless laser-induced thermal acoustics (LITA). The static pressure of the gas flow is determined with a novel differential absorption measurement of the ultrasonic sound produced by the LITA pump process. Simultaneously, stream-wise velocity and static gas temperature of the same spatially-resolved sample volume were measured with this nonresonant time-averaged LITA technique. Mach number, temperature and pressure have 0.2%, 0.4%, and 4% rms agreement, respectively, in comparison with known free-stream conditions.

  13. Pressure measurement in supersonic air flow by differential absorptive laser-induced thermal acoustics.

    PubMed

    Hart, Roger C; Herring, G C; Balla, R Jeffrey

    2007-06-15

    Nonintrusive, off-body flow barometry in Mach 2 airflow has been demonstrated in a large-scale supersonic wind tunnel using seedless laser-induced thermal acoustics (LITA). The static pressure of the gas flow is determined with a novel differential absorption measurement of the ultrasonic sound produced by the LITA pump process. Simultaneously, the streamwise velocity and static gas temperature of the same spatially resolved sample volume were measured with this nonresonant time-averaged LITA technique. Mach number, temperature, and pressure have 0.2%, 0.4%, and 4% rms agreement, respectively, in comparison with known free-stream conditions.

  14. Spatially Developing Secondary Instabilities and Attachment Line Instability in Supersonic Boundary Layers

    NASA Technical Reports Server (NTRS)

    Li, Fei; Choudhari, Meelan M.

    2008-01-01

    This paper reports on progress towards developing a spatial stability code for compressible shear flows with two inhomogeneous directions, such as crossflow dominated swept-wing boundary layers and attachment line flows. Certain unique aspects of formulating a spatial, two-dimensional eigenvalue problem for the secondary instability of finite amplitude crossflow vortices are discussed. A primary test case used for parameter study corresponds to the low-speed, NLF-0415(b) airfoil configuration as tested in the ASU Unsteady Wind Tunnel, wherein a spanwise periodic array of roughness elements was placed near the leading edge in order to excite stationary crossflow modes with a specified fundamental wavelength. The two classes of flow conditions selected for this analysis include those for which the roughness array spacing corresponds to either the naturally dominant crossflow wavelength, or a subcritical wavelength that serves to reduce the growth of the naturally excited dominant crossflow modes. Numerical predictions are compared with the measured database, both as indirect validation for the spatial instability analysis and to provide a basis for comparison with a higher Reynolds number, supersonic swept-wing configuration. Application of the eigenvalue analysis to the supersonic configuration reveals that a broad spectrum of stationary crossflow modes can sustain sufficiently strong secondary instabilities as to potentially cause transition over this configuration. Implications of this finding for transition control in swept wing boundary layers are examined. Finally, extension of the spatial stability analysis to supersonic attachment line flows is also considered.

  15. Flow control of micro-ramps on supersonic forward-facing step flow

    NASA Astrophysics Data System (ADS)

    Qing-Hu, Zhang; Tao, Zhu; Shihe, Yi; Anping, Wu

    2016-05-01

    The effects of the micro-ramps on supersonic turbulent flow over a forward-facing step (FFS) was experimentally investigated in a supersonic low-noise wind tunnel at Mach number 3 using nano-tracer planar laser scattering (NPLS) and particle image velocimetry (PIV) techniques. High spatiotemporal resolution images and velocity fields of supersonic flow over the testing model were captured. The fine structures and their spatial evolutionary characteristics without and with the micro-ramps were revealed and compared. The large-scale structures generated by the micro-ramps can survive the downstream FFS flowfield. The micro-ramps control on the flow separation and the separation shock unsteadiness was investigated by PIV results. With the micro-ramps, the reduction in the range of the reversal flow zone in streamwise direction is 50% and the turbulence intensity is also reduced. Moreover, the reduction in the average separated region and in separation shock unsteadiness are 47% and 26%, respectively. The results indicate that the micro-ramps are effective in reducing the flow separation and the separation shock unsteadiness. Project supported by the National Natural Science Foundation of China (Grant Nos. 11172326 and 11502280).

  16. The Effect of Micro-ramps on Supersonic Flow over a Forward-Facing Step

    NASA Astrophysics Data System (ADS)

    Zhang, Qing-Hu; Yi, Shi-He; Zhu, Yang-Zhu; Chen, Zhi; Wu, Yu

    2013-04-01

    The effect of micro-ramp control on fully developed turbulent flow over a forward-facing step (FFS) is investigated in a supersonic low-noise wind tunnel at Mach number 3 using nano-tracer planar laser scattering (NPLS) and supersonic particle image velocimetry (PIV) techniques. High spatiotemporal resolution images and the average velocity profiles of supersonic flow over the FFS with and without the control of the micro-ramps are captured. The fine structures of both cases, including the coherent structures of fully developed boundary layer and the large-scale hairpin-like vortices originated from the micro-ramps as well as the interaction of shock waves with the large-scale structures, are revealed and compared. Based on the time-correlation images, the temporal and spatial evolutionary characteristics of the coherent structures are investigated. It is beneficial to understand the dynamic mechanisms of the separated flow and the control mechanisms of the micro-ramps. The size of the separation region is determined by the NPLS and PIV. The results indicate that the control of the micro-ramps is capable of delaying the separation and diminishing the extent of recirculation zone.

  17. Investigations and Experiments in the Guidonia Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Ferri, Antonio

    1939-01-01

    This paper is a presentation of the experiments and equipment used in investigations at the Guidonia wind tunnel. The equipment consisted of: a number of subsonic and supersonic cones, an aerodynamic balance, and optical instruments operating on the Schlieren and interferometer principle.

  18. Calculation of solar wind flows about terrestrial planets

    NASA Technical Reports Server (NTRS)

    Stahara, S. S.; Spreiter, J. R.

    1982-01-01

    A computational model was developed for the determination of the plasma and magnetic field properties of the global interaction of the solar wind with terrestrial planetary magneto/ionospheres. The theoretical method is based on an established single fluid, steady, dissipationless, magnetohydrodynamic continuum model, and is appropriate for the calculation of supersonic, super Alfvenic solar wind flow past terrestrial planets. A summary is provided of the important research results.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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

  5. 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.

  6. 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.

  7. 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

  8. 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

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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...

  15. 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.

  16. 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.

  17. 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.

  18. Nuclear winds and the narrow-line emission from active galaxies

    NASA Technical Reports Server (NTRS)

    Smith, Steven J.

    1993-01-01

    This paper discusses active galaxy narrow-line emission in the context of a supersonic wind generated by the active nucleus, a nuclear wind. Wind acceleration can naturally produce the inferred narrow-line cloud velocities as well as those of the broad-line and broad absorption line clouds with only a weak dependence on the parameters of the active nucleus. The primary obstacle to wind acceleration of emission-line clouds is the destructive effect of instabilities; however, the stability and structure of wind-confined clouds remain areas of future research. Observations, particularly of the Seyfert galaxy NGC 1068, are interpreted in the context of nuclear winds.

  19. Supersonic Aerodynamic Characteristics of Blunt Body Trim Tab Configurations

    NASA Technical Reports Server (NTRS)

    Korzun, Ashley M.; Murphy, Kelly J.; Edquist, Karl T.

    2013-01-01

    Trim tabs are aerodynamic control surfaces that can allow an entry vehicle to meet aerodynamic performance requirements while reducing or eliminating the use of ballast mass and providing a capability to modulate the lift-to-drag ratio during entry. Force and moment data were obtained on 38 unique, blunt body trim tab configurations in the NASA Langley Research Center Unitary Plan Wind Tunnel. The data were used to parametrically assess the supersonic aerodynamic performance of trim tabs and to understand the influence of tab area, cant angle, and aspect ratio. Across the range of conditions tested (Mach numbers of 2.5, 3.5, and 4.5; angles of attack from -4deg to +20deg; angles of sideslip from 0deg to +8deg), the effects of varying tab area and tab cant angle were found to be much more significant than effects from varying tab aspect ratio. Aerodynamic characteristics exhibited variation with Mach number and forebody geometry over the range of conditions tested. Overall, the results demonstrate that trim tabs are a viable approach to satisfy aerodynamic performance requirements of blunt body entry vehicles with minimal ballast mass. For a 70deg sphere-cone, a tab with 3% area of the forebody and canted approximately 35deg with no ballast mass was found to give the same trim aerodynamics as a baseline model with ballast mass that was 5% of the total entry mass.

  20. 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

  1. Assessment of CFD-based Response Surface Model for Ares I Supersonic Ascent Aerodynamics

    NASA Technical Reports Server (NTRS)

    Hanke, Jeremy L.

    2011-01-01

    The Ascent Force and Moment Aerodynamic (AFMA) Databases (DBs) for the Ares I Crew Launch Vehicle (CLV) were typically based on wind tunnel (WT) data, with increments provided by computational fluid dynamics (CFD) simulations for aspects of the vehicle that could not be tested in the WT tests. During the Design Analysis Cycle 3 analysis for the outer mold line (OML) geometry designated A106, a major tunnel mishap delayed the WT test for supersonic Mach numbers (M) greater than 1.6 in the Unitary Plan Wind Tunnel at NASA Langley Research Center, and the test delay pushed the final delivery of the A106 AFMA DB back by several months. The aero team developed an interim database based entirely on the already completed CFD simulations to mitigate the impact of the delay. This CFD-based database used a response surface methodology based on radial basis functions to predict the aerodynamic coefficients for M > 1.6 based on only the CFD data from both WT and flight Reynolds number conditions. The aero team used extensive knowledge of the previous AFMA DB for the A103 OML to guide the development of the CFD-based A106 AFMA DB. This report details the development of the CFD-based A106 Supersonic AFMA DB, constructs a prediction of the database uncertainty using data available at the time of development, and assesses the overall quality of the CFD-based DB both qualitatively and quantitatively. This assessment confirms that a reasonable aerodynamic database can be constructed for launch vehicles at supersonic conditions using only CFD data if sufficient knowledge of the physics and expected behavior is available. This report also demonstrates the applicability of non-parametric response surface modeling using radial basis functions for development of aerodynamic databases that exhibit both linear and non-linear behavior throughout a large data space.

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

    NASA Technical Reports Server (NTRS)

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

    1979-01-01

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

  3. Supersonic aeroelastic instability results for a NASP-like wing model

    NASA Technical Reports Server (NTRS)

    Cole, Stanley R.; Florance, James R.; Thomason, Lee B.; Spain, Charles V.; Bullock, Ellen P.

    1993-01-01

    Two wing-alone wind-tunnel models were tested in the NASA Langley Unitary Plan Wind Tunnel facility to study the static divergence behavior of such configurations and to provide a data base for correlation with supersonic analytical predictions. One model had a four percent maximum thickness airfoil and the other had an eight-percent maximum thickness airfoil. The wing models had low aspect ratios and highly swept leading edges. Results show that decreasing airfoil thickness, moving the wing-pivot location upstream, or increasing the pitch-pivot stiffness have the beneficial effect of increasing the divergence dynamic pressures. The calculations accurately predicted the trend of experimental divergence dynamic pressure with Mach number.

  4. Semikinetic and generalized transport models of the polar and solar winds

    NASA Technical Reports Server (NTRS)

    Demars, H. G.; Schunk, R. W.

    1992-01-01

    In order to gain a better understanding of the relative merits of the transport and kinetic approaches to modeling thermal space plasmas, a comparison is presented of a transport (bi-Maxwellian-based 16-moment equations) and a semikinetic description of supersonic flow in the solar wind and also of both supersonic and subsonic flows in the polar wind for 'steady state' conditions. The study shows: remarkable agreement between the two models for supersonic collisionless flows, even for the higher-order moments; the inadequacy of the semikinetic approach for modeling subsonic flows; and the superiority of the 16-moment transport over the semikinetic approach for modeling the solar wind. Further evidence that the bi-Maxwellian-based transport equations are a useful tool for studying 'thermal' space plasmas that develop non-Maxwellian features is provided.

  5. Wind Simulation

    SciTech Connect

    Walker, Howard Andrew

    2008-12-31

    The Software consists of a spreadsheet written in Microsoft Excel that provides an hourly simulation of a wind energy system, which includes a calculation of wind turbine output as a power-curve fit of wind speed.

  6. 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)

  7. 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.

  8. 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.

  9. 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,...

  10. 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,...

  11. 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,...

  12. 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,...

  13. 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,...

  14. 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.

  15. 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.

  16. Historical Overview and Recent Improvements at the NASA Glenn Research Center 8x6 9x15 Wind Tunnel Complex

    NASA Technical Reports Server (NTRS)

    Dussling, Joseph John

    2015-01-01

    A brief history of the 8x6 Supersonic Wind Tunnel (SWT) and 9x15 Low Speed Wind Tunnel (LSWT) at NASA Glenn Research Center, Cleveland, Ohio is presented along with current capabilities and plans for future upgrades within the facility.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. Finite-rate water condensation in combustion-heated wind tunnels

    NASA Technical Reports Server (NTRS)

    Erickson, Wayne D.; Mall, Gerald H.; Prabhu, Ramadas K.

    1988-01-01

    A quasi-one-dimensional method for computing finite rate nucleation and droplet growth of water in a supersonic expansion of combustion products is presented. Sample computations are included for the Langley 8 foot High Temperature Tunnel, but the method can also be applied to other combustion heated wind tunnels. The sample results indicate that the free stream static pressure can be in the range of 25 to 60 percent greater than that computed for isentropic nozzle flow without water condensation. The method provides a tool for examining the effects of water condensation on static state properties and velocity of the supersonic stream in combustion heated wind tunnels.

  3. 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

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. Erosion: Wind

    Technology Transfer Automated Retrieval System (TEKTRAN)

    Wind erosion refers to the detachment, transport and deposition of sediment by wind. It is a dynamic, physical process where loose, dry, bare soils are transported by strong winds. Wind erosion is a soil degrading process that affects over 500 million ha of land worldwide and creates between 500 an...

  12. Meteorology (Wind)

    Atmospheric Science Data Center

    2014-09-25

    Wind speed at 50 m (m/s) The average and percent difference minimum and ... are given.   Percent of time for ranges of wind speed at 50 m (percent) Percentage [frequency] of time that wind ... be adjusted to heights from 10 to 300 meters using the Gipe power law. Wind speeds may be adjusted for different terrain by selecting from ...

  13. Supersonic aeroelastic instability results for a NASP-like wing model

    NASA Technical Reports Server (NTRS)

    Cole, Stanley R.; Florance, James R.; Thomason, Lee B.; Spain, Charles V.; Bullock, Ellen P.

    1993-01-01

    An experimental study and an analytical study have been conducted to examine static divergence for hypersonic-vehicle wing models at supersonic conditions. A supersonic test in the Langley Unitary Plan Wind Tunnel facility was conducted for two wind-tunnel models. These models were nearly identical with the exception of airfoil shape. One model had a four-percent maximum thickness airfoil and the other model had an eight-percent maximum thickness airfoil. The wing models had low-aspect ratios and highly swept leading edges. The all-movable wing models were supported by a single-pivot mechanism along the wing root. For both of the wind-tunnel models, configuration changes could be made in the wing-pivot location along the wing root and in the wing-pivot pitch stiffness. Three divergence conditions were measured for the four-percent thick airfoil model in the Mach number range of 2.6 to 3.6 and one divergence condition was measured for the eight-percent thick airfoil model at a Mach number of 2.9. Analytical divergence calculations were made for comparison with experimental results and to evaluate the parametric effects of wing-pivot stiffness, wing-pivot location, and airfoil thickness variations. These analyses showed that decreasing airfoil thickness, moving the wing-pivot location upstream, or increasing the pitch-pivot stiffness have the beneficial effect of increasing the divergence dynamic pressures. The calculations predicted the trend of experimental divergence dynamic pressure with Mach number accurately; however, the calculations were approximately 25 percent conservative with respect to dynamic pressure.

  14. 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.

  15. Colliding Winds and Infall Confinement of Protostellar Winds

    NASA Astrophysics Data System (ADS)

    Wilkin, F. P.; Stahler, S. W.

    1996-12-01

    We consider the hydrodynamic influence of the protostellar environment on the confinement of protostellar winds by modeling the collision of a stellar wind with anisotropic infall from a rotating cloud core. The stellar wind and infall collide supersonically, yielding a thin, radiative shocked shell. We calculate the structure of the shocked layer using a quasi-steady formulation similar to that used for stellar-wind bow shocks. At any time the shell structure is determined by simple conservation laws involving the fluxes of mass, momentum, and angular momentum of the incident wind and infall material, and including the effects of rotation and the stellar gravity. The gravitational forces are critical to the shell structure because in order for the infall ram pressure to be comparable to the wind ram pressure and allow normal force balance, the shell must be deep inside the gravitational well. We find that self-consistent, bowshock-like steady-state solutions only exist when the stellar gravity is included. There are two solution families (an inner and an outer solution), which both move outward in quasi-steady fashion with time due to the evolving infall structure. These solutions will be described and compared to new analytic solutions for bow shocks[1] and colliding winds[2] from spherically-symmetric sources. Fully non-steady calculations of the evolution will be presented in the future. [1] Wilkin, F.P. 1996, ApJ, 459, L31. [2] Canto, J., Raga, A.C., & Wilkin, F.P. 1996, ApJ, 469, 729

  16. 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.

  17. Methodology for the Design of Streamline-Traced External-Compression Supersonic Inlets

    NASA Technical Reports Server (NTRS)

    Slater, John W.

    2014-01-01

    A design methodology based on streamline-tracing is discussed for the design of external-compression, supersonic inlets for flight below Mach 2.0. The methodology establishes a supersonic compression surface and capture cross-section by tracing streamlines through an axisymmetric Busemann flowfield. The compression system of shock and Mach waves is altered through modifications to the leading edge and shoulder of the compression surface. An external terminal shock is established to create subsonic flow which is diffused in the subsonic diffuser. The design methodology was implemented into the SUPIN inlet design tool. SUPIN uses specified design factors to design the inlets and computes the inlet performance, which includes the flow rates, total pressure recovery, and wave drag. A design study was conducted using SUPIN and the Wind-US computational fluid dynamics code to design and analyze the properties of two streamline-traced, external-compression (STEX) supersonic inlets for Mach 1.6 freestream conditions. The STEX inlets were compared to axisymmetric pitot, two-dimensional, and axisymmetric spike inlets. The STEX inlets had slightly lower total pressure recovery and higher levels of total pressure distortion than the axisymmetric spike inlet. The cowl wave drag coefficients of the STEX inlets were 20% of those for the axisymmetric spike inlet. The STEX inlets had external sound pressures that were 37% of those of the axisymmetric spike inlet, which may result in lower adverse sonic boom characteristics. The flexibility of the shape of the capture cross-section may result in benefits for the integration of STEX inlets with aircraft.

  18. Flow Visualization of a Scramjet Inlet - Isolator Model in Supersonic Flow

    NASA Astrophysics Data System (ADS)

    Seckin, S.; Yuceil, K. B.

    2013-04-01

    Understanding the physical mechanisms and having insight to the complex flowfield involving unstart phenomena in supersonic inlets has gained considerable attention especially in the area of scramjet inlet/isolator aerothermodynamics. In this study, Schlieren visualization and computational analysis of shock wave structures in ramjet/scramjet inlet/isolator models in supersonic flow have been performed. Experiments were performed in the supersonic wind tunnel at the Trisonic Research Laboratory in Istanbul Technical University. The test section floor and the existing mechanism underneath have been modified to be able to mount the designed inlet/isolator model on the floor of the test section. The inlet/isolator model with a 12- degree compression ramp is investigated at Mach 2 both computationally and experimentally. Computations were performed using Star-CCM+ software to investigate shock wave structures in and around the three dimensional inlet/isolator model as mounted on the test section floor as a guide for designing the experimental model. In the results, the effects of shock wave - boundary layer interactions with flow separations with were observed. Ensemble average of the density distributions on a series of planes from one side wall to the other from the CFD results agreed well with the Schlieren images obtained experimentally. The structure of the shock waves and angles obtained from the Schlieren images agree quite well with those obtained from the CFD results. The effects of lambda-shock formations which indicate possible boundary layer separations, reflections of shock waves, and shock wave - boundary layer interactions on inlet unstart phenomena have been discussed. In order to investigate inlet unstart mechanism further, different experimental setups have been suggested for future work.

  19. Behavior of Boundary Layer in Supersonic Flow with Applied Lorentz Force

    NASA Astrophysics Data System (ADS)

    Udagawa, Keisuke; Saito, Shinya; Kawaguchi, Kenji; Tomioka, Sadatake; Yamasaki, Hiroyuki

    Experimental study on behavior of boundary layer in supersonic flow with applied Lorentz force was carried out. In the experiment, Mach 1.5 supersonic wind tunnel driven by a shock-tube was used. At the test section, the current from the external DC power supply and the magnetic field of 2.4 Tesla were applied to the boundary layer developing on the bottom wall. Argon seeded with cesium was used as an electrically conducting gas. Effect of the direction of the Lorentz force on static pressure distribution was investigated, and the remarkable increase of static pressure at the test section was observed for the decelerating Lorentz force. It is noted that the acceleration of the flow inside the boundary layer was demonstrated for the first time without accelerating the main flow when the accelerating Lorentz force was applied. At the same time, the acceleration efficiency defined by a ratio of work done by the Lorentz force to energy input into the flow was found 54-61%. These results have suggested the possibility of the boundary layer separation control by applying the accelerating Lorentz force. In the case of the decelerating Lorentz force, the significant reduction of Mach number was observed not only inside the boundary layer but also in the main flow. The reduction of Mach number could be ascribed to the growth of the boundary layer due to gas heating inside the boundary layer. When the direction of the current was changed, the difference of light emission from the discharge inside the boundary layer was observed, and this was due to the difference of the electromotive force induced in the supersonic flow.

  20. Vortex Generators in a Streamline-Traced, External-Compression Supersonic Inlet

    NASA Technical Reports Server (NTRS)

    Baydar, Ezgihan; Lu, Frank K.; Slater, John W.; Trefny, Charles J.

    2017-01-01

    Vortex generators within a streamline-traced, external-compression supersonic inlet for Mach 1.66 were investigated to determine their ability to increase total pressure recovery and reduce total pressure distortion. The vortex generators studied were rectangular vanes arranged in counter-rotating and co-rotating arrays. The vane geometric factors of interest included height, length, spacing, angle-of-incidence, and positions upstream and downstream of the inlet terminal shock. The flow through the inlet was simulated numerically through the solution of the steady-state, Reynolds-averaged Navier-Stokes equations on multi-block, structured grids using the Wind-US flow solver. The vanes were simulated using a vortex generator model. The inlet performance was characterized by the inlet total pressure recovery and the radial and circumferential total pressure distortion indices at the engine face. Design of experiments and statistical analysis methods were applied to quantify the effect of the geometric factors of the vanes and search for optimal vane arrays. Co-rotating vane arrays with negative angles-of-incidence positioned on the supersonic diffuser were effective in sweeping low-momentum flow from the top toward the sides of the subsonic diffuser. This distributed the low-momentum flow more evenly about the circumference of the subsonic diffuser and reduced distortion. Co-rotating vane arrays with negative angles-of-incidence or counter-rotating vane arrays positioned downstream of the terminal shock were effective in mixing higher-momentum flow with lower-momentum flow to increase recovery and decrease distortion. A strategy of combining a co-rotating vane array on the supersonic diffuser with a counter-rotating vane array on the subsonic diffuser was effective in increasing recovery and reducing distortion.

  1. 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

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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)

  19. A special method for finding body distortions that reduce the wave drag of wing and body combinations at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Lomax, Harvard; Heaslet, Max A

    1956-01-01

    For a given wing and supersonic Mach number, the problem of shaping an adjoining fuselage so that the combination will have a low wave drag is considered. Only fuselages that can be simulated by singularities (multipoles) distributed along the body axis are studied. However, the optimum variations of such singularities are completely specified in terms of the given wing geometry. An application is made to an elliptic wing having a biconvex section, a thickness-chord ratio equal to 0.05 at the root, and an aspect ratio equal to 3. A comparison of the theoretical results with a wind-tunnel experiment is also presented.

  20. Data Driven Model Development for the SuperSonic SemiSpan Transport (S(sup 4)T)

    NASA Technical Reports Server (NTRS)

    Kukreja, Sunil L.

    2011-01-01

    In this report, we will investigate two common approaches to model development for robust control synthesis in the aerospace community; namely, reduced order aeroservoelastic modelling based on structural finite-element and computational fluid dynamics based aerodynamic models, and a data-driven system identification procedure. It is shown via analysis of experimental SuperSonic SemiSpan Transport (S4T) wind-tunnel data that by using a system identification approach it is possible to estimate a model at a fixed Mach, which is parsimonious and robust across varying dynamic pressures.

  1. 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.

  2. Improved optical techniques for studying sonic and supersonic injection into Mach 3 flow

    NASA Astrophysics Data System (ADS)

    Buggele, Alvin E.; Seasholtz, Richard G.

    1997-11-01

    Filtered Rayleigh Scattering and shadowgraph flow visualization were used to characterize the penetration of helium or moist air injected transversely at several pressures to a Mach 3 flow in the NASA Lewis 3.81 inch by 10 inch continuous flow supersonic wind tunnel. This work is in support of the LOX augmented nuclear thermal rocket program. The present study used an injection-seeded, frequency doubled Nd:YAG pulsed laser to illuminate a transverse section of the injectant plume. Rayleigh scattered light was passed through an iodine absorption cell to suppress stray laser light and was imaged onto a cooled CCD camera. The scattering was based on condensation of water vapor in the injectant flow. Results are presented for various configurations of sonic and supersonic injector designs mounted primarily in the floor of the tunnel. Injectors studied include a single 0.25 inch diameter hole, five 0.112 inch diameter holes on 0.177 inch spacing, and a 7 degree half angle wedge. High speed shadowgraph flow visualization images were obtained with several video camera systems. Roof and floor static pressure data are presented several ways for the three configurations of injection designs with and without helium and/or air injection into Mach 3 flow.

  3. Aerodynamic characteristics of the orbital reentry vehicle experimental probe fins in a supersonic flow

    NASA Astrophysics Data System (ADS)

    Watanabe, Mitsunori; Sekine, Hideo; Tate, Atsushi; Noda, Junichi

    1994-04-01

    The aerodynamic characteristics of probe fins with a sweep angle of 60 deg, which are equipped on the Orbital Reentry Experiment (OREX) vehicle to measure the surrounding ionized gas temperature and electron number density distributions in the high temperature communication black out regions, have been measured in the supersonic wind tunnel of the National Aerospace Laboratory and compared with those of the fins of 0 deg sweep angles. Since the probes are to be embedded in the boundary layer where the local Mach number is less than 2.5 over the OREX surface at a hypersonic flight speed, the aerodynamic characteristics in supersonic regions are needed to estimate the rolling moments of fins caused by the error of the installation angles. The lift coefficient slope of the probe fins decreases as the Mach number increases, being less than the values for the 0 deg sweep fins. The drag coefficient depends highly on the sweep angle of the fins in Mach number regions less than 2.5.

  4. Observations of Shock Diffusion and Interactions in Supersonic Freestreams with Counterflowing Jets

    NASA Technical Reports Server (NTRS)

    Daso, Endwell O.; Pritchett, Victor E.; Wang, Ten-See; Blankson, Isiah M.; Auslender, Aaron H.

    2006-01-01

    One of the technical challenges in long-duration space exploration and interplanetary missions is controlled entry and re-entry into planetary and Earth atmospheres, which requires the dissipation of considerable kinetic energy as the spacecraft decelerates and penetrates the atmosphere. Efficient heat load management of stagnation points and acreage heating remains a technological challenge and poses significant risk, particularly for human missions. An innovative approach using active flow control concept is proposed to significantly modify the external flow field about the spacecraft in planetary atmospheric entry and re-entry in order to mitigate the harsh aerothermal environments, and significantly weaken and disperse the shock-wave system to reduce aerothermal loads and wave drag, as well as improving aerodynamic performance. To explore the potential benefits of this approach, we conducted fundamental experiments in a trisonic blow down wind tunnel to investigate the effects of counterflowing sonic and supersonic jets against supersonic freestreams to gain a better understanding of the flow physics of the interactions of the opposing flows and the resulting shock structure.

  5. A Quantitative Comparison of Leading-edge Vortices in Incompressible and Supersonic Flows

    NASA Technical Reports Server (NTRS)

    Wang, F. Y.; Milanovic, I. M.; Zaman, K. B. M. Q.

    2002-01-01

    When requiring quantitative data on delta-wing vortices for design purposes, low-speed results have often been extrapolated to configurations intended for supersonic operation. This practice stems from a lack of database owing to difficulties that plague measurement techniques in high-speed flows. In the present paper an attempt is made to examine this practice by comparing quantitative data on the nearwake properties of such vortices in incompressible and supersonic flows. The incompressible flow data are obtained in experiments conducted in a low-speed wind tunnel. Detailed flow-field properties, including vorticity and turbulence characteristics, obtained by hot-wire and pressure probe surveys are documented. These data are compared, wherever possible, with available data from a past work for a Mach 2.49 flow for the same wing geometry and angles-of-attack. The results indicate that quantitative similarities exist in the distributions of total pressure and swirl velocity. However, the streamwise velocity of the core exhibits different trends. The axial flow characteristics of the vortices in the two regimes are examined, and a candidate theory is discussed.

  6. Effects of plasma aerodynamic actuation on oblique shock wave in a cold supersonic flow

    NASA Astrophysics Data System (ADS)

    Wang, Jian; Li, Yinghong; Cheng, Bangqin; Su, Changbing; Song, Huimin; Wu, Yun

    2009-08-01

    Wedge oblique shock wave control using an arc discharge plasma aerodynamic actuator was investigated both experimentally and theoretically. Schlieren photography measurements in a small-scale short-duration supersonic wind tunnel indicated that the shock wave angle decreased and its start point shifted upstream with the plasma aerodynamic actuation. Also the shock wave intensity weakened, as shown by the decrease in the gas static pressure ratio of flow downstream and upstream of the shock wave. Moreover, the shock wave control effect was intensified when a static magnetic field was applied. Under test conditions of Mach 2.2, magnetic control and input voltage 3 kV, the start point of the shock wave shifted 4 mm upstream, while its angle and intensity decreased 8.6% and 8.8%, respectively. A thermal choking model was proposed to deduce the change laws of oblique shock wave control by surface arc discharge. The theoretical result was consistent with the experimental result, which demonstrated that the thermal choking model can effectively forecast the effect of plasma actuation on an oblique shock wave in a cold supersonic flow.

  7. Coherent Anti-Stokes Raman Spectroscopy (CARS) Measurements in Supersonic Combustors at NASA Langley Research Center

    NASA Technical Reports Server (NTRS)

    Danehy, Paul M.; OByrne, Sean B.; Tedder, Sarah A.; Cutler, Andrew D.

    2005-01-01

    This paper describes the recent use of coherent anti-Stokes Raman spectroscopy (CARS) to study supersonic combustion at NASA Langley Research Center. CARS is a nonlinear optical measurement technique used to measure temperature and species mole fractions remotely in harsh environments. A CARS system has been applied to two different combustor geometries at NASA Langley. Both experiments used the same vitiated wind-tunnel facility to create an air flow that simulates flight at Mach numbers of 6 and 7 for the combustor inlet and both experiments used hydrogen fuel. In the first experiment, the hydrogen was injected supersonically at a 30-degree angle with respect to the incoming flow. In the second experiment, the hydrogen was injected sonically at normal incidence. While these injection schemes produced significantly different flow features, the CARS method provided mean temperature, N2, O2 and H2 maps at multiple downstream locations for both. The primary aim of these measurements was to provide detailed flowfield information for computational fluid dynamics (CFD) code validation.

  8. Numerical Investigation of Aerodynamics of Canard-Controlled Missile Using Planar and Grid Tail Fins. Part 1. Supersonic Flow

    NASA Astrophysics Data System (ADS)

    DeSpirito, James; Vaughn, Milton E., Jr.; Washington, W. D.

    2002-09-01

    Viscous computational fluid dynamic simulations were used to predict the aerodynamic coefficients and flowfield around a generic canard-controlled missile configuration in supersonic flow. Computations were performed for Mach 1.5 and 3.0, at six angles of attack between 0 and 10, with 0 and 10 canard deflection, and with planar and grid tail fins, for a total of 48 cases. Validation of the computed results was demonstrated by the very good agreement between the computed aerodynamic coefficients and those obtained from wind tunnel measurements. Visualizations of the flowfield showed that the canard trailing vortices and downwash produced a low-pressure region on the starboard side of the missile that in turn produced an adverse side force. The pressure differential on the leeward fin produced by the interaction with the canard trailing vortices is primarily responsible for the adverse roll effect observed when planar fins are used. Grid tail fins improved the roll effectiveness of the canards at low supersonic speed. No adverse rolling moment was observed with no canard deflection, or at the higher supersonic speed for either tail fin type due to the lower intensity of the canard trailing vortices in these cases. Flow visualizations from the simulations performed in this study help in the understanding of the flow physics and can lead to improved canard and tail fin designs for missiles and rockets.

  9. 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...

  10. 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.

  11. 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.

  12. 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

  13. 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.

  14. Experimental study at low supersonic speeds of a missile concept having opposing wraparound tails

    NASA Technical Reports Server (NTRS)

    Allen, Jerry M.; Watson, Carolyn B.

    1994-01-01

    A wind-tunnel investigation has been performed at low supersonic speeds (at Mach numbers of 1.60, and 2.16) to evaluate the aerodynamic characteristics of a missile concept capable of being tube launched and controlled with a simple one-axis canard controller. This concept, which features an axisymmetric body with two planar canards and four wraparound tail fins arranged in opposing pairs, must be in rolling motion to be controllable in any radial plane with the planar canards. Thus, producing a constant rolling moment that is invariant with speed and attitude to provide the motion is desirable. Two tail-fin shaping designs, one shaved and one beveled, were evaluated for their efficiency in producing the needed rolling moments, and the results showed that the shaved fins were much more desirable for this task than the beveled fins.

  15. Modeling of a supersonic flow around a cylinder with a gas-permeable porous insert

    NASA Astrophysics Data System (ADS)

    Mironov, S. G.; Maslov, A. A.; Poplavskaya, T. V.; Kirilovskiy, S. V.

    2015-07-01

    Results of an experimental and numerical study of a supersonic (M∞ = 4.85) flow around a streamwise-aligned cylinder with a gas-permeable porous insert on the frontal face in the range of Reynolds numbers Re D = (0.1-2.0) · 105 are presented. The numerical study is performed by using the Ansys Fluent software system and a porous medium model based on a quadratic law of filtration. The parameters of the quadratic dependence are calculated on the basis of experimental data for an air flow in a porous material. Flow fields are obtained, and the wave drag of the model is calculated as a function of the porous insert length and the Reynolds number. Results of numerical simulations are compared with wind tunnel measurements.

  16. Vortex structure and breakup mechanism of gaseous jet in supersonic crossflow with laminar boundary layer

    NASA Astrophysics Data System (ADS)

    Zhao, Yanhui; Liang, Jianhan; Zhao, Yuxin

    2016-11-01

    Employing nano-particle planar laser scattering and particle image velocimetry technology, underexpanded jet in supersonic crossflow with laminar boundary layer is experimental investigated in a low noise wind tunnel. Instantaneous flow structures and average velocity distribution of jet plume are captured in experimental images. Horseshoe vortex system is dominated by oscillating and coalescing regime, contributing to vortex generation of jet shear layer. The "tilting-stretching-tearing" mechanism dominating in near field raises average fractal dimension. But vortex structures generated on the windward side of jet plume scatter in jet plume and dissipate gradually, which makes the vortexes break up from outside in near field and break down into small turbulence completely in far field.

  17. Development of an experiment for measuring film cooling performance in supersonic flows

    NASA Astrophysics Data System (ADS)

    Maqbool, Daanish

    This thesis describes the development of an experiment for acquiring supersonic film cooling performance data in canonical configurations suitable for code validation. A methodology for selecting appropriate experimental conditions is developed and used to select test conditions in the UMD atmospheric pressure wind tunnel that are relevant to film cooling conditions encountered in the J-2X rocket engine. A new technique for inferring wall heat flux with 10% uncertainty from temperature-time histories of embedded sensors is developed and implemented. Preliminary heat flux measurements on the uncooled upper wall and on the lower wall with the film cooling flow turned off suggest that RANS solvers using Menter's SST model are able to predict heat flux within 15% in the far-field (> 10 injection slot heights) but are very inaccurate in the near-field. However, more experiments are needed to confirm this finding. Preliminary Schlieren images showing the shear layer growth rate are also presented.

  18. Supersonic Flow Field Investigation Using a Fiber-optic based Doppler Global Velocimeter

    NASA Technical Reports Server (NTRS)

    Meyers, James F.; Lee, Joseph W.; Fletcher, Mark T.; Cavone, Angelo A.; AscencionGuerreroViramontes, J.

    2006-01-01

    A three-component fiber-optic based Doppler Global Velocimeter was constructed, evaluated and used to measure shock structures about a low-sonic boom model in a Mach 2 flow. The system was designed to have maximum flexibility in its ability to measure flows with restricted optical access and in various facilities. System layout is described along with techniques developed for production supersonic testing. System evaluation in the Unitary Plan Wind Tunnel showed a common acceptance angle of f4 among the three views with velocity measurement resolutions comparable with free-space systems. Flow field measurements of shock structures above a flat plate with an attached ellipsoid-cylinder store and a low-sonic boom model are presented to demonstrate the capabilities of the system during production testing.

  19. 3-D supersonic combustion experiments with hydrogen in V-trough

    NASA Technical Reports Server (NTRS)

    Friedberg, R. A.; Ahmed, A.

    1982-01-01

    Both supersonic blowdown and free jet wind tunnels were used in oil flow visualization and hydrogen gas burning tests of 60- and 90-deg V-shaped trough combustion characteristics. Oil dots showed that mass outflow from the reattachment region at the bottom of the V-trough dominated flow behavior, and occupied the trough's entire length and 10% of its cross-section. Intermingled vortex pairs were also found to superimpose themselves upon the outflow, and hydrogen burning within troughs reduced missile base drag by 25%. The troughs were aligned so that their long dimensions were parallel to the airflow, and the two top edges of the V section were mounted flush with respect to the upstream flat surfaces. Step heights tested ranged from 3.4 mm to 6.25 cm.

  20. The Partition Between Terminal Speed and Mass Loss: Thin, Thick, and Rotating Line-Driven Winds

    NASA Astrophysics Data System (ADS)

    Gayley, K. G.; Onifer, A. J.

    2003-01-01

    Steady-state supersonic line-driven winds are important contributors to wind-blown bubbles in star forming regions. The key input to the bubble in the energy-conserving phase is the wind kinetic-energy flux, which involves both the mass-loss rate and the terminal speed. However, these quantities are themselves self-consistent parameters of the line-driving process, so relate to each other and to the resulting wind optical depth. This complex interrelation between optical depth, mass-loss, and wind speed lies at the heart of line-driven wind theory. Drawing on the successes and insights of ``CAK'' theory, I will convey a simplified view of how to unite these processes using the concept of effective opacity, with attention to the ramifications for nonspherical nebular and wind-blown structures. Recent extensions to nongray optically thick environments such as Wolf-Rayet winds and supernovae are also discussed.

  1. Equatorial disk formation around rotating stars due to ram pressure confinement by the stellar wind

    NASA Technical Reports Server (NTRS)

    Bjorkman, J. E.; Cassinelli, J. P.

    1993-01-01

    The axisymmetric 2D supersonic solution of a rotating, radiation-driven stellar wind presently obtained by a simple approximation predicts the formation of a dense equatorial disk, when the star's rotation rate lies above a threshold value that depends on the ratio of the wind's terminal speed to the escape speed of the star. The disk is formed because the trajectories of the wind leaving the stellar surface at high latitudes carry it down to the equatorial plane; there, the material passes through a standing oblique shock atop the disk; it is therefore the ram pressure of the polar wind that compresses and confines the disk.

  2. WIND Validation Cases: Computational Study of Thermally-perfect Gases

    NASA Technical Reports Server (NTRS)

    DalBello, Teryn; Georgiadis, Nick (Technical Monitor)

    2002-01-01

    The ability of the WIND Navier-Stokes code to predict the physics of multi-species gases is investigated in support of future high-speed, high-temperature propulsion applications relevant to NASA's Space Transportation efforts. Three benchmark cases are investigated to evaluate the capability of the WIND chemistry model to accurately predict the aerodynamics of multi-species chemically non-reacting (frozen) gases. Case 1 represents turbulent mixing of sonic hydrogen and supersonic vitiated air. Case 2 consists of heated and unheated round supersonic jet exiting to ambient. Case 3 represents 2-D flow through a converging-diverging Mach 2 nozzle. For Case 1, the WIND results agree fairly well with experimental results and that significant mixing occurs downstream of the hydrogen injection point. For Case 2, the results show that the Wilke and Sutherland viscosity laws gave similar results, and the available SST turbulence model does not predict round supersonic nozzle flows accurately. For Case 3, results show that experimental, frozen, and 1-D gas results agree fairly well, and that frozen, homogeneous, multi-species gas calculations can be approximated by running in perfect gas mode while specifying the mixture gas constant and Ratio of Specific Heats.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. Supersonic Flow Control by Magnetic Field

    DTIC Science & Technology

    2005-12-01

    39 3. NUMERICAL SIMULATION ...in cable sheath, this synchronous pulse feeds the thyratron grid via an isolating transformer (1:2) wounded on a ferrite ring (primary winding of 20...conductivity with time obtained at different voltages on capacitor C1. The fact that that all the curves <σeff>(t) are close to each other indicates

  16. Wind sensor

    NASA Technical Reports Server (NTRS)

    Stephens, J. B.; Laue, E. G. (Inventor)

    1976-01-01

    An apparatus is described for sensing the temperature, velocity, and direction of the wind, including four temperature-dependent crystal oscillators spaced about an axis, a heater centered on the axis, and a screen through which the wind blows to pass over the crystals. In one method of operation, the frequency of the oscillators is taken when the heater is not energized, to obtain the temperature of the wind, and the frequencies of the oscillators are taken after the heater is energized to determine the direction and velocity of the wind. When the heater is energized, the wind causes the downwind crystals to achieve a higher temperature than the upwind crystals, and with the magnitude of the difference indicating the velocity of the wind.

  17. 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.

  18. Wind Energy

    SciTech Connect

    Ganley, Jason; Zhang, Jie; Hodge, Bri-Mathias

    2016-03-15

    Wind energy is a variable and uncertain renewable resource that has long been used to produce mechanical work, and has developed into a large producer of global electricity needs. As renewable sources of energy and feedstocks become more important globally to produce sustainable products, many different processes have started adopting wind power as an energy source. Many times this is through a conversion to hydrogen through electrolysis that allows for a more continuous process input. Other important pathways include methanol and ammonia. As the demand for sustainable products and production pathways increases, and wind power capital costs decrease, the role of wind power in chemical and energy production seems poised to increase significantly.

  19. 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...

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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

  9. 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.

  10. Flow Control for Supersonic Inlet Applications

    DTIC Science & Technology

    2014-06-10

    N., Babinsky, H., “ Shock wave/boundary-layer interaction control using a combination of vortex generators and bleed”, AIAA J., Vol.51, No.5, pp...Vol.51, No.9, pp. 2208-2217, September 2013 4. Titchener, N., Babinsky, H., “ Control of a Shock -Wave/Boundary-Layer Interaction and Subsequent...Approved for public release; distribution is unlimited. 1. Ogawa, H., Babinsky, H., “Wind- Tunnel Setup for Investigations of Normal Shock Wave

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. Wind energy.

    PubMed

    Leithead, W E

    2007-04-15

    From its rebirth in the early 1980s, the rate of development of wind energy has been dramatic. Today, other than hydropower, it is the most important of the renewable sources of power. The UK Government and the EU Commission have adopted targets for renewable energy generation of 10 and 12% of consumption, respectively. Much of this, by necessity, must be met by wind energy. The US Department of Energy has set a goal of 6% of electricity supply from wind energy by 2020. For this potential to be fully realized, several aspects, related to public acceptance, and technical issues, related to the expected increase in penetration on the electricity network and the current drive towards larger wind turbines, need to be resolved. Nevertheless, these challenges will be met and wind energy will, very likely, become increasingly important over the next two decades. An overview of the technology is presented.

  6. 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.

  7. Plans and Status of Wind-Tunnel Testing Employing an Aeroservoelastic Semispan Model

    NASA Technical Reports Server (NTRS)

    Perry, Boyd, III; Silva, Walter A.; Florance, James R.; Wieseman, Carol D.; Pototzky, Anthony S.; Sanetrik, Mark D.; Scott, Robert C.; Keller, Donald F.; Cole, Stanley R.; Coulson, David A.

    2007-01-01

    This paper presents the research objectives, summarizes the pre-wind-tunnel-test experimental results to date, summarizes the analytical predictions to date, and outlines the wind-tunnel-test plans for an aeroservoelastic semispan wind-tunnel model. The model is referred to as the Supersonic Semispan Transport (S4T) Active Controls Testbed (ACT) and is based on a supersonic cruise configuration. The model has three hydraulically-actuated surfaces (all-movable horizontal tail, all-movable ride control vane, and aileron) for active controls. The model is instrumented with accelerometers, unsteady pressure transducers, and strain gages and will be mounted on a 5-component sidewall balance. The model will be tested twice in the Langley Transonic Dynamics Tunnel (TDT). The first entry will be an "open-loop" model-characterization test; the second entry will be a "closed-loop" test during which active flutter suppression, gust load alleviation and ride quality control experiments will be conducted.

  8. Heat-flux gage measurements on a flat plate at a Mach number of 4.6 in the VSD high speed wind tunnel, a feasibility test (LA28). [wind tunnel tests of measuring instruments for boundary layer flow

    NASA Technical Reports Server (NTRS)

    1975-01-01

    The feasibility of employing thin-film heat-flux gages was studied as a method of defining boundary layer characteristics at supersonic speeds in a high speed blowdown wind tunnel. Flow visualization techniques (using oil) were employed. Tabulated data (computer printouts), a test facility description, and photographs of test equipment are given.

  9. The Dynamics of Shock Dispersion and Interactions in Supersonic Freestreams with Counterflowing Jets

    NASA Technical Reports Server (NTRS)

    Daso, Endwell O.; Pritchett, Victor E.; Wang, Ten-See; Ota, Dale K.; Blankson, Isaiah M.; Auslender, Aaron H.

    2007-01-01

    An active flow control concept using counterflowing jets to significantly modify the external flowfields and strongly weaken or disperse the shock-waves of supersonic and hypersonic vehicles to reduce the aerothermal loads and wave drag was investigated. Experiments were conducted in a trisonic blow-down wind-tunnel, complemented by pre-test computational fluid dynamics (CFD) analysis of a 2.6% scale model of Apollo capsule, with and without counterflowing jets, in Mach 3.48 and 4.0 freestreams, to assess the potential aerothermal and aerodynamic benefits of this concept. The model was instrumented with heat flux gauges, thermocouples and pressure taps, and employed five counterflowing jet nozzles (three sonic and other two supersonic with design Mach numbers of 2.44 and 2.94) and nozzle exit diameters ranging from 0.25 to 0.5 inch. Schlieren data show that at low jet flow rates of 0.05 and 0.1lb(sub m)/sec, the interactions result in a long penetration mode (LPM) jet, while the short penetration mode (SPM) jet is observed at flow rates greater than 0.1 lb(sub m)/sec., consistent with the pre-test CFD predictions. For the LPM, the jet appears to be nearly fully-expanded, resulting in a very unsteady and oscillatory flow structure in which the bow shock becomes highly dispersed such that it is no longer discernable. Higher speed camera Schlieren data reveal the shock to be dispersed into striations of compression waves, which suddenly coalesce to a weaker bow shock with a larger standoff distance as the flow rate reached a critical value. The pronounced shock dispersion could significantly impact the aerodynamic performance (L/D) and heat flux reduction of spacecraft in atmospheric entry and re-entry, and could also attenuate the entropy layer in hypersonic blunt body flows. For heat transfer, the results show significant reduction in heat flux, even giving negative heat flux for some of the SPM interactions, indicating that the flow wetting the model is cooling

  10. Boom and the problems related to the supersonic flights of military aircraft

    NASA Astrophysics Data System (ADS)

    Thery, Christian; Lecomte, Claude

    1992-04-01

    The flight of a supersonic projectile is accompanied by a clapping called ballistic detonation. At the same time an aircraft in supersonic flight creates a boom which travels to the ground. The boom represents an impulsional characteristic which provokes a startled reaction in living beings and makes structures vibrate. Because of the elevated altitude of the flight of aircraft the boom is felt along a band of terrain situated along each side of the trajectory of the aircraft, as large as many tens of kilometers. The intensity of the boom essentially depends on the size of the aircraft that creates it, on its altitude, and on the maneuver that it is executing; the maneuvers which are executed by military aircraft provoke locally a very great intensification of booms, these are focalizations and superfocalizations. These phenomena appear especially at the time of transonic acceleration and when turning. The annoyance resulting from these focalized booms is therefore certain and related damage to aging or poorly constructed structures can occur; the sole means of limiting the nuisance which results will be to situate these focalized boom zones over uninhabited regions. A good preview of the affected zones requires precise a priori knowledge of the actual trajectory of the aircraft and the meteorological conditions; in practice therefore, the zones susceptible to being affected by these phenomena are fairly large. Above all, the effects of boom must not be overestimated. The first temporary damages to the auditory system are observed for levels of wave intensity greatly superior to the levels of booms (3000 to 5000 Pa as opposed to 50 or 70 Pa for a boom). The structural effects of normal booms are comparable to those resulting from natural inclemency (wind) or related to modern life (highway traffic). The focalized boom cannot have an impact on sound and well-constructed buildings except by cumulative effect (fatigue).

  11. Experimental and numerical investigation of hydrogen combustion in a supersonic flow

    SciTech Connect

    Segal, C.

    1991-01-01

    Supersonic combustion ramjet, or SCRAMJET, engines are currently being evaluated for the propulsion of hypersonic vehicles. A unique supersonic wind tunnel facility has been built at the Aerospace Research Laboratory to simulate the operation of a SCRAMJET over a range of Mach numbers of 5 to 6.5 and altitudes of 40,000 to 150,000 ft. The tunnel provides high stagnation temperature, clean air in a continuous Mach 2 flow to the combustor where hydrogen is injected and burned. One of the major parameters in the design of the engine is the combustion efficiency, a quantity which is extremely difficult to evaluate directly. Wall pressure and temperature measurements were made on a model combustor in the ARL facility and a one-dimensional, chemical equilibrium, finite difference model was used to infer combustion efficiency. The initial stagnation temperature was maintained around or below 850 K and an inlet static pressure of 1/2 atm, or less. At these relative low temperatures, thermal choking occurred for relatively low equivalence ratios, limiting this parameter to a maximum of 0.1, depending on the injection configuration. A detailed validation of the analytical model requires extensive knowledge of the physical properties of the flowfield. Since reliable, non-intrusive measurements methods are still under development, the results of the combustion efficiency evaluation were compared with a detailed numerical simulation of the flowfield of interest. The numerical simulation used a 3-D full Navier-Stokes program, which includes a finite rate chemistry model to duplicate one of the experimental cases. The calculations were performed on a CRAY - 2S supercomputer at the National Supercomputer Applications Center at NASA Langley. The solution required 35 CPU hours.

  12. 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

  13. 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.; Wilcox, Peter A.; Fugal, Spencer R.; Acheson, Kurt E.; Adamson, Eric E.; Bidwell, Alicia L.; Shaw, Stephen G.

    2013-01-01

    This report describes the work conducted by The Boeing Company under American Recovery and Reinvestment Act (ARRA) and NASA funding to experimentally validate the conceptual design of a supersonic airliner feasible for entry into service in the 2018 to 2020 timeframe (NASA N+2 generation). The report discusses the design, analysis and development of a low-boom concept that meets aggressive sonic boom and performance goals for a cruise Mach number of 1.8. The design is achieved through integrated multidisciplinary optimization tools. The report also describes the detailed design and fabrication of both sonic boom and performance wind tunnel models of the low-boom concept. Additionally, a description of the detailed validation wind tunnel testing that was performed with the wind tunnel models is provided along with validation comparisons with pretest Computational Fluid Dynamics (CFD). Finally, the report describes the evaluation of existing NASA sonic boom pressure rail measurement instrumentation and a detailed description of new sonic boom measurement instrumentation that was constructed for the validation wind tunnel testing.

  14. 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.

  15. 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.

  16. 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.

  17. 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

  18. 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.

  19. 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.

  20. 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.