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Sample records for alfven mach numbers

  1. Anomalous flow deflection at planetary bow shocks in the low Alfven Mach number regime

    NASA Astrophysics Data System (ADS)

    Nishino, Masaki N.; Fujimoto, Masaki; Tai, Phan-Duc; Mukai, Toshifumi; Saito, Yoshifumi; Kuznetsova, Masha M.; Rastaetter, Lutz

    A planetary magnetosphere is an obstacle to the super-sonic solar wind and the bow shock is formed in the front-side of it. In ordinary hydro-dynamics, the flow decelerated at the shock is diverted around the obstacle symmetrically about the planet-Sun line, which is indeed observed in the magnetosheath most of the time. Here we show a case under a very low density solar wind in which duskward flow was observed in the dawnside magnetosheath of the Earth's magnetosphere. A Rankine-Hugoniot test across the bow shock shows that the magnetic effect is crucial for this "wrong flow" to appear. A full three-dimensional Magneto- Hydro-Dynamics (MHD) simulation of the situation in this previously unexplored parameter regime is also performed. It is illustrated that in addition to the "wrong flow" feature, various peculiar characteristics appear in the global picture of the MHD flow interaction with the obstacle. The magnetic effect at the bow shock should become more conspicuously around the Mercury's magnetosphere, because stronger interplanetary magnetic field and slower solar wind around the Mercury let the Alfven Mach number low. Resultant strong deformation of the magnetosphere induced by the "wrong flow" will cause more complex interaction between the solar wind and the Mercury.

  2. The Alfven Mach Number Control of the Solar Wind-Magnetosphere Coupling Efficiency and the Saturation of the Geomagnetic Indices

    NASA Astrophysics Data System (ADS)

    Myllys, M. E.; Kilpua, E.; Lavraud, B.

    2015-12-01

    We have investigated the effect of key solar wind driving parameters on the solar wind-magnetosphere coupling efficiency and saturation of the cross polar cap potential (CPCP) during sheath and magnetic cloud driven storms. The particular focus of the study was on the coupling efficiency dependence with Alfven Mach number (MA).Since we are studying the instantaneous coupling efficiency instead of the average efficiency over the whole solar wind structure, we needed to take into account the communication time between the solar wind and the magnetosphere. We present the results of the time delay analysis between geomagnetic indices (PCN, AE and SYM-H) and the interplanetary electric field y-component (EY, GSM coordinate system) and Newell and Borovsky functions. The study shows that the MA has a clear effect to the saturation of the PCN index, which can be used as a proxy of the polar cap potential. The higher the MA the higher the limit EY value after which the saturation starts to occur. Thus, the coupling efficiency increases as a function of MA. Also, the AE index saturates during high solar wind driving but the saturation is not MA depended. However, the results also suggest that the MA it is not the primary cause for the PCN saturation.

  3. Chaotic behaviour of high Mach number flows

    NASA Technical Reports Server (NTRS)

    Varvoglis, H.; Ghosh, S.

    1985-01-01

    The stability of the super-Alfvenic flow of a two-fluid plasma model with respect to the Mach number and the angle between the flow direction and the magnetic field is investigated. It is found that, in general, a large scale chaotic region develops around the initial equilibrium of the laminar flow when the Mach number exceeds a certain threshold value. After reaching a maximum the size of this region begins shrinking and goes to zero as the Mach number tends to infinity. As a result high Mach number flows in time independent astrophysical plasmas may lead to the formation of 'quasi-shocks' in the presence of little or no dissipation.

  4. Quasiperpendicular High Mach Number Shocks

    NASA Astrophysics Data System (ADS)

    Sulaiman, A. H.; Masters, A.; Dougherty, M. K.; Burgess, D.; Fujimoto, M.; Hospodarsky, G. B.

    2015-09-01

    Shock waves exist throughout the Universe and are fundamental to understanding the nature of collisionless plasmas. Reformation is a process, driven by microphysics, which typically occurs at high Mach number supercritical shocks. While ongoing studies have investigated this process extensively both theoretically and via simulations, their observations remain few and far between. In this Letter we present a study of very high Mach number shocks in a parameter space that has been poorly explored and we identify reformation using in situ magnetic field observations from the Cassini spacecraft at 10 AU. This has given us an insight into quasiperpendicular shocks across 2 orders of magnitude in Alfvén Mach number (MA ) which could potentially bridge the gap between modest terrestrial shocks and more exotic astrophysical shocks. For the first time, we show evidence for cyclic reformation controlled by specular ion reflection occurring at the predicted time scale of ˜0.3 τc , where τc is the ion gyroperiod. In addition, we experimentally reveal the relationship between reformation and MA and focus on the magnetic structure of such shocks to further show that for the same MA , a reforming shock exhibits stronger magnetic field amplification than a shock that is not reforming.

  5. AMR for low Mach number reacting flow

    SciTech Connect

    Bell, John B.

    2004-01-16

    We present a summary of recent progress on the development and application of adaptive mesh refinement algorithms for low Mach number reacting flows. Our approach uses a form of the low Mach number equations based on a general equation of state that discretely conserves both mass and energy. The discretization methodology is based on a robust projection formulation that accommodates large density contrasts. The algorithm supports modeling of multicomponent systems and incorporates an operator-split treatment of stiff reaction terms. The basic computational approach is embedded in an adaptive projection framework that uses structured hierarchical grids with subcycling in time that preserves the discrete conservation properties of the underlying single-grid algorithm. We present numerical examples illustrating the application of the methodology to turbulent premixed combustion and nuclear flames in type Ia supernovae.

  6. Critical Mach Numbers of Thin Airfoil Sections with Plain Flaps

    NASA Technical Reports Server (NTRS)

    Pardee, Otway O'm.; Heaslet, Max A.

    1946-01-01

    Critical Mach number as function of lift coefficient is determined for certain moderately thick NACA low-drag airfoils. Results, given graphically, included calculations on same airfoil sections with plain flaps for small flap deflections. Curves indicate optimum critical conditions for airfoils with flaps in such form that they can be compared with corresponding results for zero flap deflections. Plain flaps increase life-coefficient range for which critical Mach number is in region of high values characteristic of low-drag airfoils.

  7. Low Mach Number Modeling of Type Ia Supernovae

    SciTech Connect

    Almgren, Ann S.; Bell, John B.; Rendleman, Charles A.; Zingale,Michael

    2005-08-05

    We introduce a low Mach number equation set for the large-scale numerical simulation of carbon-oxygen white dwarfs experiencing a thermonuclear deflagration. Since most of the interesting physics in a Type Ia supernova transpires at Mach numbers from 0.01 to 0.1, such an approach enables both a considerable increase in accuracy and savings in computer time compared with frequently used compressible codes. Our equation set is derived from the fully compressible equations using low Mach number asymptotics, but without any restriction on the size of perturbations in density or temperature. Comparisons with simulations that use the fully compressible equations validate the low Mach number model in regimes where both are applicable. Comparisons to simulations based on the more traditional an elastic approximation also demonstrate the agreement of these models in the regime for which the anelastic approximation is valid. For low Mach number flows with potentially finite amplitude variations in density and temperature, the low Mach number model overcomes the limitations of each of the more traditional models and can serve as the basis for an accurate and efficient simulation tool.

  8. Experiments with Turbulent Jets at Mach Number 0.9

    NASA Technical Reports Server (NTRS)

    Agui, Juan; Andreopoulos, Yiannis; Davis, David O. (Technical Monitor)

    2001-01-01

    A systematic investigation of the structure of turbulent jets before their interaction with shock or expansion waves was undertaken during the last year. In particular compressibility and density effects in circular jets issuing in still air were investigated experimentally. Jets with nitrogen, helium, and krypton gases at 0.3, 0.6, and 0.9 Mach numbers were investigated in detail. Particle Image Velocimetry technique was developed, tested, and used to obtain qualitative information of the two-dimensional velocity field on a plane inside the flow field, which was illuminated by a laser sheet. The motion of particles was recorded by a CCD camera, which was appropriately triggered to capture two images within a fraction of a microsecond. Statistical averaging of the data at each location reduced the large amount of acquired data. It was found that the spreading rate of the jets was reduced with increased Mach numbers or increased density ratio. It was also found that decay rates of centerline Mach numbers are higher in gases with reduced density ratio. Mach number fluctuations appear to decrease with increasing Mach number of the flow. It has been proposed that the reason for this behavior is the reduction of vortex stretching activities with increased Mach number.

  9. Adaptive low Mach number simulations of nuclear flame microphysics

    SciTech Connect

    Bell, J.B.; Day, M.S.; Rendleman, C.A.; Woosley, S.E.; Zingale, M.A.

    2003-03-20

    We introduce a numerical model for the simulation of nuclear flames in Type Ia supernovae. This model is based on a low Mach number formulation that analytically removes acoustic wave propagation while retaining the compressibility effects resulting from nuclear burning. The formulation presented here generalizes low Mach number models used in combustion that are based on an ideal gas approximation to the arbitrary equations of state such as those describing the degenerate matter found in stellar material. The low Mach number formulation permits time steps that are controlled by the advective time scales resulting in a substantial improvement in computational efficiency compared to a compressible formulation. We briefly discuss the basic discretization methodology for the low Mach number equations and their implementation in an adaptive projection framework. We present validation computations in which the computational results from the low Mach number model are compared to a compressible code and present an application of the methodology to the Landau-Darrieus instability of a carbon flame.

  10. Enthalpy damping for high Mach number Euler solutions

    NASA Technical Reports Server (NTRS)

    Moitra, Anutosh

    1990-01-01

    An improvement on the enthalpy damping procedure currently in use in solving supersonic flow fields is described. A correction based on entropy values is shown to produce a very efficient scheme for simulation of high Mach number three-dimensional flows. Substantial improvements in convergence rates have been achieved by incorporating this enthalpy damping scheme in a finite-volume Runge-Kutta method for solving the Euler equations. Results obtained for blended wing-body geometries at very high Mach numbers are presented.

  11. Overestimation of Mach number due to probe shadow

    NASA Astrophysics Data System (ADS)

    Gosselin, J. J.; Thakur, S. C.; Sears, S. H.; McKee, J. S.; Scime, E. E.; Tynan, G. R.

    2016-07-01

    Comparisons of the plasma ion flow speed measurements from Mach probes and laser induced fluorescence were performed in the Controlled Shear Decorrelation Experiment. We show the presence of the probe causes a low density geometric shadow downstream of the probe that affects the current density collected by the probe in collisional plasmas if the ion-neutral mean free path is shorter than the probe shadow length, Lg = w2 Vdrift/D⊥, resulting in erroneous Mach numbers. We then present a simple correction term that provides the corrected Mach number from probe data when the sound speed, ion-neutral mean free path, and perpendicular diffusion coefficient of the plasma are known. The probe shadow effect must be taken into account whenever the ion-neutral mean free path is on the order of the probe shadow length in linear devices and the open-field line region of fusion devices.

  12. Mach Number Effects on Turbine Blade Transition Length Prediction

    NASA Technical Reports Server (NTRS)

    Boyle, R. J.; Simon, F. F.

    1998-01-01

    The effect of a Mach number correction on a model for predicting the length of transition was investigated. The transition length decreases as the turbulent spot production rate increases. Much of the data for predicting the spot production rate comes from low speed flow experiments. Recent data and analysis showed that the spot production rate is affected by Mach number. The degree of agreement between analysis and data for turbine blade heat transfer without film cooling is strongly dependent of accurately predicting the length of transition. Consequently, turbine blade heat transfer data sets were used to validate a transition length turbulence model. A method for modifying models for the length of transition to account for Mach number effects is presented. The modification was made to two transition length models. The modified models were incorporated into the two-dimensional Navier-Stokes code, RVCQ3D. Comparisons were made between predicted and measured midspan surface heat transfer for stator and rotor turbine blades. The results showed that accounting for Mach number effects significantly improved the agreement with the experimental data.

  13. Mach number effect on jet impingement heat transfer.

    PubMed

    Brevet, P; Dorignac, E; Vullierme, J J

    2001-05-01

    An experimental investigation of heat transfer from a single round free jet, impinging normally on a flat plate is described. Flow at the exit plane of the jet is fully developed and the total temperature of the jet is equal to the ambient temperature. Infrared measurements lead to the characterization of the local and averaged heat transfer coefficients and Nusselt numbers over the impingement plate. The adiabatic wall temperature is introduced as the reference temperature for heat transfer coefficient calculation. Various nozzle diameters from 3 mm to 15 mm are used to make the injection Mach number M vary whereas the Reynolds number Re is kept constant. Thus the Mach number influence on jet impingement heat transfer can be directly evaluated. Experiments have been carried out for 4 nozzle diameters, for 3 different nozzle-to-target distances, with Reynolds number ranging from 7200 to 71,500 and Mach number from 0.02 to 0.69. A correlation is obtained from the data for the average Nusselt number. PMID:11460655

  14. Statistical error in particle simulations of low mach number flows

    SciTech Connect

    Hadjiconstantinou, N G; Garcia, A L

    2000-11-13

    We present predictions for the statistical error due to finite sampling in the presence of thermal fluctuations in molecular simulation algorithms. The expressions are derived using equilibrium statistical mechanics. The results show that the number of samples needed to adequately resolve the flowfield scales as the inverse square of the Mach number. Agreement of the theory with direct Monte Carlo simulations shows that the use of equilibrium theory is justified.

  15. Very high Mach number shocks - Theory. [in space plasmas

    NASA Technical Reports Server (NTRS)

    Quest, Kevin B.

    1986-01-01

    The theory and simulation of collisionless perpendicular supercritical shock structure is reviewed, with major emphasis on recent research results. The primary tool of investigation is the hybrid simulation method, in which the Newtonian orbits of a large number of ion macroparticles are followed numerically, and in which the electrons are treated as a charge neutralizing fluid. The principal results include the following: (1) electron resistivity is not required to explain the observed quasi-stationarity of the earth's bow shock, (2) the structure of the perpendicular shock at very high Mach numbers depends sensitively on the upstream value of beta (the ratio of the thermal to magnetic pressure) and electron resistivity, (3) two-dimensional turbulence will become increasingly important as the Mach number is increased, and (4) nonadiabatic bulk electron heating will result when a thermal electron cannot complete a gyrorbit while transiting the shock.

  16. Boundary conditions and the simulation of low Mach number flows

    NASA Technical Reports Server (NTRS)

    Hagstrom, Thomas; Lorenz, Jens

    1993-01-01

    The problem of accurately computing low Mach number flows, with the specific intent of studying the interaction of sound waves with incompressible flow structures, such as concentrations of vorticity is considered. This is a multiple time (and/or space) scales problem, leading to various difficulties in the design of numerical methods. Concentration is on one of these difficulties - the development of boundary conditions at artificial boundaries which allow sound waves and vortices to radiate to the far field. Nonlinear model equations are derived based on assumptions about the scaling of the variables. Then these are linearized about a uniform flow and exact boundary conditions are systematically derived using transform methods. Finally, useful approximations to the exact conditions which are valid for small Mach number and small viscosity are computed.

  17. Numerical simulation of low Mach number reacting flows

    SciTech Connect

    Bell, John B.; Aspden, Andrew J.; Day, Marcus S.; Lijewski,Michael J.

    2007-06-20

    Using examples from active research areas in combustion andastrophysics, we demonstrate a computationally efficient numericalapproach for simulating multiscale low Mach number reacting flows. Themethod enables simulations that incorporate an unprecedented range oftemporal and spatial scales, while at the same time, allows an extremelyhigh degree of reaction fidelity. Sample applications demonstrate theefficiency of the approach with respect to a traditional time-explicitintegration method, and the utility of the methodology for studying theinteraction of turbulence with terrestrial and astrophysical flamestructures.

  18. Courant Number and Mach Number Insensitive CE/SE Euler Solvers

    NASA Technical Reports Server (NTRS)

    Chang, Sin-Chung

    2005-01-01

    It has been known that the space-time CE/SE method can be used to obtain ID, 2D, and 3D steady and unsteady flow solutions with Mach numbers ranging from 0.0028 to 10. However, it is also known that a CE/SE solution may become overly dissipative when the Mach number is very small. As an initial attempt to remedy this weakness, new 1D Courant number and Mach number insensitive CE/SE Euler solvers are developed using several key concepts underlying the recent successful development of Courant number insensitive CE/SE schemes. Numerical results indicate that the new solvers are capable of resolving crisply a contact discontinuity embedded in a flow with the maximum Mach number = 0.01.

  19. Hysteresis phenomenon of hypersonic inlet at high Mach number

    NASA Astrophysics Data System (ADS)

    Jiao, Xiaoliang; Chang, Juntao; Wang, Zhongqi; Yu, Daren

    2016-11-01

    When the hypersonic inlet works at a Mach number higher than the design value, the hypersonic inlet is started with a regular reflection of the external compression shock at the cowl, whereas a Mach reflection will result in the shock propagating forwards to cause a shock detachment at the cowl lip, which is called "local unstart of inlet". As there are two operation modes of hypersonic inlet at high Mach number, the mode transition may occur with the operation condition of hypersonic inlet changing. A cowl-angle-variation-induced hysteresis and a downstream-pressure-variation-induced hysteresis in the hypersonic inlet start↔local unstart transition are obtained by viscous numerical simulations in this paper. The interaction of the external compression shock and boundary layer on the cowl plays a key role in the hysteresis phenomenon. Affected by the transition of external compression shock reflection at the cowl and the transition between separated and attached flow on the cowl, a hysteresis exists in the hypersonic inlet start↔local unstart transition. The hysteresis makes the operation of a hypersonic inlet very difficult to control. In order to avoid hysteresis phenomenon and keep the hypersonic inlet operating in a started mode, the control route should never pass through the local unstarted boundary.

  20. Low Mach number fluctuating hydrodynamics of multispecies liquid mixtures

    SciTech Connect

    Donev, Aleksandar Bhattacharjee, Amit Kumar; Nonaka, Andy; Bell, John B.; Garcia, Alejandro L.

    2015-03-15

    We develop a low Mach number formulation of the hydrodynamic equations describing transport of mass and momentum in a multispecies mixture of incompressible miscible liquids at specified temperature and pressure, which generalizes our prior work on ideal mixtures of ideal gases [Balakrishnan et al., “Fluctuating hydrodynamics of multispecies nonreactive mixtures,” Phys. Rev. E 89 013017 (2014)] and binary liquid mixtures [Donev et al., “Low mach number fluctuating hydrodynamics of diffusively mixing fluids,” Commun. Appl. Math. Comput. Sci. 9(1), 47-105 (2014)]. In this formulation, we combine and extend a number of existing descriptions of multispecies transport available in the literature. The formulation applies to non-ideal mixtures of arbitrary number of species, without the need to single out a “solvent” species, and includes contributions to the diffusive mass flux due to gradients of composition, temperature, and pressure. Momentum transport and advective mass transport are handled using a low Mach number approach that eliminates fast sound waves (pressure fluctuations) from the full compressible system of equations and leads to a quasi-incompressible formulation. Thermal fluctuations are included in our fluctuating hydrodynamics description following the principles of nonequilibrium thermodynamics. We extend the semi-implicit staggered-grid finite-volume numerical method developed in our prior work on binary liquid mixtures [Nonaka et al., “Low mach number fluctuating hydrodynamics of binary liquid mixtures,” http://arxiv.org/abs/1410.2300 (2015)] and use it to study the development of giant nonequilibrium concentration fluctuations in a ternary mixture subjected to a steady concentration gradient. We also numerically study the development of diffusion-driven gravitational instabilities in a ternary mixture and compare our numerical results to recent experimental measurements [Carballido-Landeira et al., “Mixed-mode instability of a

  1. High Order Difference Method for Low Mach Number Aeroacoustics

    NASA Technical Reports Server (NTRS)

    Mueller, B.; Yee, H. C.; Mansour, Nagi (Technical Monitor)

    2001-01-01

    A high order finite difference method with improved accuracy and stability properties for computational aeroacoustics (CAA) at low Mach numbers is proposed. The Euler equations are split into a conservative and a symmetric non- conservative portion to allow the derivation of a generalized energy estimate. Since the symmetrization is based on entropy variables, that splitting of the flux derivatives is referred to as entropy splitting. Its discretization by high order central differences was found to need less numerical dissipation than conventional conservative schemes. Owing to the large disparity of acoustic and stagnation quantities in low Mach number aeroacoustics, the split Euler equations are formulated in perturbation form. The unknowns are the small changes of the conservative variables with respect to their large stagnation values. All nonlinearities and the conservation form of the conservative portion of the split flux derivatives can be retained, while cancellation errors are avoided with its discretization opposed to the conventional conservative form. The finite difference method is third-order accurate at the boundary and the conventional central sixth-order accurate stencil in the interior. The difference operator satisfies the summation by parts property analogous to the integration by parts in the continuous energy estimate. Thus, strict stability of the difference method follows automatically. Spurious high frequency oscillations are suppressed by a characteristic-based filter similar to but without limiter. The time derivative is approximated by a 4-stage low-storage second-order explicit Runge-Kutta method. The method has been applied to simulate vortex sound at low Mach numbers. We consider the Kirchhoff vortex, which is an elliptical patch of constant vorticity rotating with constant angular frequency in irrotational flow. The acoustic pressure generated by the Kirchhoff vortex is governed by the 2D Helmholtz equation, which can be solved

  2. A moving frame algorithm for high Mach number hydrodynamics

    NASA Astrophysics Data System (ADS)

    Trac, Hy; Pen, Ue-Li

    2004-07-01

    We present a new approach to Eulerian computational fluid dynamics that is designed to work at high Mach numbers encountered in astrophysical hydrodynamic simulations. Standard Eulerian schemes that strictly conserve total energy suffer from the high Mach number problem and proposed solutions to additionally solve the entropy or thermal energy still have their limitations. In our approach, the Eulerian conservation equations are solved in an adaptive frame moving with the fluid where Mach numbers are minimized. The moving frame approach uses a velocity decomposition technique to define local kinetic variables while storing the bulk kinetic components in a smoothed background velocity field that is associated with the grid velocity. Gravitationally induced accelerations are added to the grid, thereby minimizing the spurious heating problem encountered in cold gas flows. Separately tracking local and bulk flow components allows thermodynamic variables to be accurately calculated in both subsonic and supersonic regions. A main feature of the algorithm, that is not possible in previous Eulerian implementations, is the ability to resolve shocks and prevent spurious heating where both the pre-shock and post-shock fluid are supersonic. The hybrid algorithm combines the high-resolution shock capturing ability of the second-order accurate Eulerian TVD scheme with a low-diffusion Lagrangian advection scheme. We have implemented a cosmological code where the hydrodynamic evolution of the baryons is captured using the moving frame algorithm while the gravitational evolution of the collisionless dark matter is tracked using a particle-mesh N-body algorithm. Hydrodynamic and cosmological tests are described and results presented. The current code is fast, memory-friendly, and parallelized for shared-memory machines.

  3. Finite Mach number spherical shock wave, application to shock ignition

    NASA Astrophysics Data System (ADS)

    Vallet, A.; Ribeyre, X.; Tikhonchuk, V.

    2013-08-01

    A converging and diverging spherical shock wave with a finite initial Mach number Ms0 is described by using a perturbative approach over a small parameter Ms-2. The zeroth order solution is the Guderley's self-similar solution. The first order correction to this solution accounts for the effects of the shock strength. Whereas it was constant in the Guderley's asymptotic solution, the amplification factor of the finite amplitude shock Λ(t)∝dUs/dRs now varies in time. The coefficients present in its series form are iteratively calculated so that the solution does not undergo any singular behavior apart from the position of the shock. The analytical form of the corrected solution in the vicinity of singular points provides a better physical understanding of the finite shock Mach number effects. The correction affects mainly the flow density and the pressure after the shock rebound. In application to the shock ignition scheme, it is shown that the ignition criterion is modified by more than 20% if the fuel pressure prior to the final shock is taken into account. A good agreement is obtained with hydrodynamic simulations using a Lagrangian code.

  4. Finite Mach number spherical shock wave, application to shock ignition

    SciTech Connect

    Vallet, A.; Ribeyre, X.; Tikhonchuk, V.

    2013-08-15

    A converging and diverging spherical shock wave with a finite initial Mach number M{sub s0} is described by using a perturbative approach over a small parameter M{sub s}{sup −2}. The zeroth order solution is the Guderley's self-similar solution. The first order correction to this solution accounts for the effects of the shock strength. Whereas it was constant in the Guderley's asymptotic solution, the amplification factor of the finite amplitude shock Λ(t)∝dU{sub s}/dR{sub s} now varies in time. The coefficients present in its series form are iteratively calculated so that the solution does not undergo any singular behavior apart from the position of the shock. The analytical form of the corrected solution in the vicinity of singular points provides a better physical understanding of the finite shock Mach number effects. The correction affects mainly the flow density and the pressure after the shock rebound. In application to the shock ignition scheme, it is shown that the ignition criterion is modified by more than 20% if the fuel pressure prior to the final shock is taken into account. A good agreement is obtained with hydrodynamic simulations using a Lagrangian code.

  5. The Variation of Slat Noise with Mach and Reynolds Numbers

    NASA Technical Reports Server (NTRS)

    Lockhard, David P.; Choudhari, Meelan M.

    2011-01-01

    The slat noise from the 30P30N high-lift system has been computed using a computational fluid dynamics code in conjunction with a Ffowcs Williams-Hawkings solver. By varying the Mach number from 0.13 to 0.25, the noise was found to vary roughly with the 5th power of the speed. Slight changes in the behavior with directivity angle could easily account for the different speed dependencies reported in the literature. Varying the Reynolds number from 1.4 to 2.4 million resulted in almost no differences, and primarily served to demonstrate the repeatability of the results. However, changing the underlying hybrid Reynolds-averaged-Navier-Stokes/Large-Eddy-Simulation turbulence model significantly altered the mean flow because of changes in the flap separation. However, the general trends observed in both the acoustics and near-field fluctuations were similar for both models.

  6. An experimental investigation of turbulent boundary layers at high Mach number and Reynolds numbers

    NASA Technical Reports Server (NTRS)

    Holden, M. S.

    1972-01-01

    Skin friction, heat transfer and pressure measurements were obtained in laminar, transitional and turbulent boundary layers on flat plates at Mach numbers from 7 to 13 at wall-to-free stream stagnation temperature ratios from 0.1 to 0.3. Measurements in laminar flows were in excellent agreement with the theory of Cheng. Correlations of the transition measurements with measurements on flight vehicles and in ballistic ranges show good agreement. Our transition measurements do not correlate well with those of Pate and Schueler. Comparisons have been made between the skin friction and heat transfer measurements and the theories of Van Driest, Eckert and Spalding and Chi. These comparisons reveal in general that at the high end of our Mach number range (10-13) the theory of Van Driest is in best agreement with the data, whereas at lower Mach numbers (6.5-10) the Spalding Chi theory is in better agreement with the measurements.

  7. A Global Existence Result for a Zero Mach Number System

    NASA Astrophysics Data System (ADS)

    Liao, Xian

    2014-03-01

    In this paper we study the global-in-time existence of weak solutions to a zero Mach number system that derives from the Navier-Stokes-Fourier system, under a special relationship between the viscosity coefficient and the heat conductivity coefficient. Roughly speaking, this relation implies that the source term in the equation for the newly introduced divergence-free velocity vector field vanishes. In dimension two, thanks to a local-in-time existence result of a unique strong solution in critical Besov spaces given by Danchin and Liao (Commun Contemp Math 14:1250022, 2012), for arbitrary large initial data, we show that this unique strong solution exists globally in time, as a consequence of a weak-strong uniqueness argument.

  8. Turbomachinery for Low-to-High Mach Number Flight

    NASA Technical Reports Server (NTRS)

    Tan, Choon S.; Shah, Parthiv N.

    2004-01-01

    The thrust capability of turbojet cycles is reduced at high flight Mach number (3+) by the increase in inlet stagnation temperature. The 'hot section' temperature limit imposed by materials technology sets the maximum heat addition and, hence, sets the maximum flight Mach number of the operating envelope. Compressor pre-cooling, either via a heat exchanger or mass-injection, has been suggested as a means to reduce compressor inlet temperature and increase mass flow capability, thereby increasing thrust. To date, however, no research has looked at compressor cooling (i.e., using a compressor both to perform work on the gas path air and extract heat from it simultaneously). We wish to assess the feasibility of this novel concept for use in low-to-high Mach number flight. The results to-date show that an axial compressor with cooling: (1) relieves choking in rear stages (hence opening up operability), (2) yields higher-pressure ratio and (3) yields higher efficiency for a given corrected speed and mass flow. The performance benefit is driven: (i) at the blade passage level, by a decrease in the total pressure reduction coefficient and an increase in the flow turning; and (ii) by the reduction in temperature that results in less work required for a given pressure ratio. The latter is a thermodynamic effect. As an example, calculations were performed for an eight-stage compressor with an adiabatic design pressure ratio of 5. By defining non-dimensional cooling as the percentage of compressor inlet stagnation enthalpy removed by a heat sink, the model shows that a non-dimensional cooling of percent in each blade row of the first two stages can increase the compressor pressure ratio by as much as 10-20 percent. Maximum corrected mass flow at a given corrected speed may increase by as much as 5 percent. In addition, efficiency may increase by as much as 5 points. A framework for characterizing and generating the performance map for a cooled compressor has been developed

  9. Effects of Mach Number and Wall-Temperature Ratio on Turbulent Heat Transfer at Mach Numbers from 3 to 5

    NASA Technical Reports Server (NTRS)

    Tendeland, Thorval

    1959-01-01

    Heat-transfer data were evaluated from temperature time histories measured on a cooled cylindrical model with a cone-shaped nose and with turbulent flow at Mach numbers 3.00, 3.44, 4.08, 4.56, and 5.04. The experimental data were compared with calculated values using a modified Reynold's analogy between skin-friction and heat-transfer. Theoretical skin- friction coefficients were calculated using the method of Van Driest the method of Sommer and Short. The heat-transfer data obtained from the model were found to correlate when the 'T' method of Sommer and Short was used. The increase in turbulent heat-transfer rate with a reduction in wall to freestream temperature ratio was of the same order of magnitude as has been found for the turbulent skin-friction coefficient.

  10. Structure of medium Mach number quasi-parallel shocks - Upstream and downstream waves

    NASA Technical Reports Server (NTRS)

    Krauss-Varban, D.; Omidi, N.

    1991-01-01

    The transition from steady low-Mach-number to unsteady high-Mach-number quasi-parallel shocks was investigated by performing large-scale 1D hybrid code simulations at increasing Mach numbers. It was found that only at very low Mach number shocks the steepening is limited by upstream phase-standing whistlers, as predicted by the classical theory (Tidman and Northrop, 1968). In the intermediate region of Mach numbers between 1.5 and 3.5, a very diverse behavior is observed. Backstreaming ions generate fast magnetosonic waves which dominate the upstream, with wavelengths longer than phase-standing whistlers. At increasing Mach numbers, the phase and group velocities of the dominant waves are reduced until they point back toward the shock; when there is sufficient energy flux in these waves, they lead to unsteady shock behavior and eventually to shock reformation.

  11. Multiple toroidal Alfven eigenmodes with a single toroidal mode number in KSTAR plasmas

    NASA Astrophysics Data System (ADS)

    Rizvi, H.; Ryu, C. M.; Lin, Z.

    2016-11-01

    Simultaneous excitation of multiple discrete toroidal Alfven eigenmodes (TAEs) for a single toroidal mode number have been observed in KSTAR plasmas. Excitation and characteristics of these modes are studied by using a global gyrokinetic particle-in-cell simulation code. It is shown that compared to a single core-localized mode, excitation of two modes is difficult. The frequency difference between the double TAEs studied from simulation seems to agree well with the experimental value. Details of studies on the frequency, growth rate, mode structures, etc, using the GTC simulation are presented.

  12. Effect of Reynolds Number and Mach Number on flow angularity probe sensitivity

    NASA Technical Reports Server (NTRS)

    Smith, L. A.; Adcock, J. B.

    1986-01-01

    Preliminary calibrations were performed on nine flow angularity probes in the Langley 7- by 10-Foot High-Speed Tunnel (7 x 10 HST) and the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT). These probes will be used in surveying the test section flows of the National Transonic Facility (NTF). The probes used in this study have a pyramid head with five pressure orifices. The calibrations consisted of both isolated probe measurements and rake-mounted multiprobe measurements that covered a range of subsonic Mach numbers up to 0.90 and Reynolds numbers per foot up to 40 X 10 to the 6th power. The preliminary calibration in the 7 x 10 HST included testing the probes both individually and in a rake. The 0.3-m TCT calibration tested two probes singly at varying Reynolds numbers. The results from these tests include Mach number, Reynolds number, and rake-mounting effects. The results of these tests showed probe sensitivity to be slightly affected by Mach number. At Reynolds numbers per foot above 10 x 10 to the 6th power, the probe did not exhibit a Reynolds number sensitivity.

  13. Evaluation of a Quartz Bourdon Pressure Gage of Wind Tunnel Mach Number Control System Application

    NASA Technical Reports Server (NTRS)

    Chapin, W. G.

    1986-01-01

    A theoretical and experimental study was undertaken to determine the feasibility of using the National Transonic Facility's high accuracy Mach number measurement system as part of a closed loop Mach number control system. The theoretical and experimental procedures described are applicable to the engineering design of pressure control systems. The results show that the dynamic response characteristics of the NTF Mach number gage (a Ruska DDR-6000 quartz absolute pressure gage) coupled to a typical length of pressure tubing were only marginally acceptable within a limited range of the facility's total pressure envelope and could not be used in the Mach number control system.

  14. Attenuation of sound in a low Mach number nozzle flow

    NASA Technical Reports Server (NTRS)

    Howe, M. S.

    1979-01-01

    The energy conversion mechanisms which govern the emission of low frequency sound from an axisymmetric jet pipe of arbitrary nozzle contraction ratio in the case of low Mach number nozzle flow are discussed. The energy of the incident sound which flows through the nozzle is used to maintain two distinct characteristic disturbances in the exterior fluid. First, there is the emitted radiation which has the directivity equivalent to that of a monopole-dipole combination. Second, essentially incompressible vortex waves are induced on the jet by vortex shedding from the lip of the nozzle and may involve the excitation of instability modes. Two linearized analytical models are examined to determine the partition of the emitted energy between the radiation field and the vortex waves. One of these is an exact linear theory in which the jet boundary is treated as a vortex sheet. The second model assumes that the width of the mean shear layer of the jet cannot be neglected. The results are discussed with reference to recent nozzle attenuation measurements.

  15. Low Mach Number Modeling of Core Convection in Massive Stars

    NASA Astrophysics Data System (ADS)

    Gilet, C.; Almgren, A. S.; Bell, J. B.; Nonaka, A.; Woosley, S. E.; Zingale, M.

    2013-08-01

    This work presents three-dimensional simulations of core convection in a 15 M ⊙ star halfway through its main sequence lifetime. To perform the necessary long-time calculations, we use the low Mach number code MAESTRO, with initial conditions taken from a one-dimensional stellar model. We first identify several key factors that the one-dimensional initial model must satisfy to ensure efficient simulation of the convection process. We then use the three-dimensional simulations to examine the effects of two common modeling choices on the resulting convective flow: using a fixed composition approximation and using a reduced domain size. We find that using a fixed composition model actually increases the computational cost relative to using the full multi-species model because the fixed composition system takes longer to reach convection that is in a quasi-static state. Using a reduced (octant rather than full sphere) simulation domain yields flow with statistical properties that are within a factor of two of the full sphere simulation values. Both the octant and full sphere simulations show similar mixing across the convection zone boundary that is consistent with the turbulent entrainment model. However, the global character of the flow is distinctly different in the octant simulation, showing more rapid changes in the large-scale structure of the flow and thus a more isotropic flow on average.

  16. LOW MACH NUMBER MODELING OF CORE CONVECTION IN MASSIVE STARS

    SciTech Connect

    Gilet, C.; Almgren, A. S.; Bell, J. B.; Nonaka, A.; Woosley, S. E.; Zingale, M.

    2013-08-20

    This work presents three-dimensional simulations of core convection in a 15 M{sub Sun} star halfway through its main sequence lifetime. To perform the necessary long-time calculations, we use the low Mach number code MAESTRO, with initial conditions taken from a one-dimensional stellar model. We first identify several key factors that the one-dimensional initial model must satisfy to ensure efficient simulation of the convection process. We then use the three-dimensional simulations to examine the effects of two common modeling choices on the resulting convective flow: using a fixed composition approximation and using a reduced domain size. We find that using a fixed composition model actually increases the computational cost relative to using the full multi-species model because the fixed composition system takes longer to reach convection that is in a quasi-static state. Using a reduced (octant rather than full sphere) simulation domain yields flow with statistical properties that are within a factor of two of the full sphere simulation values. Both the octant and full sphere simulations show similar mixing across the convection zone boundary that is consistent with the turbulent entrainment model. However, the global character of the flow is distinctly different in the octant simulation, showing more rapid changes in the large-scale structure of the flow and thus a more isotropic flow on average.

  17. Interaction of upstream flow distortions with high Mach number cascades

    NASA Technical Reports Server (NTRS)

    Englert, G. W.

    1981-01-01

    Features of the interaction of flow distortions, such as gusts and wakes with blade rows of advance type fans and compressors having high tip Mach numbers are modeled. A typical disturbance was assumed to have harmonic time dependence and was described, at a far upstream location, in three orthogonal spatial coordinates by a double Fourier series. It was convected at supersonic relative to a linear cascade described as an unrolled annulus. Conditions were selected so that the component of this velocity parallel to the axis of the turbomachine was subsonic, permitting interaction between blades through the upstream as well as downstream flow media. A strong, nearly normal shock was considered in the blade passages which was allowed curvature and displacement. The flows before and after the shock were linearized relative to uniform mean velocities in their respective regions. Solution of the descriptive equations was by adaption of the Wiener-Hopf technique, enabling a determination of distortion patterns through and downstream of the cascade as well as pressure distributions on the blade and surfaces. Details of interaction of the disturbance with the in-passage shock were discussed. Infuences of amplitude, wave length, and phase of the disturbance on lifts and moments of cascade configurations are presented. Numerical results are clarified by reference to an especially orderly pattern of upstream vertical motion in relation to the cascade parameters.

  18. Analytic MHD Theory for Earth's Bow Shock at Low Mach Numbers

    NASA Technical Reports Server (NTRS)

    Grabbe, Crockett L.; Cairns, Iver H.

    1995-01-01

    A previous MHD theory for the density jump at the Earth's bow shock, which assumed the Alfven M(A) and sonic M(s) Mach numbers are both much greater than 1, is reanalyzed and generalized. It is shown that the MHD jump equation can be analytically solved much more directly using perturbation theory, with the ordering determined by M(A) and M(s), and that the first-order perturbation solution is identical to the solution found in the earlier theory. The second-order perturbation solution is calculated, whereas the earlier approach cannot be used to obtain it. The second-order terms generally are important over most of the range of M(A) and M(s) in the solar wind when the angle theta between the normal to the bow shock and magnetic field is not close to 0 deg or 180 deg (the solutions are symmetric about 90 deg). This new perturbation solution is generally accurate under most solar wind conditions at 1 AU, with the exception of low Mach numbers when theta is close to 90 deg. In this exceptional case the new solution does not improve on the first-order solutions obtained earlier, and the predicted density ratio can vary by 10-20% from the exact numerical MHD solutions. For theta approx. = 90 deg another perturbation solution is derived that predicts the density ratio much more accurately. This second solution is typically accurate for quasi-perpendicular conditions. Taken together, these two analytical solutions are generally accurate for the Earth's bow shock, except in the rare circumstance that M(A) is less than or = 2. MHD and gasdynamic simulations have produced empirical models in which the shock's standoff distance a(s) is linearly related to the density jump ratio X at the subsolar point. Using an empirical relationship between a(s) and X obtained from MHD simulations, a(s) values predicted using the MHD solutions for X are compared with the predictions of phenomenological models commonly used for modeling observational data, and with the predictions of a

  19. Role of Turbulent Prandtl Number on Heat Flux at Hypersonic Mach Numbers

    NASA Technical Reports Server (NTRS)

    Xiao, X.; Edwards, J. R.; Hassan, H. A.; Gaffney, R. L., Jr.

    2007-01-01

    A new turbulence model suited for calculating the turbulent Prandtl number as part of the solution is presented. The model is based on a set of two equations: one governing the variance of the enthalpy and the other governing its dissipation rate. These equations were derived from the exact energy equation and thus take into consideration compressibility and dissipation terms. The model is used to study two cases involving shock wave/boundary layer interaction at Mach 9.22 and Mach 5.0. In general, heat transfer prediction showed great improvement over traditional turbulence models where the turbulent Prandtl number is assumed constant. It is concluded that using a model that calculates the turbulent Prandtl number as part of the solution is the key to bridging the gap between theory and experiment for flows dominated by shock wave/boundary layer interactions.

  20. Role of Turbulent Prandtl Number on Heat Flux at Hypersonic Mach Numbers

    NASA Technical Reports Server (NTRS)

    Gaffney, R. L., Jr.; Xiao, X.; Edwards, J. R.; Hassan, H. A.

    2005-01-01

    A new turbulence model suited for calculating the turbulent Prandtl number as part of the solution is presented. The model is based on a set of two equations: one governing the variance of the enthalpy and the other governing its dissipation rate. These equations were derived from the exact energy equation and thus take into consideration compressibility and dissipation terms. The model is used to study two cases involving shock wave/boundary layer interaction at Mach 9.22 and Mach 5.0. In general, heat transfer prediction showed great improvement over traditional turbulence models where the turbulent Prandtl number is assumed constant. It is concluded that using a model that calculates the turbulent Prandtl number as part of the solution is the key to bridging the gap between theory and experiment for flows dominated by shock wave/boundary layer interactions.

  1. The hypersonic Mach number independence principle in the case of viscous flow

    NASA Astrophysics Data System (ADS)

    Kliche, D.; Mundt, Ch.; Hirschel, E. H.

    2011-08-01

    The hypersonic Mach number independence principle of Oswatitsch is important for hypersonic vehicle design. It explains why, above a certain flight Mach number ( M ∞ ≈ 4-6, depending on the body shape), some aerodynamic properties become independent of the flight Mach number. For ground test facilities this means that it is sufficient for the Mach number in the test section to be high enough, that Mach number independence exists. However, the principle was derived for calorically perfect gas and inviscid flow only. In this paper a theoretical study for blunt bodies in the case of viscous flow is presented. We provide numerical results which give insight into how attached viscous flow behaves at high Mach numbers. The flow past an axisymmetric configuration is analysed by applying a coupled Euler/second-order boundary-layer method. Wall boundaries are treated by assuming an adiabatic or radiation-adiabatic wall for laminar flow. Calorically perfect or equilibrium air is accounted for. Lift, drag, and moment coefficients, and lift-to-drag ratios are given for several combinations of flight Mach number and altitude, i.e. Reynolds number. For blunt bodies considered here, which are pressure dominated, Mach number independence occurs for the adiabatic wall, but not for the radiation-adiabatic wall assumption.

  2. Low Mach Number Simulation of Core Convection in Massive Stars

    NASA Astrophysics Data System (ADS)

    Gilet, Candace Elise

    This work presents three-dimensional simulations of core convection in a 15 solar mass star halfway through its main sequence lifetime. We examine the effects of two common modeling choices on the resulting convective flow: using a reduced domain size and using a monatomic, or single species, approximation. We compare a multi-species simulation on a full sphere (360 degree) domain with a multi-species simulation on an octant domain and also with a single species simulation on a full sphere domain. To perform the long-time calculations, we use the new low Mach number code MAESTRO. The first part of this work deals with numerical aspects of using MAESTRO for the core convection system, a new application for MAESTRO. We extend MAESTRO to include two new models, a single species model and a simplified two-dimensional planar model, to aid in the exploration of using MAESTRO for core convection in massive stars. We discuss using MAESTRO with a novel spherical geometry domain configuration, namely, with the outer boundary located in the interior of the star, and show how this can create spurious velocities that must be numerically damped using a sponging layer. We describe the preparation of the initial model for the simulation. We find that assuring neutral stratification in the convective core and reasonable resolution of the gravity waves in the stable layer are key factors in generating suitable initial conditions for the simulation. Further, we examine a numerical aspect of the velocity constraint that is part of the low Mach number formulation of the Euler equations. In particular, we investigate the numerical procedure for computing beta0, the density-like variable that captures background stratification in the velocity constraint, and find that the original method of computation remains a good choice. The three-dimensional simulation results show that using a single species model actually increases the computational cost of the simulation because the single

  3. Numerical Simulation of a High Mach Number Jet Flow

    NASA Technical Reports Server (NTRS)

    Hayder, M. Ehtesham; Turkel, Eli; Mankbadi, Reda R.

    1993-01-01

    The recent efforts to develop accurate numerical schemes for transition and turbulent flows are motivated, among other factors, by the need for accurate prediction of flow noise. The success of developing high speed civil transport plane (HSCT) is contingent upon our understanding and suppression of the jet exhaust noise. The radiated sound can be directly obtained by solving the full (time-dependent) compressible Navier-Stokes equations. However, this requires computational storage that is beyond currently available machines. This difficulty can be overcome by limiting the solution domain to the near field where the jet is nonlinear and then use acoustic analogy (e.g., Lighthill) to relate the far-field noise to the near-field sources. The later requires obtaining the time-dependent flow field. The other difficulty in aeroacoustics computations is that at high Reynolds numbers the turbulent flow has a large range of scales. Direct numerical simulations (DNS) cannot obtain all the scales of motion at high Reynolds number of technological interest. However, it is believed that the large scale structure is more efficient than the small-scale structure in radiating noise. Thus, one can model the small scales and calculate the acoustically active scales. The large scale structure in the noise-producing initial region of the jet can be viewed as a wavelike nature, the net radiated sound is the net cancellation after integration over space. As such, aeroacoustics computations are highly sensitive to errors in computing the sound sources. It is therefore essential to use a high-order numerical scheme to predict the flow field. The present paper presents the first step in a ongoing effort to predict jet noise. The emphasis here is in accurate prediction of the unsteady flow field. We solve the full time-dependent Navier-Stokes equations by a high order finite difference method. Time accurate spatial simulations of both plane and axisymmetric jet are presented. Jet Mach

  4. Numerical Simulation of Low Mach Number Fluid - Phenomena.

    NASA Astrophysics Data System (ADS)

    Reitsma, Scott H.

    A method for the numerical simulation of low Mach number (M) fluid-acoustic phenomena is developed. This computational fluid-acoustic (CFA) methodology is based upon a set of conservation equations, termed finite-compressible, derived from the unsteady Navier-Stokes equations. The finite-compressible and more familiar pseudo-compressible equations are compared. The impact of derivation assumptions are examined theoretically and through numerical experimentation. The error associated with these simplifications is shown to be of O(M) and proportional to the amplitude of unsteady phenomena. A computer code for the solution of the finite -compressible equations is developed from an existing pseudo -compressible code. Spatial and temporal discretization issues relevant in the context of near field fluid-acoustic simulations are discussed. The finite volume code employs a MUSCL based third order upwind biased flux difference splitting algorithm for the convective terms. An explicit, three stage, second order Runge-Kutta temporal integration is employed for time accurate simulations while an implicit, approximately factored time quadrature is available for steady state convergence acceleration. The CFA methodology is tested in a series of problems which examine the appropriateness of the governing equations, the exacerbation of spatial truncation errors and the degree of temporal accuracy. Characteristic based boundary conditions employing a spatial formulation are developed. An original non-reflective boundary condition based upon the generalization and extension of existing methods is derived and tested in a series of multi-dimensional problems including those involving viscous shear flows and propagating waves. The final numerical experiment is the simulation of boundary layer receptivity to acoustic disturbances. This represents the first simulation of receptivity at a surface inhomogeneity in which the acoustic phenomena is modeled using physically appropriate

  5. Aerodynamic characteristics at Mach numbers from 0.6 to 2.16 of a supersonic cruise fighter configuration with a design Mach number of 1.8

    NASA Technical Reports Server (NTRS)

    Shrout, B. L.

    1977-01-01

    An investigation was made in the Langley 8-foot transonic tunnel and the Langley Unitary Plan wind tunnel, over a Mach number range of 0.6 to 2.16, to determine the static longitudinal and lateral aerodynamic characteristics of a model of a supersonic-cruise fighter. The configuration, which is designed for efficient cruise at Mach number 1.8, is a twin-engine tailless arrow-wing concept with a single rectangular inlet beneath the fuselage and outboard vertical tails and ventral fins. It had untrimmed values of lift-drage ratio ranging from 10 at subsonic speeds to 6.4 at the design Mach number. The configuration was statically stable both longitudinally and laterally.

  6. Sidewall Mach Number Distributions for the NASA Langley Transonic Dynamics Tunnel

    NASA Technical Reports Server (NTRS)

    Florance, James R.; Rivera, Jose A., Jr.

    2001-01-01

    The Transonic Dynamics Tunnel(TDT) was recalibrated due to the conversion of the heavy gas test medium from R-12 to R-134a. The objectives of the tests were to determine the relationship between the free-stream Mach number and the measured test section Mach number, and to quantify any necessary corrections. Other tests included the measurement of pressure distributions along the test-section walls, test-section centerline, at certain tunnel stations via a rake apparatus, and in the tunnel settling chamber. Wall boundary layer, turbulence, and flow angularity measurements were also performed. This paper discusses the determination of sidewall Mach number distributions.

  7. The development of cambered airfoil sections having favorable lift characteristics at supercritical Mach numbers

    NASA Technical Reports Server (NTRS)

    Graham, Donald J

    1949-01-01

    Several groups of new airfoil sections, designated as the NACA 8-series, are derived analytically to have lift characteristics at supercritical Mach numbers which are favorable in the sense that the abrupt loss of lift, characteristic of the usual airfoil section at Mach numbers above the critical, is avoided. Aerodynamic characteristics determined from two-dimensional wind-tunnel tests at Mach numbers up to approximately 0.9 are presented for each of the derived airfoils. Comparisons are made between the characteristics of these airfoils and the corresponding characteristics of representative NACA 6-series airfoils.

  8. Numerical Analysis of the Trailblazer Inlet Flowfield for Hypersonic Mach Numbers

    NASA Technical Reports Server (NTRS)

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

    1999-01-01

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

  9. Low Mach Number Modeling of Type Ia Supernovae. II. EnergyEvolution

    SciTech Connect

    Almgren, Ann S.; Bell, John B.; Rendleman, Charles A.; Zingale,Mike

    2006-03-28

    The convective period leading up to a Type Ia supernova (SNIa) explosion is characterized by very low Mach number flows, requiringhydrodynamical methods well-suited to long-time integration. We continuethe development of the low Mach number equation set for stellar scaleflows by incorporating the effects of heat release due to externalsources. Low Mach number hydrodynamics equations with a time-dependentbackground state are derived, and a numerical method based on theapproximate projection formalism is presented. We demonstrate throughvalidation with a fully compressible hydrodynamics code that this lowMach number model accurately captures the expansion of the stellaratmosphere as well as the local dynamics due to external heat sources.This algorithm provides the basis for an efficient simulation tool forstudying the ignition of SNe Ia.

  10. Mach-Number Measurement with Laser and Pressure Probes in Humid Supersonic Flow

    NASA Technical Reports Server (NTRS)

    Herring, G. C.

    2008-01-01

    Mach-number measurements using a nonintrusive optical technique, laser-induced thermal acoustics (LITA), are compared to pressure probes in humid supersonic airflow. The two techniques agree well in dry flow (-35 C dew point), but LITA measurements show about five times larger fractional change in Mach number than that of the pressure-probe when water is purposefully introduced into the flow. Possible reasons for this discrepancy are discussed.

  11. Role of Turbulent Prandtl Number on Heat Flux at Hypersonic Mach Number

    NASA Technical Reports Server (NTRS)

    Xiao, X.; Edwards, J. R.; Hassan, H. A.

    2004-01-01

    Present simulation of turbulent flows involving shock wave/boundary layer interaction invariably overestimates heat flux by almost a factor of two. One possible reason for such a performance is a result of the fact that the turbulence models employed make use of Morkovin's hypothesis. This hypothesis is valid for non-hypersonic Mach numbers and moderate rates of heat transfer. At hypersonic Mach numbers, high rates of heat transfer exist in regions where shock wave/boundary layer interactions are important. As a result, one should not expect traditional turbulence models to yield accurate results. The goal of this investigation is to explore the role of a variable Prandtl number formulation in predicting heat flux in flows dominated by strong shock wave/boundary layer interactions. The intended applications involve external flows in the absence of combustion such as those encountered in supersonic inlets. This can be achieved by adding equations for the temperature variance and its dissipation rate. Such equations can be derived from the exact Navier-Stokes equations. Traditionally, modeled equations are based on the low speed energy equation where the pressure gradient term and the term responsible for energy dissipation are ignored. It is clear that such assumptions are not valid for hypersonic flows. The approach used here is based on the procedure used in deriving the k-zeta model, in which the exact equations that governed k, the variance of velocity, and zeta, the variance of vorticity, were derived and modeled. For the variable turbulent Prandtl number, the exact equations that govern the temperature variance and its dissipation rate are derived and modeled term by term. The resulting set of equations are free of damping and wall functions and are coordinate-system independent. Moreover, modeled correlations are tensorially consistent and invariant under Galilean transformation. The final set of equations will be given in the paper.

  12. Nonlinear standing Alfven wave current system at Io: Theory

    SciTech Connect

    Neubauer, F.M.

    1980-03-01

    We present a nonlinear analytical model of the Alfven current tubes continuing the currents through Io (or rather its ionosphere) generated by the unipolar inductor effect due to Io's motion relative to the magnetospheric plasma. We thereby extend the linear work by Drell et al. (1965) to the fully nonlinear, sub-Alfvenic situation also including flow which is not perpendicular to the background magnetic field. The following principal results have been obtained: (1) The portion of the currents feeding Io is aligned with the Alfven characteristics at an angle theta/sub A/ is the Alfven Mach number. (2) The Alfven tubes act like an external conductance ..sigma../sub A/=1/(..mu../sub 0/V/sub A/(1+M/sub A//sup 2/+2M/sub A/ sin theta)/sup 1/2/ where V/sub A/ is the Alfven wave propagation. Hence the Jovian ionospheric conductivity is not necessary for current closure. (3) In addition, the Alfven tubes may be reflected from either the torus boundary or the Jovian ionosphere. The efficiency of the resulting interaction with these boundaries varies with Io position. The interaction is particularly strong at extreme magnetic latitudes, thereby suggesting a mechanism for the Io control of decametric emissions. (4) The reflected Alfven waves may heat both the torus plasma and the Jovian ionosphere as well as produce increased diffusion of high-energy particles in the torus. (5) From the point of view of the electrodynamic interaction, Io is unique among the Jovian satellites for several reasons: these include its ionosphere arising from ionized volcanic gases, a high external Alfvenic conductance ..sigma../sub A/, and a high corotational voltage in addition to the interaction phenomenon with a boundary. (6) We find that Amalthea is probably strongly coupled to Jupiter's ionosphere while the outer Galilean satellites may occasionally experience super-Alfvenic conditions.

  13. Sensitivity of transport aircraft performance and economics to advanced technology and cruise Mach number

    NASA Technical Reports Server (NTRS)

    Ardema, M. D.

    1974-01-01

    Sensitivity data for advanced technology transports has been systematically collected. This data has been generated in two separate studies. In the first of these, three nominal, or base point, vehicles designed to cruise at Mach numbers .85, .93, and .98, respectively, were defined. The effects on performance and economics of perturbations to basic parameters in the areas of structures, aerodynamics, and propulsion were then determined. In all cases, aircraft were sized to meet the same payload and range as the nominals. This sensitivity data may be used to assess the relative effects of technology changes. The second study was an assessment of the effect of cruise Mach number. Three families of aircraft were investigated in the Mach number range 0.70 to 0.98: straight wing aircraft from 0.70 to 0.80; sweptwing, non-area ruled aircraft from 0.80 to 0.95; and area ruled aircraft from 0.90 to 0.98. At each Mach number, the values of wing loading, aspect ratio, and bypass ratio which resulted in minimum gross takeoff weight were used. As part of the Mach number study, an assessment of the effect of increased fuel costs was made.

  14. On the instabilities of supersonic mixing layers - A high-Mach-number asymptotic theory

    NASA Technical Reports Server (NTRS)

    Balsa, Thomas F.; Goldstein, M. E.

    1990-01-01

    The stability of a family of tanh mixing layers is studied at large Mach numbers using perturbation methods. It is found that the eigenfunction develops a multilayered structure, and the eigenvalue is obtained by solving a simplified version of the Rayleigh equation (with homogeneous boundary conditions) in one of these layers which lies in either of the external streams. This analysis leads to a simple hypersonic similarity law which explains how spatial and temporal phase speeds and growth rates scale with Mach number and temperature ratio. Comparisons are made with numerical results, and it is found that this similarity law provides a good qualitative guide for the behavior of the instability at high Mach numbers. In addition to this asymptotic theory, some fully numerical results are also presented (with no limitation on the Mach number) in order to explain the origin of the hypersonic modes (through mode splitting) and to discuss the role of oblique modes over a very wide range of Mach number and temperature ratio.

  15. Mach number study of supersonic turbulence: the properties of the density field

    NASA Astrophysics Data System (ADS)

    Konstandin, L.; Schmidt, W.; Girichidis, P.; Peters, T.; Shetty, R.; Klessen, R. S.

    2016-08-01

    We analyse the scaling properties of turbulent flows using a suite of three-dimensional numerical simulations. We model driven, compressible, isothermal, turbulence with Mach numbers ranging from the subsonic ({M} ≈ 0.5) to the highly supersonic regime ({M}≈ 16). The forcing scheme consists of both solenoidal (transverse) and compressive (longitudinal) modes in equal parts. We confirm the relation σ s^2 = ln {(1+b^2{M}^2)} between the Mach number and the standard deviation of the logarithmic density with b = 0.33. We find increasing deviations with higher Mach number from the predicted lognormal shape in the high-density wing of the density probability density function. The density spectra follow {D}(k, {M}) ∝ k^{ζ ({M})} with scaling exponents depending on the Mach number. We find ζ ({M}) = α {M}^{β } with coefficients α = -2.1 and β = -0.33. The dependence of the scaling exponent on the Mach number implies a fractal dimension D=2+1.05 {M}^{-0.33}.

  16. Mach number effects on conical surface features of swept shock boundary-layer interactions

    NASA Technical Reports Server (NTRS)

    Lu, F. K.; Settles, G. S.; Horstman, C. C.

    1987-01-01

    A joint experimental and computational study is made of the shock-wave turbulent boundary-layer interaction generated by sharp fins, with emphasis on Mach-number effects. The Mach number range is from 2 to 4 and the unit Reynolds number is from 50 to 80 million per meter. Fin angles are varied from 4 to 22 deg. Surface-flow patterns are obtained using a color surface-flow-visualization technique. The results show that the upstream-influence response in the conical far-field region is a function of the freestream Mach number and the shock strength. A new interpretation of the behavior of the upstream influence with changes of the inviscid shock angle is given. Agreement between the experimental and the computed upstream-influence lines becomes poorer for stronger interactions, with the computations underpredicting the upstream-influence line.

  17. A time-accurate implicit method for chemically reacting flows at all Mach numbers

    NASA Technical Reports Server (NTRS)

    Withington, J. P.; Yang, V.; Shuen, J. S.

    1991-01-01

    The objective of this work is to develop a unified solution algorithm capable of treating time-accurate chemically reacting flows at all Mach numbers, ranging from molecular diffusion velocities to supersonic speeds. A rescaled pressure term is used in the momentum equation to circumvent the singular behavior of pressure at low Mach numbers. A dual time-stepping integration procedure is established. The system eigenvalues become well behaved and have the same order of magnitude, even in the very low Mach number regime. The computational efficiency for moderate and high speed flow is competitive with the conventional density-based scheme. The capabilities of the algorithm are demonstrated by applying it to selected model problems including nozzle flows and flame dynamics.

  18. Detailed noise measurements on the SR-7A propeller: Tone behavior with helical tip Mach number

    NASA Technical Reports Server (NTRS)

    Dittmar, James H.; Hall, David G.

    1991-01-01

    Detailed noise measurements were taken on the SR-7A propeller to investigate the behavior of the noise with helical tip Mach number and then to level off as Mach number was increased further. This behavior was further investigated by obtaining detailed pressure-time histories of data. The pressure-time histories indicate that a portion of the primary pressure pulse is progressively cancelled by a secondary pulse which results in the noise leveling off as the helical tip Mach number is increased. This second pulse appears to originate on the same blade as the primary pulse and is in some way connected to the blade itself. This leaves open the possibility of redesigning the blade to improve the cancellation; thereby, the propeller noise is reduced.

  19. Guderley reflection for higher Mach numbers in a standard shock tube

    NASA Astrophysics Data System (ADS)

    Cachucho, A.; Skews, B. W.

    2012-03-01

    An experimental study shows that the Guderley reflection (GR) of shock waves can be produced in a standard shock tube. A new technique was utilised which comprises triple point of a developed weak Mach reflection undergoing a number of reflections off the ceiling and floor of the shock tube before arriving at the test section. Both simple perturbation sources and diverging ramps were used to generate a transverse wave in the tube which then becomes the weak reflected wave of the reflection pattern. Tests were conducted for three ramp angles (10°, 15°, and 20°) and two perturbation sources for a range of Mach numbers (1.10-1.40) and two shock tube expansion chamber lengths (2.0 and 4.0 m). It was found that the length of the Mach stem of the reflection pattern is the overall vertical distance traveled by the triple point. Images with equivalent Mach stem lengths in the order of 2.0 m were produced. All tests showed evidence of the fourth wave of the GR, namely the expansion wave behind the reflected shock wave. A shocklet terminating the expansion wave was also identified in a few cases mainly for incident wave Mach numbers of approximately 1.20.

  20. Dispersive nature of high mach number collisionless plasma shocks: Poynting flux of oblique whistler waves.

    PubMed

    Sundkvist, David; Krasnoselskikh, V; Bale, S D; Schwartz, S J; Soucek, J; Mozer, F

    2012-01-13

    Whistler wave trains are observed in the foot region of high Mach number quasiperpendicular shocks. The waves are oblique with respect to the ambient magnetic field as well as the shock normal. The Poynting flux of the waves is directed upstream in the shock normal frame starting from the ramp of the shock. This suggests that the waves are an integral part of the shock structure with the dispersive shock as the source of the waves. These observations lead to the conclusion that the shock ramp structure of supercritical high Mach number shocks is formed as a balance of dispersion and nonlinearity.

  1. 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.

  2. 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.

  3. Comparison of jet Mach number decay data with a correlation and jet spreading contours for a large variety of nozzles

    NASA Technical Reports Server (NTRS)

    Groesbeck, D. E.; Huff, R. G.; Vonglahn, U. H.

    1977-01-01

    Small-scale circular, noncircular, single- and multi-element nozzles with flow areas as large as 122 sq cm were tested with cold airflow at exit Mach numbers from 0.28 to 1.15. The effects of multi-element nozzle shape and element spacing on jet Mach number decay were studied in an effort to reduce the noise caused by jet impingement on externally blown flap (EBF) STOL aircraft. The jet Mach number decay data are well represented by empirical relations. Jet spreading and Mach number decay contours are presented for all configurations tested.

  4. A NEW DENSITY VARIANCE-MACH NUMBER RELATION FOR SUBSONIC AND SUPERSONIC ISOTHERMAL TURBULENCE

    SciTech Connect

    Konstandin, L.; Girichidis, P.; Federrath, C.; Klessen, R. S.

    2012-12-20

    The probability density function of the gas density in subsonic and supersonic, isothermal, driven turbulence is analyzed using a systematic set of hydrodynamical grid simulations with resolutions of up to 1024{sup 3} cells. We perform a series of numerical experiments with root-mean-square (rms) Mach number M ranging from the nearly incompressible, subsonic (M=0.1) to the highly compressible, supersonic (M=15) regime. We study the influence of two extreme cases for the driving mechanism by applying a purely solenoidal (divergence-free) and a purely compressive (curl-free) forcing field to drive the turbulence. We find that our measurements fit the linear relation between the rms Mach number and the standard deviation (std. dev.) of the density distribution in a wide range of Mach numbers, where the proportionality constant depends on the type of forcing. In addition, we propose a new linear relation between the std. dev. of the density distribution {sigma}{sub {rho}} and that of the velocity in compressible modes, i.e., the compressible component of the rms Mach number, M{sub comp}. In this relation the influence of the forcing is significantly reduced, suggesting a linear relation between {sigma}{sub {rho}} and M{sub comp}, independent of the forcing, and ranging from the subsonic to the supersonic regime.

  5. Collisionless relaxation of downstream ion distributions in low-Mach number shocks

    SciTech Connect

    Gedalin, M.; Friedman, Y.; Balikhin, M.

    2015-07-15

    Collisionlessly formed downstream distributions of ions in low-Mach number shocks are studied. General expressions for the asymptotic value of the ion density and pressure are derived for the directly transmitted ions. An analytical approximation for the overshoot strength is suggested for the low-β case. Spatial damping scale of the downstream magnetic oscillations is estimated.

  6. Tests of a Hermes A-2 Missile Body at Mach Number 4.04

    NASA Technical Reports Server (NTRS)

    Ulmann, Edward F.; Lord, Douglas R.

    1950-01-01

    Force tests on a proposed body shape of the Hermes A-2 missile with and without longitudinal spoilers were made at Mach number 4.04. Values of normal force coefficient, pitching-moment coefficient, and center-of-pressure position were obtained.

  7. Mach Number Dependence of Near Wall Structure in Compressible Channel Flows

    NASA Astrophysics Data System (ADS)

    Pei, J.; Chen, J.; She, Z. S.; Hussain, F.

    2011-09-01

    A newly developed statistical correlation structure is used to analyze compressible channel flows up to M = 3.0. Using velocity-vorticity correlation structure (VVCS), the Mach number dependence of the characteristic scales of near wall structure are analyzed. The detailed results show that the length scale and the spanwise spacing of VVCS exponentially increase with Mach number in the near wall region. For example, for VVCSuωx, the length scale of the statistical streamwise structure is Lxuωx = e6.5 + M/2.8 + (M/4.1)2, and spacing between the structure is Dxuωx = 60eM/2.2 + 13.3, where the parameters 2.8, 4.1 and 2.2 are characteristic Mach numbers to be explained further. The geometrical features of the statistical structure are consistent with the observations of Coleman et al., and it is also argued that the quantitative relationship between the characteristic scales of VVCS and Mach number is important to consider in performing numerical computation of compressible flows. This study also suggests that a set of geometrical structures should be invoked for modeling inhomogeneous compressible shear flows.

  8. A two-dimensional, TVD numerical scheme for inviscid, high Mach number flows in chemical equilibrium

    NASA Technical Reports Server (NTRS)

    Eberhardt, S.; Palmer, G.

    1986-01-01

    A new algorithm has been developed for hypervelocity flows in chemical equilibrium. Solutions have been achieved for Mach numbers up to 15 with no adverse effect on convergence. Two methods of coupling an equilibrium chemistry package have been tested, with the simpler method proving to be more robust. Improvements in boundary conditions are still required for a production-quality code.

  9. Effects of Aspect Ratio on Air Flow at High Subsonic Mach Numbers

    NASA Technical Reports Server (NTRS)

    Lindsey, W F; Humphreys, Milton D

    1952-01-01

    Schlieren photographs were used in an investigation to determine the effects of changing the aspect ratio from infinity to 2 on the air flow past a wing at high subsonic Mach numbers. The results indicated that the decreased effects of compressibility on drag coefficients for the finite wing are produced by a reduction in the compression shock and flow separation.

  10. Turbulent mixing of a slightly supercritical van der Waals fluid at low-Mach number

    SciTech Connect

    Battista, F.; Casciola, C. M.; Picano, F.

    2014-05-15

    Supercritical fluids near the critical point are characterized by liquid-like densities and gas-like transport properties. These features are purposely exploited in different contexts ranging from natural products extraction/fractionation to aerospace propulsion. Large part of studies concerns this last context, focusing on the dynamics of supercritical fluids at high Mach number where compressibility and thermodynamics strictly interact. Despite the widespread use also at low Mach number, the turbulent mixing properties of slightly supercritical fluids have still not investigated in detail in this regime. This topic is addressed here by dealing with Direct Numerical Simulations of a coaxial jet of a slightly supercritical van der Waals fluid. Since acoustic effects are irrelevant in the low Mach number conditions found in many industrial applications, the numerical model is based on a suitable low-Mach number expansion of the governing equation. According to experimental observations, the weakly supercritical regime is characterized by the formation of finger-like structures – the so-called ligaments – in the shear layers separating the two streams. The mechanism of ligament formation at vanishing Mach number is extracted from the simulations and a detailed statistical characterization is provided. Ligaments always form whenever a high density contrast occurs, independently of real or perfect gas behaviors. The difference between real and perfect gas conditions is found in the ligament small-scale structure. More intense density gradients and thinner interfaces characterize the near critical fluid in comparison with the smoother behavior of the perfect gas. A phenomenological interpretation is here provided on the basis of the real gas thermodynamics properties.

  11. Turbulent mixing of a slightly supercritical van der Waals fluid at low-Mach number

    NASA Astrophysics Data System (ADS)

    Battista, F.; Picano, F.; Casciola, C. M.

    2014-05-01

    Supercritical fluids near the critical point are characterized by liquid-like densities and gas-like transport properties. These features are purposely exploited in different contexts ranging from natural products extraction/fractionation to aerospace propulsion. Large part of studies concerns this last context, focusing on the dynamics of supercritical fluids at high Mach number where compressibility and thermodynamics strictly interact. Despite the widespread use also at low Mach number, the turbulent mixing properties of slightly supercritical fluids have still not investigated in detail in this regime. This topic is addressed here by dealing with Direct Numerical Simulations of a coaxial jet of a slightly supercritical van der Waals fluid. Since acoustic effects are irrelevant in the low Mach number conditions found in many industrial applications, the numerical model is based on a suitable low-Mach number expansion of the governing equation. According to experimental observations, the weakly supercritical regime is characterized by the formation of finger-like structures - the so-called ligaments - in the shear layers separating the two streams. The mechanism of ligament formation at vanishing Mach number is extracted from the simulations and a detailed statistical characterization is provided. Ligaments always form whenever a high density contrast occurs, independently of real or perfect gas behaviors. The difference between real and perfect gas conditions is found in the ligament small-scale structure. More intense density gradients and thinner interfaces characterize the near critical fluid in comparison with the smoother behavior of the perfect gas. A phenomenological interpretation is here provided on the basis of the real gas thermodynamics properties.

  12. Application of a transitional boundary-layer theory in the low hypersonic Mach number regime

    NASA Technical Reports Server (NTRS)

    Shamroth, S. J.; Mcdonald, H.

    1975-01-01

    An investigation is made to assess the capability of a finite-difference boundary-layer procedure to predict the mean profile development across a transition from laminar to turbulent flow in the low hypersonic Mach-number regime. The boundary-layer procedure uses an integral form of the turbulence kinetic-energy equation to govern the development of the Reynolds apparent shear stress. The present investigation shows the ability of this procedure to predict Stanton number, velocity profiles, and density profiles through the transition region and, in addition, to predict the effect of wall cooling and Mach number on transition Reynolds number. The contribution of the pressure-dilatation term to the energy balance is examined and it is suggested that transition can be initiated by the direct absorption of acoustic energy even if only a small amount (1 per cent) of the incident acoustic energy is absorbed.

  13. Assessment of a transitional boundary layer theory at low hypersonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Shamroth, S. J.; Mcdonald, H.

    1972-01-01

    An investigation was carried out to assess the accuracy of a transitional boundary layer theory in the low hypersonic Mach number regime. The theory is based upon the simultaneous numerical solution of the boundary layer partial differential equations for the mean motion and an integral form of the turbulence kinetic energy equation which controls the magnitude and development of the Reynolds stress. Comparisions with experimental data show the theory is capable of accurately predicting heat transfer and velocity profiles through the transitional regime and correctly predicts the effects of Mach number and wall cooling on transition Reynolds number. The procedure shows promise of predicting the initiation of transition for given free stream disturbance levels. The effects on transition predictions of the pressure dilitation term and of direct absorption of acoustic energy by the boundary layer were evaluated.

  14. The Design of Variable Mach Number Asymmetric Super-Sonic Nozzles by Two Procedures Employing Inclined and Curved Sonic Lines

    NASA Technical Reports Server (NTRS)

    Syvertson, Clarence A.; Savin, Raymond C.

    1951-01-01

    Two theoretical procedures are developed for designing asymmetric supersonic nozzles for which the calculated exit flow is nearly uniform over a range of Mach numbers. One procedure is applicable at Mach numbers less than approximately 3. This approach yields, without iteration, a nozzle for which the calculated exit flow is uniform at two Mach numbers and, with proper design, is nearly uniform at Mach numbers between, slightly above, and slightly below these two. The use of an inclined and curved sonic line is an essential feature of this approach, The second procedure requires iteration and is used far designs at Mach numbers exceeding 3. Although it is not a necessary feature, an inclined and curved sonic line is also used in this procedure. In both approaches the flow field dawn stream of the sonic line is determined using the method of characteristics.

  15. Investigation of the Influence of Convective Mach Number on Compressible Plane Jet Exhausting into Parallel Streams

    NASA Astrophysics Data System (ADS)

    Datta, Abanti; Sinhamahapatra, Kalyan Prasad

    2016-08-01

    The effects of convective Mach number on plane jets with parallel co-flow streams are investigated in the present study. Two-dimensional viscous compressible plane jet issuing to co-flowing parallel streams is numerically simulated using higher order spatial and temporal integration schemes. To remove mesh induced non-uniformities explicit tridiagonal spatial filter is applied. The mean flow field and turbulence statistics are captured from the simulated results using time-average over streamwise direction. The captured mean velocity profiles in appropriate non-dimensional form exhibit self-similar behaviour and do not change considerably with convective Mach number. The streamwise mean excess velocity profiles collapse very well with previous experimental data. The turbulent intensity profiles and Reynolds shear stress profiles do not show self-similar characteristic and increase slowly in farther downstream. The two-dimensional simulation is found capable of capturing correct jet spreading and decay.

  16. Statistics of the cosmic Mach number from numerical simulations of a cold dark matter universe

    NASA Technical Reports Server (NTRS)

    Suto, Yasushi; Cen, Renyue; Ostriker, Jeremiah P.

    1992-01-01

    Results are presented of an analysis of the cosmic Mach number, M, the ratio of the streaming velocity, v, to the random velocity dispersion, sigma, of galaxies in a given patch of the universe, which was performed on the basis of hydrodynamical simulations of the cold dark matter scenario. Galaxy formation is modeled by application of detailed physical processes rather than by the ad hoc assumption of 'bias' between dark matter and galaxy fluctuations. The correlation between M and sigma is found to be very weak for both components. No evidence is found for a physical 'velocity bias' in the quantities which appear in the definition of M. Standard cold-dark-matter-dominated universes are in conflict, at a statistically significant level, with the available observation, in that they predict a Mach number considerably lower than is observed.

  17. Mach number scaling of helicopter rotor blade/vortex interaction noise

    NASA Technical Reports Server (NTRS)

    Leighton, Kenneth P.; Harris, Wesley L.

    1985-01-01

    A parametric study of model helicopter rotor blade slap due to blade vortex interaction (BVI) was conducted in a 5 by 7.5-foot anechoic wind tunnel using model helicopter rotors with two, three, and four blades. The results were compared with a previously developed Mach number scaling theory. Three- and four-bladed rotor configurations were found to show very good agreement with the Mach number to the sixth power law for all conditions tested. A reduction of conditions for which BVI blade slap is detected was observed for three-bladed rotors when compared to the two-bladed baseline. The advance ratio boundaries of the four-bladed rotor exhibited an angular dependence not present for the two-bladed configuration. The upper limits for the advance ratio boundaries of the four-bladed rotors increased with increasing rotational speed.

  18. Identification of the aeroacoustic response of a low Mach number flow through a T-joint.

    PubMed

    Martínez-Lera, P; Schram, C; Föller, S; Kaess, R; Polifke, W

    2009-08-01

    A methodology to study numerically the aeroacoustic response of low Mach number confined flows to acoustic excitations is presented. The approach combines incompressible flow computations, vortex sound theory, and system identification techniques, and is applied here to study the behavior of a two-dimensional laminar flow through a T-joint. Comparison with experimental results available in literature shows that the computed source models capture the main physical mechanisms of the sound production in the shear layer of the T-joint.

  19. Laminar friction and heat transfer at Mach numbers from 1 to 10

    NASA Technical Reports Server (NTRS)

    Klunker, E B; Mclean, F Edward

    1951-01-01

    Velocity and temperature profiles and laminar boundary-layer characteristics have been computed for Mach numbers from 1 to 10, utilizing experimental values of the heat capacity, viscosity, and conductivity. The analysis shows that effective temperature, which is a function of the surface temperature and stream conditions, arises naturally and is the proper reference temperature to be used in heat-transfer calculations. The effective temperature and the recovery temperature become identical for the condition of zero heat transfer.

  20. Nearfield Unsteady Pressures at Cruise Mach Numbers for a Model Scale Counter-Rotation Open Rotor

    NASA Technical Reports Server (NTRS)

    Stephens, David B.

    2012-01-01

    An open rotor experiment was conducted at cruise Mach numbers and the unsteady pressure in the nearfield was measured. The system included extensive performance measurements, which can help provide insight into the noise generating mechanisms in the absence of flow measurements. A set of data acquired at a constant blade pitch angle but various rotor speeds was examined. The tone levels generated by the front and rear rotor were found to be nearly equal when the thrust was evenly balanced between rotors.

  1. A half-explicit, non-split projection method for low Mach number flows.

    SciTech Connect

    Pousin, Jerome G.; Najm, Habib N.; Pebay, Philippe Pierre

    2004-02-01

    In the context of the direct numerical simulation of low MACH number reacting flows, the aim of this article is to propose a new approach based on the integration of the original differential algebraic (DAE) system of governing equations, without further differentiation. In order to do so, while preserving a possibility of easy parallelization, it is proposed to use a one-step index 2 DAE time-integrator, the Half Explicit Method (HEM). In this context, we recall why the low MACH number approximation belongs to the class of index 2 DAEs and discuss why the pressure can be associated with the constraint. We then focus on a fourth-order HEM scheme, and provide a formulation that makes its implementation more convenient. Practical details about the consistency of initial conditions are discussed, prior to focusing on the implicit solve involved in the method. The method is then evaluated using the Modified KAPS Problem, since it has some of the features of the low MACH number approximation. Numerical results are presented, confirming the above expectations. A brief summary of ongoing efforts is finally provided.

  2. Mach Number Effects on Ignition and Mixing Processes in a Reacting Shock-Bubble Interaction

    NASA Astrophysics Data System (ADS)

    Hickel, Stefan; Diegelmann, Felix; Tritschler, Volker

    2015-11-01

    We investigate reacting shock-bubble interactions (RSBI) by direct numerical simulations (DNS) with detailed chemical reaction kinetics. The bubble contains a stoichiometric H2-O2 gas mixture and is surrounded by pure N2. The interaction with a planar shock wave induces Richtmyer-Meshkov instability. Secondary instabilities develop into a turbulent mixing zone at the bubble interface. The transmitted shock focuses at the downstream pole of the bubble and may ignite the bubble gas. To trigger different reaction wave types, we performed DNS of RSBI for shock Mach numbers in the range of Ma = 2 . 13 - 2 . 50 at a constant initial pressure of p0 = 0 . 50 atm. Deflagration, dominated by H, O and OH production, is observed for a shock Mach number of Ma = 2 . 13 . Increasing the shock Mach number reduces the induction time and eventually leads to deflagration-detonation transition. Ignition by a Ma = 2 . 50 shock wave directly leads to a detonation wave, driven by HO2 and H2O2 high-pressure chemistry. Richtmyer-Meshkov instability, subsequent Kelvin Helmholtz instabilities, and bubble expansion are highly affected by the reaction wave. Mixing is significantly decreased by both reaction waves types. In particular detonation waves reduce the mixing distinctly.

  3. A study of multi-body aerodynamic interference at transonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Cottrell, Charles J.; Martinez, Agusto; Chapman, Gary T.

    1987-01-01

    A wind tunnel experiment involving single, double, and triple combinations of mutually interfering generic, unfinned aircraft stores has been conducted. Each combination of stores was tested at Mach numbers from 0.60 to 1.20 and at angles of attack from 0 to 25 deg for the single store and from 0 to 6 deg for the double and triple store configurations. Extensive axial and circumferential pressure and flow visualization data at each store location were obtained. Euler solutions for each configuration at 0 deg incidence have been generated and compared with experimental data. This comparison indicates an Euler flow solver can yield accurate predictions of the location and magnitude of multibody interference provided an appropriate grid is used and the viscous effects associated with these configurations remain small. The data indicate multibody interference in the transonic region increases as the freestream Mach number approaches 1 from either direction, and subsides as the Mach number moves away from sonic conditions. This interference is characterized by a large, localized reduction in pressure on the inboard surfaces of the bodies which results in forces that draw the configuration closer together.

  4. An investigation of several NACA 1-series nose inlets with and without protruding central bodies at high-subsonic Mach numbers and at a Mach number of 1.2

    NASA Technical Reports Server (NTRS)

    Pendley, Robert E; Robinson, Harold L

    1950-01-01

    An investigation of three NACA 1-series nose inlets, two of which were fitted with protruded central bodies, was conducted in the Langley 8-foot high-speed tunnel. An elliptical-nose body, which had a critical Mach number approximately equal to that of one of the nose inlets, was also tested. Tests were made near zero angle of attack for a Mach number range from 0.4 to 0.925 and for the supersonic Mach number of 1.2. The inlet-velocity-ratio range extended from zero to a maximum value of 1.34. Measurements included pressure distribution, external drag, and total-pressure loss of the internal flow near the inlet. Drag was not measured for the tests at the supersonic Mach number. Over the range of inlet-velocity ratio investigated, the calculated external pressure-drag coefficient at a Mach number of 1.2 was consecutively lower for the nose inlets of higher critical Mach number, and the pressure-drag coefficient of the longest nose inlet was in the range of pressure-drag coefficient for two solid noses of fineness ratio 2.4 and 6.0. For Mach numbers below the Mach number of the supercritical drag rise, extrapolation of the test data indicated that the external drag of the nose inlets was little affected by the addition of central bodies at or slightly below the minimum inlet-velocity ratio for unseparated central-body flow. The addition of central bodies to the nose inlets also led to no appreciable effects on either the Mach number of the supercritical drag rise, or, for inlet-velocity ratios high enough to avoid a pressure peak at the inlet lip, on the critical Mach number. The total-pressure recovery of the inlets tested, which were of a subsonic type, was sensibly unimpaired at the supersonic Mach number of 1.2 Low-speed measurements of the minimum inlet-velocity ratio for unseparated central-body flow appear to be applicable for Mach numbers extending to 1.2.

  5. LOW MACH NUMBER MODELING OF CONVECTION IN HELIUM SHELLS ON SUB-CHANDRASEKHAR WHITE DWARFS. I. METHODOLOGY

    SciTech Connect

    Zingale, M.; Orvedahl, R. J.; Nonaka, A.; Almgren, A. S.; Bell, J. B.; Malone, C. M.

    2013-02-10

    We assess the robustness of a low Mach number hydrodynamics algorithm for modeling helium shell convection on the surface of a white dwarf in the context of the sub-Chandrasekhar model for Type Ia supernovae. We use the low Mach number stellar hydrodynamics code, MAESTRO, to perform three-dimensional, spatially adaptive simulations of convection leading up to the point of the ignition of a burning front. We show that the low Mach number hydrodynamics model provides a robust description of the system.

  6. CURRENT - A Computer Code for Modeling Two-Dimensional, Chemically Reaccting, Low Mach Number Flows

    SciTech Connect

    Winters, W.S.; Evans, G.H.; Moen, C.D.

    1996-10-01

    This report documents CURRENT, a computer code for modeling two- dimensional, chemically reacting, low Mach number flows including the effects of surface chemistry. CURRENT is a finite volume code based on the SIMPLER algorithm. Additional convergence acceleration for low Peclet number flows is provided using improved boundary condition coupling and preconditioned gradient methods. Gas-phase and surface chemistry is modeled using the CHEMKIN software libraries. The CURRENT user-interface has been designed to be compatible with the Sandia-developed mesh generator and post processor ANTIPASTO and the post processor TECPLOT. This report describes the theory behind the code and also serves as a user`s manual.

  7. Aeropropulsive characteristics of isolated combined turbojet/ramjet nozzles at Mach numbers from 0 to 1.20

    NASA Technical Reports Server (NTRS)

    Carson, George T., Jr.; Lamb, Milton

    1988-01-01

    An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the aeropropulsive performance characteristics (the aerodynamic quantities affected by propulsion) of 13 isolated combined turbojet/ramjet nozzle configurations. These configurations simulated the variable-geometry features of two nozzle designs designated as the multiple-expansion ramp nozzle (MERN) and the composite contour nozzle (CCN). Test data were obtained at static conditions and at Mach numbers of 0.60, 0.90, and 1.20 with jet exhaust simulated by high-pressure air. The results showed that the CCN had the higher performance over the Mach number range than the MERN, as indicated by the difference of thrust minus drag divided by ideal thrust. Increasing the ramjet throat area for the MERN resulted in an increase in performance that increased with Mach number. For the CCN at Mach numbers less than 1.20, increasing the ramjet throat area resulted in a loss in performance.

  8. A tabulation of pipe length to diameter ratios as a function of Mach number and pressure ratios for compressible flow

    NASA Technical Reports Server (NTRS)

    Dixon, G. V.; Barringer, S. R.; Gray, C. E.; Leatherman, A. D.

    1975-01-01

    Computer programs and resulting tabulations are presented of pipeline length-to-diameter ratios as a function of Mach number and pressure ratios for compressible flow. The tabulations are applicable to air, nitrogen, oxygen, and hydrogen for compressible isothermal flow with friction and compressible adiabatic flow with friction. Also included are equations for the determination of weight flow. The tabulations presented cover a wider range of Mach numbers for choked, adiabatic flow than available from commonly used engineering literature. Additional information presented, but which is not available from this literature, is unchoked, adiabatic flow over a wide range of Mach numbers, and choked and unchoked, isothermal flow for a wide range of Mach numbers.

  9. Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Vol. 4: Large-radius leading edge

    NASA Technical Reports Server (NTRS)

    Chu, Julio; Luckring, James M.

    1996-01-01

    An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 120 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6) and 60 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

  10. Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Volume 2; Small-Radius Leading Edge

    NASA Technical Reports Server (NTRS)

    Chu, Julio; Luckring, James M.

    1996-01-01

    An experimental wind tunnel test of a 65 deg. delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 84 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6) and 60 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

  11. Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Vol. 3: Medium-radius leading edge

    NASA Technical Reports Server (NTRS)

    Chu, Julio; Luckring, James M.

    1996-01-01

    An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 120 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6), 60 x 10(exp 6), and 120 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

  12. Flow noise induced by small gaps in low-Mach-number turbulent boundary layers

    NASA Astrophysics Data System (ADS)

    Hao, Jin; Wang, Meng; Ji, Minsuk; Wang, Kan

    2013-11-01

    The flow-noise induced by small gaps underneath low-Mach-number turbulent boundary layers at Reθ = 4755 is studied using large-eddy simulation and Lighthill's theory. The gap leading-edge height is 13% of the boundary-layer thickness, and the gap width and trailing-edge height are varied to investigate their effects on surface-pressure fluctuations and sound generation. The maximum surface pressure fluctuations, which increase with gap width and trailing-edge height, occur at the trailing edge or near the reattachment point if there is separation from the trailing edge. The downstream recovery towards an equilibrium boundary layer is significantly faster for gap flows compared to step flows, and the recovery distance scales with the reattachment length for gaps with trailing-edge separation. The acoustic field is dominated by the forward-facing step in the gap and resembles forward-step sound for wide gaps and/or asymmetric gaps with trailing edge higher than leading edge. In these cases, the dominant acoustic source mechanisms are the impingement of the separated shear layer from the leading edge onto the trailing edge and the unsteady separation from the trailing edge, coupled with edge diffraction. For narrow and symmetric gaps, the destructive interference of sound from the leading and trailing edges causes a significant decline in low-frequency sound and thereby creates a broad spectral peak in the mid-frequency range. The effects of gap acoustic non-compactness and free-stream convection are investigated by comparing solutions based on a compact gap Green's function with those from a boundary-element calculation. They are found to be negligible at the typical hydroacoustc Mach number of 0.01, but become significant at Mach numbers as low as 0.1 and moderately high frequencies.

  13. A simulation study of multiple ion wave generation downstream of low Mach number quasiperpendicular shocks

    NASA Technical Reports Server (NTRS)

    Motschmann, Uwe; Raeder, Joachim

    1992-01-01

    The behavior of minor ions just downstream of a low Mach number quasi-perpendicular shock is investigated both theoretically and by computer simulations. Because all ions see the same cross shock electric field their deceleration depends on their charge to mass ratio, yielding different downstream velocities. It is shown that these differences in velocity can lead to coherent wave structures in the downstream region of quasi-perpendicular shocks with a narrow transition layer. These waves are shown to be multi ion hybrid waves in contrast to mirror waves and ion cyclotron waves. Under favorable conditions these waves should be observable both at interplanetary shocks and at planetary bowshocks.

  14. Bumblebee program, aerodynamic data. Part 2: Flow fields at Mach number 2.0. [supersonic missiles

    NASA Technical Reports Server (NTRS)

    Barnes, G. A.; Cronvich, L. L.

    1979-01-01

    Available flow field data which can be used in validating theoretical procedures for computing flow fields around supersonic missiles are presented. Tabulated test data are given which define the flow field around a conical-nosed cylindrical body in a crossflow plane corresponding to a likely tail location. The data were obtained at a Mach number of 2.0 for an angle of attack of 0 to 23 degrees. The data define the flow field for cases both with and without a forward wing present.

  15. A fluctuating surface pressure test technique utilizing Mach number sweeps at transonic speeds

    NASA Technical Reports Server (NTRS)

    Hanly, R. D.

    1974-01-01

    A multichannel on-line RMS data acquisition and reduction system has been developed using commercial RMS computing modules and a programmable calculator. Details of this system, which has the capability of acquiring 96 channels of RMS data and computing and printing desired parameters in near real-time, are presented. In addition, raw data can be recorded at a much higher rate for computation and printing later. Results are presented showing the benefits of this system in 'sweep' tests where one parameter such as Mach number or angle of attack is slowly varied with time.

  16. Profile of a low-Mach-number shock in two-fluid plasma theory

    NASA Astrophysics Data System (ADS)

    Gedalin, M.; Kushinsky, Y.; Balikhin, M.

    2015-08-01

    Magnetic profiles of low-Mach-number collisionless shocks in space plasmas are studied within the two-fluid plasma theory. Particular attention is given to the upstream magnetic oscillations generated at the ramp. By including weak resistive dissipation in the equations of motion for electrons and protons, the dependence of the upstream wave train features on the ratio of the dispersion length to the dissipative length is established quantitatively. The dependence of the oscillation amplitude and spatial damping scale on the shock normal angle θ is found.

  17. Laminar-turbulent transition calculations of heat transfer at hypersonic Mach numbers over sharp cones

    NASA Technical Reports Server (NTRS)

    Kaul, U. K.

    1988-01-01

    Computations of the hypersonic flow around sharp cones were carried out using the PNS code with attention given to the heat transfer predictions around the transition region. Results of calculations performed over 5, 8, and 10 deg half-angle sharp cones in the Mach number range of 7 to 10 are presented. It is noted that calculations of this type have become an integral part of the general design procedure for hypersonic vehicles such as the National Aerospace Plane and the Space Shuttle.

  18. A low Mach number preconditioned scheme for a two-phase liquid-gas compressible flow model

    NASA Astrophysics Data System (ADS)

    Pelanti, Marica

    2015-11-01

    The simulation of liquid-gas flows such as cavitating flows demands numerical methods efficient for a wide range of Mach number regimes, due to the large and rapid variation of the speed of sound in these two-phase flows. When classical upwind finite volume discretizations for compressible flow models are employed, suitable strategies are needed to overcome the well known difficulty of loss of accuracy encountered at low Mach number by these methods. In this work we present a novel finite volume wave propagation scheme with low Mach number preconditioning for the numerical approximation of a six-equation two-phase liquid-gas compressible flow model with stiff mechanical relaxation. A Turkel-type preconditioner is designed to correct the acoustic fields at low Mach number, by altering the numerical dissipation tensor of the scheme. We present numerical results for two-dimensional liquid-gas nozzle flow tests both for low Mach number regimes and for transonic regimes with shock formation, which show the effectiveness and accuracy of the proposed preconditioned method. In particular, in the low Mach number limit the order of pressure perturbations at the discrete level agrees with the theoretical results for the continuous two-phase flow model.

  19. A study of sonic boom overpressure trends with respect to weight, altitude, Mach number, and vehicle shaping

    NASA Technical Reports Server (NTRS)

    Needleman, Kathy E.; Mack, Robert J.

    1990-01-01

    This paper presents and discusses trends in nose shock overpressure generated by two conceptual Mach 2.0 configurations. One configuration was designed for high aerodynamic efficiency, while the other was designed to produce a low boom, shaped-overpressure signature. Aerodynamic lift, sonic boom minimization, and Mach-sliced/area-rule codes were used to analyze and compute the sonic boom characteristics of both configurations with respect to cruise Mach number, weight, and altitude. The influence of these parameters on the overpressure and the overpressure trends are discussed and conclusions are given.

  20. Entropy Splitting for High Order Numerical Simulation of Vortex Sound at Low Mach Numbers

    NASA Technical Reports Server (NTRS)

    Mueller, B.; Yee, H. C.; Mansour, Nagi (Technical Monitor)

    2001-01-01

    A method of minimizing numerical errors, and improving nonlinear stability and accuracy associated with low Mach number computational aeroacoustics (CAA) is proposed. The method consists of two levels. From the governing equation level, we condition the Euler equations in two steps. The first step is to split the inviscid flux derivatives into a conservative and a non-conservative portion that satisfies a so called generalized energy estimate. This involves the symmetrization of the Euler equations via a transformation of variables that are functions of the physical entropy. Owing to the large disparity of acoustic and stagnation quantities in low Mach number aeroacoustics, the second step is to reformulate the split Euler equations in perturbation form with the new unknowns as the small changes of the conservative variables with respect to their large stagnation values. From the numerical scheme level, a stable sixth-order central interior scheme with a third-order boundary schemes that satisfies the discrete analogue of the integration-by-parts procedure used in the continuous energy estimate (summation-by-parts property) is employed.

  1. Low Mach number analysis of idealized thermoacoustic engines with numerical solution.

    PubMed

    Hireche, Omar; Weisman, Catherine; Baltean-Carlès, Diana; Le Quéré, Patrick; Bauwens, Luc

    2010-12-01

    A model of an idealized thermoacoustic engine is formulated, coupling nonlinear flow and heat exchange in the heat exchangers and stack with a simple linear acoustic model of the resonator and load. Correct coupling results in an asymptotically consistent global model, in the small Mach number approximation. A well-resolved numerical solution is obtained for two-dimensional heat exchangers and stack. The model assumes that the heat exchangers and stack are shorter than the overall length by a factor of the order of a representative Mach number. The model is well-suited for simulation of the entire startup process, whereby as a result of some excitation, an initially specified temperature profile in the stack evolves toward a near-steady profile, eventually reaching stationary operation. A validation analysis is presented, together with results showing the early amplitude growth and approach of a stationary regime. Two types of initial excitation are used: Random noise and a small periodic wave. The set of assumptions made leads to a heat-exchanger section that acts as a source of volume but is transparent to pressure and to a local heat-exchanger model characterized by a dynamically incompressible flow to which a locally spatially uniform acoustic pressure fluctuation is superimposed. PMID:21218877

  2. The influence of incident shock Mach number on radial incident shock wave focusing

    NASA Astrophysics Data System (ADS)

    Chen, Xin; Tan, Sheng; He, Liming; Rong, Kang; Zhang, Qiang; Zhu, Xiaobin

    2016-04-01

    Experiments and numerical simulations were carried out to investigate radial incident shock focusing on a test section where the planar incident shock wave was divided into two identical ones. A conventional shock tube was used to generate the planar shock. Incident shock Mach number of 1.51, 1.84 and 2.18 were tested. CCD camera was used to obtain the schlieren photos of the flow field. Third-order, three step strong-stability-preserving (SSP) Runge-Kutta method, third-order weighed essential non-oscillation (WENO) scheme and adaptive mesh refinement (AMR) algorithm were adopted to simulate the complicated flow fields characterized by shock wave interaction. Good agreement between experimental and numerical results was observed. Complex shock wave configurations and interactions (such as shock reflection, shock-vortex interaction and shock focusing) were observed in both the experiments and numerical results. Some new features were observed and discussed. The differences of structure of flow field and the variation trends of pressure were compared and analyzed under the condition of different Mach numbers while shock wave focusing.

  3. MAESTRO: An Adaptive Low Mach Number Hydrodynamics Algorithm for Stellar Flows

    NASA Astrophysics Data System (ADS)

    Nonaka, Andrew; Almgren, A. S.; Bell, J. B.; Malone, C. M.; Zingale, M.

    2010-01-01

    Many astrophysical phenomena are highly subsonic, requiring specialized numerical methods suitable for long-time integration. We present MAESTRO, a low Mach number stellar hydrodynamics code that can be used to simulate long-time, low-speed flows that would be prohibitively expensive to model using traditional compressible codes. MAESTRO is based on an equation set that we have derived using low Mach number asymptotics; this equation set does not explicitly track acoustic waves and thus allows a significant increase in the time step. MAESTRO is suitable for two- and three-dimensional local atmospheric flows as well as three-dimensional full-star flows, and uses adaptive mesh refinement (AMR) to locally refine grids in regions of interest. Our initial scientific applications include the convective phase of Type Ia supernovae and Type I X-ray Bursts on neutron stars. The work at LBNL was supported by the SciDAC Program of the DOE Office of Advanced Scientific Computing Research under the DOE under contract No. DE-AC02-05CH11231. The work at Stony Brook was supported by the DOE/Office of Nuclear Physics, grant No. DE-FG02-06ER41448. We made use of the Jaguar via a DOE INCITE allocation at the OLCF at ORNL and Franklin at NERSC at LBNL.

  4. Revised tables of airspeed, altitude, and Mach number presented in the International system of units

    NASA Technical Reports Server (NTRS)

    Benner, M. S.; Sawyer, R. H.

    1973-01-01

    Because inception of a national program to implement the International System of Units (SI) appears to be inevitable and imminent, the tables of airspeed, altitude, and Mach number prepared by Livingston and Gracey to serve for airspeed meter and altimeter calibrations and for the conversion of flight measurements of these quantities to related parameters - Mach number, true airspeed, equivalent airspeed, etc. - have been revised to the SI. Tables of airspeed in knots are also included because of the significance of this quantity in navigation. In addition, the data in the altitude tables have been revised to the U.S. Standard Atmosphere of 1962. The latter data reflect increased knowledge of the higher atmosphere and more precise determination of basic quantities, including the redefinition of the absolute thermodynamic temperature scale by the Tenth General Conference on Weights and Measures in 1954. The U.S. Standard Atmosphere, 1962, corresponds to the International Civil Aviation Organization (ICAO) Standard Atmosphere up to 20 kilometers (geopotential altitude). A table of conversion factors for various pressure units is presented in SI Units.

  5. On the high Mach number shock structure singularity caused by overreach of Maxwellian molecules

    SciTech Connect

    Myong, R. S.

    2014-05-15

    The high Mach number shock structure singularity arising in moment equations of the Boltzmann equation was investigated. The source of the singularity is shown to be the unbalanced treatment between two high order kinematic and dissipation terms caused by the overreach of Maxwellian molecule assumption. In compressive gaseous flow, the high order stress-strain coupling term of quadratic nature will grow far faster than the strain term, resulting in an imbalance with the linear dissipation term and eventually a blow-up singularity in high thermal nonequilibrium. On the other hand, the singularity arising from unbalanced treatment does not occur in the case of velocity shear and expansion flows, since the high order effects are cancelled under the constraint of the free-molecular asymptotic behavior. As an alternative method to achieve the balanced treatment, Eu's generalized hydrodynamics, consistent with the second law of thermodynamics, was revisited. After introducing the canonical distribution function in exponential form and applying the cumulant expansion to the explicit calculation of the dissipation term, a natural platform suitable for the balanced treatment was derived. The resulting constitutive equation with the nonlinear factor was then shown to be well-posed for all regimes, effectively removing the high Mach number shock structure singularity.

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

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

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

  7. Experimental aerodynamic characteristics at Mach numbers from 0.60 to 2.70 of two supersonic cruise fighter configurations

    NASA Technical Reports Server (NTRS)

    Dollyhigh, S. M.

    1979-01-01

    Two 0.085-scale full span wind-tunnel models of a Mach 1.60 design supercruiser configuration were tested at Mach numbers from 0.60 to 2.70. One model incorporated a varying dihedral (swept-up) wing to obtain the desired lateral-directional characteristics; the other incorporated more conventional twin vertical tails. The data from the wind-tunnel tests are presented without analysis.

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

    NASA Technical Reports Server (NTRS)

    Wittliff, C. E.

    1975-01-01

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

  9. The least-squares finite element method for low-mach-number compressible viscous flows

    NASA Technical Reports Server (NTRS)

    Yu, Sheng-Tao

    1994-01-01

    The present paper reports the development of the Least-Squares Finite Element Method (LSFEM) for simulating compressible viscous flows at low Mach numbers in which the incompressible flows pose as an extreme. Conventional approach requires special treatments for low-speed flows calculations: finite difference and finite volume methods are based on the use of the staggered grid or the preconditioning technique; and, finite element methods rely on the mixed method and the operator-splitting method. In this paper, however, we show that such difficulty does not exist for the LSFEM and no special treatment is needed. The LSFEM always leads to a symmetric, positive-definite matrix through which the compressible flow equations can be effectively solved. Two numerical examples are included to demonstrate the method: first, driven cavity flows at various Reynolds numbers; and, buoyancy-driven flows with significant density variation. Both examples are calculated by using full compressible flow equations.

  10. On the precise implications of acoustic analogies for aerodynamic noise at low Mach numbers

    NASA Astrophysics Data System (ADS)

    Spalart, Philippe R.

    2013-05-01

    We seek a clear statement of the scaling which may be expected with rigour for transportation or other noise at low Mach numbers M, based on Lighthill's and Curle's theories of 1952 and 1955. In the presence of compact solid bodies, the leading term in the acoustic intensity is of order M6. Contrary to the belief held since that time that it is of order M8, the contribution of quadrupoles, in the presence of dipoles, is of order only M7. Retarded-time-difference effects are also of order M7. Curle's widely used approximation based on unsteady forces neglects both effects. Its order of accuracy is thus lower than was thought, and the common estimates of the value of M below which it applies appear precarious. The M6 leading term is modified by powers up to the fourth of (1-Mr), where Mr is the relative Mach number between source and observer; at speeds of interest the effect is several dB. However, this is only one of the corrections of order M7, which makes its value debatable. The same applies to the difference between emission distance and reception distance. The scaling with M6 is theoretically correct to leading order, but this prediction may be so convincing, like the M8 scaling for jet noise, that some authors rush to confirm it when their measurements are in conflict with it. We survey experimental studies of landing-gear noise, and argue that the observed power of M is often well below 6. We also object to comparisons across Mach numbers at fixed frequency; they should be made at fixed Strouhal number St instead. Finally, the compact-source argument does not only require M≪1; it requires MSt≪1. This is more restrictive if the relevant St is well above 1, a situation which can be caused by interference with a boundary or by wake impingement, among other effects. The best length scales to define St for this purpose are discussed.

  11. Effects of Mach Number, Leading-Edge Bluntness, and Sweep on Boundary-Layer Transition on a Flat Plate

    NASA Technical Reports Server (NTRS)

    Jillie, Don W.; Hopkins, Edward J.

    1961-01-01

    The effects of leading-edge bluntness and sweep on boundary-layer transition on flat plate models were investigated at Mach numbers of 2.00, 2.50, 3.00, and 4.00. The effect of sweep on transition was also determined on a flat plate model equipped with an elliptical nose at a Mach number of 0.27. Models used for the supersonic investigation had leading-edge radii varying from 0.0005 to 0.040 inch. The free-stream unit Reynolds number was held constant at 15 million per foot for the supersonic tests and the angle of attack was 0 deg. Surface flow conditions were determined by visual observation and recorded photographically. The sublimation technique was used to indicate transition, and the fluorescent-oil technique was used to indicate flow separation. Measured Mach number and sweep effects on transition are compared with those predicted from shock-loss considerations as described in NACA Rep. 1312. For the models with the blunter leading edges, the transition Reynolds number (based on free-stream flow conditions) was approximately doubled by an increase in Mach number from 2.50 to 4.00; and nearly the same result was predicted from shock-loss considerations. At all super- sonic Mach numbers, increases in sweep reduced the transition Reynolds number and the amount of reduction increased with increases in bluntness. The shock-loss method considerably underestimated- the sweep effects, possibly because of the existence of crossflow instability associated with swept wings. At a Mach number of 0.27, no reduction in the transition Reynolds number with sweep was measured (as would be expected with no shock loss) until the sweep angle was attained where crossflow instability appeared.

  12. Parametric investigation of single-expansion-ramp nozzles at Mach numbers from 0.60 to 1.20

    NASA Technical Reports Server (NTRS)

    Capone, Francis J.; Re, Richard J.; Bare, E. Ann

    1992-01-01

    An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of varying six nozzle geometric parameters on the internal and aeropropulsive performance characteristics of single-expansion-ramp nozzles. This investigation was conducted at Mach numbers from 0.60 to 1.20, nozzle pressure ratios from 1.5 to 12, and angles of attack of 0 deg +/- 6 deg. Maximum aeropropulsive performance at a particular Mach number was highly dependent on the operating nozzle pressure ratio. For example, as the nozzle upper ramp length or angle increased, some nozzles had higher performance at a Mach number of 0.90 because of the nozzle design pressure was the same as the operating pressure ratio. Thus, selection of the various nozzle geometric parameters should be based on the mission requirements of the aircraft. A combination of large upper ramp and large lower flap boattail angles produced greater nozzle drag coefficients at Mach number greater than 0.80, primarily from shock-induced separation on the lower flap of the nozzle. A static conditions, the convergent nozzle had high and nearly constant values of resultant thrust ratio over the entire range of nozzle pressure ratios tested. However, these nozzles had much lower aeropropulsive performance than the convergent-divergent nozzle at Mach number greater than 0.60.

  13. Flight tests of Viking parachute system in three Mach number regimes. 2: Parachute test results

    NASA Technical Reports Server (NTRS)

    Bendura, R. J.; Lundstrom, R. R.; Renfroe, P. G.; Lecroy, S. R.

    1974-01-01

    Tests of the Viking 16.15-meter nominal-diameter disk-gap-band parachute were conducted at Mach number and dynamic pressure conditions which bracketed the range postulated for the Viking '75 mission to Mars. Parachutes were deployed at supersonic, transonic, and subsonic speeds behind a simulated Viking entry capsule. All parachutes successfully deployed, inflated, and exhibited sufficient drag and stability for mission requirements. Basic parachute data including loads, drag coefficients, pull-off angles, and canopy area ratios are presented. Trajectory reconstruction and onboard camera data methods were combined to yield continuous histories of both parachute and test-vehicle angular motions which are presented for the period from parachute deployment through steady inflation.

  14. Tests of Full-Scale Helicopter Rotors at High Advancing Tip Mach Numbers and Advance Ratios

    NASA Technical Reports Server (NTRS)

    Biggers, James C.; McCloud, John L., III; Stroub, Robert H.

    2015-01-01

    As a continuation of the studies of reference 1, three full-scale helicopter rotors have been tested in the Ames Research Center 40- by SO-foot wind tunnel. All three of them were two-bladed, teetering rotors. One of the rotors incorporated the NACA 0012 airfoil section over the entire length of the blade. This rotor was tested at advance ratios up to 1.05. Both of the other rotors were tapered in thickness and incorporated leading-edge camber over the outer 20 percent of the blade radius. The larger of these rotors was tested at advancing tip Mach numbers up to 1.02. Data were obtained for a wide range of lift and propulsive force, and are presented without discussion.

  15. A high-energy-density, high-Mach number single jet experiment

    SciTech Connect

    Hansen, J. F.; Dittrich, T. R.; Elliott, J. B.; Glendinning, S. G.; Cotrell, D. L.

    2011-08-15

    A high-energy-density, x-ray-driven, high-Mach number (M{>=} 17) single jet experiment shows constant propagation speeds of the jet and its bowshock into the late time regime. The jet assumes a characteristic mushroom shape with a stalk and a head. The width of the head and the bowshock also grow linearly in time. The width of the stalk decreases exponentially toward an asymptotic value. In late time images, the stalk kinks and develops a filamentary nature, which is similar to experiments with applied magnetic fields. Numerical simulations match the experiment reasonably well, but ''exterior'' details of the laser target must be included to obtain a match at late times.

  16. Sonic-box method employing local Mach number for oscillating wings with thickness

    NASA Technical Reports Server (NTRS)

    Ruo, S. Y.

    1978-01-01

    A computer program was developed to account approximately for the effects of finite wing thickness in the transonic potential flow over an oscillating wing of finite span. The program is based on the original sonic-box program for planar wing which was previously extended to include the effects of the swept trailing edge and the thickness of the wing. Account for the nonuniform flow caused by finite thickness is made by application of the local linearization concept. The thickness effect, expressed in terms of the local Mach number, is included in the basic solution to replace the coordinate transformation method used in the earlier work. Calculations were made for a delta wing and a rectangular wing performing plunge and pitch oscillations, and the results were compared with those obtained from other methods. An input quide and a complete listing of the computer code are presented.

  17. Preconditioning for Numerical Simulation of Low Mach Number Three-Dimensional Viscous Turbomachinery Flows

    NASA Technical Reports Server (NTRS)

    Tweedt, Daniel L.; Chima, Rodrick V.; Turkel, Eli

    1997-01-01

    A preconditioning scheme has been implemented into a three-dimensional viscous computational fluid dynamics code for turbomachine blade rows. The preconditioning allows the code, originally developed for simulating compressible flow fields, to be applied to nearly-incompressible, low Mach number flows. A brief description is given of the compressible Navier-Stokes equations for a rotating coordinate system, along with the preconditioning method employed. Details about the conservative formulation of artificial dissipation are provided, and different artificial dissipation schemes are discussed and compared. The preconditioned code was applied to a well-documented case involving the NASA large low-speed centrifugal compressor for which detailed experimental data are available for comparison. Performance and flow field data are compared for the near-design operating point of the compressor, with generally good agreement between computation and experiment. Further, significant differences between computational results for the different numerical implementations, revealing different levels of solution accuracy, are discussed.

  18. Ballistic range experiments on the superboom generated at increasing flight Mach numbers

    NASA Technical Reports Server (NTRS)

    Sanai, M.; Toong, T.-Y.; Pierce, A. D.

    1976-01-01

    Ballistic range experiments for the study of the propagation of converging shocks are described and the similarity between the observed phenomenon and that expected for superbooms created by accelerating supersonic aircraft is discussed. For weak shocks (shock Mach numbers of about 1.03), a structure resembling that of a folded shock predicted by geometrical acoustics theory is observed while for stronger shocks, a concave front with enhanced overpressure is recorded. Other results are in general accord with the basic concepts of shock propagation and in conjunction with some theoretical scaling laws indicate that the peak magnification of sonic booms due to aircraft flight acceleration in the real atmosphere should be in the range of 6 to 13.

  19. Cosmic Mach Number: a sensitive probe for the growth of structure

    NASA Astrophysics Data System (ADS)

    Ma, Yin-Zhe; Ostriker, Jeremiah P.; Zhao, Gong-Bo

    2012-06-01

    We investigate the potential power of the Cosmic Mach Number (CMN), which is the ratio between the mean velocity and the velocity dispersion of galaxies as a function of cosmic scales, to constrain cosmologies. We first measure the CMN from 4 catalogs of galaxy peculiar velocity surveys at low redshift (zin[0.002,0.03]), and use them to contrast cosmological models. Overall, current data is consistent with the WMAP7 ΛCDM model. We find that the CMN is highly sensitive to the growth of structure on scales kin[0.01,0.1] h/Mpc in Fourier space. Therefore, modified gravity models, and models with massive neutrinos, in which the structure growth generically deviates from that of the ΛCDM model in a scale-dependent way, can be well differentiated from the ΛCDM model by using future CMN data.

  20. On the proper Mach number and ratio of specific heats for modeling the Venus bow shock

    NASA Technical Reports Server (NTRS)

    Tatrallyay, M.; Russell, C. T.; Luhmann, J. G.; Barnes, A.; Mihalov, J. D.

    1984-01-01

    Observational data from the Pioneer Venus Orbiter are used to investigate the physical characteristics of the Venus bow shock, and to explore some general issues in the numerical simulation of collisionless shocks. It is found that since equations from gas-dynamic (GD) models of the Venus shock cannot in general replace MHD equations, it is not immediately obvious what the optimum way is to describe the desired MHD situation with a GD code. Test case analysis shows that for quasi-perpendicular shocks it is safest to use the magnetospheric Mach number as an input to the GD code. It is also shown that when comparing GD predicted temperatures with MHD predicted temperatures total energy should be compared since the magnetic energy density provides a significant fraction of the internal energy of the MHD fluid for typical solar wind parameters. Some conclusions are also offered on the properties of the terrestrial shock.

  1. Stability Analysis of a mortar cover ejected at various Mach numbers and angles of attack

    NASA Astrophysics Data System (ADS)

    Schwab, Jane; Carnasciali, Maria-Isabel; Andrejczyk, Joe; Kandis, Mike

    2011-11-01

    This study utilized CFD software to predict the aerodynamic coefficient of a wedge-shaped mortar cover which is ejected from a spacecraft upon deployment of its Parachute Recovery System (PRS). Concern over recontact or collision between the mortar cover and spacecraft served as the impetus for this study in which drag and moment coefficients were determined at Mach numbers from 0.3 to 1.6 at 30-degree increments. These CFD predictions were then used as inputs to a two-dimensional, multi-body, three-DoF trajectory model to calculate the relative motion of the mortar cover and spacecraft. Based upon those simulations, the study concluded a minimal/zero risk of collision with either the spacecraft or PRS. Sponsored by Pioneer Aerospace.

  2. Some Experimental Studies of Panel Flutter at Mach Number 1.3

    NASA Technical Reports Server (NTRS)

    Sylvester, Maurice A; Baker, John E

    1957-01-01

    Experimental studies of panel flutter using thin metal plates were conducted at a Mach number of 1.3 to verify its existence and to study the effects of some structural parameters on the flutter characteristics. The effects of tensile forces and buckling were studied on panels clamped front and rear, in addition to initially buckled panels clamped on all four edges. Panel flutter was obtained under controlled laboratory conditions and it was found that tensile forces, shortening the panels, and increasing the bending stiffness were effective means for eliminating flutter. Buckled panels were more susceptible to flutter than unbuckled panels. No apparent systematic trends in the flutter modes or frequencies could be observed.

  3. The Mach number of the cosmic flow - A critical test for current theories

    NASA Technical Reports Server (NTRS)

    Ostriker, Jeremiah P.; Suto, Yusushi

    1990-01-01

    A new cosmological, self-contained test using the ratio of mean velocity and the velocity dispersion in the mean flow frame of a group of test objects is presented. To allow comparison with linear theory, the velocity field must first be smoothed on a suitable scale. In the context of linear perturbation theory, the Mach number M(R) which measures the ratio of power on scales larger than to scales smaller than the patch size R, is independent of the perturbation amplitude and also of bias. An apparent inconsistency is found for standard values of power-law index n = 1 and cosmological density parameter Omega = 1, when comparing values of M(R) predicted by popular models with tentative available observations. Nonstandard models based on adiabatic perturbations with either negative n or small Omega value also fail, due to creation of unacceptably large microwave background fluctuations.

  4. Aerodynamic Characteristics of a Circular Cylinder at Mach Number 6.86 and Angles of Attack up to 90 Degrees

    NASA Technical Reports Server (NTRS)

    Penland, Jim A

    1957-01-01

    Pressure-distribution and force tests of a circular cylinder have been made in the Langley 11-inch hypersonic tunnel at a Mach number of 6.88, a Reynolds number of 129,000, and angles of attack up to 90 degrees. The results are compared with the hypersonic approximation of Grimminger, Williams, and Young and a simple modification of the Newtonian flow theory. An evaluation of the crossflow theory is made through comparison of present results with available crossflow Mach number drag coefficients.

  5. Analysis of Turbulent Flow and Heat Transfer on a Flat Plate at High Mach Numbers with Variable Fluid Properties

    NASA Technical Reports Server (NTRS)

    Deissler, R. G.; Loeffler, A. L., Jr.

    1959-01-01

    A previous analysis of turbulent heat transfer and flow with variable fluid properties in smooth passages is extended to flow over a flat plate at high Mach numbers, and the results are compared with experimental data. Velocity and temperature distributions are calculated for a boundary layer with appreciative effects of frictional heating and external heat transfer. Viscosity and thermal conductivity are assumed to vary as a power or the temperature, while Prandtl number and specific heat are taken as constant. Skin-friction and heat-transfer coefficients are calculated and compared with the incompressible values. The rate of boundary-layer growth is obtained for various Mach numbers.

  6. A fan pressure ratio correlation in terms of Mach number and Reynolds number for the Langley 0.3 meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Lawing, P. L.; Adcock, J. B.; Ladson, C. L.

    1980-01-01

    Calibration data for the two dimensional test section of the Langley 0.3 Meter Transonic Cryogenic Tunnel were used to develop a Mach number-Reynolds number correlation for the fan pressure ratio in terms of test section conditions. Well established engineering relationships combined to form an equation which is functionally analogous to the correlation. A geometric loss coefficient which is independent of Reynolds number or Mach number was determined. Present and anticipated uses of this concept include improvement of tunnel control schemes, comparison of efficiencies for operationally similar wind tunnels, prediction of tunnel test conditions and associated energy usage, and determination of Reynolds number scaling laws for similar fluid flow systems.

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

    NASA Technical Reports Server (NTRS)

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

    2007-01-01

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

  8. Variation in Heat Transfer During Transient Heating of a Hemisphere at a Mach Number of 2

    NASA Technical Reports Server (NTRS)

    English, Roland D.; Carter, Howard S.

    1960-01-01

    Convective heat-transfer tests were made on a 5-inch-diameter hemisphere to determine the variation of Stanton number with the ratio of wall temperature to total temperature. The tests were made at a nominal Mach number of 2 for stagnation temperatures of 760 deg R, 1,030 deg R, and 1,380 deg R. The model was constructed so that radiation effects and also streamwise conduction effects within the model skin were minimized. The results of the tests verified that these effects were small. Tests which were made with different masses of air inside the model to check for conduction effects to the internal air cavity showed these effects to be negligible. For laminar flow on the hemisphere, the Stanton number remained essentially constant as the ratio of wall temperature to total temperature increased. However, for fully established turbulent flow, the Stanton number at some stations decreased on the order of 50 percent as the ratio of wall temperature to total temperature increased. A theory which agreed fairly well with the trend of this decrease is shown for comparison.

  9. Analysis of gas turbine engines using water and oxygen injection to achieve high Mach numbers and high thrust

    NASA Technical Reports Server (NTRS)

    Henneberry, Hugh M.; Snyder, Christopher A.

    1993-01-01

    An analysis of gas turbine engines using water and oxygen injection to enhance performance by increasing Mach number capability and by increasing thrust is described. The liquids are injected, either separately or together, into the subsonic diffuser ahead of the engine compressor. A turbojet engine and a mixed-flow turbofan engine (MFTF) are examined, and in pursuit of maximum thrust, both engines are fitted with afterburners. The results indicate that water injection alone can extend the performance envelope of both engine types by one and one-half Mach numbers at which point water-air ratios reach 17 or 18 percent and liquid specific impulse is reduced to some 390 to 470 seconds, a level about equal to the impulse of a high energy rocket engine. The envelope can be further extended, but only with increasing sacrifices in liquid specific impulse. Oxygen-airflow ratios as high as 15 percent were investigated for increasing thrust. Using 15 percent oxygen in combination with water injection at high supersonic Mach numbers resulted in thrust augmentation as high as 76 percent without any significant decrease in liquid specific impulse. The stoichiometric afterburner exit temperature increased with increasing oxygen flow, reaching 4822 deg R in the turbojet engine at a Mach number of 3.5. At the transonic Mach number of 0.95 where no water injection is needed, an oxygen-air ratio of 15 percent increased thrust by some 55 percent in both engines, along with a decrease in liquid specific impulse of 62 percent. Afterburner temperature was approximately 4700 deg R at this high thrust condition. Water and/or oxygen injection are simple and straightforward strategies to improve engine performance and they will add little to engine weight. However, if large Mach number and thrust increases are required, liquid flows become significant, so that operation at these conditions will necessarily be of short duration.

  10. OPACITY BROADENING OF {sup 13}CO LINEWIDTHS AND ITS EFFECT ON THE VARIANCE-SONIC MACH NUMBER RELATION

    SciTech Connect

    Correia, C.; De Medeiros, J. R.; Burkhart, B.; Lazarian, A.; Ossenkopf, V.; Stutzki, J.; Kainulainen, J.; Kowal, G.

    2014-04-10

    We study how the estimation of the sonic Mach number (M{sub s} ) from {sup 13}CO linewidths relates to the actual three-dimensional sonic Mach number. For this purpose we analyze MHD simulations that include post-processing to take radiative transfer effects into account. As expected, we find very good agreement between the linewidth estimated sonic Mach number and the actual sonic Mach number of the simulations for optically thin tracers. However, we find that opacity broadening causes M{sub s} to be overestimated by a factor of ≈1.16-1.3 when calculated from optically thick {sup 13}CO lines. We also find that there is a dependence on the magnetic field: super-Alfvénic turbulence shows increased line broadening compared with sub-Alfvénic turbulence for all values of optical depth for supersonic turbulence. Our results have implications for the observationally derived sonic Mach number-density standard deviation (σ{sub ρ/(ρ)}) relationship, σ{sub ρ/〈ρ〉}{sup 2}=b{sup 2}M{sub s}{sup 2}, and the related column density standard deviation (σ {sub N/(N)}) sonic Mach number relationship. In particular, we find that the parameter b, as an indicator of solenoidal versus compressive driving, will be underestimated as a result of opacity broadening. We compare the σ {sub N/(N)}-M{sub s} relation derived from synthetic dust extinction maps and {sup 13}CO linewidths with recent observational studies and find that solenoidally driven MHD turbulence simulations have values of σ {sub N/(N)}which are lower than real molecular clouds. This may be due to the influence of self-gravity which should be included in simulations of molecular cloud dynamics.

  11. Application of Pressure Sensitive Paint to Confined Flow at Mach Number 2.5

    NASA Technical Reports Server (NTRS)

    Lepicovsky, J.; Bencic, T. J.; Bruckner, R. J.

    1998-01-01

    Pressure sensitive paint (PSP) is a novel technology that is being used frequently in external aerodynamics. For internal flows in narrow channels, and applications at elevated nonuniform temperatures, however, there are still unresolved problems that complicate the procedures for calibrating PSP signals. To address some of these problems, investigations were carried out in a narrow channel with supersonic flows of Mach 2.5. The first set of tests focused on the distribution of the wall pressure in the diverging section of the test channel downstream of the nozzle throat. The second set dealt with the distribution of wall static pressure due to the shock/wall interaction caused by a 25 deg. wedge in the constant Mach number part of the test section. In addition, the total temperature of the flow was varied to assess the effects of temperature on the PSP signal. Finally, contamination of the pressure field data, caused by internal reflection of the PSP signal in a narrow channel, was demonstrated. The local wall pressures were measured with static taps, and the wall pressure distributions were acquired by using PSP. The PSP results gave excellent qualitative impressions of the pressure field investigated. However, the quantitative results, specifically the accuracy of the PSP data in narrow channels, show that improvements need to be made in the calibration procedures, particularly for heated flows. In the cases investigated, the experimental error had a standard deviation of +/- 8.0% for the unheated flow, and +/- 16.0% for the heated flow, at an average pressure of 11 kpa.

  12. Performance Characteristics of Flush and Shielded Auxiliary Exits at Mach Numbers of 1.5 to 2.0

    NASA Technical Reports Server (NTRS)

    Abdalla, Kaleel L.

    1959-01-01

    The performance characteristics of several flush and shielded auxiliary exits were investigated at Mach numbers of 1.5 to 2.0, and jet pressure ratios from jet off to 10. The results indicate that the shielded configurations produced better overall performance than the corresponding flush exits over the Mach-number and pressure-ratio ranges investigated. Furthermore, the full-length shielded exit was highest in performance of all the configurations. The flat-exit nozzle block provided considerably improved performance compared with the curved-exit nozzle block.

  13. Sagdeev potential approach for large amplitude compressional Alfvenic double layers in viscous plasmas

    SciTech Connect

    Panwar, Anuraj; Rizvi, H.; Ryu, C. M.

    2013-11-15

    Sagdeev’s technique is used to study the large amplitude compressional Alfvenic double layers in a magnetohydrodynamic plasma taking into account the small plasma β and small values of kinematic viscosity. Dispersive effect raised by non-ideal electron inertia currents perpendicular to the ambient magnetic field. The range of allowed values of the soliton speed, M (Mach number), plasma β (ratio of the plasma thermal pressure to the pressure in the confining magnetic field), and viscosity coefficient, wherein double layer may exist, are determined. In the absence of collisions, viscous dissipation modifies the Sagdeev potential and results in large amplitude compressional Alfvenic double layers. The depth of Sagdeev potential increases with the increasing Mach number and plasma β, however, decreases with the increasing viscosity. The double layer structure increases with the increasing plasma β, but decreases with increasing viscous dissipation μ(tilde sign)

  14. Galileo probe parachute test program: Wake properties of the Galileo probe at Mach numbers from 0.25 to 0.95

    NASA Astrophysics Data System (ADS)

    Canning, Thomas N.; Edwards, Thomas M.

    1988-04-01

    The results of surveys of the near and far wake of the Galileo Probe are presented for Mach numbers from 0.25 tp 0.95. The trends in the data resulting from changes in Mach number, radial and axial distance, angle of attack, and a small change in model shape are shown in crossplots based on the data. A rationale for selecting an operating volume suitable for parachute inflation based on low Mach number flight results is outlined.

  15. Galileo probe parachute test program: Wake properties of the Galileo probe at Mach numbers from 0.25 to 0.95

    NASA Technical Reports Server (NTRS)

    Canning, Thomas N.; Edwards, Thomas M.

    1988-01-01

    The results of surveys of the near and far wake of the Galileo Probe are presented for Mach numbers from 0.25 tp 0.95. The trends in the data resulting from changes in Mach number, radial and axial distance, angle of attack, and a small change in model shape are shown in crossplots based on the data. A rationale for selecting an operating volume suitable for parachute inflation based on low Mach number flight results is outlined.

  16. Heat transfer predictions for two turbine nozzle geometries at high Reynolds and Mach numbers

    NASA Technical Reports Server (NTRS)

    Boyle, R. J.; Jackson, R.

    1995-01-01

    Predictions of turbine vane and endwall heat transfer and pressure distributions are compared with experimental measurements for two vane geometries. The differences in geometries were due to differences in the hub profile, and both geometries were derived from the design of a high rim speed turbine (HRST). The experiments were conducted in the Isentropic Light Piston Facility (ILPF) at Pyestock at a Reynolds number of 5.3 x 10(exp 6), a Mach number of 1.2, and a wall-to-gas temperature ratio of 0.66. Predictions are given for two different steady-state three-dimensional Navier-Stokes computational analyses. C-type meshes were used, and algebraic models were employed to calculate the turbulent eddy viscosity. The effects of different turbulence modeling assumptions on the predicted results are examined. Comparisons are also given between predicted and measured total pressure distributions behind the vane. The combination of realistic engine geometries and flow conditions proved to be quite demanding in terms of the convergence of the CFD solutions. An appropriate method of grid generation, which resulted in consistently converged CFD solutions, was identified.

  17. Revisiting Turbulence Model Validation for High-Mach Number Axisymmetric Compression Corner Flows

    NASA Technical Reports Server (NTRS)

    Georgiadis, Nicholas J.; Rumsey, Christopher L.; Huang, George P.

    2015-01-01

    Two axisymmetric shock-wave/boundary-layer interaction (SWBLI) cases are used to benchmark one- and two-equation Reynolds-averaged Navier-Stokes (RANS) turbulence models. This validation exercise was executed in the philosophy of the NASA Turbulence Modeling Resource and the AIAA Turbulence Model Benchmarking Working Group. Both SWBLI cases are from the experiments of Kussoy and Horstman for axisymmetric compression corner geometries with SWBLI inducing flares of 20 and 30 degrees, respectively. The freestream Mach number was approximately 7. The RANS closures examined are the Spalart-Allmaras one-equation model and the Menter family of kappa - omega two equation models including the Baseline and Shear Stress Transport formulations. The Wind-US and CFL3D RANS solvers are employed to simulate the SWBLI cases. Comparisons of RANS solutions to experimental data are made for a boundary layer survey plane just upstream of the SWBLI region. In the SWBLI region, comparisons of surface pressure and heat transfer are made. The effects of inflow modeling strategy, grid resolution, grid orthogonality, turbulent Prandtl number, and code-to-code variations are also addressed.

  18. An experimental investigation of propfan installations on an upswept supercritical wing at transonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Bartlett, G. R.

    1985-01-01

    An investigation has been conducted in the Langley 16 Foot Transonic Tunnel to determine propfan installation and slipstream interference effects on an unswept supercritical wing. This data can be used for verification of existing and developing theoretical codes as well as giving an understanding of the flow interactions associated with propeller/nacelle/wing integration. The investigation was conducted over a Mach number range of 0.5 to 0.8 and at angles of attack from 0 deg to 3 deg. The propeller was powered by an air turbine simulator and the exhaust from the air turbine was used to simulate the exhaust from the propfan nacelle. Reynolds number based on wing chord varied from 3 to 4 million. Results indicate that the propfan causes an increase in the wing lift coefficient. It was found that most of the propeller induced swirl is recovered by the wing. The propeller slipstream also causes a large favorable leading edge suction peak on the upwash side and a smaller unfavorable decrease on the downwash side.

  19. An evaluation of three two-dimensional computational fluid dynamics codes including low Reynolds numbers and transonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Hicks, Raymond M.; Cliff, Susan E.

    1991-01-01

    Full-potential, Euler, and Navier-Stokes computational fluid dynamics (CFD) codes were evaluated for use in analyzing the flow field about airfoils sections operating at Mach numbers from 0.20 to 0.60 and Reynolds numbers from 500,000 to 2,000,000. The potential code (LBAUER) includes weakly coupled integral boundary layer equations for laminar and turbulent flow with simple transition and separation models. The Navier-Stokes code (ARC2D) uses the thin-layer formulation of the Reynolds-averaged equations with an algebraic turbulence model. The Euler code (ISES) includes strongly coupled integral boundary layer equations and advanced transition and separation calculations with the capability to model laminar separation bubbles and limited zones of turbulent separation. The best experiment/CFD correlation was obtained with the Euler code because its boundary layer equations model the physics of the flow better than the other two codes. An unusual reversal of boundary layer separation with increasing angle of attack, following initial shock formation on the upper surface of the airfoil, was found in the experiment data. This phenomenon was not predicted by the CFD codes evaluated.

  20. Aerodynamic Characteristics of a Revised Target Drone Vehicle at Mach Numbers from 1.60 to 2.86

    NASA Technical Reports Server (NTRS)

    Blair, A. B., Jr.; Babb, C. Donald

    1968-01-01

    An investigation has been conducted in the Langley Unitary Plan wind tunnel to determine the aerodynamic characteristics of a revised target drone vehicle through a Mach number range from 1.60 to 2.86. The vehicle had canard surfaces and a swept clipped-delta wing with twin tip-mounted vertical tails.

  1. Wind-Tunnel Results of Advanced High-Speed Propellers at Takeoff, Climb, and Landing Mach Numbers

    NASA Technical Reports Server (NTRS)

    Stefko, George L.; Jeracki, Robert J.

    1985-01-01

    Low-speed wind-tunnel performance tests of two advanced propellers have been completed at the NASA Lewis Research Center as part of the NASA Advanced Turboprop Program. The 62.2 cm (24.5 in.) diameter adjustable-pitch models were tested at Mach numbers typical of takeoff, initial climbout, and landing speeds (i.e., from Mach 0.10 to 0.34) at zero angle of attack in the NASA Lewis 10 by 10 Foot Supersonic Wind Tunnel. Both models had eight blades and a cruise-design-point operating condition of Mach 0.80, and 10.668 km (35,000 ft) I.S.A. altitude, a 243.8 m/s (800 ft/sec) tip speed, and a high power loading of 301 kW/sq m (37.5 shp/sq ft). Each model had its own integrally designed area-ruled spinner, but used the same specially contoured nacelle. These features reduced blade-section Mach numbers and relieved blade-root choking at the cruise condition. No adverse or unusual low-speed operating conditions were found during the test with either the straight blade SR-2 or the 45 deg swept SR-3 propeller. Typical efficiencies of the straight and 45 deg swept propellers were 50.2 and 54.9 percent, respectively, at a takeoff condition of Mach 0.20 and 53.7 and 59.1 percent, respectively, at a climb condition of Mach 0.34.

  2. Arbitrary amplitude kinetic Alfven solitary waves in two temperature electron superthermal plasma

    NASA Astrophysics Data System (ADS)

    Singh, Manpreet; Singh Saini, Nareshpal; Ghai, Yashika

    2016-07-01

    Through various satellite missions it is observed that superthermal velocity distribution for particles is more appropriate for describing space and astrophysical plasmas. So it is appropriate to use superthermal distribution, which in the limiting case when spectral index κ is very large ( i.e. κ→∞), shifts to Maxwellian distribution. Two temperature electron plasmas have been observed in auroral regions by FAST satellite mission, and also by GEOTAIL and POLAR satellite in the magnetosphere. Kinetic Alfven waves arise when finite Larmor radius effect modifies the dispersion relation or characteristic perpendicular wavelength is comparable to electron inertial length. We have studied the kinetic Alfven waves (KAWs) in a plasma comprising of positively charged ions, superthermal hot electrons and Maxwellian distributed cold electrons. Sagdeev pseudo-potential has been employed to derive an energy balance equation. The critical Mach number has been determined from the expression of Sagdeev pseudo-potential to see the existence of solitary structures. It is observed that sub-Alfvenic compressive solitons and super-Alfvenic rarefactive solitons exist in this plasma model. It is also observed that various parameters such as superthermality of hot electrons, relative concentration of cold and hot electron species, Mach number, plasma beta, ion to cold electron temperature ratio and ion to hot electron temperature ratio have significant effect on the amplitude and width of the KAWs. Findings of this investigation may be useful to understand the dynamics of coherent non-linear structures (i.e. KAWs) in space and astrophysical plasmas.

  3. Nonlinear effects on sound propagation through high subsonic Mach number flows in variable area ducts

    NASA Technical Reports Server (NTRS)

    Callegari, A. J.

    1979-01-01

    A nonlinear theory for sound propagation in variable area ducts carrying a nearly sonic flow is presented. Linear acoustic theory is shown to be singular and the detailed nature of the singularity is used to develop the correct nonlinear theory. The theory is based on a quasi-one dimensional model. It is derived by the method of matched asymptotic expansions. In a nearly chocked flow, the theory indicates the following processes to be acting: a transonic trapping of upstream propagating sound causing an intensification of this sound in the throat region of the duct; generation of superharmonics and an acoustic streaming effect; development of shocks in the acoustic quantities near the throat. Several specific problems are solved analytically and numerical parameter studies are carried out. Results indicate that appreciable acoustic power is shifted to higher harmonics as shocked conditions are approached. The effect of the throat Mach number on the attenuation of upstream propagating sound excited by a fixed source is also determined.

  4. The effect of varying Mach number on crossing, glancing shocks/turbulent boundary-layer interactions

    NASA Technical Reports Server (NTRS)

    Hingst, W. R.; Williams, K. E.

    1991-01-01

    Two crossing side-wall shocks interacting with a supersonic tunnel wall boundary layer have been investigated over a Mach number range of 2.5 to 4.0. The investigation included a range of equal shock strengths produced by shock generators at angles from 4.0 to 12.0 degrees. Results of flow visualization show that the interaction is unseparated at the low shock generator angles. With increasing shock strength, the flow begins to form a separated region that grows in size and moves forward and eventually the model unstarts. The wall static pressures show a symmetrical compression that merges on the centerline upstream of the inviscid shock locations and becomes more 1D downstream. The region of the 1D pressure gradient moves upstream with increasing shock strengths until it coincides with the leading edge of the shock generators at the limit before model unstart. At the limiting conditions the wall pressure gradients are primarily in the axial direction throughout.

  5. Measurement and analysis of the noise radiated by low Mach numbers centrifugal blowers

    NASA Astrophysics Data System (ADS)

    Yeager, D. M.; Lauchle, G. C.

    1987-11-01

    The broad band, aerodynamically generated noise in low tip-speed Mach number, centrifugal air moving devices is investigated. An interdisciplinary approach was taken which involved investigation of the aerodynamic and acoustic fields, and their mutual relationship. The noise generation process was studied using two experimental vehicles: (1) a scale model of a homologous family of centrifugal blowers typical of those used to cool computer and business equipment, and (2) a single blade from a centrifugal blower impeller which was placed in a known, controllable flow field. The radiation characteristics of the model blower were investigated by measuring the acoustic intensity distribution near the blower inlet and comparing it with the intensity near the inlet to an axial flow fan. Aerodynamic studies of the flow field in the inlet and at the discharge to the rotating impeller were used to assess the mean flow distribution through the impeller blade channels and to identify regions of excessive turbulence near the rotating blade row. New frequency-domain expressions for the correlation area and dipole source strength per unit area on a surface immersed in turbulence were developed which can be used to characterize the noise generation process over a rigid surface immersed in turbulence. An investigation of the noise radiated from the single, isolated airfoil (impeller blade) was performed using modern correlation and spectral analysis techniques.

  6. Noise Sources in a Low-Reynolds-Number Turbulent Jet at Mach 0.9

    NASA Technical Reports Server (NTRS)

    Freund, Jonathan B.

    2001-01-01

    The mechanisms of sound generation in a Mach 0.9, Reynolds number 3600 turbulent jet are investigated by direct numerical simulation. Details of the numerical method are briefly outlined and results are validated against an experiment at the same flow conditions. Lighthill's theory is used to define a nominal acoustic source in the jet, and a numerical solution of Lighthill's equation is compared to the simulation to verify the computational procedures. The acoustic source is Fourier transformed in the axial coordinate and time and then filtered in order to identify and separate components capable of radiating to the far field. This procedure indicates that the peak radiating component of the source is coincident with neither the peak of the full unfiltered source nor that of the turbulent kinetic energy. The phase velocities of significant components range from approximately 5% to 50% of the ambient sound speed which calls into question the commonly made assumption that the noise sources convect at a single velocity. Space-time correlations demonstrate that the sources are not acoustically compact in the streamwise direction and that the portion of the source that radiates at angles greater than 45 deg. is stationary. Filtering non-radiating wavenumber components of the source at single frequencies reveals that a simple modulated wave forms for the source, as might be predicted by linear stability analysis. At small angles from the jet axis the noise from these modes is highly directional, better described by an exponential than a standard Doppler factor.

  7. Semi-implicit iterative methods for low Mach number turbulent reacting flows

    NASA Astrophysics Data System (ADS)

    Macart, Jonathan F.; Mueller, Michael E.

    2015-11-01

    A formally second-order accurate Strang splitting approach has been developed and applied to the solution of scalar transport/reaction equations for Direct Numerical Simulation (DNS) of low Mach number turbulent reacting flows. The temporal discretization errors of this scheme are analyzed and compared with a formally first-order accurate Lie splitting approach and variations on a second-order accurate monolithic preconditioned scheme, utilizing two different preconditioners: the full Jacobian of the chemical source term and its diagonal approximation. The effect of chemical mechanism size on the relative performance of these schemes is assessed with a simple one-dimensional unsteady test case. The improved stability of the full Jacobian preconditioner is found to outpace its increased cost per time step compared to the diagonal approximation, and this advantage is found to increase with mechanism size. Likewise, the Strang splitting scheme is demonstrated to achieve better performance than the other approaches due to greater stability at larger time steps, despite greater cost per step. Finally, the schemes are evaluated with a three-dimensional unsteady turbulent planar jet flame, and similar conclusions are found as for the one-dimensional test case for relative performance.

  8. Pressure distributions on a cambered wing body configuration at subsonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Henderson, W. P.

    1975-01-01

    An investigation was conducted in the Langley high-speed 7- by 10-foot tunnel at Mach numbers of 0.20 and 0.40 and angles of attack up to about 22 deg to measure the pressure distributions on two cambered-wing configurations. The wings had the same planform (aspect ratio of 2.5 and a leading-edge-sweep angle of 44 deg) but differed in amounts of camber and twist (wing design lift coefficient of 0.35 and 0.70). The effects of wing strake on the wing pressure distributions were also studied. The results indicate that the experimental chordwise pressure distribution agrees reasonably well with the design distribution over the forward 60 percent of nearly all the airfoil sections for the lower cambered wing. The measured lifting pressures are slightly less than the design pressures over the aft part of the airfoil. For the highly cambered wing, there is a significant difference between the experimental and the design pressure level. The experimental distribution, however, is still very similar to the prescribed distribution. At angles of attack above 12 deg, the addition of a wing-fuselage strake results in a significant increase in lifting pressure coefficient at all wing stations outboard of the strake-wing intersection.

  9. Flow vector, Mach number and abundance of the Warm Breeze of neutral He observed by IBEX

    NASA Astrophysics Data System (ADS)

    Kubiak, Marzena A.; McComas, David; Galli, Andre; Kucharek, Harald; Wurz, Peter; Schwadron, Nathan; Sokol, Justyna M.; Bzowski, Maciej; Heirtzler, David M.; Möbius, Eberhard; Fuselier, Stephen; Swaczyna, Paweł; Leonard, Trevor; Park, Jeewoo

    2016-07-01

    With the velocity vector and temperature of the pristine interstellar neutral (ISN) He recently obtained with high precision from a coordinated analysis by the IBEX Science Team, we analyzed the IBEX observations of neutral He left out from this analysis. These observations were collected during the interstellar neutral observation seasons 2010---2014 and cover the region in the Earth's orbit where the Warm Breeze persists. The Warm Breeze is a newly discovered population of neutral He in the heliosphere. We search for the inflow velocity vector and the temperature of the Warm Breeze and used the same simulation model and a very similar parameter fitting method to that used for the analysis of ISN He. We approximate the parent population of the Warm Breeze in front of the heliosphere with a homogeneous Maxwell-Boltzmann distribution function and find a temperature of ~9 500 K, an inflow speed of ~11.3 km/s, and an inflow longitude and latitude in the J2000 ecliptic coordinates 251.6°, 12.0°. The abundance of the Warm Breeze relative to the interstellar neutral He is 5.6% and the Mach number of the flow is 1.97. We discuss implications of this result for the heliospheric physics and an insight into the behavior of interstellar plasma in the outer heliosheath.

  10. Towards a generalized computational fluid dynamics technique for all Mach numbers

    NASA Astrophysics Data System (ADS)

    Walters, R. W.; Slack, D. C.; Godfrey, A. G.

    1993-07-01

    Currently there exists no single unified approach for efficiently and accurately solving computational fluid dynamics (CFD) problems across the Mach number regime, from truly low speed incompressible flows to hypersonic speeds. There are several CFD codes that have evolved into sophisticated prediction tools with a wide variety of features including multiblock capabilities, generalized chemistry and thermodynamics models among other features. However, as these codes evolve, the demand placed on the end user also increases simply because of the myriad of features that are incorporated into these codes. In order for a user to be able to solve a wide range of problems, several codes may be needed requiring the user to be familiar with the intricacies of each code and their rather complicated input files. Moreover, the cost of training users and maintaining several codes becomes prohibitive. The objective of the current work is to extend the compressible, characteristic-based, thermochemical nonequilibrium Navier-Stokes code GASP to very low speed flows and simultaneously improve convergence at all speeds. Before this work began, the practical speed range of GASP was Mach numbers on the order of 0.1 and higher. In addition, a number of new techniques have been developed for more accurate physical and numerical modeling. The primary focus has been on the development of optimal preconditioning techniques for the Euler and the Navier-Stokes equations with general finite-rate chemistry models and both equilibrium and nonequilibrium thermodynamics models. We began with the work of Van Leer, Lee, and Roe for inviscid, one-dimensional perfect gases and extended their approach to include three-dimensional reacting flows. The basic steps required to accomplish this task were a transformation to stream-aligned coordinates, the formulation of the preconditioning matrix, incorporation into both explicit and implicit temporal integration schemes, and modification of the numerical

  11. Towards a generalized computational fluid dynamics technique for all Mach numbers

    NASA Technical Reports Server (NTRS)

    Walters, R. W.; Slack, D. C.; Godfrey, A. G.

    1993-01-01

    Currently there exists no single unified approach for efficiently and accurately solving computational fluid dynamics (CFD) problems across the Mach number regime, from truly low speed incompressible flows to hypersonic speeds. There are several CFD codes that have evolved into sophisticated prediction tools with a wide variety of features including multiblock capabilities, generalized chemistry and thermodynamics models among other features. However, as these codes evolve, the demand placed on the end user also increases simply because of the myriad of features that are incorporated into these codes. In order for a user to be able to solve a wide range of problems, several codes may be needed requiring the user to be familiar with the intricacies of each code and their rather complicated input files. Moreover, the cost of training users and maintaining several codes becomes prohibitive. The objective of the current work is to extend the compressible, characteristic-based, thermochemical nonequilibrium Navier-Stokes code GASP to very low speed flows and simultaneously improve convergence at all speeds. Before this work began, the practical speed range of GASP was Mach numbers on the order of 0.1 and higher. In addition, a number of new techniques have been developed for more accurate physical and numerical modeling. The primary focus has been on the development of optimal preconditioning techniques for the Euler and the Navier-Stokes equations with general finite-rate chemistry models and both equilibrium and nonequilibrium thermodynamics models. We began with the work of Van Leer, Lee, and Roe for inviscid, one-dimensional perfect gases and extended their approach to include three-dimensional reacting flows. The basic steps required to accomplish this task were a transformation to stream-aligned coordinates, the formulation of the preconditioning matrix, incorporation into both explicit and implicit temporal integration schemes, and modification of the numerical

  12. Discrete sonic jets used as boundary-layer trips at Mach numbers of 6 and 8.5

    NASA Technical Reports Server (NTRS)

    Stone, D. R.; Cary, A. M., Jr.

    1972-01-01

    The effect of discrete three-dimensional sonic jets used to promote transition on a sharp-leading-edge flat plate at Mach numbers of 6 and 8.5 and unit Reynolds numbers as high as 2.5 x 100,000 per cm in the Langley 20-inch hypersonic tunnels is discussed. An examination of the downstream flow-field distortions associated with the discrete jets for the Mach 8.5 flow was also conducted. Jet trips are found to produce lengths of turbulent flow comparable to those obtained for spherical-roughness-element trips while significantly reducing the downstream flow distortions. A Reynolds number based upon secondary jet penetration into a supersonic main flow is used to correlate jet-trip effectiveness just as a Reynolds number based upon roughness height is used to correlate spherical-trip effectiveness. Measured heat-transfer data are in agreement with the predictions.

  13. Convective heat transfer studies at high temperatures with pressure gradient for inlet flow Mach number of 0.45

    NASA Technical Reports Server (NTRS)

    Pedrosa, A. C. F.; Nagamatsu, H. T.; Hinckel, J. A.

    1984-01-01

    Heat transfer measurements were determined for a flat plate with and without pressure gradient for various free stream temperatures, wall temperature ratios, and Reynolds numbers for an inlet flow Mach number of 0.45, which is a representative inlet Mach number for gas turbine rotor blades. A shock tube generated the high temperature and pressure air flow, and a variable geometry test section was used to produce inlet flow Mach number of 0.45 and accelerate the flow over the plate to sonic velocity. Thin-film platinum heat gages recorded the local heat flux for laminar, transition, and turbulent boundary layers. The free stream temperatures varied from 611 R (339 K) to 3840 R (2133 K) for a T(w)/T(r,g) temperature ratio of 0.87 to 0.14. The Reynolds number over the heat gages varied from 3000 to 690,000. The experimental heat transfer data were correlated with laminar and turbulent boundary layer theories for the range of temperatures and Reynolds numbers and the transition phenomenon was examined.

  14. In-flight boundary-layer measurements on a hollow cylinder at a Mach number of 3.0

    NASA Technical Reports Server (NTRS)

    Quinn, R. D.; Gong, L.

    1980-01-01

    Skin temperatures, shear forces, surface static pressures, boundary layer pitot pressures, and boundary layer total temperatures were measured on the external surface of a hollow cylinder that was 3.04 meters long and 0.437 meter in diameter and was mounted beneath the fuselage of the YF-12A airplane. The data were obtained at a nominal free stream Mach number of 3.0 (a local Mach number of 2.9) and at wall to recovery temperature ratios of 0.66 to 0.91. The local Reynolds number had a nominal value of 4,300,000 per meter. Heat transfer coefficients and skin friction coefficients were derived from skin temperature time histories and shear force measurements, respectively. In addition, boundary layer velocity profiles were derived from pitot pressure measurements, and a Reynolds analogy factor was obtained from the heat transfer and skin friction measurements. The measured data are compared with several boundary layer prediction methods.

  15. Low Mach number two-dimensional hydrodynamic turbulence - Energy budgets and density fluctuations in a polytropic fluid

    NASA Technical Reports Server (NTRS)

    Ghosh, S.; Matthaeus, W. H.

    1992-01-01

    Theory suggests that three distinct types of turbulence can occur in the low Mach number limit of polytropic flow: nearly incompressible flows dominated by vorticity, nearly pure acoustic turbulence dominated by compression, and flows characterized by near statistical equipartition of vorticity and compressions. Distinctions between these kinds of turbulence are investigated here by direct numerical simulation of two-dimensional compressible hydrodynamic turbulence. Dynamical scalings of density fluctuations, examination of the ratio of transverse to longitudinal velocity fluctuations, and spectral decomposition of the fluctuations are employed to distinguish the nature of these low Mach number solutions. A strong dependence on the initial data is observed, as well as a tendency for enhanced effects of compressibility at later times and at higher wave numbers, as suggested by theories of nearly incompressible flows.

  16. Measurements in Flight of the Pressure Distribution on the Right Wing of a Pursuit-Type Airplane at Several Values of Mach Number

    NASA Technical Reports Server (NTRS)

    Clousing, Lawrence A; Turner, William N; Rolls, L Stewart

    1946-01-01

    Pressure-distribution measurements were made on the right wing of a pursuit-type airplane at values of Mach number up to 0.80. The results showed that a considerable portion of the lift was carried by components of the airplane other than the wings, and that the proportion of lift carried by the wings may vary considerably with Mach number, thus changing the bending moment at the wing root whether or not there is a shift in the lateral position of the center of pressure. It was also shown that the center of pressure does not necessarily move outward at high Mach numbers, even though the wing-thickness ratio decreases toward the wing tip. The wing pitching-moment coefficient increased sharply in a negative direction at a Mach lift-curve slope increased with Mach number up to values of above the critical value. Pressures inside the wing were small and negative.

  17. Wind tunnel investigation of the aerodynamic characteristics of five forebody models at high angles of attack at Mach numbers from 0.25 to 2

    NASA Technical Reports Server (NTRS)

    Keener, E. R.; Taleghani, J.

    1975-01-01

    Five forebody models of various shapes were tested in the Ames 6- by 6-Foot Wind Tunnel to determine the aerodynamic characteristics at Mach numbers from 0.25 to 2 at a Reynolds number of 800000. At a Mach number of 0.6 the Reynolds number was varied from 0.4 to 1.8 mil. Angle of attack was varied from -2 deg to 88 deg at zero sideslip. The purpose of the investigation was to determine the effect of Mach number of the side force that develops at low speeds and zero sideslip for all of these forebody models when the nose is pointed. Test results show that with increasing Mach number the maximum side forces decrease to zero between Mach numbers of 0.8 and 1.5, depending on the nose angle; the smaller the nose angle of the higher the Mach number at which the side force exists. At a Mach number of 0.6 there is some variation of side force with Reynolds number, the variation being the largest for the more slender tangent ogive.

  18. Magnetic Helicity of Alfven Simple Waves

    NASA Technical Reports Server (NTRS)

    Webb, Gary M.; Hu, Q.; Dasgupta, B.; Zank, G. P.; Roberts, D.

    2010-01-01

    The magnetic helicity of fully nonlinear, multi-dimensional Alfven simple waves are investigated, by using relative helicity formulae and also by using an approach involving poloidal and toroidal decomposition of the magnetic field and magnetic vector potential. Different methods to calculate the magnetic vector potential are used, including the homotopy and Biot-Savart formulas. Two basic Alfven modes are identified: (a) the plane 1D Alfven simple wave given in standard texts, in which the Alfven wave propagates along the z-axis, with wave phase varphi=k_0(z-lambda t), where k_0 is the wave number and lambda is the group velocity of the wave, and (b)\\ the generalized Barnes (1976) simple Alfven wave in which the wave normal {bf n} moves in a circle in the xy-plane perpendicular to the mean field, which is directed along the z-axis. The plane Alfven wave (a) is analogous to the slab Alfven mode and the generalized Barnes solution (b) is analogous to the 2D mode in Alfvenic, incompressible turbulence. The helicity characteristics of these two basic Alfven modes are distinct. The helicity characteristics of more general multi-dimensional simple Alfven waves are also investigated. Applications to nonlinear Aifvenic fluctuations and structures observed in the solar wind are discussed.

  19. Magnetic Helicity of Alfven Simple Waves

    NASA Astrophysics Data System (ADS)

    Webb, G. M.; Hu, Q.; Dasgupta, B.; Zank, G. P.; Roberts, D.

    2010-12-01

    The magnetic helicity of fully nonlinear, multi-dimensional Alfven simple waves are investigated, by using relative helicity formulae and also by using an approach involving poloidal and toroidal decomposition of the magnetic field and magnetic vector potential. Different methods to calculate the magnetic vector potential are used, including the homotopy and Biot-Savart formulas. Two basic Alfven modes are identified: (a) the plane 1D Alfven simple wave given in standard texts, in which the Alfven wave propagates along the z-axis, with wave phase \\varphi=k0(z-λ t), where k0 is the wave number and λ is the group velocity of the wave, and (b) the generalized Barnes (1976) simple Alfvén wave in which the wave normal n moves in a circle in the xy-plane perpendicular to the mean field, which is directed along the z-axis. The plane Alfven wave (a) is analogous to the slab Alfven mode and the generalized Barnes solution (b) is analogous to the 2D mode in Alfvenic, incompressible turbulence. The helicity characteristics of these two basic Alfven modes are distinct. The helicity characteristics of more general multi-dimensional simple Alfven waves are also investigated. Applications to nonlinear Alfvenic fluctuations and structures observed in the solar wind are discussed.

  20. Bumblebee Program: Aerodynamic data. Part 4: Wing loads at Mach numbers 1.5 and 2.0. [missile configurations

    NASA Technical Reports Server (NTRS)

    Barnes, G. A.; Cronvich, L. L.

    1979-01-01

    Individual wing panel aerodynamic characteristics are provided for rectangular wings with aspect ratios of 0.25, 0.75, and 1.00 each panel at Mach numbers if 1.5 and 2.0 for angles of attack to 23 degrees. Data plots produced from reports of wind tunnel tests show normal force coefficients, and the spanwise and chordwise center of pressure locations.

  1. Modification of the formation of high-Mach number electrostatic shock-like structures by the ion acoustic instability

    SciTech Connect

    Dieckmann, M. E.; Sarri, G.; Doria, D.; Borghesi, M.; Pohl, M.

    2013-10-15

    The formation of unmagnetized electrostatic shock-like structures with a high Mach number is examined with one- and two-dimensional particle-in-cell (PIC) simulations. The structures are generated through the collision of two identical plasma clouds, which consist of equally hot electrons and ions with a mass ratio of 250. The Mach number of the collision speed with respect to the initial ion acoustic speed of the plasma is set to 4.6. This high Mach number delays the formation of such structures by tens of inverse ion plasma frequencies. A pair of stable shock-like structures is observed after this time in the 1D simulation, which gradually evolves into electrostatic shocks. The ion acoustic instability, which can develop in the 2D simulation but not in the 1D one, competes with the nonlinear process that gives rise to these structures. The oblique ion acoustic waves fragment their electric field. The transition layer, across which the bulk of the ions change their speed, widens and their speed change is reduced. Double layer-shock hybrid structures develop.

  2. Modification of the formation of high-Mach number electrostatic shock-like structures by the ion acoustic instability

    NASA Astrophysics Data System (ADS)

    Dieckmann, M. E.; Sarri, G.; Doria, D.; Pohl, M.; Borghesi, M.

    2013-10-01

    The formation of unmagnetized electrostatic shock-like structures with a high Mach number is examined with one- and two-dimensional particle-in-cell (PIC) simulations. The structures are generated through the collision of two identical plasma clouds, which consist of equally hot electrons and ions with a mass ratio of 250. The Mach number of the collision speed with respect to the initial ion acoustic speed of the plasma is set to 4.6. This high Mach number delays the formation of such structures by tens of inverse ion plasma frequencies. A pair of stable shock-like structures is observed after this time in the 1D simulation, which gradually evolves into electrostatic shocks. The ion acoustic instability, which can develop in the 2D simulation but not in the 1D one, competes with the nonlinear process that gives rise to these structures. The oblique ion acoustic waves fragment their electric field. The transition layer, across which the bulk of the ions change their speed, widens and their speed change is reduced. Double layer-shock hybrid structures develop.

  3. Measurement and Analysis of the Noise Radiated by Low Mach Number Centrifugal Blowers.

    NASA Astrophysics Data System (ADS)

    Yeager, David Marvin

    An investigation was performed of the broad band, aerodynamically generated noise in low tip-speed Mach number, centrifugal air moving devices. An interdisciplinary experimental approach was taken which involved investigation of the aerodynamic and acoustic fields, and their mutual relationship. The noise generation process was studied using two experimental vehicles: (1) a scale model of a homologous family of centrifugal blowers typical of those used to cool computer and business equipment, and (2) a single blade from a centrifugal blower impeller placed in a known, controllable flow field. The radiation characteristics of the model blower were investigated by measuring the acoustic intensity distribution near the blower inlet and comparing it with the intensity near the inlet to an axial flow fan. Results showed that the centrifugal blower is a distributed, random noise source, unlike an axial fan which exhibited the effects of a coherent, interacting source distribution. Aerodynamic studies of the flow field in the inlet and at the discharge to the rotating impeller were used to assess the mean flow distribution through the impeller blade channels and to identify regions of excessive turbulence near the rotating blade row. Both circumferential and spanwise mean flow nonuniformities were identified along with a region of increased turbulence just downstream of the scroll cutoff. The fluid incidence angle, normally taken as an indicator of blower performance, was estimated from mean flow data as deviating considerably from an ideal impeller design. An investigation of the noise radiated from the single, isolated airfoil was performed using modern correlation and spectral analysis techniques. Radiation from the single blade in flow was characterized using newly developed expressions for the correlation area and the dipole source strength per unit area, and from the relationship between the blade surface pressure and the incident turbulent flow field. Results

  4. Effects of body shape on the aerodynamics of a body of revolution at Mach numbers from 1.6 to 4.6

    NASA Technical Reports Server (NTRS)

    Spearman, M. L.

    1985-01-01

    The aerodnamic characteristics for several bodies of revolution have been determined from wind tunnel tests at Mach numbers from 1.6 to 4.63. Six bodies, each having a length-to-diameter ratio of 6.67, were investigated. Geometric modifications included forebody shape, afterbody shape, and midsection slope. Significant aerodynamic changes were observed to be functions of geometric change and Mach number. Because of the aerodynamic dependence on geometry as well as Mach number, it is obvious that a number of trades must be considered in selecting a projectile shape.

  5. On the use of freon-12 for increasing Reynolds number in wind-tunnel testing of three dimensional aircraft models at subcritical and supercritical Mach numbers

    NASA Technical Reports Server (NTRS)

    Treon, S. L.; Hofstetter, W. R.; Abbott, F. T.

    1971-01-01

    The aerodynamic suitability of Freon-12 for general wind-tunnel testing was investigated at low and high subsonic speeds. Static aerodynamic characteristics of two transport airplane models were determined from strain gage balance measurements in both air and Freon-12 at several Reynolds numbers. A low-speed high-lift configuration was evaluated at Mach number 0.25, and a high-speed cruise wing-fuselage combination was tested at Mach numbers up to 0.825. The data obtained in air and in Freon-12 agree well, even in stalled flow, until compressibility effects evidently become significant in air and in Freon-12 agree well, even in stalled flow, until compressibility effects evidently become significant in air.

  6. Isolated Performance at Mach Numbers From 0.60 to 2.86 of Several Expendable Nozzle Concepts for Supersonic Applications

    NASA Technical Reports Server (NTRS)

    Re, Richard J.; Berrier, Bobby L.; Abeyounis, William K.

    2001-01-01

    Investigations have been conducted in the Langley 16-Foot Transonic Tunnel (at Mach numbers from 0.60 to 1.25) and in the Langley Unitary Plan Wind Tunnel (at Mach numbers from 2.16 to 2.86) at an angle of attack of 0 deg to determine the isolated performance of several expendable nozzle concepts for supersonic nonaugmented turbojet applications. The effects of centerbody base shape, shroud length, shroud ventilation, cruciform shroud expansion ratio, and cruciform shroud flap vectoring were investigated. The nozzle pressure ratio range, which was a function of Mach number, was between 1.9 and 11.8 in the 16-Foot Transonic Tunnel and between 7.9 and 54.9 in the Unitary Plan Wind Tunnel. Discharge coefficient, thrust-minus-drag, and the forces and moments generated by vectoring the divergent shroud flaps (for Mach numbers of 0.60 to 1.25 only) of a cruciform nozzle configuration were measured. The shortest nozzle had the best thrust-minus-drag performance at Mach numbers up to 0.95 but was approached in performance by other configurations at Mach numbers of 1.15 and 1.25. At Mach numbers above 1.25, the cruciform nozzle configuration having the same expansion ratio (2.64) as the fixed geometry nozzles had the best thrust-minus-drag performance. Ventilation of the fixed geometry divergent shrouds to the nozzle external boattail flow generally improved thrust-minus-drag performance at Mach numbers from 0.60 to 1.25, but decreased performance above a Mach number of 1.25.

  7. Alfven cascades with downward frequency sweeping

    SciTech Connect

    Marchenko, V. S.; Reznik, S. N.

    2011-04-15

    It is suggested that relatively rare, but challenging for the existing theory Alfven cascades with downward frequency sweeping are actually the infernal Alfven eigenmodes (IAEs). Such modes exist in discharges with flat or weakly reversed q-profile in the broad central region, when the value of the safety factor in this region is slightly above the integer or low-order rational. Similar to the toroidal Alfven eigenmode, but in contrast to the ''conventional'' Alfven cascade with upward frequency sweeping, the spectrum of IAE is almost degenerate with respect to the mode numbers. Both features mentioned above are consistent with experimental observations.

  8. Flight-measured afterbody pressure coefficients from an airplane having twin side-by-side jet engines for Mach numbers from 0.6 to 1.6

    NASA Technical Reports Server (NTRS)

    Steers, L. L.

    1979-01-01

    Afterbody pressure distribution data were obtained in flight from an airplane having twin side-by-side jet exhausts. The data were obtained in level flight at Mach numbers from 0.60 to 1.60 and at elevated load factors for Mach numbers of 0.60, 0.90, and 1.20. The test altitude varied from 2300 meters (7500 feet) to 15,200 meters (50,000 feet) over a speed range that provided a matrix of constant Mach number and constant unit Reynolds number test conditions. The results of the full-scale flight afterbody pressure distribution program are presented in the form of plotted pressure distributions and tabulated pressure coefficients with Mach number, angle of attack, engine nozzle pressure ratio, and unit Reynolds number as controlled parameters.

  9. Results obtained during accelerated transonic tests of the Bell XS-1 airplane in flights to a Mach number of 0.92

    NASA Technical Reports Server (NTRS)

    Drake, Hubert M; Mclaughlin, Milton D; Goodman, Harold R

    1948-01-01

    Results are presented of tests up to a Mach number of 0.92 at altitudes around 30,000 feet. The data obtained show that the airplane can be flown to this Mach number above 30,000 feet. Longitudinal trim changes have been experienced but the forces involved have been small. The elevator effectiveness decreased about one-half with increase of Mach number from 0.70 to 0.87. Buffeting has been experienced in level flight but it has been mild and the associated tail loads have been small. No aileron buzz or other flutter phenomena have been noted.

  10. A study of the effects of Reynolds number and Mach number on constant pressure coefficient jump for shock-induced trailing-edge separation

    NASA Technical Reports Server (NTRS)

    Cunningham, Atlee M., Jr.; Spragle, Gregory S.

    1987-01-01

    The influence of Mach and Reynolds numbers as well as airfoil and planform geometry on the phenomenon of constant shock jump pressure coefficient for conditions of shock induced trailing edge separation (SITES) was studied. It was demonstrated that the phenomenon does exist for a wide variety of two and three dimensional flow cases and that the influence of free stream Mach number was not significant. The influence of Reynolds number was found to be important but was not strong. Airfoil and planform geometric characteristics were found to be very important where the pressure coefficient jump was shown to vary with the sum of: (1) airfoil curvature at the upper surface crest, and (2) camber surface slope at the trailing edge. It was also determined that the onset of SITES could be defined as a function of airfoil geometric parameters and Mach number normal to the leading edge. This onset prediction was shown to predict the angle of onset to within + or - 1 deg accuracy or better for about 90% of the cases studied.

  11. Aerodynamic Characteristics of a Slender Cone-cylinder Body of Revolution at a Mach Number of 3.85

    NASA Technical Reports Server (NTRS)

    Jack, John R

    1951-01-01

    An experimental investigation of the aerodynamics of a slender cone-cylinder body of revolution was conducted at a Mach number of 3.85 for angles of attack of 0 degree to 10 degrees and a Reynolds number of 3.85x10(exp 6). Boundary-layer measurements at zero angle of attack are compared with the compressible-flow formulations for predicting laminar boundary-layer characteristics. Comparison of experimental pressure and force values with theoretical values showed relatively good agreement for small angles of attack. The measured mean skin-friction coefficients agreed well with theoretical values obtained for laminar flow over cones.

  12. Flight-determined derivatives and dynamic characteristics for the HL-10 lifting body vehicle at subsonic and transonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Strutz, L. W.

    1972-01-01

    The HL-10 lifting body stability and control derivatives were determined by using an analog-matching technique and compared with derivatives obtained from wind-tunnel results. The flight derivatives were determined as a function of angle of attack for a subsonic configuration at Mach 0.7 and for a transonic configuration at Mach 0.7, 0.9, and 1.2. At an angle of attack of 14 deg, data were obtained for a Mach number range from 0.6 to 1.4. The flight and wind-tunnel derivatives were in general agreement, with the possible exception of the longitudinal and lateral damping derivatives. Some differences were noted between the vehicle dynamic response characteristics calculated from flight-determined derivatives and those predicted by the wind-tunnel results. However, the only difference the pilots noted between the response of the vehicle in flight and the response of a simulator programed with wind-tunnel-predicted data was that the damping generally was higher in the flight vehicle.

  13. Non-thermal electron acceleration in low Mach number collisionless shocks. I. Particle energy spectra and acceleration mechanism

    SciTech Connect

    Guo, Xinyi; Narayan, Ramesh; Sironi, Lorenzo

    2014-10-20

    Electron acceleration to non-thermal energies in low Mach number (M{sub s} ≲ 5) shocks is revealed by radio and X-ray observations of galaxy clusters and solar flares, but the electron acceleration mechanism remains poorly understood. Diffusive shock acceleration, also known as first-order Fermi acceleration, cannot be directly invoked to explain the acceleration of electrons. Rather, an additional mechanism is required to pre-accelerate the electrons from thermal to supra-thermal energies, so they can then participate in the Fermi process. In this work, we use two- and three-dimensional particle-in-cell plasma simulations to study electron acceleration in low Mach number shocks. We focus on the particle energy spectra and the acceleration mechanism in a reference run with M{sub s} = 3 and a quasi-perpendicular pre-shock magnetic field. We find that about 15% of the electrons can be efficiently accelerated, forming a non-thermal power-law tail in the energy spectrum with a slope of p ≅ 2.4. Initially, thermal electrons are energized at the shock front via shock drift acceleration (SDA). The accelerated electrons are then reflected back upstream where their interaction with the incoming flow generates magnetic waves. In turn, the waves scatter the electrons propagating upstream back toward the shock for further energization via SDA. In summary, the self-generated waves allow for repeated cycles of SDA, similarly to a sustained Fermi-like process. This mechanism offers a natural solution to the conflict between the bright radio synchrotron emission observed from the outskirts of galaxy clusters and the low electron acceleration efficiency usually expected in low Mach number shocks.

  14. Comparison of reacting and non-reacting shear layers at a high subsonic Mach number

    NASA Technical Reports Server (NTRS)

    Chang, C. T.; Marek, C. J.; Wey, C.; Jones, R. A.; Smith, M. J.

    1993-01-01

    The flow field in a hydrogen-fueled planar reacting shear layer was measured with an LDV system and is compared with a similar air to air case without combustion. Measurements were made with a speed ratio of 0.34 with the highspeed stream at Mach 0.71. They show that the shear layer with reaction grows faster than one without, and both cases are within the range of data scatter presented by the established database. The coupling between the streamwise and the cross-stream turbulence components inside the shear layer is slow, and reaction only increased it slightly. However, a more organized pattern of the Reynolds stress is present in the reacting shear layer, possibly as a result of larger scale structure formation in the layer associated with heat release.

  15. Longitudinal Stability and Drag Characteristics at Mach Numbers from 0.70 to 1.37 of Rocket-propelled Models Having a Modified Triangular Wing

    NASA Technical Reports Server (NTRS)

    Chapman, Rowe, Jr; Morrow, John D

    1952-01-01

    A modified triangular wing of aspect ratio 2.53 having an airfoil section 3.7 percent thick at the root and 5.98 percent thick at the tip was designed in an attempt to improve the lift and drag characteristics of triangular wings. Free-flight drag and stability tests were made using rocket-propelled models equipped with the modified wing. The Mach number range of the test was from 0.70 to 1.37. Test results indicated the following: The lift-curve slope of wing plus fuselage approaches the theoretical value of wing alone at supersonic Mach numbers. The drag coefficient, based on total wing area, for wing plus interference was approximately 0.0035 at subsonic Mach numbers and 0.0080 at supersonic Mach numbers. The maximum shift in aerodynamic center for the complete configuration was 14 percent in the rearward direction from the forward position of 51.5 percent of mean aerodynamic chord at subsonic Mach numbers. The variation of lift and moment with angle of attack was linear at supersonic Mach numbers for the range of coefficients covered in the test. The high value of lift-curve slope was considered to be a significant result attributable to the wing modifications.

  16. Effects of forebody strakes and Mach number on overall aerodynamic characteristics of configuration with 55 deg cropped delta wing

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.; Rogers, Lawrence W.

    1992-01-01

    A wind tunnel data base was established for the effects of chine-like forebody strakes and Mach number on the longitudinal and lateral-directional characteristics of a generalized 55 degree cropped delta wing-fuselage-centerline vertical tail configuration. The testing was conducted in the 7- by 10-Foot Transonic Tunnel at the David Taylor Research Center at free-stream Mach numbers of 0.40 to 1.10 and Reynolds numbers based on the wing mean aerodynamic chord of 1.60 x 10(exp 6) to 2.59 x 10(exp 6). The best matrix included angles of attack from 0 degree to a maximum of 28 degree, angles of sidesip of 0, +5, and -5 degrees, and wing leading-edge flat deflection angles of 0 and 30 degrees. Key flow phenomena at subsonic and transonic conditions were identified by measuring off-body flow visualization with a laser screen technique. These phenomena included coexisting and interacting vortex flows and shock waves, vortex breakdown, vortex flow interactions with the vertical tail, and vortices induced by flow separation from the hinge line of the deflected wing flap. The flow mechanisms were correlated with the longitudinal and lateral-directional aerodynamic data trends.

  17. Study of turbulent boundary layers over rough surfaces with emphasis n the effects of roughness character and Mach number

    NASA Astrophysics Data System (ADS)

    Finson, M. L.

    1982-02-01

    A Reynolds stress model for turbulent boundary layers on rough walls is used to investigate the effects of roughness character and compressibility. The flow around roughness elements is treated as form drag. A method is presented for deriving the required roughness shape and spacing from profilometer surface measurements. Calculations based on the model compare satisfactorily with low speed data on roughness character and hypersonic measurements with grit roughness. The computer model is exercised systematically over a wide range of parameters to derive a practical scaling law for the equivalent roughness. In contrast to previous correlations, for most roughness element shapes the effective roughness is not predicted to show a pronounced maximum as the element spacing decreases. The effect of roughness tends to be reduced with increasing edge Mach number, primarily due to decreasing density in the vicinity of the roughness elements. It is further shown that the required roughness Reynolds number for fully rough behavior increases with increasing Mach number, explaining the small roughness effects observed in some hypersonic tests.

  18. Helicopter far-field acoustic levels as a function of reduced main-rotor advancing blade-tip Mach number

    NASA Technical Reports Server (NTRS)

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

    1990-01-01

    During the design of a helicopter, the weight, engine, rotor speed, and rotor geometry are given significant attention when considering the specific operations for which the helicopter will be used. However, the noise radiated from the helicopter and its relationship to the design variables is currently not well modeled with only a limited set of full-scale field test data to study. In general, limited field data have shown that reduced main-rotor advancing blade-tip Mach numbers result in reduced far-field noise levels. The status of a recent helicopter noise research project is reviewed. It is designed to provide flight experimental data which may be used to further understand helicopter main-rotor advancing blade-tip Mach number effects on far-field acoustic levels. Preliminary results are presented relative to tests conducted with a Sikorsky S-76A helicopter operating with both the rotor speed and the flight speed as the control variable. The rotor speed was operated within the range of 107 to 90 percent NR at nominal forward speeds of 35, 100, and 155 knots.

  19. Performance characteristics of two multiaxis thrust-vectoring nozzles at Mach numbers up to 1.28

    NASA Technical Reports Server (NTRS)

    Wing, David J.; Capone, Francis J.

    1993-01-01

    The thrust-vectoring axisymmetric (VA) nozzle and a spherical convergent flap (SCF) thrust-vectoring nozzle were tested along with a baseline nonvectoring axisymmetric (NVA) nozzle in the Langley 16-Foot Transonic Tunnel at Mach numbers from 0 to 1.28 and nozzle pressure ratios from 1 to 8. Test parameters included geometric yaw vector angle and unvectored divergent flap length. No pitch vectoring was studied. Nozzle drag, thrust minus drag, yaw thrust vector angle, discharge coefficient, and static thrust performance were measured and analyzed, as well as external static pressure distributions. The NVA nozzle and the VA nozzle displayed higher static thrust performance than the SCF nozzle throughout the nozzle pressure ratio (NPR) range tested. The NVA nozzle had higher overall thrust minus drag than the other nozzles throughout the NPR and Mach number ranges tested. The SCF nozzle had the lowest jet-on nozzle drag of the three nozzles throughout the test conditions. The SCF nozzle provided yaw thrust angles that were equal to the geometric angle and constant with NPR. The VA nozzle achieved yaw thrust vector angles that were significantly higher than the geometric angle but not constant with NPR. Nozzle drag generally increased with increases in thrust vectoring for all the nozzles tested.

  20. Performance of a short annular dump diffuser using suction-stabilized vortices at inlet Mach numbers to 0.41

    NASA Technical Reports Server (NTRS)

    Smith, J. M.; Juhasz, A. J.

    1978-01-01

    A short, annular dump diffuser was designed to use suction to establish stabilized vortices on both walls for improved flow expansion in the region of an abrupt area change. The diffuser was tested at near ambient inlet pressure and temperature. The overall diffuser area ratio was 4.0. The inlet height was 2.54 cm and the exit pitot-static rakes were located at a distance from the vortex fence equal to two or six times the inlet height. Performance data were taken at near ambient temperature and pressure for nominal inlet Mach numbers of 0.18 to 0.41 with suction rates of 0 to 18 percent of the total inlet airflow. The exit velocity profile could be shifted toward either wall by adjusting the inner- or outer-wall suction rate. Symmetrical exit velocity profiles were unstable, with a tendency to shift back to hub- or tip-weighted profile. Diffuser effectiveness was increased from about 47 percent without suction to over 85 percent at a total suction rate of about 14 percent. The diffuser total pressure losses at inlet Mach numbers of 0.18 and 0.41 decreased from 1.1 and 5.6 percent without suction to 0.48 and 5.2 percent at total suction rates of 14.4 and 5.6 percent, respectively.

  1. Preliminary Base Pressures Obtained from the X-15 Airplane at Mach Numbers from 1.1 to 3.2

    NASA Technical Reports Server (NTRS)

    Saltzman, Edwin J.

    1961-01-01

    Base pressure measurements have been made on the fuselage, 10 deg.-wedge vertical fin, and side fairing of the X-15 airplane. Data are presented for Mach numbers between 1.1 and 3.2 for both powered and unpowered flight. Comparisons are made with data from small-scale-model tests, semiempirical estimates, and theory. The results of this preliminary study show that operation of the interim rocket engines (propellant flow rate approximately 70 lb/sec) reduces the base drag of the X-15 by 25 to 35 percent throughout the test Mach number range. Values of base drag coefficient for the side fairing and fuselage obtained from X-15 wind-tunnel models were adequate for predicting the overall full-scale performance of the test airplane. The leading-edge sweep of the upper movable vertical fin was not an important factor affecting the fin base pressure. The power-off base pressure coefficients of the upper movable vertical fin (a 10 deg. wedge with chord-to-thickness ratio of 5.5 and semispan-to-thickness ratio of 3.2) are in general agreement with the small-scale blunt-trailing-edge-wing data of several investigators and with two-dimensional theory.

  2. Wind-tunnel investigation of a flush airdata system at Mach numbers from 0.7 to 1.4

    NASA Technical Reports Server (NTRS)

    Larson, Terry J.; Moes, Timothy R.; Siemers, Paul M., III

    1990-01-01

    Flush pressure orifices installed on the nose section of a 1/7-scale model of the F-14 airplane were evaluated for use as a flush airdata system (FADS). Wing-tunnel tests were conducted in the 11- by 11-ft Unitary Wind Tunnel at NASA Ames Research Center. A full-scale FADS of the same configuration was previously tested using an F-14 aircraft at the Dryden Flight Research Facility of NASA Ames Research Center (Ames-Dryden). These tests, which were published, are part of a NASA program to assess accuracies of FADS for use on aircraft. The test program also provides data to validate algorithms for the shuttle entry airdata system developed at the NASA Langley Research Center. The wind-tunnel test Mach numbers were 0.73, 0.90, 1.05, 1.20, and 1.39. Angles of attack were varied in 2 deg increments from -4 deg to 20 deg. Sideslip angles were varied in 4 deg increments from -8 deg to 8 deg. Airdata parameters were evaluated for determination of free-stream values of stagnation pressure, static pressure, angle of attack, angle of sideslip, and Mach number. These parameters are, in most cases, the same as the parameters investigated in the flight test program. The basic FADS wind-tunnel data are presented in tabular form. A discussion of the more accurate parameters is included.

  3. Aerodynamic Performance and Static Stability and Control of Flat-Top Hypersonic Gliders at Mach Numbers from 0.6 to 18

    NASA Technical Reports Server (NTRS)

    Syvertson, Clarence A; Gloria, Hermilo R; Sarabia, Michael F

    1958-01-01

    A study is made of aerodynamic performance and static stability and control at hypersonic speeds. In a first part of the study, the effect of interference lift is investigated by tests of asymmetric models having conical fuselages and arrow plan-form wings. The fuselage of the asymmetric model is located entirely beneath the wing and has a semicircular cross section. The fuselage of the symmetric model was centrally located and has a circular cross section. Results are obtained for Mach numbers from 3 to 12 in part by application of the hypersonic similarity rule. These results show a maximum effect of interference on lift-drag ratio occurring at Mach number of 5, the Mach number at which the asymmetric model was designed to exploit favorable lift interference. At this Mach number, the asymmetric model is indicated to have a lift-drag ratio 11 percent higher than the symmetric model and 15 percent higher than the asymmetric model when inverted. These differences decrease to a few percent at a Mach number of 12. In the course of this part of the study, the accuracy to the hypersonic similarity rule applied to wing-body combinations is demonstrated with experimental results. These results indicate that the rule may prove useful for determining the aerodynamic characteristics of slender configurations at Mach numbers higher than those for which test equipment is really available. In a second part of the study, the aerodynamic performance and static stability and control characteristics of a hypersonic glider are investigated in somewhat greater detail. Results for Mach numbers from 3 to 18 for performance and 0.6 to 12 for stability and control are obtained by standard text techniques, by application of the hypersonic stability rule, and/or by use of helium as a test medium. Lift-drag ratios of about 5 for Mach numbers up to 18 are shown to be obtainable. The glider studied is shown to have acceptable longitudinal and directional stability characteristics through the

  4. 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.

  5. Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Volume 1; Sharp Leading Edge; [conducted in the Langley National Transonic Facility (NTF)

    NASA Technical Reports Server (NTRS)

    Chu, Julio; Luckring, James M.

    1996-01-01

    An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 36 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at a Reynolds number of 6 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

  6. Experimental Reacting Hydrogen Shear Layer Data at High Subsonic Mach Number

    NASA Technical Reports Server (NTRS)

    Chang, C. T.; Marek, C. J.; Wey, C.; Wey, C. C.

    1996-01-01

    The flow in a planar shear layer of hydrogen reacting with hot air was measured with a two-component laser Doppler velocimeter (LDV) system, a schlieren system, and OH fluorescence imaging. It was compared with a similar air-to-air case without combustion. The high-speed stream's flow speed was about 390 m/s, or Mach 0.71, and the flow speed ratio was 0.34. The results showed that a shear layer with reaction grows faster than one without; both cases are within the range of data scatter presented by the established data base. The coupling between the streamwise and the cross-stream turbulence components inside the shear layers was low, and reaction only increased it slightly. However, the shear layer shifted laterally into the lower speed fuel stream, and a more organized pattern of Reynolds stress was present in the reaction shear layer, likely as a result of the formation of a larger scale structure associated with shear layer corrugation from heat release. Dynamic pressure measurements suggest that coherent flow perturbations existed inside the shear layer and that this flow became more chaotic as the flow advected downstream. Velocity and thermal variable values are listed in this report for a computational fluid dynamics (CFD) benchmark.

  7. Aerodynamic characteristics of a canard-controlled missile at Mach numbers of 1.5 and 2.0.

    NASA Technical Reports Server (NTRS)

    Kassner, D. L.; Wettlaufer, B.

    1977-01-01

    A typical missile model with nose mounted canards and cruciform tail surfaces was tested in the Ames 6- by 6-Foot Wind Tunnel to determine the contributions of the component aerodynamic surfaces to the static aerodynamic characteristics at Mach numbers of 1.5 and 2.0 and Reynolds number of 1 million based on body diameter. Data were obtained at angles of attack ranging from -3 deg to 12 deg for various stages of model build-up (i.e., with and without canard and/or tail surfaces). Results were obtained both with the model unrolled and rolled 45 deg. For the canard and tail arrangements investigated, the model was trimmable at angles of attack up to about 10 deg with canard deflections of 9 deg. Also, the tail arrangements studied provided ample pitch stability. there were no appreciable effects of model roll orientation.

  8. Aerothermal tests of quilted dome models on a flat plate at a Mach number of 6.5

    NASA Technical Reports Server (NTRS)

    Glass, Christopher E.; Hunt, L. Roane

    1988-01-01

    Aerothermal tests were conducted in the NASA Langley 8 Foot High Temperature Tunnel (8'HTT) at a Mach number of 6.5 on simulated arrays of thermally bowed metallic thermal protection system (TPS) tiles at an angle of attack of 5 deg. Detailed surface pressures and heating rates were obtained for arrays aligned with the flow and skewed 45 deg diagonally to the flow with nominal bowed heights of 0.1, 0.2, and 0.4 inch submerged in both laminar and turbulent boundary layers. Aerothermal tests were made at a nominal total temperature of 3300 R, a total pressure of 400 psia, a total enthalpy of 950 Btu/lbm, a dynamic pressure of 2.7 psi, and a unit Reynolds number of 400,000 per foot. The experimental results form a data base that can be used to help protect aerothermal load increases from bowed arrays of TPS tiles.

  9. Wave merging mechanism: formation of low-frequency Alfven and magnetosonic waves in cosmic plasmas

    SciTech Connect

    Tishchenko, V N; Shaikhislamov, I F

    2014-02-28

    We investigate the merging mechanism for the waves produced by a pulsating cosmic plasma source. A model with a separate background/source description is used in our calculations. The mechanism was shown to operate both for strong and weak source – background interactions. We revealed the effect of merging of individual Alfven waves into a narrow low-frequency wave, whose amplitude is maximal for a plasma expansion velocity equal to 0.5 – 1 of the Alfven Mach number. This wave is followed along the field by a narrow low-frequency magnetosonic wave, which contains the bulk of source energy. For low expansion velocities the wave contains background and source particles, but for high velocities it contains only the background particles. The wave lengths are much greater than their transverse dimension. (letters)

  10. Aerodynamic Characteristics of a Circular Cylinder at Mach Number of 6.86 and Angles of Attack up to 90 Degrees

    NASA Technical Reports Server (NTRS)

    Penland, Jim A

    1954-01-01

    Pressure-distribution and force tests of a circular cylinder have been made in the Langley 11-inch hypersonic tunnel at a Mach number of 6.86, a Reynolds number of 129,000 based on diameter, and angles of attack up to 90 degrees. The results are compared with the hypersonic approximation of Grimminger, Williams, and Young and with a simple modification of the Newtonian flow theory. The comparison of experimental results shows that either theory gives adequate general aerodynamic characteristics but that the modified Newtonian theory gives a more accurate prediction of the pressure distribution. The calculated crossflow drag coefficients plotted as a function of crossflow Mach number were found to be in reasonable agreement with similar results obtained from other investigations at lower supersonic Mach numbers. Comparison of the results of this investigation with data obtained at a lower Mach number indicates that the drag coefficient of a cylinder normal to the flow is relatively constant for Mach numbers above about 4.

  11. An experimental and theoretical study of the aerodynamic characteristics of some generic missile concepts at Mach numbers from 2 to 6.8

    NASA Technical Reports Server (NTRS)

    Spearman, M. Leroy; Braswell, Dorothy O.

    1994-01-01

    A study has been made of the experimental and theoretical aerodynamic characteristics for some generic high-speed missile concepts at Mach numbers from 2 to 6.8. The basic body for this study had a length-to-diameter ratio of 10 with the forward half being a modified blunted ogive and the rear half being a cylinder. Modifications made to the basic body included the addition of an after body flare, the addition of highly swept cruciform wings and the addition of highly swept aft tails. The effects of some controls were also investigated with all-moving wing controls on the flared body and trailing-edge flap controls on the winged body. The results indicated that the addition of a flare, wings, or tails to the basic body all provided static longitudinal stability with varying amounts of increased axial force. The control arrangements were effective in producing increments of normal-force and pitching-moment at the lower Mach numbers. At the highest Mach number, the flap control on the winged body was ineffective in producing normal-force or pitching-moment but the all-moving wing control on the flared body, while losing pitch effectiveness, still provided normal-force increments. Calculated results obtained through the use of hypersonic impact theory were in generally good agreement with experiment at the higher Mach numbers but were not accurate at the lower Mach numbers.

  12. A Reynolds Number Study of Wing Leading-Edge Effects on a Supersonic Transport Model at Mach 0.3

    NASA Technical Reports Server (NTRS)

    Williams, M. Susan; Owens, Lewis R., Jr.; Chu, Julio

    1999-01-01

    A representative supersonic transport design was tested in the National Transonic Facility (NTF) in its original configuration with small-radius leading-edge flaps and also with modified large-radius inboard leading-edge flaps. Aerodynamic data were obtained over a range of Reynolds numbers at a Mach number of 0.3 and angles of attack up to 16 deg. Increasing the radius of the inboard leading-edge flap delayed nose-up pitching moment to a higher lift coefficient. Deflecting the large-radius leading-edge flap produced an overall decrease in lift coefficient and delayed nose-up pitching moment to even higher angles of attack as compared with the undeflected large- radius leading-edge flap. At angles of attack corresponding to the maximum untrimmed lift-to-drag ratio, lift and drag coefficients decreased while lift-to-drag ratio increased with increasing Reynolds number. At an angle of attack of 13.5 deg., the pitching-moment coefficient was nearly constant with increasing Reynolds number for both the small-radius leading-edge flap and the deflected large-radius leading-edge flap. However, the pitching moment coefficient increased with increasing Reynolds number for the undeflected large-radius leading-edge flap above a chord Reynolds number of about 35 x 10 (exp 6).

  13. Shadowgraphs of air flow over prospective space shuttle configurations at Mach numbers from 0.8 to 1.4

    NASA Technical Reports Server (NTRS)

    Dods, J. B., Jr.; Hanly, R. D.; Efting, J. H.

    1975-01-01

    Shadowgraphs of five space shuttle launch configurations are presented. The model was a 4 percent-scale space shuttle vehicle, tested in the 11- by 11-foot Transonic Wind Tunnel at Ames Research Center. The Mach number was varied from 0.8 to 1.4 with three angles of sideslip (0 deg, 5 deg and -5 deg) that were used in conjunction with three angles of attack (4 deg, -4 deg, and 0 deg). The model configurations included both series-burn and parallel-burn configurations, two canopy configurations, two positions of the orbiter nose relative to the HO tank nose, and two HO tank nose-cones angles (15 deg and 20 deg). The data consist entirely of shadowgraph photographs.

  14. Effect of initial conditions and Mach number on the Richtmyer-Meshkov instability in ICF like conditions

    NASA Astrophysics Data System (ADS)

    Rao, Pooja; She, Dan; Lim, Hyunkyung; Glimm, James

    2015-11-01

    The qualitative and quantitative effect of initial conditions (linear and non-linear) and high Mach number (1.3 and 1.45) is studied on the turbulent mixing induced by the Richtmyer-Meshkov instability in idealized ICF conditions. The Richtmyer-Meshkov instability seeds Rayleigh-taylor instabilities in ICF experiments and is one of the factors that contributes to reduced performance of ICF experiments. Its also found in collapsing cores of stars and supersonic combustion. We use the Stony Brook University code, FronTier, which is verified via a code comparison study against the AMR multiphysics code FLASH, and validated against vertical shock tube experiments done by the LANL Extreme Fluids Team. These simulations are designed as a step towards simulating more realistic ICF conditions and quantifying the detrimental effects of mixing on the yield.

  15. 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.

  16. Airframe-propulsion system aerodynamic interference predictions at high transonic Mach numbers including off-design engine airflow effects

    NASA Technical Reports Server (NTRS)

    Kulfan, R. M.; Sigalla, A.

    1981-01-01

    The transonic speed regime for airplanes at conditions where inlet spillage takes place is discussed. A wind tunnel test program to evaluate aerodynamic performance penalties associated with propulsion system installation and operation at subsonic through low supersonic speeds was conducted. The accuracy of analytic methods for predicting transonic engine airframe interference effects was assessed. Study variables included Mach number, angle of attack, relative nacelle location, and nacelle mass flow ratio. Results include test theory comparisons of forces as well as induced pressure fields. Prediction capability of induced shock wave strength and locations is assessed. It was found that large interference forces due to engine location and flow spillage occur at transonic speeds, that theory explains these effects; and that theory can predict quantitatively these effects.

  17. Analysis of turbine stator adjustment required for compressor design-point operation in high Mach number supersonic turbojet engines

    NASA Technical Reports Server (NTRS)

    English, Robert E; Cavicchi, Richard H

    1953-01-01

    For turbojet engines designed for flight Mach numbers of 2.5 and 3.0, use of turbine stator adjustment to maintain compressor design-point operation was evaluated analytically to determine the effect on the aerodynamics of the turbine. Since the effect of turbine stator adjustment is to make the turbine design sensitive to the particular engine design conditions selected, in some cases the turbine must be conservatively designed for the high-speed flight condition to assure satisfactory turbine performance at take-off. A new concept, the break-even point, is introduced to provide quick evaluation of the proximity of turbines to the blade-loading limit at any off-design operation.

  18. Phase Averaged Measurements of the Coherent Structure of a Mach Number 0.6 Jet. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Emami, S.

    1983-01-01

    The existence of a large scale structure in a Mach number 0.6, axisymmetric jet of cold air was proven. In order to further characterize the coherent structure, phase averaged measurements of the axial mass velocity, radial velocity, and the product of the two were made. These measurements yield information about the percent of the total fluctuations contained in the coherent structure. These measured values were compared to the total fluctuation levels for each quantity and the result expressed as a percent of the total fluctuation level contained in the organized structure at a given frequency. These measurements were performed for five frequencies (St=0.16, 0.32, 0.474, 0.95, and 1.26). All of the phase averaged measurements required that the jet be artificially excited.

  19. Drag and stability characteristics of a variety of reefed and unreefed parachute configurations at Mach 1.80 with an empirical correlation for supersonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Couch, L. M.

    1975-01-01

    An investigation was conducted at Mach 1.80 in the Langley 4-foot supersonic pressure tunnel to determine the effects of variation in reefing ratio and geometric porosity on the drag and stability characteristics of four basic canopy types deployed in the wake of a cone-cylinder forebody. The basic designs included cross, hemisflo, disk-gap-band, and extended-skirt canopies; however, modular cross and standard flat canopies and a ballute were also investigated. An empirical correlation was determined which provides a fair estimation of the drag coefficients in transonic and supersonic flow for parachutes of specified geometric porosity and reefing ratio.

  20. Flight Test of a 30-Foot Nominal Diameter Cross Parachute Deployed at a Mach Number of 1.57

    NASA Technical Reports Server (NTRS)

    1968-01-01

    Flight Test of a 30-Foot Nominal Diameter Cross Parachute Deployed at a Mach Number of 1.57 and a Dynamic Pressure of 9.7 Pounds per Square Foot. A 30-foot (9.1-meter) nominal-diameter cross-type parachute with a cloth area (reference area) of 709 square feet (65.9 square meters) was flight tested in the rocket-launched portion of the NASA Planetary Entry Parachute Program (PEPP). The test parachute was ejected from an instrumented payload by means of a mortar when the system was at a Mach number of 1.57 and a dynamic pressure of 9.7 psf. The parachute deployed to suspension-line stretch in 0.44 second with a resulting snatch-force loading of 1100 pounds (4900 newtons), Canopy inflation began at 0.58 second and a first full inflation was achieved at approximately 0.77 second. The maximum opening load occurred at 0.81 second and was 4255 pounds (18,930 newtons). Thereafter, the test item exhibited a canopy-shape instability in that the four panel arms experienced fluctuations, a 'scissoring' type of motion predominating throughout the test period. Calculated values of axial-force coefficient during the deceleration portion of the test varied between 0.35 and 1.05, with an average value of 0.69. During descent, canopy-shape variations had reduced to small amplitudes and resultant pitch-yaw angles of the payload with respect to the local vertical averaged less than 10 degrees. The effective drag coefficient, based on the vertical components of velocity and acceleration during system descent, was 0.78. [Entire movie available on DVD from CASI as Doc ID 20070030994. Contact help@sti.nasa.gov

  1. Forces and Moments on Pointed Blunt-nosed Bodies of Revolution at Mach Numbers from 2.75 to 5.00

    NASA Technical Reports Server (NTRS)

    Dennis, David H; Cunningham, Bernard E

    1952-01-01

    Results of tests to determine the aerodynamic forces and moments on bodies of revolution at angles of attack from 0 degrees to 25 degrees are presented and compared with theory. Cones and ogives of fineness ratios 3 to 7 and two blunt-nosed body shapes with fineness ratios 3 and 5 were tested at Mach numbers from 2.75 to 5.00. Reynolds numbers were from 0.5 million to 6.4 million, depending on Mach number and body fineness ratio.

  2. Space shuttle: Aerodynamic stability, control effectiveness and drag characteristics of a shuttle orbiter configuration at Mach numbers from 0.6 to 4.96

    NASA Technical Reports Server (NTRS)

    Ramsey, P. E.

    1972-01-01

    Experimental aerodynamic investigations were conducted in the NASA/MSFC 14-inch Trisonic Wind Tunnel from Sept. 27 to Oct. 7, 1972 on a 0.004 scale model of the NR ATP baseline shuttle orbiter configuration. Six component aerodynamic force and moment data were recorded at 0 deg sideslip angle over an angle of attack range from 0 to 20 deg for Mach numbers of 0.6 to 4.96, 20 to 40 deg for Mach numbers of 0.6, 0.9, 2.99, and 4.96, and 40 to 60 deg for Mach numbers of 2.99 and 4.96. Data were obtained over a sideslip range of -10 to 10 deg at 0, 10, and 20 deg angles of attack over the Mach range and 30 and 50 deg at Mach numbers of 2.99 and 4.96. The purpose of the test was to define the buildup, performance, stability, and control characteristics of the orbiter configuration. The model parameters, were: body alone; body-wing; body-wing-tail; elevon deflections of 0, 10, -20, and -40 deg both full and split); aileron deflections of plus or minus 10 deg (full and split); rudder flares of 10 and 40 deg, and a rudder deflection of 15 deg about the 10 and 40 deg flare positions.

  3. Radio Bridge Structure and Its Application to Estimate the Mach Number and Ambient Gas Temperature of Powerful Sources

    NASA Astrophysics Data System (ADS)

    Wellman, Greg F.; Daly, Ruth A.; Wan, Lin

    1997-05-01

    little reacceleration of relativistic electrons within the radio bridge and that the backflow velocity of relativistic plasma within the bridge is small compared with the lobe advance velocity. These results are consistent with implications based on the bridge shape and structure discussed by Alexander & Leahy since we consider only very powerful FR II sources here. The Mach number with which the radio lobe propagates into the ambient medium can be estimated using the structure of the radio bridge; this Mach number is the ratio of the lobe propagation velocity to the sound speed of the ambient gas. The lateral expansion of the bridge is driven initially by a blast wave. When the velocity of the blast wave falls to a value of the order of the sound speed of the ambient medium, the character of the expansion changes, and the functional form of the bridge width as a function of position exhibits a break, which may be used to estimate the ratio of the lobe advance velocity to the sound speed of the ambient gas. We observe this break in several sources studied here. The Mach number of lobe advance depends only upon the ratio of the width to the length of the bridge as a function of position, which is purely geometric. Typical Mach numbers obtained range from about 2 to 10 and seem to be roughly independent of redshift and the total size (core-lobe separation) of the radio source. The Mach number can be used to estimate the temperature of the ambient gas if an independent estimate of the lobe propagation velocity is available. Lobe propagation velocities estimated using the effects of synchrotron and inverse Compton aging of the relativistic electrons that produce the radio emission are combined with the Mach numbers in order to estimate ambient gas temperatures. The temperature obtained for Cygnus A matches that indicated by X-ray data for this source. Typical temperatures obtained range from about 1 to 20 keV. This temperature is characteristic of gas in clusters of galaxies

  4. Aerodynamic characteristics of an all-body hypersonic aircraft configuration at Mach numbers from 0.65 to 10.6

    NASA Technical Reports Server (NTRS)

    Nelms, W. P., Jr.; Thomas, C. L.

    1971-01-01

    Aerodynamic characteristics of a model designed to represent an all body, hypersonic cruise aircraft are presented for Mach numbers from 0.65 to 10.6. The configuration had a delta planform with an elliptic cone forebody and an afterbody of elliptic cross section. Detailed effects of varying angle of attack (-2 to +15 deg), angle of sideslip (-2 to +8 deg), Mach number, and configuration buildup were considered. In addition, the effectiveness of horizontal tail, vertical tail, and canard stabilizing and control surfaces was investigated. The results indicate that all configurations were longitudinally stable near maximum lift drag ratio. The configurations with vertical tails were directionally stable at all angles of attack. Trim penalties were small at hypersonic speeds for a center of gravity location representative of the airplane, but because of the large rearward travel of the aerodynamic center, trim penalties were severe at transonic Mach numbers.

  5. Effect of wall suction on performance of a short annular diffuser at inlet Mach numbers up to 0.5. [gas turbine engines

    NASA Technical Reports Server (NTRS)

    Juhasz, A. J.

    1975-01-01

    A short annular diffuser equipped with wall bleed (suction)capability was evaluated at inlet Mach numbers of 0.186 to 0.5. The diffuser had an area ratio of 4.0 and a length-to-inlet height ratio of 1.6. Test results show that the exit velocity profiles, typical of annular jet flow without suction, could be considerably flattened by application of wall suction. This improved performance was also reflected in diffuser effectiveness (static-pressure recovery) and total-pressure loss results. At the inlet Mach number of 0.5 diffuser static-pressure recovery is equal to or better than at lower inlet Mach numbers for comparable suction rates.

  6. Effect of inlet-air humidity, temperature, pressure, and reference Mach number on the formation of oxides of nitrogen in a gas turbine combustor

    NASA Technical Reports Server (NTRS)

    Marchionna, N. R.; Diehl, L. A.; Trout, A. M.

    1973-01-01

    Tests were conducted to determine the effect of inlet air humidity on the formation of oxides of nitrogen (NOx) from a gas turbine combustor. Combustor inlet air temperature ranged from 506 K (450 F) to 838 K (1050 F). The tests were primarily run at a constant pressure of 6 atmospheres and reference Mach number of 0.065. The NOx emission index was found to decrease with increasing inlet air humidity at a constant exponential rate: NOx = NOx0e-19H (where H is the humidity and the subscript 0 denotes the value at zero humidity). the emission index increased exponentially with increasing normalized inlet air temperature to the 1.14 power. Additional tests made to determine the effect of pressure and reference Mach number on NOx showed that the NOx emission index varies directly with pressure to the 0.5 power and inversely with reference Mach number.

  7. The Longitudinal Aerodynamic Characteristics of a Sweptback Wing-Body Combination With and Without End Plates at Mach Numbers from 0.40 to 0.93

    NASA Technical Reports Server (NTRS)

    Henderson, William P.

    1960-01-01

    An investigation was made at high subsonic speeds in the Langley high-speed 7- by 10-foot tunnel to determine the effect of end plates on the longitudinal aerodynamic characteristics of a sweptback wing-body combination with and without drooped chord-extensions. The wing had 45 deg sweepback of the quarter-chord line, an aspect ratio of 4, a taper ratio of 0.3, and NACA 65AO06 airfoil sections parallel to the plane of symmetry, and was mounted near the rear of a body of revolution having a fineness ratio of approximately 8. The results indicated that the addition of the end plates to either the wing with drooped chord-extensions or to the wing without drooped chord-extensions slightly increased the lift in the low angle-of-attack range but slightly decreased the lift at moderate and high angles of attack. The addition of the end plates to the wing without the chord-extensions caused a small increase in the maximum lift-drag ratio at Mach numbers below 0.65 and a slight decrease at the higher Mach numbers; however, for the addition of the end plates to the wing with the chord- extensions the maximum lift-drag ratio was slightly decreased below a Mach number of 0.88, while a slight increase occurred for the higher Mach numbers. The addition of the end plates to the wings with and without the chord-extensions caused the static longitudinal stability to increase considerably for all Mach numbers; however, only a slight reduction in the aerodynamic-center variation with Mach number was observed.

  8. Parametric Study of Afterbody/nozzle Drag on Twin Two-dimensional Convergent-divergent Nozzles at Mach Numbers from 0.60 to 1.20

    NASA Technical Reports Server (NTRS)

    Pendergraft, Odis C., Jr.; Burley, James R., II; Bare, E. Ann

    1986-01-01

    An investigation has been conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of upper and lower external nozzle flap geometry on the external afterbody/nozzle drag of nonaxisymmetric two-dimensional convergent-divergent exhaust nozzles having parallel external sidewalls installed on a generic twin-engine, fighter-aircraft model. Tests were conducted over a Mach number range from 0.60 to 1.20 and over an angle-of-attack range from -5 to 9 deg. Nozzle pressure ratio was varied from jet off (1.0) to approximately 10.0, depending on Mach number.

  9. Turbulent boundary-layer velocity profiles on a nonadiabatic at Mach number 6.5

    NASA Technical Reports Server (NTRS)

    Keener, E. R.; Hopkins, E. J.

    1972-01-01

    Velocity profiles were obtained from pitot-pressure and total-temperature measurements within a turbulent boundary layer on a large sharp-edged flat plate. Momentum-thickness Reynolds number ranged from 2590 to 8860 and wall-to-adiabatic-wall temperature ratios ranged from 0.3 to 0.5. Measurements were made both with and without boundary layer trips. Five methods are evaluated for correlating the measured velocity profiles with the incompressible law-of-the-wall and the velocity defect law. The mixing-length generalization of Van Driest gives the best correlation.

  10. Control of a high Reynolds number Mach 0.9 heated jet using plasma actuators

    SciTech Connect

    Kearney-Fischer, M.; Kim, J.-H.; Samimy, M.

    2009-09-15

    The results of particle image velocimetry (PIV) measurements in a high subsonic, heated, jet forced using localized arc filament plasma actuators (LAFPAs) show that LAFPAs can consistently produce significant mixing enhancement over a wide range of temperatures. These actuators have been used successfully in high Reynolds number, high-speed unheated jets. The facility consists of an axisymmetric jet with different nozzle blocks of exit diameter of 2.54 cm and variable jet temperature in an anechoic chamber. The focus of this paper is on a high subsonic (M{sub j}=0.9) jet. Twelve experiments with various forcing azimuthal modes (m=0, 1, and {+-}1) and temperatures (T{sub o}/T{sub a}=1.0, 1.4, and 2.0) at a fixed forcing Strouhal number (St{sub DF}=0.3) have been conducted and PIV results compared with the baseline results to characterize the effectiveness of LAFPAs for mixing enhancement. Centerline velocity and turbulent kinetic energy as well as jet width are used for determining the LAFPAs' effectiveness. The characteristics of large-scale structures are analyzed through the use of Galilean streamlines and swirling strength. Across the range of temperatures collected, the effectiveness of LAFPAs improves as temperature increases. Possible reasons for the increase in effectiveness are discussed.

  11. Observational Evidence for High-Mach Number Regime of Coronal Shock Waves During Powerful Solar Particle Events

    NASA Astrophysics Data System (ADS)

    Rouillard, A. P.; Illya, P.; Zucca, P.; Tylka, A. J.; Vainio, R. O.; Vourlidas, A.

    2015-12-01

    Identifying the physical mechanisms that produce the most energetic particles is a long-standing observational and theoretical challenge in astrophysics. Strong shock waves have been proposed as efficient accelerators both in the solar physics and astrophysical contexts via various acceleration mechanisms. The proposed processes rely on shock waves being super-critical or moving several times faster than the characteristic speed of the medium they propagate through (a high MA). Using recent imaging of the NASA STEREO, SOHO and SDO spacecraft, we provide the first observations of the time-dependent 3-dimensional distribution of the expansion speed and MA of a coronal shock wave. These observations show that the high-energy particles measured near Earth are produced at the time of the sharp rise in the shock Mach number (>10) magnetically connected to Earth. These findings provide direct evidence to energetic particles being accelerated during the formation of a strong coronal shock. Using our new technique, we study the longitudinal spread and timing of a number of other energetic particle events during cycle 24.

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

    NASA Technical Reports Server (NTRS)

    Kelly, Thomas C.

    1961-01-01

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

  13. An experimental/computational study of heat transfer in sharp fin induced shock wave/turbulent boundary layer interactions at low hypersonic Mach numbers

    NASA Astrophysics Data System (ADS)

    Rodi, Patrick Elroy

    1992-01-01

    A combined experimental and computational study has been performed of sharp fin induced shock wave/turbulent boundary layer interactions at Mach numbers of 5, 6, and 11. New experimental data were obtained at Mach 5 and include mean surface heat transfer and pressure distributions and surface flow visualization for fin angles of attack of 6, 8, 10, 12, 14 and 16 degrees. Detailed heat transfer measurements were also taken radially at 8 and 16 degrees to study the flowfield's conical nature. Conical Navier-Stokes calculations have been performed using the Baldwin/Lomax turbulence model. Computations were made for two angles of attack at each of the three Mach numbers. The Mach 6 and 11 data used for comparison with the computations, were from the earlier experimental studies of Law and Holden. Careful evaluation of the performance of this numerical approach has been carried out with an emphasis on the surface heat transfer predictions. The new experimental results are described in detail and compared with existing empirical correlations. Comparisons of the experimental data with the computations reveal that the conical Navier-Stokes/Baldwin-Lomax approach underpredicts the physical extent of the interactions for the lower two Mach numbers. However, this trend is reversed at Mach 11. The numerical results overpredict peak heat transfer, although the outer scaling of the Baldwin-Lomax model prevent a grid independent solution. The computed flowfields reveal a large primary vortex located adjacent to the surface just beneath the inviscid shock wave and a smaller corner vortex located very near the fin/surface junction.

  14. Investigation at Mach Numbers of 0.20 to 3.50 of Blended Wing-Body Combinations of Sonic Design with Diamond, Delta, and Arrow Plan Forms

    NASA Technical Reports Server (NTRS)

    Holdaway, George H.; Mellenthin, Jack A.

    1960-01-01

    The models had aspect-ratio-2 diamond, delta, and arrow wings with the leading edges swept 45.00 deg, 59.04 deg, and 70.82 deg, respectively. The wing sections were computed by varying the section shape along with the body radii (blending process) to match the prescribed area distribution and wing plan form. The wing sections had an average value of maximum thickness ratio of about 4 percent of the local chords in a streamwise direction. The models were tested with transition fixed at Reynolds numbers of about 4,000,000 to 9,000,0000, based on the mean aerodynamic chord of the wings. The effect of varying Reynolds number was checked at both subsonic and supersonic speeds. The diamond model was superior to the other plan forms at transonic speeds ((L/D)max = 11.00 to 9.52) because of its higher lift-curve slope and near optimum wave drag due to the blending process. For the wing thickness tested with the diamond model, the marked body and wing contouring required for transonic conditions resulted in a large wave-drag penalty at the higher supersonic Mach numbers where the leading and trailing edges of the wing were supersonic. Because of the low sweep of the trailing edge of the delta model, this configuration was less adaptable to the blending process. Removing a body bump prescribed by the Mach number 1.00 design resulted in a good supersonic design. This delta model with 10 percent less volume was superior to the other plan forms at Mach numbers of 1.55 to 2.35 ((L/D)max = 8.65 to 7.24), but it and the arrow model were equally good at Mach numbers of 2.50 to 3.50 ((L/D)max - 6.85 to O.39). At transonic speeds the arrow model was inferior because of the reduced lift-curve slope associated with its increased sweep and also because of the wing base drag. The wing base-drag coefficients of the arrow model based on the wing planform area decreased from a peak value of 0.0029 at Mach number 1.55 to 0.0003 at Mach number 3.50. Linear supersonic theory was satisfactory

  15. Inviscid Flow Computations of the Orbital Sciences X-34 Over a Mach Number Range of 1.25 to 6.0

    NASA Technical Reports Server (NTRS)

    Prabhu, Ramadas K.

    2001-01-01

    This report documents the results of an inviscid computational study conducted on the Orbital Sciences X-34 vehicle to compute its inviscid longitudinal aerodynamic characteristics over a Mach number range of 1.25 to 6.0. The unstructured grid software FELISA was used and th e aerodynamic characteristics were computed at Mach numbers 1.25, 1.6, 2.5, 4.0, 4.63, and 6.0, and an angle of attack range of -4 to 32 degrees. These results were compared with available aerodynamic data from wind tunnel test on X-34 models. The comparison showed excellent agreement in C(sub N). The computed pitching moment compared well at Mach numbers 2.5 and higher, and at angles of attack of up to 12 deg. The agreement was not good at higher angles of attack possibly due to viscous effects. At lower Mach numbers there were significant differences between computed and measured C(sub m) values. This could not be explained. Since the present computations are inviscid, the computed C(sub A) was consistently lower than the measured values as expected.

  16. Exploratory wind tunnel tests of a shock-swallowing air data sensor at a Mach number of approximately 1.83

    NASA Technical Reports Server (NTRS)

    Nugent, J.; Couch, L. M.; Webb, L. D.

    1975-01-01

    The test probe was designed to measure free-stream Mach number and could be incorporated into a conventional airspeed nose boom installation. Tests were conducted in the Langley 4-by 4-foot supersonic pressure tunnel with an approximate angle of attack test range of -5 deg to 15 deg and an approximate angle of sideslip test range of + or - 4 deg. The probe incorporated a variable exit area which permitted internal flow. The internal flow caused the bow shock to be swallowed. Mach number was determined with a small axially movable internal total pressure tube and a series of fixed internal static pressure orifices. Mach number error was at a minimum when the total pressure tube was close to the probe tip. For four of the five tips tested, the Mach number error derived by averaging two static pressures measured at horizontally opposed positions near the probe entrance were least sensitive to angle of attack changes. The same orifices were also used to derive parameters that gave indications of flow direction.

  17. Effect of variation of length-to-depth ratio and Mach number on the performance of a typical double cavity scramjet combustor

    NASA Astrophysics Data System (ADS)

    Mahto, Navin Kumar; Choubey, Gautam; Suneetha, Lakka; Pandey, K. M.

    2016-11-01

    The two equation standard k-ɛ turbulence model and the two-dimensional compressible Reynolds-Averaged Navier-Stokes (RANS) equations have been used to computationally simulate the double cavity scramjet combustor. Here all the simulations are performed by using ANSYS 14-FLUENT code. At the same time, the validation of the present numerical simulation for double cavity has been performed by comparing its result with the available experimental data which is in accordance with the literature. The results are in good agreement with the schlieren image and the pressure distribution curve obtained experimentally. However, the pressure distribution curve obtained numerically is under-predicted in 5 locations by numerical calculation. Further, investigations on the variations of the effects of the length-to-depth ratio of cavity and Mach number on the combustion characteristics has been carried out. The present results show that there is an optimal length-to-depth ratio for the cavity for which the performance of combustor significantly improves and also efficient combustion takes place within the combustor region. Also, the shifting of the location of incident oblique shock took place in the downstream of the H2 inlet when the Mach number value increases. But after achieving a critical Mach number range of 2-2.5, the further increase in Mach number results in lower combustion efficiency which may deteriorate the performance of combustor.

  18. A critical shock mach number for particle acceleration in the absence of pre-existing cosmic rays: M=√5

    SciTech Connect

    Vink, Jacco

    2014-01-10

    It is shown that, under some generic assumptions, shocks cannot accelerate particles unless the overall shock Mach number exceeds a critical value M>√5. The reason is that for M≤√5 the work done to compress the flow in a particle precursor requires more enthalpy flux than the system can sustain. This lower limit applies to situations without significant magnetic field pressure. In case that the magnetic field pressure dominates the pressure in the unshocked medium, i.e., for low plasma beta, the resistivity of the magnetic field makes it even more difficult to fulfill the energetic requirements for the formation of shock with an accelerated particle precursor and associated compression of the upstream plasma. We illustrate the effects of magnetic fields for the extreme situation of a purely perpendicular magnetic field configuration with plasma beta β = 0, which gives a minimum Mach number of M = 5/2. The situation becomes more complex, if we incorporate the effects of pre-existing cosmic rays, indicating that the additional degree of freedom allows for less strict Mach number limits on acceleration. We discuss the implications of this result for low Mach number shock acceleration as found in solar system shocks, and shocks in clusters of galaxies.

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

    NASA Technical Reports Server (NTRS)

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

    1950-01-01

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

  20. Wind tunnel test results for the direction controlled antitank DCAT missile at Mach numbers from 0.64 to 2.50

    NASA Technical Reports Server (NTRS)

    Martin, T. A.; Spring, D. J.

    1973-01-01

    Wind tunnel test results are presented to show aerodynamic characteristics over the Mach number range of 0.64 to 2.50 of the DCAT missile. Data are presented showing the interference created by the rear mounted reaction control system. Two candidate fins were installed on the model during tests: a flat folding fin and a curved wrap around fin.

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

    NASA Technical Reports Server (NTRS)

    Grow, R. Bruce; Preisser, John S.

    1971-01-01

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

  2. Effects of leading edge sweep angle and design lift coefficient on performance of a modified arrow wing at a design Mach number of 2.6

    NASA Technical Reports Server (NTRS)

    Mack, R. J.

    1974-01-01

    Wing models were tested in the high-speed section of the Langley Unitary Plan wind tunnel to study the effects of the leading-edge sweep angle and the design lift coefficient on aerodynamic performance and efficiency. The models had leading-edge sweep angles of 69.44 deg, 72.65 deg, and 75.96 deg which correspond to values of the design Mach-number-sweep-angle parameter (beta cotangent A) sub DES of 0.6, 0.75, and 0.9, respectively. For each sweep angle, camber surfaces having design lift coefficients of 0,0.08, and 0.12 at a design Mach number of 2.6 were generated. The wind-tunnel tests were conducted at Mach numbers of 2.3, 2.6, and 2.96 with a stagnation temperature of 338.7 K (150 F) and a Reynolds number per meter of 9.843 times 10 to the 6th power. The results of the tests showed that only a moderate sweeping of the wing leading edge aft of the Mach line along with a small-to-moderate amount of camber and twist was needed to significantly improve the zero-lift (flat camber surface) wing performance and efficiency.

  3. Internal pressure distributions for a two-dimensional thrust-reversing nozzle operating at a free-stream Mach number of zero

    NASA Technical Reports Server (NTRS)

    Putnam, L. E.; Strong, E. G.

    1983-01-01

    An investigation was conducted in the static test facility of the Langley 16-Foot Transonic Tunnel to measure static pressure distributions inside a nonaxisymmetric thrust reversing nozzle. The tests were made at nozzle total pressures ranging from ambient to about eight times ambient pressure at a free stream Mach number of zero. Tabulated pressure data are presented.

  4. The Effect of the Inlet Mach Number and Inlet-boundary-layer Thickness on the Performance of a 23 Degree Conical-diffuser-tail-pipe Combination

    NASA Technical Reports Server (NTRS)

    Persh, Jerome

    1950-01-01

    An investigation was conducted to determine the effect of the inlet Mach number and entrance-boundary-layer thickness on the performance of a 23 degree 21-inch conical-diffuser - tail-pipe combination with a 2:1 area ratio. The air flows used in this investigation covered an inlet Mach number range from 0.17 to 0.89 and corresponding Reynolds numbers of 1,700,000 to 7,070,000. Results are reported for two inlet-boundary-layer thicknesses. Over the entire range of flows, the mean value of the inlet displacement thickness is about 0.034 inch for the thinner inlet boundary layer and about 0.170 inch for the case of the thicker inlet boundary layer. The performance of the diffuser - tail-pipe combination is presented together with examples of longitudinal static-pressure distribution and the results of boundary-layer pressure surveys made at six points along the diffuser wall. The results indicated a progressive diminution of the static-pressure recovery and a steady increase in the total-pressure losses as the inlet Mach number was increased for both inlet-boundary-layer thicknesses. The ratio of actual static-pressure rise to that theoretically possible was much less and the total-pressure losses were greater for the case of the thicker inlet boundary layer throughout the speed range investigated. With the thinner inlet boundary layer, flow separation occurred at the diffuser exit at all inlet Mach numbers.Unseparated flow alternating with separated flow was observed near the inlet at the higher velocities. For the case of the thicker inlet boundary layer, the origin of the separated region occurred in the vicinity of the inlet-duct-diffuser junction section at all Mach numbers.

  5. Global Alfven modes: Theory and experiment

    SciTech Connect

    Turnbull, A.D.; Strait, E.J.; Heidbrink, W.W.; Chu, M.S.; Duong, H.H.; Greene, J.M.; Lao, L.L.; Taylor, T.S.; Thompson, S.J. )

    1993-07-01

    It is shown that the theoretical predictions and experimental observations of toroidicity-induced Alfven eigenmodes (TAE's) are now in good agreement, with particularly detailed agreement in the mode frequencies. Calculations of the driving and damping rates predict the importance of continuum damping for low toroidal mode numbers and this is confirmed experimentally. However, theoretical calculations in finite-[beta], shaped discharges predict the existence of other global Alfven modes, in particular the ellipticity-induced Alfven eigenmode (EAE) and a new mode, the beta-induced Alfven eigenmode (BAE). The BAE mode is calculated to be in or below the same frequency range as the TAE mode and may contribute to the experimental observations at high [beta]. Experimental evidence and complementary analyses are presented confirming the presence of the EAE mode at higher frequencies.

  6. Design and experimental investigation of a single-stage turbine with a rotor entering relative Mach number of 2

    NASA Technical Reports Server (NTRS)

    Moffitt, Thomas P

    1958-01-01

    The design and experimental investigation of a single-stage supersonic turbine are presented herein. The turbine was designed for a rotor entering relative Mach number of 2. The maximum equivalent specific work output of the turbine at design speed and approximately design over-all pressure ratio was 32.9 BTU per pound at a static efficiency of 0.414. This static efficiency gave good verification to an independent reference that indicated theoretical static efficiencies for similar single-stage turbines within the range 0.40 to 0.45. An experimental ratio of effective rotor blade momentum thickness to mean camber length was determined to be 0.0014, which compares favorably with the results obtained from several transonic and subsonic turbines. The design procedure for this turbine would have been improved by allowing for more rotor losses by assuming a value of this momentum parameter comparable with those obtained from transonic turbines. Removing a large portion of the rotor suction surface enabled a lower static pressure to be felt at the stator exit, at the expense of higher rotor losses. The net result was an improvement in turbine work output of about 3 percent at design setting conditions. No problems associated with supersonic starting were encountered even under the worst conditions of turbine operation with respect to this problem.

  7. Analysis of an inviscid zero-Mach number system in endpoint Besov spaces for finite-energy initial data

    NASA Astrophysics Data System (ADS)

    Fanelli, Francesco; Liao, Xian

    2015-11-01

    The present paper is the continuation of work [18], devoted to the study of an inviscid zero-Mach number system in the framework of endpoint Besov spaces of type B∞,rs (Rd), r ∈ [ 1, ∞ ], d ≥ 2, which can be embedded in the Lipschitz class C 0, 1. In particular, the largest case B∞,11 and the case of Hölder spaces C 1, α are taken into account. The local in time well-posedness of this system is proved, under an additional finite-energy hypothesis on the initial data. The key to get this result is new a priori estimates for parabolic equations with variable coefficients in endpoint spaces B∞,rs (Rd), which are of independent interest. In the special case of space dimension d = 2, we are able to give a lower bound for the lifespan, such that the solutions tend to be globally defined when the initial inhomogeneity is small. There, we will show refined a priori estimates in endpoint Besov spaces for transport equations.

  8. Concurrent identification of aero-acoustic scattering and noise sources at a flow duct singularity in low Mach number flow

    NASA Astrophysics Data System (ADS)

    Sovardi, Carlo; Jaensch, Stefan; Polifke, Wolfgang

    2016-09-01

    A numerical method to concurrently characterize both aeroacoustic scattering and noise sources at a duct singularity is presented. This approach combines Large Eddy Simulation (LES) with techniques of System Identification (SI): In a first step, a highly resolved LES with external broadband acoustic excitation is carried out. Subsequently, time series data extracted from the LES are post-processed by means of SI to model both acoustic propagation and noise generation. The present work studies the aero-acoustic characteristics of an orifice placed in a duct at low flow Mach numbers with the "LES-SI" method. Parametric SI based on the Box-Jenkins mathematical structure is employed, with a prediction error approach that utilizes correlation analysis of the output residuals to avoid overfitting. Uncertainties of model parameters due to the finite length of times series are quantified in terms of confidence intervals. Numerical results for acoustic scattering matrices and power spectral densities of broad-band noise are validated against experimental measurements over a wide range of frequencies below the cut-off frequency of the duct.

  9. Experimental investigation of low Mach number flow past a rectangular cavity using dual-camera Cinematographic PIV system

    NASA Astrophysics Data System (ADS)

    Bian, Shiyao; Driscoll, James

    2005-11-01

    Flow past cavity has been of interest due to its geometrical simplicity and complex flow characteristics. A dual-camera Cinematographic Particle Image Velocimetry (CPIV) system has been developed to study low Mach number flow over a rectangular cavity. This system consists of two high-repetition rate Nd:YAG lasers and two high-speed CMOS cameras registered to have sub-pixel alignment errors. A rectangular cavity with a length-to-depth ratio of 2 was mounted in the test section of a recirculating water tunnel providing free-stream flow speeds between 5˜26 m/s. Consecutive CPIV images with a spatial resolution of 1632 x 800 pixels and 20 μs time delay were obtained at frame rate of 1.5 KHz. Time traces of surface pressures at the bottom of the cavity are acquired simultaneously by using flush-mounted dynamic pressure transducers. The temporal evolution of velocity and vortical fields reveals the time-dependence of the mixing and mass transport between the shear layer and the cavity. The simultaneous velocity and pressure measurements also show the unsteady interaction between vortical structures and the trailing edge of the cavity under resonating and non-resonating conditions. [Sponsored by National Science Foundation Grant: CTM 0203140

  10. Inviscid Flow Computations of Two '07 Mars Lander Aeroshell Configurations Over a Mach Number Range of 2 to 24

    NASA Technical Reports Server (NTRS)

    Prabhu, Ramadas K.

    2001-01-01

    This report documents the results of an inviscid computational study conducted on two aeroshell configurations for a proposed '07 Mars Lander. The aeroshell configurations are asymmetric due to the presence of the tabs at the maximum diameter location. The purpose of these tabs was to change the pitching moment characteristics so that the aeroshell will trim at a non-zero angle of attack and produce a lift-to-drag ratio of approximately -0.25. This is required in the guidance of the vehicle on its trajectory. One of the two configurations is called the shelf and the other is called the tab. The unstructured grid software FELISA with the equilibrium Mars gas option was used for these computations. The computations were done for six points on a preliminary trajectory of the '07 Mars Lander at nominal Mach numbers of 2, 3, 5, 10, 15, and 24. Longitudinal aerodynamic characteristics namely lift, drag, and pitching moment were computed for 10, 15, and 20 degrees angles of attack. The results indicated that the two configurations have aerodynamic characteristics that have very similar aerodynamic characteristics, and provide the desired trim LID of approximately -0.25.

  11. Measurements of Free-Space Oscillating Pressures Near Propellers at Flight Mach Numbers to 0.72

    NASA Technical Reports Server (NTRS)

    Kurbjun, Max C; Vogeley, Arthur W

    1958-01-01

    In the course of a short flight program initiated to check the theory of Garrick and Watkins (NACA rep. 1198), a series of measurements at three stations were made of the oscillating pressures near a tapered-blade plan-form propeller and rectangular-blade plan form propeller at flight Mach numbers up to 0.72. In contradiction to the results for the propeller studied in NACA rep. 1198, the oscillating pressures in the plane ahead of the propeller were found to be higher than those immediately behind the propeller. Factors such as variation in torque and thrust distribution, since the blades of the present investigation were operating above their design forward speed, may account for this contradiction. The effect of blade plan form shows that a tapered-blade plan-form propeller will produce lower sound-pressure levels than a rectangular-blade plan-form propeller for the low blade-passage harmonics (the frequencies where structural considerations are important) and produce higher sound-pressure levels for the higher blade-passage harmonics (frequencies where passenger comfort is important).

  12. Effect of entry-lip design on aerodynamics and acoustics of high throat Mach number inlets for the quiet, clean, short-haul experimental engine

    NASA Technical Reports Server (NTRS)

    Miller, B. A.; Dastoli, B. J.; Wesoky, H. L.

    1975-01-01

    Results of scale model tests of high-throat-Mach-number inlets designed to suppress inlet-emitted engine machinery noise produced in a V/STOL wind tunnel are presented. A vacuum system was used to induce inlet airflow with a siren as a noise source. Inlet mass flow was 11.68 kilograms (25.75 lb. min) per second at a throat Mach number of 0.79. The effect of entry-lip design (contraction ratio and diameter ratio) on inlet total-pressure recovery, steady-state pressure distortion, performance at high incidence angles, and noise suppression was determined. With proper entry-lip design, total-pressure recovery in excess of 0.988 could be obtained statically at an average throat Mach number of 0.79. Total-pressure distortion was 5 percent. The reduction in the siren tone sound pressure level transmitted through the inlet was 10 to 14 db relative to that measured at throat Mach 0.6.

  13. Interaction of two glancing, crossing shock waves with a turbulent boundary-layer at various Mach numbers

    NASA Technical Reports Server (NTRS)

    Hingst, Warren R.; Williams, Kevin E.

    1991-01-01

    A preliminary experimental investigation was conducted to study two crossing, glancing shock waves of equal strengths, interacting with the boundary-layer developed on a supersonic wind tunnel wall. This study was performed at several Mach numbers between 2.5 and 4.0. The shock waves were created by fins (shock generators), spanning the tunnel test section, that were set at angles varying from 4 to 12 degrees. The data acquired are wall static pressure measurements, and qualitative information in the form of oil flow and schlieren visualizations. The principle aim is two-fold. First, a fundamental understanding of the physics underlying this flow phenomena is desired. Also, a comprehensive data set is needed for computational fluid dynamic code validation. Results indicate that for small shock generator angles, the boundary-layer remains attached throughout the flow field. However, with increasing shock strengths (increasing generator angles), boundary layer separation does occur and becomes progressively more severe as the generator angles are increased further. The location of the separation, which starts well downstream of the shock crossing point, moves upstream as shock strengths are increased. At the highest generator angles, the separation appears to begin coincident with the generator leading edges and engulfs most of the area between the generators. This phenomena occurs very near the 'unstart' limit for the generators. The wall pressures at the lower generator angles are nominally consistent with the flow geometries (i.e. shock patterns) although significantly affected by the boundary-layer upstream influence. As separation occurs, the wall pressures exhibit a gradient that is mainly axial in direction in the vicinity of the separation. At the limiting conditions the wall pressure gradients are primarily in the axial direction throughout.

  14. An improved high-order scheme for DNS of low Mach number turbulent reacting flows based on stiff chemistry solver

    NASA Astrophysics Data System (ADS)

    Yu, Rixin; Yu, Jiangfei; Bai, Xue-Song

    2012-06-01

    We present an improved numerical scheme for numerical simulations of low Mach number turbulent reacting flows with detailed chemistry and transport. The method is based on a semi-implicit operator-splitting scheme with a stiff solver for integration of the chemical kinetic rates, developed by Knio et al. [O.M. Knio, H.N. Najm, P.S. Wyckoff, A semi-implicit numerical scheme for reacting flow II. Stiff, operator-split formulation, Journal of Computational Physics 154 (2) (1999) 428-467]. Using the material derivative form of continuity equation, we enhance the scheme to allow for large density ratio in the flow field. The scheme is developed for direct numerical simulation of turbulent reacting flow by employing high-order discretization for the spatial terms. The accuracy of the scheme in space and time is verified by examining the grid/time-step dependency on one-dimensional benchmark cases: a freely propagating premixed flame in an open environment and in an enclosure related to spark-ignition engines. The scheme is then examined in simulations of a two-dimensional laminar flame/vortex-pair interaction. Furthermore, we apply the scheme to direct numerical simulation of a homogeneous charge compression ignition (HCCI) process in an enclosure studied previously in the literature. Satisfactory agreement is found in terms of the overall ignition behavior, local reaction zone structures and statistical quantities. Finally, the scheme is used to study the development of intrinsic flame instabilities in a lean H2/air premixed flame, where it is shown that the spatial and temporary accuracies of numerical schemes can have great impact on the prediction of the sensitive nonlinear evolution process of flame instability.

  15. PARTICLE ACCELERATION AND WAVE EXCITATION IN QUASI-PARALLEL HIGH-MACH-NUMBER COLLISIONLESS SHOCKS: PARTICLE-IN-CELL SIMULATION

    SciTech Connect

    Kato, Tsunehiko N.

    2015-04-01

    We herein investigate shock formation and particle acceleration processes for both protons and electrons in a quasi-parallel high-Mach-number collisionless shock through a long-term, large-scale, particle-in-cell simulation. We show that both protons and electrons are accelerated in the shock and that these accelerated particles generate large-amplitude Alfvénic waves in the upstream region of the shock. After the upstream waves have grown sufficiently, the local structure of the collisionless shock becomes substantially similar to that of a quasi-perpendicular shock due to the large transverse magnetic field of the waves. A fraction of protons are accelerated in the shock with a power-law-like energy distribution. The rate of proton injection to the acceleration process is approximately constant, and in the injection process, the phase-trapping mechanism for the protons by the upstream waves can play an important role. The dominant acceleration process is a Fermi-like process through repeated shock crossings of the protons. This process is a “fast” process in the sense that the time required for most of the accelerated protons to complete one cycle of the acceleration process is much shorter than the diffusion time. A fraction of the electrons are also accelerated by the same mechanism, and have a power-law-like energy distribution. However, the injection does not enter a steady state during the simulation, which may be related to the intermittent activity of the upstream waves. Upstream of the shock, a fraction of the electrons are pre-accelerated before reaching the shock, which may contribute to steady electron injection at a later time.

  16. Comparisons of a Three-Dimensional, Full Navier Stokes Computer Model with High Mach Number Combuster Test Data

    NASA Technical Reports Server (NTRS)

    Watkins, William B.

    1990-01-01

    Comparisons between scramjet combustor data and a three-dimensional full Navier-Stokes calculation have been made to verify and substantiate computational fluid dynamics (CFD) codes and application procedures. High Mach number scramjet combustor development will rely heavily on CFD applications to provide wind tunnel-equivalent data of quality sufficient to design, build and fly hypersonic aircraft. Therefore. detailed comparisons between CFD results and test data are imperative. An experimental case is presented, for which combustor wall static pressures were measured and flow-fieid interferograms were obtained. A computer model was done of the experiment, and counterpart parameters are compared with experiment. The experiment involved a subscale combustor designed and fabricated for the National Aero-Space Plane Program, and tested in the Calspan Corporation 96" hypersonic shock tunnel. The combustor inlet ramp was inclined at a 20 angle to the shock tunnel nozzle axis, and resulting combustor entrance flow conditions simulated freestream M=10. The combustor body and cowl walls were instrumented with static pressure transducers, and the combustor lateral walls contained windows through which flowfield holographic interferograms were obtained. The CFD calculation involved a three-dimensional time-averaged full Navier-Stokes code applied to the axial flow segment containing fuel injection and combustion. The full Navier-Stokes approach allowed for mixed supersonic and subsonic flow, downstream-upstream communication in subsonic flow regions, and effects of adverse pressure gradients. The code included hydrogen-air chemistry in the combustor segment which begins near fuel injection and continues through combustor exhaust. Combustor ramp and inlet segments on the combustor lateral centerline were modelled as two dimensional. Comparisons to be shown include calculated versus measured wall static pressures as functions of axial flow coordinate, and calculated path

  17. Effects of Sting-Support Diameter on the Base Pressures of an Elliptic Cone at Mach Numbers from 0.60 to 1.40

    NASA Technical Reports Server (NTRS)

    Stivers, Louis S., Jr.; Levy, Lionel L., Jr.

    1961-01-01

    Measurements were made to determine the effects of sting-support diameter on the base pressures of an elliptic cone with ratio of cross-section thickness to width of 1/3 and a plan-form, semi-apex angle of 15 deg. The investigation was made for model angles of attack from -2 deg to +20 deg at Mach numbers from 0.60 to 1.40, and for a constant Reynolds number of 1.4 million, based on the length of the model. The results indicated that the sting interference decreased the base axial-force coefficients by substantial amounts up to a maximum of about one-third the value of the coefficient for no sting interference. There was no practical diameter of the sting for which the effects of the sting on the base pressures would be negligible throughout the Mach number and angle-of-attack ranges of the investigation.

  18. Free-flight Performance of 16-inch-diameter Supersonic Ram-jet Units II : Five Units Designed for Combustion-chamber-inlet Mach Number of 0.16 at Free-stream Mach Number of 1.60 (units B-1, B-2, B-3, B-4, and B-5) /c Wesley E. Messing and Scott H. S

    NASA Technical Reports Server (NTRS)

    Messing, Wesley E; Simpkinson, Scott H

    1950-01-01

    Free-flight performance of five 16-inch-diameter ram-jet units was determined over range of free-stream Mach numbers of 0.50 to 1.86 and gas total-temperature ratios between 1.0 and 6.1 Time histories of performance data are presented for each unit. Correlations illustrate effect of free-stream Mach number and gas total-temperature ratio on diffuser total-pressure recovery, net-thrust coefficient, and external drag coefficient. One unit had smooth steady burning throughout the entire flight and encountered a maximum free-stream Mach number of 1.86 with a net acceleration of approximately 4.2 g's.

  19. Heat transfer on a flat plate in continuum to rarefied hypersonic flows at Mach numbers of 19.2 and 25.4

    NASA Technical Reports Server (NTRS)

    Nagamatsu, H. T.; Sheer, R. E., Jr.

    1981-01-01

    Local heat transfer rates were measured on a flat steel plate (10 in. wide and 16 in. long) with sharp and blunt leading edges (0.001 and 0.010 in.) in the transition from the strong interaction boundary layer regime with no slip at the surface to the free molecule regime. The tests were conducted in a combustion driven hypersonic shock tunnel, with the nominal free stream Mach numbers of 19.2 and 25.4, and a reflected stagnation temperature of approximately 2340 R. The twelve heat transfer gages were made of platinum sputtered on a Pyrex backing to a thickness of approximately 350 A, insulated by a silicone dioxide film. For both Mach numbers the heat transfer data agreed reasonably well with the strong interaction prediction of Li and Nagamatsu (1953, 1955) for unit Reynolds numbers greater than approximately 100,000 and leading edge Knudsen numbers less than approximately 4. At lower density conditions the rarefied flow effects began to dominate the flow phenomena near the leading edge region of the sharp flat plate. A systematic reduction in the heat transfer rate close to the leading edge was observed for both Mach number tests as the leading edge density was reduced and the mean free path was increased.

  20. Wind tunnel investigation of three axisymmetric cowls of different lengths at Mach numbers from 0.60 to 0.92

    NASA Technical Reports Server (NTRS)

    Re, Richard J.; Abeyounis, William K.

    1993-01-01

    Pressure distributions on three inlets having different cowl lengths were obtained in the Langley 16-Foot Transonic Tunnel. The cowl diameter ratio (highlight diameter to maximum diameter) was 0.85 and the cowl length ratios (cowl length to maximum diameter) were 0.337, 0.439, and 0.547. The cowls had identical nondimensionalized (with respect to cowl length) external geometry and identical internal geometry. The internal contraction ratio (highlight area to throat area) was 1.250. The inlets had longitudinal rows of static pressure orifices on the top and bottom (external) surfaces and on the contraction (internal) and diffuser surfaces. The afterbody was cylindrical in shape, and its diameter was equal to the maximum diameter of the cowl. Depending on the cowl configuration and free-stream Mach number, the mass-flow ratio varied between 0.27 and 0.87 during the tests. Angle of attack varied from 0 to 4.1 deg at selected Mach numbers and mass-flow ratios, and the Reynolds number varied with the Mach number from 3.2x10(exp 6) to 4.2x10(exp 6) per foot.

  1. Aerodynamic characteristics of an F-8 aircraft configuration with a variable camber wing at Mach numbers from 0.70 to 1.15

    NASA Technical Reports Server (NTRS)

    Boltz, F. W.; Pena, D. F.

    1977-01-01

    A 0.1-scale model of an F-8 aircraft was tested in the Ames 14-Foot Transonic Wind Tunnel at Mach numbers from 0.7 to 1.15. Angle of attack was varied from -2 deg. to 22 deg. at sideslip angles of 0 deg and -5 deg. Reynolds number, dictated by the atmospheric stagnation pressure, varied with Mach number from 3.4 to 4.0 million based on mean aerodynamic chord. The model was configured with a wing designed to simulate the downward deflection of the leading and trailing edges of an advanced-technology-conformal-variable camber wing. This wing was also equipped with conventional (simple hinge) flaps. In addition, the model was tested with the basic F-8 wing to provide a reference for extrapolating to flight data. In general, at all Mach numbers the use of conformal flap deflections at both the leading edge and trailing edge resulted in slightly higher maximum lift coefficients and lower drag coefficients than with the use of simple hinge flaps. There were also found to be small improvements in the pitching-moment characteristics with the use of conformal flaps.

  2. Experimental aerodynamic characteristics of two V/STOL fighter/attack aircraft configurations at Mach numbers from 0.4 to 1.4

    NASA Technical Reports Server (NTRS)

    Nelms, W. P.; Durston, D. A.; Lummus, J. R.

    1980-01-01

    A wind tunnel test was conducted to measure the aerodynamic characteristics of two horizontal attitude takeoff and landing V/STOL fighter/attack aircraft concepts. In one concept, a jet diffuser ejector was used for the vertical lift system; the other used a remote augmentation lift system (RALS). Wind tunnel tests to investigate the aerodynamic uncertainties and to establish a data base for these types of concepts were conducted over a Mach number range from 0.2 to 2.0. The present report covers tests, conducted in the 11 foot transonic wind tunnel, for Mach numbers from 0.4 to 1.4. Detailed effects of varying the angle of attack (up to 27 deg), angle of sideslip (-4 deg to +8 deg), Mach number, Reynolds number, and configuration buildup were investigated. In addition, the effects of wing trailing edge flap deflections, canard incidence, and vertical tail deflections were explored. Variable canard longitudinal location and different shapes of the inboard nacelle body strakes were also investigated.

  3. Numerical Modeling of Flow Control in a Boundary-Layer-Ingesting Offset Inlet Diffuser at Transonic Mach Numbers

    NASA Technical Reports Server (NTRS)

    Allan, Brian G.; Owens, Lewis R.

    2006-01-01

    This paper will investigate the validation of the NASA developed, Reynolds-averaged Navier-Stokes (RANS) flow solver, OVERFLOW, for a boundary-layer-ingesting (BLI) offset (S-shaped) inlet in transonic flow with passive and active flow control devices as well as a baseline case. Numerical simulations are compared to wind tunnel results of a BLI inlet experiment conducted at the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel. Comparisons of inlet flow distortion, pressure recovery, and inlet wall pressures are performed. The numerical simulations are compared to the BLI inlet data at a free-stream Mach number of 0.85 and a Reynolds number of approximately 2 million based on the fanface diameter. The numerical simulations with and without tunnel walls are performed, quantifying tunnel wall effects on the BLI inlet flow. A comparison is made between the numerical simulations and the BLI inlet experiment for the baseline and VG vane cases at various inlet mass flow rates. A comparison is also made to a BLI inlet jet configuration for varying actuator mass flow rates at a fixed inlet mass flow rate. Overall, the numerical simulations were able to predict the baseline circumferential flow distortion, DPCP avg, very well within the designed operating range of the BLI inlet. A comparison of the average total pressure recovery showed that the simulations were able to predict the trends but had a negative 0.01 offset when compared to the experimental levels. Numerical simulations of the baseline inlet flow also showed good agreement with the experimental inlet centerline surface pressures. The vane case showed that the CFD predicted the correct trends in the circumferential distortion levels for varying inlet mass flow but had a distortion level that was nearly twice as large as the experiment. Comparison to circumferential distortion measurements for a 15 deg clocked 40 probe rake indicated that the circumferential distortion levels are very sensitive to the symmetry of

  4. Numerical Modeling of Flow Control in a Boundary-Layer-Ingesting Offset Inlet Diffuser at Transonic Mach Numbers

    NASA Technical Reports Server (NTRS)

    Allan Brian G.; Owens, Lewis, R.

    2006-01-01

    This paper will investigate the validation of a NASA developed, Reynolds-averaged Navier-Stokes (RANS) flow solver, OVERFLOW, for a boundary-layer-ingesting (BLI) offset (S-shaped) inlet in transonic flow with passive and active flow control devices as well as the baseline case. Numerical simulations are compared to wind tunnel results of a BLI inlet conducted at the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel. Comparisons of inlet flow distortion, pressure recovery, and inlet wall pressures are performed. The numerical simulations are compared to the BLI inlet data at a freestream Mach number of 0.85 and a Reynolds number of approximately 2 million based on the length of the fan-face diameter. The numerical simulations with and without wind tunnel walls are performed, quantifying effects of the tunnel walls on the BLI inlet flow measurements. The wind tunnel test evaluated several different combinations of jet locations and mass flow rates as well as a vortex generator (VG) vane case. The numerical simulations will be performed on a single jet configuration for varying actuator mass flow rates at a fix inlet mass flow condition. Validation of the numerical simulations for the VG vane case will also be performed for varying inlet mass flow rates. Overall, the numerical simulations were able to predict the baseline circumferential flow distortion, DPCPavg, very well for comparisons made within the designed operating range of the BLI inlet. However the CFD simulations did predict a total pressure recovery that was 0.01 lower than the experiment. Numerical simulations of the baseline inlet flow also showed good agreement with the experimental inlet centerline surface pressures. The vane case showed that the CFD predicted the correct trends in the circumferential distortion for varying inlet mass flow but had a distortion level that was nearly twice as large as the experiment. Comparison to circumferential distortion measurements for a 15 deg clocked 40 probe

  5. Free-flight Performance of a Rocket-boosted, Air-launched 16-inch-diameter Ram-jet Engine at Mach Numbers up to 2.20

    NASA Technical Reports Server (NTRS)

    Disher, John H; Kohl, Robert C; Jones, Merle L

    1953-01-01

    The investigation of air-launched ram-jet engines has been extended to include a study of models with a nominal design free-stream Mach number of 2.40. These models require auxiliary thrust in order to attain a flight speed at which the ram jet becomes self-accelerating. A rocket-boosting technique for providing this auxiliary thrust is described and time histories of two rocket-boosted ram-jet flights are presented. In one flight, the model attained a maximum Mach number of 2.20 before a fuel system failure resulted in the destruction of the engine. Performance data for this model are presented in terms of thrust and drag coefficients, diffuser pressure recovery, mass-flow ratio, combustion efficiency, specific fuel consumption, and over-all engine efficiency.

  6. Longitudinal aerodynamic characteristics of an elliptical body with a horizontal tail at Mach numbers from 2.3 to 4.63

    NASA Technical Reports Server (NTRS)

    Shrout, B. L.; Robins, A. W.

    1982-01-01

    Longitudinal aerodynamic characteristics of a configuration consisting of an elliptical body with an in plane horizontal tail were investigated. The tests were conducted at Mach numbers of 2.3, 2.96, 4.0, and 4.63. In some cases, the configuration with negative tail deflections yielded higher values of maximum lift drag ratio than did the configuration with an undeflected tail. This was due to body upwash acting on the tail and producing an additional lift increment with essentially no drag penalty. Linear theory methods used to estimate some of the longitudinal aerodynamic characteristics of the model yielded results which compared well with experimental data for all Mach numbers in this investigation and for both small angles of attack and larger angles of attack where nonlinear (vortex) flow phenomena were present.

  7. Wind tunnel investigation of nacelle-airframe interference at Mach numbers of 0.9 to 1.4 - pressure data, volume 1

    NASA Technical Reports Server (NTRS)

    Bencze, D. P.

    1976-01-01

    Detailed interference force and pressure data were obtained on a representative wing-body nacelle combination at Mach numbers of 0.9 to 1.4. The model consisted of a delta wing-body aerodynamic force model with four independently supported nacelles located beneath the wing-body combination. The model was mounted on a six component force balance, and the left hand wing was pressure instrumented. Each of the two right hand nacelles was mounted on a six component force balance housed in the thickness of the nacelle, while each of the left hand nacelles was pressure instrumented. The primary variables examined included Mach number, angle of attack, nacelle position, and nacelle mass flow ratio. Nacelle axial location, relative to both the wing-body combination and to each other, was the most important variable in determining the net interference among the components.

  8. Exploratory investigation at Mach numbers from 0.40 to 0.95 of the effects of jets blown over a wing

    NASA Technical Reports Server (NTRS)

    Putnam, L. E.

    1973-01-01

    An exploratory investigation has been made at Mach numbers from 0.40 to 0.95 to determine the effects on lift, drag, and pitching moment of blowing a jet exhaust over the upper surface of a 50 deg swept leading-edge wing. Also investigated were the effects of varying the longitudinal and vertical location of the nozzle exit on the induced effects of jet blowing.

  9. Experimental and analytical investigation of axisymmetric supersonic cruise nozzle geometry at Mach numbers from 0.60 to 1.30

    NASA Technical Reports Server (NTRS)

    Carson, G. T., Jr.; Lee, E. E., Jr.

    1981-01-01

    Quantitative pressure and force data for five axisymmetric boattail nozzle configurations were examined. These configurations simulate the variable-geometry feature of a single nozzle design operating over a range of engine operating conditions. Five nozzles were tested in the Langley 16-Foot Transonic Tunnel at Mach numbers from 0.60 to 1.30. The experimental data were also compared with theoretical predictions.

  10. High Altitude Flight Test of a 40-Foot Diameter (12.2 meter) Ringsail Parachute at Deployment Mach Number of 2.95

    NASA Technical Reports Server (NTRS)

    Eckstrom, Clinton V.

    1970-01-01

    A 40-foot-nominal-diameter (12.2-meter) modified ringsail parachute was flight tested as part of the NASA Supersonic High Altitude Parachute Experiment (SHAPE) program. The 41-pound (18.6-kg) test parachute system was deployed from a 239.5-pound (108.6-kg) instrumented payload by means of a deployment mortar when the payload was at an altitude of 171,400 feet (52.3 km), a Mach number of 2.95, and a free-stream dynamic pressure of 9.2 lb/sq ft (440 N/m(exp 2)). The parachute deployed properly, suspension line stretch occurring 0.54 second after mortar firing with a resulting snatch-force loading of 932 pounds (4146 newtons). The maximum loading due to parachute opening was 5162 pounds (22 962 newtons) at 1.29 seconds after mortar firing. The first near full inflation of the canopy at 1.25 seconds after mortar firing was followed immediately by a partial collapse and subsequent oscillations of frontal area until the system had decelerated to a Mach number of about 1.5. The parachute then attained a shape that provided full drag area. During the supersonic part of the test, the average axial-force coefficient varied from a minimum of about 0.24 at a Mach number of 2.7 to a maximum of 0.54 at a Mach number of 1.1. During descent under subsonic conditions, the average effective drag coefficient was 0.62 and parachute-payload oscillation angles averaged about &loo with excursions to +/-20 degrees. The recovered parachute was found to have slight damage in the vent area caused by the attached deployment bag and mortar lid.

  11. Stability and performance characteristics of a fixed arrow wing supersonic transport configuration (SCAT 15F-9898) at Mach numbers from 0.60 to 1.20

    NASA Technical Reports Server (NTRS)

    Decker, J. P.; Jacobs, P. F.

    1978-01-01

    Tests on a 0.015 scale model of a supersonic transport were conducted at Mach numbers from 0.60 to 1.20. Tests of the complete model with three wing planforms, two different leading-edge radii, and various combinations of component parts, including both leading- and trailing-edge flaps, were made over an angle-of-attack range from about -6 deg to 13 deg and at sideslip angles of 0 deg and 2 deg.

  12. Control effectiveness and tip-fin dihedral effects for the HL-20 lifting-body configuration at Mach numbers from 1.6 to 4.5

    NASA Technical Reports Server (NTRS)

    Cruz, Christopher I.; Ware, George M.

    1995-01-01

    Wind tunnel tests were made with a scale model of the HL-20 in the Langley Unitary Plan Wind Tunnel. Pitch control was investigated by deflecting the elevon surfaces on the outboard fins and body flaps on the fuselage. Yaw control tests were made with the all movable center fin deflected 5 deg. Almost full negative body flap deflection (-30 deg) was required to trim the HL-20 (moment reference center at 0.54-percent body length from nose) to positive values of life in the Mach number range from 1.6 to 2.5. Elevons were twice as effective as body flaps as a longitudinal trim device. The elevons were effective as a roll control, but because of tip-fin dihedral angle, produced about as much adverse yawing moment as rolling moment. The body flaps were less effective in producing rolling moment, but produced little adverse yawing moment. The yaw effectiveness of the all movable center fin was essentially constant over the angle-of-attack range at each Mach number. The value of yawing moment, however, was small. Center-fin deflection produced almost no rolling moments. The model was directionally unstable over most of the Mach number range with tip-fin dihedral angles less than the baseline value of 50 deg.

  13. A STABLE, ACCURATE METHODOLOGY FOR HIGH MACH NUMBER, STRONG MAGNETIC FIELD MHD TURBULENCE WITH ADAPTIVE MESH REFINEMENT: RESOLUTION AND REFINEMENT STUDIES

    SciTech Connect

    Li, Pak Shing; Klein, Richard I.; Martin, Daniel F.; McKee, Christopher F. E-mail: klein@astron.berkeley.edu E-mail: cmckee@astro.berkeley.edu

    2012-02-01

    Performing a stable, long-duration simulation of driven MHD turbulence with a high thermal Mach number and a strong initial magnetic field is a challenge to high-order Godunov ideal MHD schemes because of the difficulty in guaranteeing positivity of the density and pressure. We have implemented a robust combination of reconstruction schemes, Riemann solvers, limiters, and constrained transport electromotive force averaging schemes that can meet this challenge, and using this strategy, we have developed a new adaptive mesh refinement (AMR) MHD module of the ORION2 code. We investigate the effects of AMR on several statistical properties of a turbulent ideal MHD system with a thermal Mach number of 10 and a plasma {beta}{sub 0} of 0.1 as initial conditions; our code is shown to be stable for simulations with higher Mach numbers (M{sub rms}= 17.3) and smaller plasma beta ({beta}{sub 0} = 0.0067) as well. Our results show that the quality of the turbulence simulation is generally related to the volume-averaged refinement. Our AMR simulations show that the turbulent dissipation coefficient for supersonic MHD turbulence is about 0.5, in agreement with unigrid simulations.

  14. Performance of an isolated two-dimensional wedge nozzle with fixed cowl and variable wedge centerbody at Mach numbers up to 2.01

    NASA Technical Reports Server (NTRS)

    Maiden, D. L.

    1976-01-01

    A wind tunnel investigation has been conducted to determine the aeropropulsion performance (thrust minus drag) of an isolated, two-dimensional wedge nozzle with a simulated variable-wedge mechanism and a fixed cowl. The investigation was conducted statically and at Mach numbers from 0.60 to 1.20 in the Langley 16-foot transonic tunnel and at a Mach number of 2.01 in the Langley 4-foot supersonic pressure tunnel. The ratio of exhaust jet total pressure to free-stream static pressure was varied up to 27 depending on free-stream Mach number. The results indicate that the aeropropulsion performance of the two-dimensional fixed-cowl variable-wedge nozzle is slightly lower (0.7 to 1.4 percent of ideal thrust) than that achieved for a two-dimensional wedge nozzle with a translating shroud, although part of the difference in performance is attributed to internal-performance differences. The effects of cowl boattail angle, internal expansion area ratio, and wedge half-angle on the performance of the two-dimensional wedge nozzle are discussed.

  15. The effects of winglets on low aspect ratio wings at supersonic Mach numbers. M.S. Thesis Report Feb. 1989 - Apr. 1991

    NASA Technical Reports Server (NTRS)

    Keenan, James A.; Kuhlman, John M.

    1991-01-01

    A computational study was conducted on two wings, of aspect ratios 1.244 and 1.865, each having 65 degree leading edge sweep angles, to determine the effects of nonplanar winglets at supersonic Mach numbers. A Mach number of 1.62 was selected as the design value. The winglets studied were parametrically varied in alignment, length, sweep, camber, thickness, and dihedral angle to determine which geometry had the best predicted performance. For the computational analysis, an available Euler marching technique was used. The results indicated that the possibility existed for wing-winglet geometries to equal the performance of wing-alone bodies in supersonic flows with both bodies having the same semispan. The first wing with winglet used NACA 1402 airfoils for the base wing and was shown to have lift-to-pressure drag ratios within 0.136 percent to 0.360 percent of the NACA 1402 wing-alone. The other base wing was a natural flow wing which was previously designed specifically for a Mach number of 1.62. The results obtained showed that the natural wing-alone had a slightly higher lift-to-pressure drag than the natural wing with winglets.

  16. Observation of mode conversion of m = minus 1 fast waves on the Alfven resonance layer

    SciTech Connect

    Amagishi, Y. )

    1990-03-12

    Fast waves or MHD surface waves of {ital m}={minus}1 (poloidal mode number of left-hand rotation) have been observed to be mode converted on the Alfven resonance layer. The converted waves are a quasielectrostatic form of the shear Alfven waves, i.e., kinetic Alfven wave and/or the resistive mode.

  17. Drag and Longitudinal Trim at Low Lift of the North American YF-100A Airplane at Mach Numbers from 0.76 to 1.77 as Determined from the Flight Test of a 0.11-Scale Rocket Model

    NASA Technical Reports Server (NTRS)

    Blanchard, Willard S.

    1953-01-01

    Drag and longitudinal trim at low lift of the North American YF-100A airplane at Mach numbers from 0.76 to 1.77 as determined from the flight test of a 0.11-scale rocket model are presented herein. Also included are some longitudinal stability and some qualitative pitch-damping data. The subsonic external-drag-coefficient level was about 0.012, and the supersonic level was about 0.043. The drag rise occurred at a Mach number of 0.95. The longitudinal trim change at low lift consisted basically of a mild nose-up tendency at a Mach number of 0.90. An indication of wing flutter was present at Mach numbers from 0.95 to 1.11. However, the full-scale airplane wing has approximately twice the scaled first-bending frequency as the model tested and, hence, will probably be free of this type of flutter. The aerodynamic-center location was 71 percent behind the leading edge of the mean aerodynamic chord at a Mach number of 1.03 and 62 percent at a Mach number of 1.74. Qualitative measurement of damping in pitch indicates that at low lift coefficients damping will be low at a Mach number of 1.03.

  18. Altitude-Test-Chamber Investigation of the Endurance and Performance Characteristics of the J65-W-7 Engine at a Mach Number of 2.0

    NASA Technical Reports Server (NTRS)

    Biermann, A.E.; Braithwaite, Willis M.

    1955-01-01

    An investigation of the endurance characteristics, at high Mach number, of the J65-W-7 engine was made in an altitude chamber at the Lewis laboratory. The investigation was made to determine whether this engine can be operated at flight conditions of Mach 2 at 35,000-feet altitude (inlet temperature, 250 F) as a limited-service-life engine Failure of the seventh-stage aluminum compressor blades occurred in both engines tested and was attributed to insufficient strength of the blade fastenings at the elevated temperatures. For the conditions of these tests, the results showed that it is reasonable to expect 10 to 15 minutes of satisfactory engine operation before failure. The high temperatures and pressures imposed upon the compressor housing caused no permanent deformation. In general, the performance of the engines tested was only slightly affected by the high ram conditions of this investigation. There was no discernible depreciation of performance with time prior to failure.

  19. Reynolds Number and Cowl Position Effects for a Generic Sidewall Compression Scramjet Inlet at Mach 10: A Computational and Experimental Investigation

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.

    1992-01-01

    Reynolds number and cowl position effects on the internal shock structure and the resulting performance of a generic three-dimensional sidewall compression scramjet inlet with a leading edge sweep of 45 degrees at Mach 10 have been examined both computationally and experimentally. Prior to the experiment, a three-dimensional Navier-Stokes code was adapted to perform preliminary parametric studies leading to the design of the present configuration. Following this design phase, the code was then utilized as an analysis tool to provide a better understanding of the flow field and the experimental static pressure data for the final experimental configuration. The wind tunnel model possessed 240 static pressure orifices distributed on the forebody plane, sidewalls, and cowl and was tested in the NASA Langley 31 Inch Mach 10 Tunnel.

  20. Longitudinal Aerodynamic Characteristics to Large Angles of Attack of a Cruciform Missile Configuration at a Mach Number of 2

    NASA Technical Reports Server (NTRS)

    Spahr, J. R.

    1954-01-01

    The lift, pitching-moment, and drag characteristics of a missile configuration having a body of fineness ratio 9.33 and a cruciform triangular wing and tail of aspect ratio 4 were measured at a Mach number of 1.99 and a Reynolds number of 6.0 million, based on the body length. The tests were performed through an angle-of-attack range of -5 deg to 28 deg to investigate the effects on the aerodynamic characteristics of roll angle, wing-tail interdigitation, wing deflection, and interference among the components (body, wing, and tail). Theoretical lift and moment characteristics of the configuration and its components were calculated by the use of existing theoretical methods which have been modified for application to high angles of attack, and these characteristics are compared with experiment. The lift and drag characteristics of all combinations of the body, wing, and tail were independent of roll angle throughout the angle-of-attack range. The pitching-moment characteristics of the body-wing and body-wing-tail combinations, however, were influenced significantly by the roll angle at large angles of attack (greater than 10 deg). A roll from 0 deg (one pair of wing panels horizontal) to 45 deg caused a forward shift in the center of pressure which was of the same magnitude for both of these combinations, indicating that this shift originated from body-wing interference effects. A favorable lift-interference effect (lift of the combination greater than the sum of the lifts of the components) and a rearward shift in the center of pressure from a position corresponding to that for the components occurred at small angles of attack when the body was combined with either the exposed wing or tail surfaces. These lift and center-of-pressure interference effects were gradually reduced to zero as the angle of attack was increased to large values. The effect of wing-tail interference, which influenced primarily the pitching-moment characteristics, is dependent on the distance

  1. Longitudinal Aerodynamic Characteristics to Large Angles of Attack of a Cruciform Missile Configuration at a Mach Number of 2

    NASA Technical Reports Server (NTRS)

    Spahr, J Richard

    1954-01-01

    The lift, pitching-moment, and drag characteristics of a missile configuration having a body of fineness ratio 9.33 and a cruciform triangular wing and tail of aspect ratio 4 were measured at a Mach number of 1.99 and a Reynolds number of 6.0 million, based on the body length. The tests were performed through an angle-of-attack range of -5 deg to 28 deg to investigate the effects on the aerodynamic characteristics of roll angle, wing-tail interdigitation, wing deflection, and interference among the components (body, wing, and tail). Theoretical lift and moment characteristics of the configuration and its components were calculated by the use of existing theoretical methods which have been modified for application to high angles of attack, and these characteristics are compared with experiment. The lift and drag characteristics of all combinations of the body, wing, and tail were independent of roll angle throughout the angle-of-attack range. The pitching-moment characteristics of the body-wing and body-wing- tail combinations, however, were influenced significantly by the roll angle at large angles of attack (greater than 10 deg). A roll from 0 deg (one pair of wing panels horizontal) to 45 deg caused a forward shift in the center of pressure which was of the same magnitude for both of these combinations, indicating that this shift originated from body-wing interference effects. A favorable lift - interference effect (lift of the combination greater than the sum of the lifts of the components) and a rearward shift in the center of pressure from a position corresponding to that for the components occurred at small angles of attack when the body was combined with either the exposed wing or tail surfaces. These lift and center-of-pressure interference effects were gradually reduced to zero as the angle of attack was increased to large values. The effect of wing-tail interference, which influenced primarily the pitching-moment characteristics, is dependent on the

  2. Aerodynamic characteristics of three helicopter rotor airfoil sections at Reynolds number from model scale to full scale at Mach numbers from 0.35 to 0.90. [conducted in Langley 6 by 28 inch transonic tunnel

    NASA Technical Reports Server (NTRS)

    Noonan, K. W.; Bingham, G. J.

    1980-01-01

    An investigation was conducted in the Langely 6 by 28 inch transonic tunnel to determine the two dimensional aerodynamic characteristics of three helicopter rotor airfoils at Reynolds numbers from typical model scale to full scale at Mach numbers from about 0.35 to 0.90. The model scale Reynolds numbers ranged from about 700,00 to 1,500,000 and the full scale Reynolds numbers ranged from about 3,000,000 to 6,600,000. The airfoils tested were the NACA 0012 (0 deg Tab), the SC 1095 R8, and the SC 1095. Both the SC 1095 and the SC 1095 R8 airfoils had trailing edge tabs. The results of this investigation indicate that Reynolds number effects can be significant on the maximum normal force coefficient and all drag related parameters; namely, drag at zero normal force, maximum normal force drag ratio, and drag divergence Mach number. The increments in these parameters at a given Mach number owing to the model scale to full scale Reynolds number change are different for each of the airfoils.

  3. NUMERICAL SIMULATIONS OF CONVERSION TO ALFVEN WAVES IN SUNSPOTS

    SciTech Connect

    Khomenko, E.; Cally, P. S. E-mail: paul.cally@monash.edu

    2012-02-10

    We study the conversion of fast magnetoacoustic waves to Alfven waves by means of 2.5D numerical simulations in a sunspot-like magnetic configuration. A fast, essentially acoustic, wave of a given frequency and wave number is generated below the surface and propagates upward through the Alfven/acoustic equipartition layer where it splits into upgoing slow (acoustic) and fast (magnetic) waves. The fast wave quickly reflects off the steep Alfven speed gradient, but around and above this reflection height it partially converts to Alfven waves, depending on the local relative inclinations of the background magnetic field and the wavevector. To measure the efficiency of this conversion to Alfven waves we calculate acoustic and magnetic energy fluxes. The particular amplitude and phase relations between the magnetic field and velocity oscillations help us to demonstrate that the waves produced are indeed Alfven waves. We find that the conversion to Alfven waves is particularly important for strongly inclined fields like those existing in sunspot penumbrae. Equally important is the magnetic field orientation with respect to the vertical plane of wave propagation, which we refer to as 'field azimuth'. For a field azimuth less than 90 Degree-Sign the generated Alfven waves continue upward, but above 90 Degree-Sign downgoing Alfven waves are preferentially produced. This yields negative Alfven energy flux for azimuths between 90 Degree-Sign and 180 Degree-Sign . Alfven energy fluxes may be comparable to or exceed acoustic fluxes, depending upon geometry, though computational exigencies limit their magnitude in our simulations.

  4. Drag Interference Between a Pointed Cylindrical Body and Triangular Wings of Various Aspect Ratios at Mach Numbers of 1.50 and 2.02

    NASA Technical Reports Server (NTRS)

    Katzen, Elliott D; Kaattari, George E

    1956-01-01

    The drag of a body alone, six triangular wings of various aspect ratios, and the combinations were measured at Mach numbers of 1.50 and 2.02 at a Reynolds number of 5.5 million (based on the body length). The experimental drag-interference results were in accordance with calculations based on NACA RM A9E19, 1949, with skin-friction effects taken into account, the interference effect being principally the result of fixing transition on the body by adding a wing.

  5. Pressure and heat flux results from the space shuttle/external fuel tank interaction test at Mach numbers 16 and 19

    NASA Technical Reports Server (NTRS)

    Brewer, E. B.; Haberman, D. R.

    1974-01-01

    Heat transfer rates and pressures were measured on a 0.0175-scale model of the space shuttle external tank (ET), model MCR0200. Tests were conducted with the ET model separately and while mated with a 0.0175-scale model of the orbiter, model 21-OT (Grumman). The tests were conducted in the AEDC-VKF Hypervelocity Wind Tunnel (F) at Mach numbers 16 and 19. The primary data consisted of the interaction heating rates experienced by the ET while mated with the orbiter in the flight configuration. Data were taken for a range of Reynolds numbers from 50,000 to 65,000 under laminar flow conditions.

  6. Completed Tabulation in the United States of Tests of 24 Airfoils at High Mach Numbers (Derived from Interrupted Work at Guidonia, Italy in the 1.31- by 1.74-Foot High-Speed Tunnel)

    NASA Technical Reports Server (NTRS)

    Ferri, Antonio

    1945-01-01

    Two-dimensional data were obtained in Mach range of from 0.40 to 0.94 and Reynolds Number range of (3.4 - 4.2) X 10 Degrees. Results indicate that thickness ratio is dominating shape parameter at high Mach numbers and that aerodynamic advantages are attainable by using thinnest possible sections. Effects of jet boundaries, Reynolds Number, and Data presented are free from jet-boundary and humidity effects.

  7. An Experimental Parametric Study of Geometric, Reynolds Number, and Ratio of Specific Heats Effects in Three-Dimensional Sidewall Compression Scramjet Inlets at Mach 6

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.; Murphy, Kelly J.

    1993-01-01

    Since mission profiles for airbreathing hypersonic vehicles such as the National Aero-Space Plane include single-stage-to-orbit requirements, real gas effects may become important with respect to engine performance. The effects of the decrease in the ratio of specific heats have been investigated in generic three-dimensional sidewall compression scramjet inlets with leading-edge sweep angles of 30 and 70 degrees. The effects of a decrease in ratio of specific heats were seen by comparing data from two facilities in two test gases: in the Langley Mach 6 CF4 Tunnel in tetrafluoromethane (where gamma=1.22) and in the Langley 15-Inch Mach 6 Air Tunnel in perfect gas air (where gamma=1.4). In addition to the simulated real gas effects, the parametric effects of cowl position, contraction ratio, leading-edge sweep, and Reynolds number were investigated in the 15-Inch Mach 6 Air Tunnel. The models were instrumented with a total of 45 static pressure orifices distributed on the sidewalls and baseplate. Surface streamline patterns were examined via oil flow, and schlieren videos were made of the external flow field. The results of these tests have significant implications to ground based testing of inlets in facilities which do not operate at flight enthalpies.

  8. An experimental parametric study of geometric, Reynolds number, and ratio of specific heats effects in three-dimensional sidewall compression scramjet inlets at Mach 6

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.; Murphy, Kelly J.

    1993-01-01

    The effects of the decrease in the ratio of specific heats have been investigated in generic 3D sidewall compression scramjet inlets with leading-edge sweep angles of 30 and 70 degrees. The effects of a decrease in ratio of specific heats were seen by comparing data from two facilities in two test gases: in the Langley Mach 6 CF4 Tunnel in tetrafluoromethane and in the Langley 15-Inch Mach 6 Air Tunnel in perfect gas air. In addition to the simulated real gas effects, the parametric effects of cowl position, contraction ratio, leading-edge sweep, and Reynolds number were investigated in the 15-Inch Mach 6 Air Tunnel. The models were instrumented with a total of 45 static pressure orifices distributed on the sidewalls and baseplate. Surface streamline patterns were examined via oil flow, and schlieren videos were made of the external flow field. The results of these tests have significant implications to ground based testing of inlets in facilities which do not operate at flight enthalpies.

  9. Aerodynamic performance of transonic and subsonic airfoils: Effects of surface roughness, turbulence intensity, Mach number, and streamline curvature-airfoil shape

    NASA Astrophysics Data System (ADS)

    Zhang, Qiang

    The effects of surface roughness, turbulence intensity, Mach number, and streamline curvature-airfoil shape on the aerodynamic performance of turbine airfoils are investigated in compressible, high speed flows. The University of Utah Transonic Wind Tunnel is employed for the experimental part of the study. Two different test sections are designed to produce Mach numbers, Reynolds numbers, passage mass flow rates, and physical dimensions, which match values along turbine blades in operating engines: (i) a nonturning test section with a symmetric airfoil, and (ii) a cascade test section with a cambered turbine vane. The nonuniform, irregular, three-dimensional surface roughness is characterized using the equivalent sand grain roughness size. Changing the airfoil surface roughness condition has a substantial effect on wake profiles of total pressure loss coefficients, normalized Mach number, normalized kinetic energy, and on the normalized and dimensional magnitudes of Integrated Aerodynamic Losses produced by the airfoils. Comparisons with results for a symmetric airfoil and a cambered vane show that roughness has more substantial effects on losses produced by the symmetric airfoil than the cambered vane. Data are also provided that illustrate the larger loss magnitudes are generally present with flow turning and cambered airfoils, than with symmetric airfoils. Wake turbulence structure of symmetric airfoils and cambered vanes are also studied experimentally. The effects of surface roughness and freestream turbulence levels on wake distributions of mean velocity, turbulence intensity, and power spectral density profiles and vortex shedding frequencies are quantified one axial chord length downstream of the test airfoils. As the level of surface roughness increases, all wake profile quantities broaden significantly and nondimensional vortex shedding frequencies decrease. Wake profiles produced by the symmetric airfoil are more sensitive to variations of surface

  10. High-dynamic-range extinction mapping of infrared dark clouds. Dependence of density variance with sonic Mach number in molecular clouds

    NASA Astrophysics Data System (ADS)

    Kainulainen, J.; Tan, J. C.

    2013-01-01

    Context. Measuring the mass distribution of infrared dark clouds (IRDCs) over the wide dynamic range of their column densities is a fundamental obstacle in determining the initial conditions of high-mass star formation and star cluster formation. Aims: We present a new technique to derive high-dynamic-range, arcsecond-scale resolution column density data for IRDCs and demonstrate the potential of such data in measuring the density variance - sonic Mach number relation in molecular clouds. Methods: We combine near-infrared data from the UKIDSS/Galactic Plane Survey with mid-infrared data from the Spitzer/GLIMPSE survey to derive dust extinction maps for a sample of ten IRDCs. We then examine the linewidths of the IRDCs using 13CO line emission data from the FCRAO/Galactic Ring Survey and derive a column density - sonic Mach number relation for them. For comparison, we also examine the relation in a sample of nearby molecular clouds. Results: The presented column density mapping technique provides a very capable, temperature independent tool for mapping IRDCs over the column density range equivalent to AV ≃ 1-100 mag at a resolution of 2″. Using the data provided by the technique, we present the first direct measurement of the relationship between the column density dispersion, σN/⟨N⟩, and sonic Mach number, ℳs, in molecular clouds. We detect correlation between the variables with about 3-σ confidence. We derive the relation σN/⟨N⟩ ≈ (0.047 ± 0.016)ℳs, which is suggestive of the correlation coefficient between the volume density and sonic Mach number, σρ/⟨ρ⟩ ≈ (0.20-0.22+0.37)ℳs, in which the quoted uncertainties indicate the 3-σ range. When coupled with the results of recent numerical works, the existence of the correlation supports the picture of weak correlation between the magnetic field strength and density in molecular clouds (i.e., B ∝ ρ0.5). While our results remain suggestive because of the small number of clouds in our

  11. Sub-Alfvenic Non-Ideal MHD Turbulence Simulations with Ambipolar Diffusion: I. Turbulence Statistics

    SciTech Connect

    Klein, R I; Li, P S; McKee, C F; Fisher, R

    2008-04-10

    Most numerical investigations on the role of magnetic fields in turbulent molecular clouds (MCs) are based on ideal magneto-hydrodynamics (MHD). However, MCs are weakly ionized, so that the time scale required for the magnetic field to diffuse through the neutral component of the plasma by ambipolar diffusion (AD) can be comparable to the dynamical time scale. We have performed a series of 256{sup 3} and 512{sup 3} simulations on supersonic but sub-Alfvenic turbulent systems with AD using the Heavy-Ion Approximation developed in Li et al. (2006). Our calculations are based on the assumption that the number of ions is conserved, but we show that these results approximately apply to the case of time-dependent ionization in molecular clouds as well. Convergence studies allow us to determine the optimal value of the ionization mass fraction when using the heavy-ion approximation for low Mach number, sub-Alfvenic turbulent systems. We find that ambipolar diffusion steepens the velocity and magnetic power spectra compared to the ideal MHD case. Changes in the density PDF, total magnetic energy, and ionization fraction are determined as a function of the AD Reynolds number. The power spectra for the neutral gas properties of a strongly magnetized medium with a low AD Reynolds number are similar to those for a weakly magnetized medium; in particular, the power spectrum of the neutral velocity is close to that for Burgers turbulence.

  12. Non-thermal electron acceleration in low Mach number collisionless shocks. II. Firehose-mediated Fermi acceleration and its dependence on pre-shock conditions

    SciTech Connect

    Guo, Xinyi; Narayan, Ramesh; Sironi, Lorenzo

    2014-12-10

    Electron acceleration to non-thermal energies is known to occur in low Mach number (M{sub s} ≲ 5) shocks in galaxy clusters and solar flares, but the electron acceleration mechanism remains poorly understood. Using two-dimensional (2D) particle-in-cell (PIC) plasma simulations, we showed in Paper I that electrons are efficiently accelerated in low Mach number (M{sub s} = 3) quasi-perpendicular shocks via a Fermi-like process. The electrons bounce between the upstream region and the shock front, with each reflection at the shock resulting in energy gain via shock drift acceleration. The upstream scattering is provided by oblique magnetic waves that are self-generated by the electrons escaping ahead of the shock. In the present work, we employ additional 2D PIC simulations to address the nature of the upstream oblique waves. We find that the waves are generated by the shock-reflected electrons via the firehose instability, which is driven by an anisotropy in the electron velocity distribution. We systematically explore how the efficiency of wave generation and of electron acceleration depend on the magnetic field obliquity, the flow magnetization (or equivalently, the plasma beta), and the upstream electron temperature. We find that the mechanism works for shocks with high plasma beta (≳ 20) at nearly all magnetic field obliquities, and for electron temperatures in the range relevant for galaxy clusters. Our findings offer a natural solution to the conflict between the bright radio synchrotron emission observed from the outskirts of galaxy clusters and the low electron acceleration efficiency usually expected in low Mach number shocks.

  13. Non-thermal Electron Acceleration in Low Mach Number Collisionless Shocks. II. Firehose-mediated Fermi Acceleration and its Dependence on Pre-shock Conditions

    NASA Astrophysics Data System (ADS)

    Guo, Xinyi; Sironi, Lorenzo; Narayan, Ramesh

    2014-12-01

    Electron acceleration to non-thermal energies is known to occur in low Mach number (Ms <~ 5) shocks in galaxy clusters and solar flares, but the electron acceleration mechanism remains poorly understood. Using two-dimensional (2D) particle-in-cell (PIC) plasma simulations, we showed in Paper I that electrons are efficiently accelerated in low Mach number (Ms = 3) quasi-perpendicular shocks via a Fermi-like process. The electrons bounce between the upstream region and the shock front, with each reflection at the shock resulting in energy gain via shock drift acceleration. The upstream scattering is provided by oblique magnetic waves that are self-generated by the electrons escaping ahead of the shock. In the present work, we employ additional 2D PIC simulations to address the nature of the upstream oblique waves. We find that the waves are generated by the shock-reflected electrons via the firehose instability, which is driven by an anisotropy in the electron velocity distribution. We systematically explore how the efficiency of wave generation and of electron acceleration depend on the magnetic field obliquity, the flow magnetization (or equivalently, the plasma beta), and the upstream electron temperature. We find that the mechanism works for shocks with high plasma beta (gsim 20) at nearly all magnetic field obliquities, and for electron temperatures in the range relevant for galaxy clusters. Our findings offer a natural solution to the conflict between the bright radio synchrotron emission observed from the outskirts of galaxy clusters and the low electron acceleration efficiency usually expected in low Mach number shocks.

  14. Aeropropulsive characteristics of Mach numbers up to 2.2 of axisymmetric and nonaxisymmetric nozzles installed on an F-18 model

    NASA Technical Reports Server (NTRS)

    Capone, F. J.

    1982-01-01

    An investigation to determine the aeropropulsive characteristics of nonaxisymmetric nozzles on an F-18 jet effects model was conducted in the Langley 16-foot transonic tunnel and the AEDC 16-foot supersonic wind tunnel. The performance of a two dimensional convergent-divergent nozzle, a single expansion ramp nozzle, and a wedge nozzle was compared with that of the baseline axisymmetric nozzle. Test data were obtained at static conditions and at Mach numbers from 0.60 to 2.20 at an angle of attack of 0 deg. Nozzle pressure ratio was varied from jet-off to about 20.

  15. Experimental investigation of water injection in subsonic diffuser of a conical inlet operation at free-stream Mach number of 2.5

    NASA Technical Reports Server (NTRS)

    Beke, Andrew

    1957-01-01

    A spike-type nose inlet with sharp-lip cowl was investigated at a free-stream Mach number of 2.5 with water injection in its 16-inch diameter, 11-foot-long subsonic diffuser section. Inlet total temperature of exit with liquid-air ratios of about 0.04 with no apparent change in the critical pressure recovery. The observed temperature drops were less than the theoretically predicted values, and the amount of water evaporated was 35 to 50 percent less than that theoretically possible.

  16. An investigation of the stability of the Bondi-Hoyle model of accretion flow. [onto massive astronomical bodies at high Mach number

    NASA Technical Reports Server (NTRS)

    Cowie, L. L.

    1977-01-01

    The Bondi-Hoyle-Lyttleton (1944) accretion model is considered which involves accretion onto a massive body moving at a high Mach number with respect to the ambient medium and the production of a high-density accretion column along the axis where particle orbits intersect. The stability of steady-state solutions with respect to short-wavelength perturbations is analyzed using the WKB approximation, and the accretion column is shown to be unstable toward such perturbations. It is noted that this instability is not affected by the position of the stagnation point in the steady-state solution.

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

    NASA Technical Reports Server (NTRS)

    Calleja, John; Tamagno, Jose

    1993-01-01

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

  18. Fluctuating pressures measured beneath a high-temperature, turbulent boundary layer on a flat plate at Mach number of 5

    NASA Technical Reports Server (NTRS)

    Parrott, Tony L.; Jones, Michael G.; Albertson, Cindy W.

    1989-01-01

    Fluctuating pressures were measured beneath a Mach 5, turbulent boundary layer on a flat plate with an array of piezoresistive sensors. The data were obtained with a digital signal acquisition system during a test run of 4 seconds. Data sampling rate was such that frequency analysis up to 62.5 kHz could be performed. To assess in situ frequency response of the sensors, a specially designed waveguide calibration system was employed to measure transfer functions of all sensors and related instrumentation. Pressure time histories were approximated well by a Gaussian prohibiting distribution. Pressure spectra were very repeatable over the array span of 76 mm. Total rms pressures ranged from 0.0017 to 0.0046 of the freestream dynamic pressure. Streamwise, space-time correlations exhibited expected decaying behavior of a turbulence generated pressure field. Average convection speed was 0.87 of freestream velocity. The trendless behavior with sensor separation indicated possible systematic errors.

  19. Effects of Wing Sweep on Boundary-layer Transition for a Smooth F-14A Wing at Mach Numbers from 0.700 to 0.825

    NASA Technical Reports Server (NTRS)

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

    1990-01-01

    The results are discussed of the variable sweep transition flight experiment (VSTFE). The VSTFE was a natural laminar flow experiment flown on the swing wing F-14A aircraft. The main objective of the VSTFE was to determine the effects of wing sweep on boundary layer transition at conditions representative of transport aircraft. The experiment included the flight testing of two laminar flow wing gloves. Glove 1 was a cleanup of the existing F-14A wing. Glove 2, not discussed herein, was designed to provide favorable pressure distributions for natural laminar flow at Mach number (M) 0.700. The transition locations presented for glove 1 were determined primarily by using hot film sensors. Boundary layer rake data was provided as a supplement. Transition data were obtained for leading edge wing sweeps of 15, 20, 25, 30, and 35 degs, with Mach numbers ranging from 0.700 to 0.825, and altitudes ranging from 10,000 to 35,000 ft. Results show that a substantial amount of laminar flow was maintained at all the wing sweeps evaluated. The maximum transition Reynolds number of 13.7 x 10(exp 6) was obtained for the condition of 15 deg of sweep, M = 0.800, and an altitude of 20,000 ft.

  20. Gas-jet and tangent-slot film cooling tests of a 12.5 deg cone at Mach number of 6.7

    NASA Technical Reports Server (NTRS)

    Nowak, Robert J.

    1988-01-01

    Tests were conducted in the Langley 8-Foot High Temperature Tunnel to determine the aerothermal effects of gaseous nitrogen-coolant ejection on a 3-ft base-diameter, 12.5 degree half-angle cone. Free-stream Mach number, total temperature, and unit Reynolds number per foot were 6.7, 3300 deg R, and 1.4 million, respectively. Two coolant ejection noses were tested, an ogive frustum with a forward-facing 0.8-in radius gas-jet tip, and a 3-in radius hemisphere with a 0.243-in high rearward-facing tangent slot. Data include surface pressures and heating rates, shock shapes, and shock-layer profiles; results are compared with no-cooling data obtained with 1-in and 3-in radius solid noses. Surface pressures were reduced with gas-jet ejection but were affected little by tangent-slot ejection. For both gas-jet and tangent-slot ejection, high coolant flow rates reduced heating even far downstream from the region of ejection; however, low coolant rates caused transition to turbulence and increased heating. Shock-layer profiles of pitot pressure, Mach number, and total temperature were reduced for both gas-jet and tangent-slot ejection. Insight into the gas-jet heat-flux mechanisms was obtained by using shock-layer rake data and established, no-cooling, heat-transfer equations.

  1. Effect of Mach number, valve angle and length to diameter ratio on thermal performance in flow of air through Ranque Hilsch vortex tube

    NASA Astrophysics Data System (ADS)

    Devade, Kiran D.; Pise, Ashok T.

    2016-04-01

    Ranque Hilsch vortex tube is a device that can produce cold and hot air streams simultaneously from pressurized air. Performance of vortex tube is influenced by a number of geometrical and operational parameters. In this study parametric analysis of vortex tube is carried out. Air is used as the working fluid and geometrical parameters like length to diameter ratio (15, 16, 17, 18), exit valve angles (30°-90°), orifice diameters (5, 6 and 7 mm), 2 entry nozzles and tube divergence angle 4° is used for experimentation. Operational parameters like pressure (200-600 kPa), cold mass fraction (0-1) is varied and effect of Mach number at the inlet of the tube is investigated. The vortex tube is tested at sub sonic (0 < Ma < 1), sonic (Ma = 1) and supersonic (1 < Ma < 2) Mach number, and its effect on thermal performance is analysed. As a result it is observed that, higher COP and low cold end temperature is obtained at subsonic Ma. As CMF increases, COP rises and cold and temperature drops. Optimum performance of the tube is observed for CMF up to 0.5. Experimental correlations are proposed for optimum COP. Parametric correlation is developed for geometrical and operational parameters.

  2. Free-flight Performance of 16-inch-diameter Supersonic Ram-jet Units III : Four Units Designed for Combustion-chamber-inlet Mach Number of 0.245 at Free-stream Mach Number of 1.8 (units D-1, D-2, D-3, and D-4)

    NASA Technical Reports Server (NTRS)

    Disher, John H; Rabinowitz, Leonard

    1950-01-01

    Performance of four 16-inch-diameter ram-jet units was determined at free-stream Mach numbers of 0.49 to 1.78 over range of gas total-temperature ratios of 1.0 to 6.1. Time histories of each flight and data on thrust, drag, diffuser efficiency, and combustion are presented. A maximum thrust coefficient of 0.88 and a maximum net acceleration of 5.13 g's were observed for the four units.

  3. An Inviscid Computational Study of Three '07 Mars Lander Aeroshell Configurations Over a Mach Number Range of 2.3 to 4.5

    NASA Technical Reports Server (NTRS)

    Prabhu, Ramadas K.; Sutton, Kenneth (Technical Monitor)

    2001-01-01

    This report documents the results of a study conducted to compute the inviscid longitudinal aerodynamic characteristics of three aeroshell configurations of the proposed '07 Mars lander. This was done in support of the activity to design a smart lander for the proposed '07 Mars mission. In addition to the three configurations with tabs designated as the shelf, the canted, and the Ames, the baseline configuration (without tab) was also studied. The unstructured grid inviscid CFD software FELISA was used, and the longitudinal aerodynamic characteristics of the four configurations were computed for Mach number of 2.3, 2.7, 3.5, and 4.5, and for an angle of attack range of -4 to 20 degrees. Wind tunnel tests had been conducted on scale models of these four configurations in the Unitary Plan Wind Tunnel, NASA Langley Research Center. Present computational results are compared with the data from these tests. Some differences are noticed between the two results, particularly at the lower Mach numbers. These differences are attributed to the pressures acting on the aft body. Most of the present computations were done on the forebody only. Additional computations were done on the full body (forebody and afterbody) for the baseline and the Shelf configurations. Results of some computations done (to simulate flight conditions) with the Mars gas option and with an effective gamma are also included.

  4. PARTICLE-IN-CELL SIMULATIONS OF PARTICLE ENERGIZATION VIA SHOCK DRIFT ACCELERATION FROM LOW MACH NUMBER QUASI-PERPENDICULAR SHOCKS IN SOLAR FLARES

    SciTech Connect

    Park, Jaehong; Ren Chuang; Workman, Jared C.; Blackman, Eric G.

    2013-03-10

    Low Mach number, high beta fast mode shocks can occur in the magnetic reconnection outflows of solar flares. These shocks, which occur above flare loop tops, may provide the electron energization responsible for some of the observed hard X-rays and contemporaneous radio emission. Here we present new two-dimensional particle-in-cell simulations of low Mach number/high beta quasi-perpendicular shocks. The simulations show that electrons above a certain energy threshold experience shock-drift-acceleration. The transition energy between the thermal and non-thermal spectrum and the spectral index from the simulations are consistent with some of the X-ray spectra from RHESSI in the energy regime of E {approx}< 40 {approx} 100 keV. Plasma instabilities associated with the shock structure such as the modified-two-stream and the electron whistler instabilities are identified using numerical solutions of the kinetic dispersion relations. We also show that the results from PIC simulations with reduced ion/electron mass ratio can be scaled to those with the realistic mass ratio.

  5. Wind tunnel investigation of Nacelle-Airframe interference at Mach numbers of 0.9 to 1.4-force data

    NASA Technical Reports Server (NTRS)

    Bencze, D. P.

    1976-01-01

    Detailed interference force and pressure data were obtained on a representative wing-body-nacelle combination at Mach numbers of 0.9 to 1.4. The model consisted of a delta wing-body aerodynamic force model with four independently supported nacelles located beneath the wing-body combination. The model was mounted on a six-component force balance, and the left-hand wing was pressure-instrumented. Each of the two right-hand nacelles was mounted on a six-component force balance housed in the thickness of the nacelle, while each of the left-hand nacelles was pressure-instrumented. The primary variables examined included Mach number, angle of attack, nacelle position, and nacelle mass-flow ratio. Four different configurations were tested to identify various interference forces and pressures on each component; these included tests of the isolated nacelle, the isolated wing-body combination, the four nacelles as a unit, and the total wing-body-nacelle combination. Nacelle axial location, relative to both the wing-body combination and to each other, was the most important variable in determining the net interference among the components.

  6. Application of supersonic linear theory and hypersonic impact methods to three nonslender hypersonic airplane concepts at Mach numbers from 1.10 to 2.86

    NASA Technical Reports Server (NTRS)

    Pittman, J. L.

    1979-01-01

    Aerodynamic predictions from supersonic linear theory and hypersonic impact theory were compared with experimental data for three hypersonic research airplane concepts over a Mach number range from 1.10 to 2.86. The linear theory gave good lift prediction and fair to good pitching-moment prediction over the Mach number (M) range. The tangent-cone theory predictions were good for lift and fair to good for pitching moment for M more than or equal to 2.0. The combined tangent-cone theory predictions were good for lift and fair to good for pitching moment for M more than or equal to 2.0. The combined tangent-cone/tangent-wedge method gave the least accurate prediction of lift and pitching moment. The zero-lift drag was overestimated, especially for M less than 2.0. The linear theory drag prediction was generally poor, with areas of good agreement only for M less than or equal to 1.2. For M more than or equal to 2.), the tangent-cone method predicted the zero-lift drag most accurately.

  7. Performance Characteristics of a Normal-shock Side Inlet Located Downstream of a Canard Control Surface at Mach Numbers of 1.5 and 1.8/

    NASA Technical Reports Server (NTRS)

    Dryer, Murray; Beke, Andrew

    1952-01-01

    The performance characteristics of a downward canted normal-shock side (scoop) inlet located downstream of a triangular control surface are presented for free-stream Mach numbers of 1.5 and 1.8 in terms of total pressure recovery and mass flow ratio for various boundary-layer removal systems,angles of attack, control surface deflections and adverse yaw. An engine operating condition for a hypothetical turbojet engine is established, and the match point characteristics of the engine-inlet configuration are summarized. 520::It is shown that the diffuser performance increases with increased boundary-layer removal and decreases because of the presence of the wake from the forward control surface. At the higher angles of attack the wake passes over the inlet and does not affect the inlet performance. Adverse yaw reduces the total pressure recovery values below those for the unawed case. Magnitudes of the total pressure recovery were below the theoretical normal-shock recovery for the respective test Mach numbers.

  8. Wind tunnel investigation of Nacelle-Airframe interference at Mach numbers of 0.9 to 1.4-pressure data, volume 2

    NASA Technical Reports Server (NTRS)

    Bencze, D. P.

    1976-01-01

    Detailed interference force and pressure data were obtained on a representative wing-body nacelle combination at Mach numbers of 0.9 to 1.4. The model consisted of a delta wing-body aerodynamic force model with four independently supported nacelles located beneath the wing-body combination. The primary variables examined included Mach number, angle of attack, nacelle position, and nacelle mass flow ratio. Four different configurations were tested to identify various interference forces and pressures on each component; these included tests of the isolated nacelle, the isolated wing-body combination, the four nacelles as a unit, and the total wing-body-nacelle combination. Nacelle axial location, relative to both the wing-body combination and to each other, was the most important variable in determining the net interference among the components. The overall interference effects were found to be essentially constant over the operating angle-of-attack range of the configuration, and nearly independent of nacelle mass flow ratio.

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

    NASA Technical Reports Server (NTRS)

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

    1990-01-01

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

  10. Aerodynamic characteristics of wings designed with a combined-theory method to cruise at a Mach number of 4.5

    NASA Technical Reports Server (NTRS)

    Mack, Robert J.

    1988-01-01

    A wind-tunnel study was conducted to determine the capability of a method combining linear theory and shock-expansion theory to design optimum camber surfaces for wings that will fly at high-supersonic/low-hypersonic speeds. Three force models (a flat-plate reference wing and two cambered and twisted wings) were used to obtain aerodynamic lift, drag, and pitching-moment data. A fourth pressure-orifice model was used to obtain surface-pressure data. All four wing models had the same planform, airfoil section, and centerbody area distribution. The design Mach number was 4.5, but data were also obtained at Mach numbers of 3.5 and 4.0. Results of these tests indicated that the use of airfoil thickness as a theoretical optimum, camber-surface design constraint did not improve the aerodynamic efficiency or performance of a wing as compared with a wing that was designed with a zero-thickness airfoil (linear-theory) constraint.

  11. Effect of tail-fin span on stability and control characteristics of a Canard-controlled missile at supersonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Blair, A. B., Jr.; Allen, J. M.; Hernandez, G.

    1983-01-01

    An experimental wind-tunnel investigation was conducted at Mach numbers from 1.60 to 3.50 to obtain the longitudinal and lateral-directional aerodynamic characteristics of a circular, cruciform, canard-controlled missile with variations in tail-fin span. In addition, comparisons were made with the experimental aerodynamic characteristics using three missile aeroprediction programs: MISSILE1, MISSILE2, and NSWCDM. The results of the investigation indicate that for the test Mach number range, canard roll control at low angles of attack is feasible on tail-fin configurations with tail-to-canard span ratios of less than or equal to 0.75. The conards are effective pitch and yaw control devices on each tail-fin span configuration tested. Programs MISSILE1 and MISSILE2 provide very good predictions of longitudinal aerodynamic characteristics and fair predictions of lateral-directional aerodynamic characteristics at low angles of attack, with MISSILE2 predictions generally in better agreement with test data. Program NSWCDM provides good longitudinal and lateral-directional aerodynamic predictions that improve with increases in tail-tin span.

  12. Longitudinal aerodynamic characteristics of a fighter model with a close-coupled canard at Mach numbers from 0.40 to 1.20

    NASA Technical Reports Server (NTRS)

    Re, R. J.; Capone, F. J.

    1978-01-01

    A Au aircraft model with a close-coupled canard mounted above the wing chord plane was considered. Model angle of attack was varied from -4 deg to 15 deg; canard incidence was varied from -5 deg to 18 deg; and selected canard and wing flap deflections were investigated. By using the canard incidence for trim, maximum trimmed lift-drag ratios of about 8.8, 7.7, and 4.7 were obtained at free-stream Mach numbers of 0.40, 0.90, and 1.20, respectively. At a lift coefficient of 0.60, model trim angle of attack could be varied over an incremental range between 3.0 deg and 3.8 deg, depending on Mach number, by different combinations of control settings. At high lift coefficients, larger trimmed lift-drag ratios were obtained by using the deflection capability of the canard leading- and trailing-edge flaps before increasing canard incidence angle.

  13. Experimental Investigation of the Heat-Transfer Rate to a Series of 20 deg Cones of Various Surface Finishes at a Mach Number of 4.95

    NASA Technical Reports Server (NTRS)

    Jones, Jim J.

    1959-01-01

    The heat-transfer rates were measured on a series of cones of various surface finishes at a Mach number of 4.95 and Reynolds numbers per foot varying from 20 x 10(exp 6) to 100 x 10(exp 6). The range of surface finish was from a very smooth polish to smooth machining with no polish (65 micro inches rms). Some laminar boundary-layer data were obtained, since transition was not artificially tripped. Emphasis, however, is centered on the turbulent boundary layer. The results indicated that the turbulent heat-transfer rate for the highest roughness tested was only slightly greater than that for the smoothest surface. The laminar-sublayer thickness was calculated to be about half the roughness height for the roughest model at the highest value of unit Reynolds number tested.

  14. Static aerodynamic characteristics of a 0.035-scale model of a modified NKC-135 airplane at a Mach number of 0.28

    NASA Technical Reports Server (NTRS)

    Hedstrom, E.; Whitcomb, W. M.

    1977-01-01

    A 0.035-scale model fo a modified NKC-135 airplane was tested in 12-foot pressure wind tunnel to determine the effects on the static aerodynamic characteristics of modifications to the basic aircraft. Modifications investigated included: nose, lower fuselage, and upper fuselage radomes; wing pylons and pods; overwing probe; and air conditioning inlets. The investigation was performed at a Mach number of 0.28 over a Reynolds number range from 6.6 to 26.2 million per meter. Angles of attack and sideslip varied from -8 deg to 20 deg and from -18 deg to 8 deg, respectively, for various combinations of flap, aileron, and rudder deflections. A limited analysis of the test results indicates that the addition of the radomes reduces lateral-directional stability and control effectiveness of the basic aircraft.

  15. Wind-Tunnel Investigation of a Balloon as a Towed Decelerator at Mach Numbers from 1.47 to 2.50

    NASA Technical Reports Server (NTRS)

    McShera, John T.; Keyes, J. Wayne

    1961-01-01

    A wind-tunnel investigation has been conducted to study the characteristics of a towed spherical balloon as a drag device at Mach numbers from 1.47 to 2.50, Reynolds numbers from 0.36 x 10(exp 6) to 1.0 x 10(exp 6) , and angles of attack from -15 to 15 deg. Towed spherical balloons were found to be stable at supersonic speeds. The drag coefficient of the balloon is reduced by the presence of a tow cable and a further reduction occurs with the addition of a payload. The balloon inflation pressure required to maintain an almost spherical shape is about equal to the free-stream dynamic pressure. Measured pressure and temperature distribution around the balloon alone were in fair agreement with predicted values. There was a pronounced decrease in the pressure coefficients on the balloon when attached to a tow cable behind a payload.

  16. Surface pressure data on a series of conical forebodies at Mach numbers from 1.70 to 4.50 and combined angles of attack and sideslip

    NASA Technical Reports Server (NTRS)

    Townsend, J. C.; Collins, I. K.; Howell, D. T.; Hayes, C.

    1979-01-01

    Tabulated surface pressure data for a series of forebodies which have analytically defined cross sections and are based on a 20 degs half-angle cone are presented without analysis. Five of the cross sections were ellipses having axis ratios of 3/1, 2/1, 1/1, 1/2, and 1/3. The sixth cross section was defined by a curve having a single lobe. The data generally cover angles of attack from -5 degs to 20 degs at angles of sideslip from 0 degs to 5 degs for Mach numbers of 1.70, 2.50, 3.95, and 4.50 at a constant Reynolds number.

  17. Tabulated Pressure Data for a Series of Controls on a 40 Deg Sweptback Wing at Mach Numbers of 1.61 and 2.01

    NASA Technical Reports Server (NTRS)

    Lord, D. R.

    1957-01-01

    An investigation has been made at Mach numbers of 1.61 and 2.01 and Reynolds numbers of 1.7 x l0(exp 6) and 3.6 x l0(exp 6) to determine the pressure distributions over a swept wing with a series of 14 control configurations. The wing had 40 deg of sweep of the quarter-chord line, an aspect ratio of 3.1, and a taper ratio of 0.4. Measurements were made at angles of attack from 0 deg to +/- 15 deg for control deflections from -60 deg to 60 deg. This report contains tabulated pressure data for the complete range of test conditions.

  18. Aerodynamic characteristics at Mach numbers of 1.5, 1.8, and 2.0 of a blended wing-body configuration with and without integral canards

    NASA Technical Reports Server (NTRS)

    Robins, A. W.; Lamb, M.; Miller, D. S.

    1979-01-01

    An exploratory, experimental, and theoretical investigation was made of a cambered, twisted, and blended wing-body concept with and without integral canard surfaces. Theoretical calculations of the static longitudinal and lateral aerodynamic characteristics of the wing-body configurations were compared with the characteristics obtained from tests of a model in the Langley Unitary Plan wind tunnel. Mach numbers of 1.5, 1.8, and 2.0 and a Reynolds number per meter of 6.56 million were used in the calculations and tests. Overall results suggest that planform selection is extremely important and that the supplemental application of new calculation techniques should provide a process for the design of supersonic wings in which spanwise distribution of upwash and leading-edge thrust might be rationally controlled and exploited.

  19. Visualization of Flow Separation Around an Atmospheric Entry Capsule at Low-Subsonic Mach Number Using Background-Oriented Schlieren (BOS)

    NASA Technical Reports Server (NTRS)

    Mizukaki, Toshiharu; Borg, Stephen E.; Danehy, Paul M.; Murman, Scott M.

    2014-01-01

    This paper presents the results of visualization of separated flow around a generic entry capsule that resembles the Apollo Command Module (CM) and the Orion Multi-Purpose Crew Vehicle (MPCV). The model was tested at flow speeds up to Mach 0.4 at a single angle of attack of 28 degrees. For manned spacecraft using capsule-shaped vehicles, certain flight operations such as emergency abort maneuvers soon after launch and flight just prior to parachute deployment during the final stages of entry, the command module may fly at low Mach number. Under these flow conditions, the separated flow generated from the heat-shield surface on both windward and leeward sides of the capsule dominates the wake flow downstream of the capsule. In this paper, flow visualization of the separated flow was conducted using the background-oriented schlieren (BOS) method, which has the capability of visualizing significantly separated wake flows without the particle seeding required by other techniques. Experimental results herein show that BOS has detection capability of density changes on the order of 10(sup-5).

  20. Experimental aerodynamic characteristics of a generic hypersonic accelerator configuration at Mach numbers 1.5 and 2.0. [conducted in the Langley Unitary Plan Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Walker, Ira J.; Covell, Peter F.; Forrest, Dana K.

    1993-01-01

    An experimental investigation of the static longitudinal and lateral-directional aerodynamic characteristics of a generic hypersonic research vehicle was conducted in the Langley Unitary Plan Wind Tunnel (UPWT). A parametric study was performed to determine the interference effects of various model components. Configuration variables included delta and trapezoidal canards; large and small centerline-mounted vertical tails, along with a set of wing-mounted vertical tails; and a set of model noses with different degrees of bluntness. Wing position was varied by changing the longitudinal location and the incidence angle. The test Mach numbers were 1.5 and 2.0 at Reynolds numbers of 1 x 10(exp 6) per foot, 2 x 10(exp 6) per foot, and 4 x 10(exp 6) per foot. Angle of attack was varied from -4 degrees to 27 degrees, and sideslip angle was varied from -8 degrees to 8 degrees. Generally, the effect of Reynolds number did not deviate from conventional trends. The longitudinal stability and lift-curve slope decreased with increasing Mach number. As the wing was shifted rearward, the lift-curve slope decreased and the longitudinal stability increased. Also, the wing-mounted vertical tails resulted in a more longitudinally stable configuration. In general, the lift-drag ratio was not significantly affected by vertical-tail arrangement. The best lateral-directional stability was achieved with the large centerline-mounted tail, although the wing-mounted vertical tails exhibited the most favorable characteristics at the higher angles of attack.

  1. Side forces on forebodies at high angles of attack and Mach numbers from 0.1 to 0.7: two tangent ogives, paraboloid and cone

    NASA Technical Reports Server (NTRS)

    Keener, E. R.; Chapman, G. T.; Taleghani, J.; Cohen, L.

    1977-01-01

    An experimental investigation was conducted in the Ames 12-Foot Wind Tunnel to determine the subsonic aerodynamic characteristics of four forebodies at high angles of attack. The forebodies tested were a tangent ogive with fineness ratio of 5, a paraboloid with fineness ratio of 3.5, a 20 deg cone, and a tangent ogive with an elliptic cross section. The investigation included the effects of nose bluntness and boundary-layer trips. The tangent-ogive forebody was also tested in the presence of a short afterbody and with the afterbody attached. Static longitudinal and lateral/directional stability data were obtained. The investigation was conducted to investigate the existence of large side forces and yawing moments at high angles of attack and zero sideslip. It was found that all of the forebodies experience steady side forces that start at angles of attack of from 20 deg to 35 deg and exist to as high as 80 deg, depending on forebody shape. The side is as large as 1.6 times the normal force and is generally repeatable with increasing and decreasing angle of attack and, also, from test to test. The side force is very sensitive to the nature of the boundary layer, as indicated by large changes with boundary trips. The maximum side force caries considerably with Reynolds number and tends to decrease with increasing Mach number. The direction of the side force is sensitive to the body geometry near the nose. The angle of attack of onset of side force is not strongly influenced by Reynolds number or Mach number but varies with forebody shape. Maximum normal force often occurs at angles of attack near 60 deg. The effect of the elliptic cross section is to reduce the angle of onset by about 10 deg compared to that of an equivalent circular forebody with the same fineness ratio. The short afterbody reduces the angle of onset by about 5 deg.

  2. Flight Test of a 40-Foot Nominal Diameter Disk-Gap-Band Parachute Deployed at a Mach Number of 2.72

    NASA Technical Reports Server (NTRS)

    1968-01-01

    Flight Test of a 40-Foot Nominal Diameter Disk-Gap-Band Parachute Deployed at a Mach Number of 2.72 and a Dynamic Pressure of 9.7 Pounds per Square Foot. A 40-foot-nominal-diameter (12.2 meter) disk-gap-band parachute was flight tested as part of the NASA Supersonic Planetary Entry Decelerator (SPED-I) Program. The test parachute was deployed from an instrumented payload by means of a deployment mortar when the payload was at an altitude of 158,500 feet (48.2 kilometers), a Mach number of 2.72, and a free-stream dynamic pressure of 9.7 pounds per foot(exp 2) (465 newtons per meter(exp 2)). Suspension line stretch occurred 0.46 second after mortar firing and the resulting snatch force loading was -8.lg. The maximum acceleration experienced by the payload due to parachute opening was -27.2g at 0.50 second after the snatch force peak for a total elapsed time from mortar firing of 0.96 second. Canopy-shape variations occurred during the higher Mach number portion of the flight test (M greater than 1.4) and the payload was subjected to large amplitude oscillatory loads. A calculated average nominal axial-force coefficient ranged from about 0.25 immediately after the first canopy opening to about 0.50 as the canopy attained a steady inflated shape. One gore of the test parachute was damaged when the deployment bag with mortar lid passed through it from behind approximately 2 seconds after deployment was initiated. Although the canopy damage caused by the deployment bag penetration had no apparent effect on the functional capability of the test parachute, it may have affected parachute performance since the average effective drag coefficient of 0.48 was 9 percent less than that of a previously tested parachute of the same configuration. [Entire movie available on DVD from CASI as Doc ID 20070031009. Contact help@sti.nasa.gov

  3. Free-Flight Skin Temperature and Pressure Measurements on a Slightly Blunted 25 Deg Cone-Cylinder-Flare Configuration to a Mach Number of 9.89

    NASA Technical Reports Server (NTRS)

    Bond, Aleck C.; Rumsey, Charles B.

    1957-01-01

    Skin temperatures and surface pressures have been measured on a slightly blunted cone-cylinder-flare configuration to a maximum Mach number of 9.89 with a rocket-propelled model. The cone had a t o t a l angle of 25 deg and the flare had a 10 deg half-angle. Temperature data were obtained at eight cone locations, four cylinder locations, and seven flare locations; pressures were measured at one cone location, one cylinder location, and three flare locations. Four stages of propulsion were utilized and a reentry type of trajectory was employed in which the high-speed portion of flight was obtained by firing the last two stages during the descent of the model from a peak altitude of 99,400 feet. The Reynolds number at peak Mach number was 1.2 x 10(exp 6) per foot of model length. The model length was 6.68 feet. During the higher speed portions of flight, temperature measurements along one element of the nose cone indicated that the boundary layer was probably laminar, whereas on the opposite side of the nose the measurements indicated transitional or turbulent flow. Temperature distributions along one meridian of the model showed the flare to have the highest temperatures and the cylinder generally to have the lowest. A maximum temperature of 970 F was measured on the cone element showing the transitional or turbulent flow; along the opposite side of the model, the maximum temperatures of the cone, cylinder, and flare were 545 F, 340 F, and 680 F, respectively, at the corresponding time.

  4. Overall fluctuating pressure levels on prospective space shuttle launch configurations at Mach numbers from 0.8 to 2.2

    NASA Technical Reports Server (NTRS)

    Dods, J. B., Jr.; Hanly, R. D.; Efting, J. H.

    1973-01-01

    Overall fluctuating pressure levels of seven space shuttle launch configurations are presented. The model was a 4-percent-scale space shuttle vehicle, tested in both a 11- by 11-foot transonic wind tunnel and a 9- by 7-foot supersonic wind tunnel. Mach numbers varied from 0.8 to 2.2, and the angle of attack range was from -8 deg to 8 deg at angles of sideslip of -5 deg, and 5 deg. The model configurations included both series-burn and parallel-burn configurations, two canopy configurations, two positions of the orbiter nose relative to the HO tank nose and two HO tank nose-cone angles (15 deg and 20 deg). The fluctuating pressure levels are presented in three forms.

  5. The effects on propulsion-induced aerodynamic forces of vectoring a partial-span rectangular jet at Mach numbers from 0.40 to 1.20

    NASA Technical Reports Server (NTRS)

    Capone, F. J.

    1975-01-01

    An investigation was conducted in the Langley 16-foot transonic tunnel to determine the induced lift characteristics of a vectored thrust concept in which a rectangular jet exhaust nozzle was located in the fuselage at the wing trailing edge. The effects of nozzle deflection angles of 0 deg to 45 deg were studied at Mach numbers from 0.4 to 1.2, at angles of attack up to 14 deg, and with thrust coefficients up to 0.35. Separate force balances were used to determine total aerodynamic and thrust forces as well as thrust forces which allowed a direct measurement of jet turning angle at forward speeds. Wing pressure loading and flow characteristics using oil flow techniques were also studied.

  6. On the calculation of the complex wavenumber of plane waves in rigid-walled low-Mach-number turbulent pipe flows

    NASA Astrophysics Data System (ADS)

    Weng, Chenyang; Boij, Susann; Hanifi, Ardeshir

    2015-10-01

    A numerical method for calculating the wavenumbers of axisymmetric plane waves in rigid-walled low-Mach-number turbulent flows is proposed, which is based on solving the linearized Navier-Stokes equations with an eddy-viscosity model. In addition, theoretical models for the wavenumbers are reviewed, and the main effects (the viscothermal effects, the mean flow convection and refraction effects, the turbulent absorption, and the moderate compressibility effects) which may influence the sound propagation are discussed. Compared to the theoretical models, the proposed numerical method has the advantage of potentially including more effects in the computed wavenumbers. The numerical results of the wavenumbers are compared with the reviewed theoretical models, as well as experimental data from the literature. It shows that the proposed numerical method can give satisfactory prediction of both the real part (phase shift) and the imaginary part (attenuation) of the measured wavenumbers, especially when the refraction effects or the turbulent absorption effects become important.

  7. Performance Comparison at Mach Numbers 1.8 and 2.0 of Full Scale and Quarter Scale Translating-Spike Inlets

    NASA Technical Reports Server (NTRS)

    Anderson, B. H.; Dryer, M.; Hearth, D. P.

    1957-01-01

    The performance of a full-scale translating-spike inlet was obtained at Mach numbers of 1.8 and 2.0 and at angles of attach from 0 deg to 6 deg. Comparisons were made between the full-scale production inlet configuration and a geometrically similar quarter-scale model. The inlet pressure-recovery, cowl pressure-distribution, and compressor-face distortion characteristics of the full-scale inlet agreed fairly well with the quarter-scale results. In addition, the results indicated that bleeding around the periphery ahead of the compressor-face station improved pressure recovery and compressor-face distortion, especially at angle of attack.

  8. Pitot pressure measurements in flow fields behind circular-arc nozzles with exhaust jets at subsonic free-stream Mach numbers. [langley 16 foot transonic tunnel

    NASA Technical Reports Server (NTRS)

    Mason, M. L.; Putnam, L. E.

    1979-01-01

    The flow field behind a circular arc nozzle with exhaust jet was studied at subsonic free stream Mach numbers. A conical probe was used to measure the pitot pressure in the jet and free stream regions. Pressure data were recorded for two nozzle configurations at nozzle pressure ratios of 2.0, 2.9, and 5.0. At each set of test conditions, the probe was traversed from the jet center line into the free stream region at seven data acquisition stations. The survey began at the nozzle exit and extended downstream at intervals. The pitot pressure data may be applied to the evaluation of computational flow field models, as illustrated by a comparison of the flow field data with results of inviscid jet plume theory.

  9. Low Mach Number Modeling of Convection in Helium Shells on Sub-Chandrasekhar White Dwarfs. II. Bulk Properties of Simple Models

    NASA Astrophysics Data System (ADS)

    Jacobs, A. M.; Zingale, M.; Nonaka, A.; Almgren, A. S.; Bell, J. B.

    2016-08-01

    The dynamics of helium shell convection driven by nuclear burning establish the conditions for runaway in the sub-Chandrasekhar-mass, double-detonation model for SNe Ia, as well as for a variety of other explosive phenomena. We explore these convection dynamics for a range of white dwarf core and helium shell masses in three dimensions using the low Mach number hydrodynamics code MAESTRO. We present calculations of the bulk properties of this evolution, including time-series evolution of global diagnostics, lateral averages of the 3D state, and the global 3D state. We find a variety of outcomes, including quasi-equilibrium, localized runaway, and convective runaway. Our results suggest that the double-detonation progenitor model is promising and that 3D dynamic convection plays a key role.

  10. A modification to linearized theory for prediction of pressure loadings on lifting surfaces at high supersonic Mach numbers and large angles of attack

    NASA Technical Reports Server (NTRS)

    Carlson, H. W.

    1979-01-01

    A new linearized-theory pressure-coefficient formulation was studied. The new formulation is intended to provide more accurate estimates of detailed pressure loadings for improved stability analysis and for analysis of critical structural design conditions. The approach is based on the use of oblique-shock and Prandtl-Meyer expansion relationships for accurate representation of the variation of pressures with surface slopes in two-dimensional flow and linearized-theory perturbation velocities for evaluation of local three-dimensional aerodynamic interference effects. The applicability and limitations of the modification to linearized theory are illustrated through comparisons with experimental pressure distributions for delta wings covering a Mach number range from 1.45 to 4.60 and angles of attack from 0 to 25 degrees.

  11. Effect of gaseous and solid simulated jet plumes on a 040A space shuttle launch configuration at Mach numbers from 1.6 to 2.2

    NASA Technical Reports Server (NTRS)

    Lanfranco, M. J.; Sparks, V. W.; Kavanaugh, A. T.

    1973-01-01

    An experimental investigation was conducted in a 9- by 7-foot supersonic wind tunnel to determine the effect of plume-induced flow separation and aspiration effects due to operation of both the orbiter and the solid rocket motors on a 0.019-scale model of the launch configuration of the space shuttle vehicle. Longitudinal and lateral-directional stability data were obtained at Mach numbers of 1.6, 2.0, and 2.2 with and without the engines operating. The plumes exiting from the engines were simulated by a cold gas jet supplied by an auxiliary 200 atmosphere air supply system, and by solid body plume simulators. Comparisons of the aerodynamic effects produced by these two simulation procedures are presented. The data indicate that the parameters most significantly affected by the jet plumes are the pitching moment, the elevon control effectiveness, the axial force, and the orbiter wing loads.

  12. Sonic-boom measurements for SR-71 aircraft operating at Mach numbers to 3.0 and altitudes to 24384 meters

    NASA Technical Reports Server (NTRS)

    Maglieri, D. J.; Huckel, V.; Henderson, H. R.

    1972-01-01

    Sonic-boom pressure signatures produced by the SR-71 aircraft at altitudes from 10,668 to 24,384 meters and Mach numbers 1.35 to 3.0 were obtained as an adjunct to the sonic boom evaluation program relating to structural and subjective response which was conducted in 1966-1967 time period. Approximately 2000 sonic-boom signatures from 33 flights of the SR-71 vehicle and two flights of the F-12 vehicle were recorded. Measured ground-pressure signatures for both on-track and lateral measuring station locations are presented and the statistical variations of the overpressure, positive impulse, wave duration, and shock-wave rise time are illustrated.

  13. Investigation at Mach Number 1.91 of Side and Base Pressure Distributions over Conical Boattails Without and with Jet Flow Issuing from Base

    NASA Technical Reports Server (NTRS)

    Cortright, Edgar M , Jr; Schroeder, Albert H

    1951-01-01

    Experimental side and bade pressure distributions over a series of conical boattails without and with jet flow from the base are presented at a Mach number of 1.91. For the case of no jet flow the methods of characteristics and linearized theory are shown to overpredict the side pressure drag. A semi-empirical theory is presented to predict the effect of boattail angle on base pressure. With the boattail extending to a sharp edge at the nozzle exit, the over-pressure jet is shown to decrease the side pressure drag. Presence of an annular base may eliminate the effect of the jet on the side pressure drag, but the jet effect on the base pressure drag may greatly increase or decrease the total boattail drag.

  14. Comparison of Experimental and Theoretical Zero-Lift Wave-Drag Results for Various Wing-Body-Tail Combinations at Mach Numbers up to 1.9

    NASA Technical Reports Server (NTRS)

    Petersen, R. B.

    1957-01-01

    Comparisons are made of experimental and theoretical zero-lift wave drag for several nose shapes, wing-body combinations, and models of current airplanes at Mach numbers up to 1.0. The experimental data were obtained from tests in the Ames 6- by6-foot supersonic wind tunnel and at the NACA Wallops Island facility. The theoretical drag was found by use of linear theory utilizing model area distributions. The agreement between theoretical and experimental zero-lift wave-drag coefficients was generally very good, especially for a fuselage or for fuselage-wing combinations that were vertically symmetrical. For other models that had rapid changes in body shape and/or were not vertically symmetrical, the agreement of theory with experiment ranged from fair to poor, depending on the severity of the change in shape.

  15. Jet Interference Effects on a Model of a Single-Engine Four Jet V/STOL Airplane at Mach Numbers from 0.60 to 1.00

    NASA Technical Reports Server (NTRS)

    Schmeer, James W.; Runckel, Jack F.

    1962-01-01

    An investigation was conducted in the Langley 16-foot transonic tunnel to determine the interference from four exhaust jets on the aerodynamic characteristics of a model of a V/STOL airplane. The single- engine four-jet turbofan power plant of the airplane was simulated by inducing tunnel airflow through two large side inlets and injecting the decomposition products of hydrogen peroxide into the internal flow. The heated gas mixture was exhausted through four nozzles located on the sides of the fuselage under the wing, two near the wing leading edge and two forward of the trailing edge; the nozzles were deflected downward 1.5 deg and outward 5.0 deg to simulate cruise conditions. The wing of the model was a clipped delta with leading-edge sweep of 40 deg, aspect ratio of 3.06, taper ratio of 0.218, thickness-chord ratio of 0.09 at the root and 0.07 at the tip, and 10 deg negative dihedral. Aerodynamic and longitudinal stability coefficients were obtained for the model with the tail removed, and for horizontal-tail incidences of 0 deg and -5 deg. Data were obtained at Mach numbers from 0.60 to 1.00, angles of attack from 0 deg to 12 deg, and with jet total-pressure ratios up to 3.1. Jet operation generally caused a decrease in lift, an increase in pitching-moment coefficient, and a decrease in longitudinal stability at subsonic speeds. The jet interference effects on drag were detrimental at a Mach number of 0.60 and favorable at higher speeds for cruising-flight attitudes.

  16. Interstellar Neutral Helium in the Heliosphere from IBEX Observations. III. Mach Number of the Flow, Velocity Vector, and Temperature from the First Six Years of Measurements

    NASA Astrophysics Data System (ADS)

    Bzowski, M.; Swaczyna, P.; Kubiak, M. A.; Sokół, J. M.; Fuselier, S. A.; Galli, A.; Heirtzler, D.; Kucharek, H.; Leonard, T. W.; McComas, D. J.; Möbius, E.; Schwadron, N. A.; Wurz, P.

    2015-10-01

    We analyzed observations of interstellar neutral helium (ISN He) obtained from the Interstellar Boundary Explorer (IBEX) satellite during its first six years of operation. We used a refined version of the ISN He simulation model, presented in the companion paper by Sokół et al. (2015b), along with a sophisticated data correlation and uncertainty system and parameter fitting method, described in the companion paper by Swaczyna et al. We analyzed the entire data set together and the yearly subsets, and found the temperature and velocity vector of ISN He in front of the heliosphere. As seen in the previous studies, the allowable parameters are highly correlated and form a four-dimensional tube in the parameter space. The inflow longitudes obtained from the yearly data subsets show a spread of ˜6°, with the other parameters varying accordingly along the parameter tube, and the minimum χ2 value is larger than expected. We found, however, that the Mach number of the ISN He flow shows very little scatter and is thus very tightly constrained. It is in excellent agreement with the original analysis of ISN He observations from IBEX and recent reanalyses of observations from Ulysses. We identify a possible inaccuracy in the Warm Breeze parameters as the likely cause of the scatter in the ISN He parameters obtained from the yearly subsets, and we suppose that another component may exist in the signal or a process that is not accounted for in the current physical model of ISN He in front of the heliosphere. From our analysis, the inflow velocity vector, temperature, and Mach number of the flow are equal to λISNHe = 255.°8 ± 0.°5, βISNHe = 5.°16 ± 0.°10, TISNHe = 7440 ± 260 K, vISNHe = 25.8 ± 0.4 km s-1, and MISNHe = 5.079 ± 0.028, with uncertainties strongly correlated along the parameter tube.

  17. Flight Test of a 40-Foot Nominal Diameter Disk-Gap-Band Parachute Deployed at a Mach Number of 3.31

    NASA Technical Reports Server (NTRS)

    1969-01-01

    Flight Test of a 40-Foot Nominal Diameter Disk-Gap-Band Parachute Deployed at a Mach Number of 3.31 and a Dynamic Pressure of 10.6 Pounds per Square Foot. A 40-foot-nominal-diameter (12.2 meter) disk-gap-band parachute was flight tested as part of the NASA supersonic high altitude parachute experiment (SHAPE) program. The test parachute (which included an experimental energy absorber in the attachment riser) was deployed from an instrumented payload by means of a deployment mortar when the payload was at a Mach number of 3.31 and a free-stream dynamic pressure of 10.6 pounds per square foot (508 newtons per square meter). The parachute deployed properly, the canopy inflating to a full-open condition at 1.03 seconds after mortar firing. The first full inflation of the canopy was immediately followed by a partial collapse with subsequent oscillations of the frontal area from about 30 to 75 percent of the full-open frontal area. After 1.07 seconds of operation, a large tear appeared in the cloth near the canopy apex. This tear was followed by two additional tears shortly thereafter. It was later determined that a section of the canopy cloth was severely weakened by the effects of aerodynamic heating. As a result of the damage to the disk area of the canopy, the parachute performance was significantly reduced; however, the parachute remained operationally intact throughout the flight test and the instrumented payload was recovered undamaged. [Entire movie available on DVD from CASI as Doc ID 20070031012. Contact help@sti.nasa.gov

  18. Exploratory Investigation of the Effects of Boundary-Layer Control on the Pressure-Recovery Characteristics of a Circular Internal-Contraction Inlet with Translating Centerbody at Mach Numbers of 2.00 and 2.35

    NASA Technical Reports Server (NTRS)

    Martin, Norman J.

    1959-01-01

    Exploratory tests of a circular internal-contraction inlet were made at Mach numbers of 2.00 and 2.35 to determine the effect of a cowl-type boundary-layer control located downstream of the inlet throat. The inlet was designed for a Mach number of 2.5. Tests were also made of the inlet modified to correspond to design Mach numbers of 2.35 and 2.25. Surveys near the minimum area section of the inlet without boundary-layer control indicated maximum averaged pressure recoveries between 0.90 and 0.92 at a free-stream Mach number, M(sub infinity), of 2.35 for the inlets. Farther downstream, after partial subsonic diffusion, a maximum pressure recovery of 0.842 was obtained with the inlet at M(sub infinity) = 2.35. The pressure recovery of the inlet was increased by 0.03 at a Mach number of 2.35 and decreased by 0.02 at a Mach number of 2.00 by the application of cowl-type boundary-layer control. Further investigation with the inlet without bleed demonstrated that an increase of angle of attack from 0 deg to 3 deg reduced the pressure recovery 0.04. The effect of Reynolds number was to increase pressure recovery 0.07 (from 0.785 to 0.855) with an increase in Reynolds number (based on inlet diameter) from 0.79 x 10(exp 6) to 3.19 x 10(exp 6).

  19. Lift, Drag, and Pitching Moments of an Arrow Wing Having 80 Degree of Sweepback at Mach Numbers from 2.48 to 3.51 and Reynolds Numbers up to 11.0 Million

    NASA Technical Reports Server (NTRS)

    Hopkins, Edward J.; Jillie, Don W.; Levin, Alan D.

    1959-01-01

    Measurements were made of the lift, drag, and pitching moments on an arrow wing (taper ratio of zero) having an aspect ratio of 1.4 and a leading-edge sweepback of 80 (degrees). The wing was designed to have a subsonic leading-edge and a Clark-Y airfoil with a thickness ratio of 12 percent of the chord perpendicular to the wing leading edge. The wing was tested both with and without the wing tips bent upward in an attempt to alleviate possible flow separation in the vicinity of the wing tips. Small jets of air were used to fix transition near the wing leading edge. Force results are presented for Mach numbers of 2.48, 2.75, 3.04, 3.28, and 3.51 at Reynolds numbers of 3.5 and 9.0 million and for a Mach number of 3.04 at a Reynolds number of 11.0 million. The measured aerodynamic characteristics are compared with those estimated by linear theory. The maximum lift-drag ratio measured was much less than that predicted. This difference is attributed to lack of full leading-edge thrust and to the experimental lift-curve slope being about 20 percent below the theoretical value.

  20. Conventional and nonconventional global Alfven eigenmodes in stellarators

    SciTech Connect

    Kolesnichenko, Ya. I.; Lutsenko, V. V.; Weller, A.; Werner, A.; Yakovenko, Yu. V.; Geiger, J.; Fesenyuk, O. P.

    2007-10-15

    Conditions of the existence of the Global Alfven Eigenmodes (GAE) and Nonconventional Global Alfven Eigenmodes (NGAE) predicted for stellarators by Ya. I. Kolesnichenko et al. [Phys. Rev. Lett. 94, 165004 (2005)] have been obtained. It is found that they depend on the nature of the rotational transform and that conditions for NGAE can be most easily satisfied in currentless stellarators. It is shown that the plasma compressibility may play an important role for the modes with the frequency about or less than that of the Toroidicity-induced Alfven Eigenmodes. It is found that features of the Alfven continuum in the vicinity of the k{sub parallel}=0 radius (k{sub parallel}) is the longitudinal wave number) can be very different, depending on a parameter which we refer to as 'the sound parameter'. Specific calculations modeling low-frequency Alfven instabilities in the stellarator Wendelstein 7-AS [A. Weller et al., Phys. Plasmas 8, 931 (2001)] are carried out, which are in reasonable agreement with the observations. It is emphasized that experimental data on low-frequency Alfvenic activity can be used for the reconstruction of the profile of the rotational transform. The mentioned results are obtained with the use of the equations derived in this paper for the GAE/NGAE modes and of the codes COBRAS and BOA-fe.

  1. Investigation of Some Wake Vortex Characteristics of an Inclined Ogive-Cylinder Body at Mach Number 2

    NASA Technical Reports Server (NTRS)

    Jorgensen, Leland H; Perkins, Edward W

    1958-01-01

    For a body consisting of a fineness-ratio-3 ogival nose tangent to a cylindrical afterbody 7.3 diameters long, pitot-pressure distributions in the flow field, pressure distributions over the body, and downwash distributions along a line through the vortex centers have been measured for angles of attack to 20 degrees. The Reynolds numbers, based on body diameter, were 0.15 x 10 to the 6th power and 0.44 x 10 to the 6th power. Comparisons of computed and measured vortex paths and downwash distributions are made. (author)

  2. Schlieren photographs and internal pressure distributions for three-dimensional sidewall-compression scramjet inlets at a Mach number of 6 in CF4

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.

    1993-01-01

    Three-dimensional sidewall-compression scramjet inlets with leading-edge sweeps of 30 deg and 70 deg were tested in the Langley Hypersonic CF4 Tunnel at a Mach number of 6 and a free-stream ratio of specific heats of 1.2. The parametric effects of leading-edge sweep, cowl position, contraction ratio, and Reynolds number were investigated. The models were instrumented with static pressure orifices distributed on the sidewalls, baseplate, and cowl. Schlieren movies were made of selected tunnel runs for flow visualization of the entrance plane and cowl region. Although these movies could not show the internal flow, the effect of the internal flow on the external flow was evident by way of spillage. The purpose is to provide a preliminary data release for the investigation. The models, facility, and testing methods are described, and the test matrix and a tabulation of tunnel runs are provided. Line plots highlighting the stated parametric effects and a representative set of schlieren photographs are presented without analysis.

  3. Longitudinal aerodynamic performance of a series of power-law and minimum wave drag bodies at Mach 6 and several Reynolds numbers

    NASA Technical Reports Server (NTRS)

    Ashby, G. C., Jr.

    1974-01-01

    Experimental data have been obtained for two series of bodies at Mach 6 and Reynolds numbers, based on model length, from 1.4 million to 9.5 million. One series consisted of axisymmetric power-law bodies geometrically constrained for constant length and base diameter with values of the exponent n of 0.25, 0.5, 0.6, 0.667, 0.75, and 1.0. The other series consisted of positively and negatively cambered bodies of polygonal cross section, each having a constant longitudinal area distribution conforming to that required for minimizing zero-lift wave drag at hypersonic speeds under the geometric constraints of given length and volume. At the highest Reynolds number, the power-law body for minimum drag is blunter (exponent n lower) than predicted by inviscid theory (n approximately 0.6 instead of n = 0.667); however, the peak value of lift-drag ratio occurs at n = 0.667. Viscous effects were present on the bodies of polygonal cross section but were less pronounced than those on the power-law bodies. The trapezoidal bodies with maximum width at the bottom were found to have the highest maximum lift-drag ratio and the lowest mimimum drag.

  4. Aerodynamic Performance and Static Stability at Mach Number 3.3 of an Aircraft Configuration Employing Three Triangular Wing Panels and a Body Equal Length

    NASA Technical Reports Server (NTRS)

    James, Carlton S.

    1960-01-01

    An aircraft configuration, previously conceived as a means to achieve favorable aerodynamic stability characteristics., high lift-drag ratio, and low heating rates at high supersonic speeds., was modified in an attempt to increase further the lift-drag ratio without adversely affecting the other desirable characteristics. The original configuration consisted of three identical triangular wing panels symmetrically disposed about an ogive-cylinder body equal in length to the root chord of the panels. This configuration was modified by altering the angular disposition of the wing panels, by reducing the area of the panel forming the vertical fin, and by reshaping the body to produce interference lift. Six-component force and moment tests of the modified configuration at combined angles of attack and sideslip were made at a Mach number of 3.3 and a Reynolds number of 5.46 million. A maximum lift-drag ratio of 6.65 (excluding base drag) was measured at a lift coefficient of 0.100 and an angle of attack of 3.60. The lift-drag ratio remained greater than 3 up to lift coefficient of 0.35. Performance estimates, which predicted a maximum lift-drag ratio for the modified configuration 27 percent greater than that of the original configuration, agreed well with experiment. The modified configuration exhibited favorable static stability characteristics within the test range. Longitudinal and directional centers of pressure were slightly aft of the respective centroids of projected plan-form and side area.

  5. Aerodynamic characteristics at Mach numbers from 0.33 to 1.20 of a wing-body design concept for a hypersonic research airplane

    NASA Technical Reports Server (NTRS)

    Dillon, J. L.; Pittman, J. L.

    1977-01-01

    An experimental investigation of the static aerodynamic characteristics of a model of one design concept for the proposed National Hypersonic Flight Research Facility was conducted in the Langley 8 foot transonic pressure tunnel. The experiment consisted of configuration buildup from the basic body by adding a wing, center vertical tail, and a three module or six module scramjet engine. The freestream test Mach numbers were 0.33, 0.80, 0.90, 0.95, 0.98, 1.10, and 1.20 at Reynolds numbers per meter ranging from 4.8 x 1 million to 10.4 x 1 million. The test angle of attack range was approximately -4 deg to 22 deg at constant angles of sideslip of 0 deg and 4 deg; the angle of sideslip ranged from about -6 deg to 6 deg at constant angles of attack of 0 deg and 17 deg. The elevons were deflected 0 deg, -10 deg, and -20 deg with rudder deflections of 0 deg and 15.6 deg.

  6. Incompressible boundary-layer stability analysis of LFC experimental data for sub-critical Mach numbers. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Berry, S. A.

    1986-01-01

    An incompressible boundary-layer stability analysis of Laminar Flow Control (LFC) experimental data was completed and the results are presented. This analysis was undertaken for three reasons: to study laminar boundary-layer stability on a modern swept LFC airfoil; to calculate incompressible design limits of linear stability theory as applied to a modern airfoil at high subsonic speeds; and to verify the use of linear stability theory as a design tool. The experimental data were taken from the slotted LFC experiment recently completed in the NASA Langley 8-Foot Transonic Pressure Tunnel. Linear stability theory was applied and the results were compared with transition data to arrive at correlated n-factors. Results of the analysis showed that for the configuration and cases studied, Tollmien-Schlichting (TS) amplification was the dominating disturbance influencing transition. For these cases, incompressible linear stability theory correlated with an n-factor for TS waves of approximately 10 at transition. The n-factor method correlated rather consistently to this value despite a number of non-ideal conditions which indicates the method is useful as a design tool for advanced laminar flow airfoils.

  7. The Aerodynamic Characteristics in Pitch of a 1/15-Scale Model of the Grumman F11F-1 Airplane at Mach Numbers of 1.41, 1.61, and 2.01, TED No. NACA DE 390

    NASA Technical Reports Server (NTRS)

    Driver, Cornelius

    1956-01-01

    Tests have been made in the Langley 4- by 4-foot supersonic pressure tunnel at Mach numbers of 1.41, 1.61, and 2.01 to determine the static longitudinal stability and control characteristics of various arrangements of the Grumman F11F-1 airplane. Tests were made of the complete model and various combinations of its component parts and, in addition, the effects of various body modifications, a revised vertical tail, and wing fences on the longitudinal characteristics were determined. The results indicate that for a horizontal-tail incidence of -10 deg the trim lift coefficient varied from 0.29 at a Mach number of 1.61 to 0.23 at a Mach number of 2.01 with a corresponding decrease in lift-drag trim from 3.72 to 3.15. Stick-position instability was indicated in the low-supersonic-speed range. A photographic-type nose modification resulted in slightly higher values of minimum drag coefficient but did not significantly affect the static stability or lift-curve slope. The minimum drag coefficient for the complete model with the production nose remained essentially constant at 0.047 throughout the Mach number range investigated.

  8. Investigation of the NACA 4-(3)(8)-045 Two-blade Propellers at Forward Mach Numbers to 0.725 to Determine the Effects of Compressibility and Solidity on Performance

    NASA Technical Reports Server (NTRS)

    Stack, John; Draley, Eugene C; Delano, James B; Feldman, Lewis

    1950-01-01

    As part of a general investigation of propellers at high forward speeds, tests of two 2-blade propellers having the NACA 4-(3)(8)-03 and NACA 4-(3)(8)-45 blade designs have been made in the Langley 8-foot high-speed tunnel through a range of blade angle from 20 degrees to 60 degrees for forward Mach numbers from 0.165 to 0.725 to establish in detail the changes in propeller characteristics due to compressibility effects. These propellers differed primarily only in blade solidity, one propeller having 50 percent and more solidity than the other. Serious losses in propeller efficiency were found as the propeller tip Mach number exceeded 0.91, irrespective of forward speed or blade angle. The magnitude of the efficiency losses varied from 9 percent to 22 percent per 0.1 increase in tip Mach number above the critical value. The range of advance ratio for peak efficiency decreased markedly with increase of forward speed. The general form of the changes in thrust and power coefficients was found to be similar to the changes in airfoil lift coefficient with changes in Mach number. Efficiency losses due to compressibility effects decreased with increase of blade width. The results indicated that the high level of propeller efficiency obtained at low speeds could be maintained to forward sea-level speeds exceeding 500 miles per hour.

  9. A high-order numerical algorithm for DNS of low-Mach-number reactive flows with detailed chemistry and quasi-spectral accuracy

    NASA Astrophysics Data System (ADS)

    Motheau, E.; Abraham, J.

    2016-05-01

    A novel and efficient algorithm is presented in this paper to deal with DNS of turbulent reacting flows under the low-Mach-number assumption, with detailed chemistry and a quasi-spectral accuracy. The temporal integration of the equations relies on an operating-split strategy, where chemical reactions are solved implicitly with a stiff solver and the convection-diffusion operators are solved with a Runge-Kutta-Chebyshev method. The spatial discretisation is performed with high-order compact schemes, and a FFT based constant-coefficient spectral solver is employed to solve a variable-coefficient Poisson equation. The numerical implementation takes advantage of the 2DECOMP&FFT libraries developed by [1], which are based on a pencil decomposition method of the domain and are proven to be computationally very efficient. An enhanced pressure-correction method is proposed to speed up the achievement of machine precision accuracy. It is demonstrated that a second-order accuracy is reached in time, while the spatial accuracy ranges from fourth-order to sixth-order depending on the set of imposed boundary conditions. The software developed to implement the present algorithm is called HOLOMAC, and its numerical efficiency opens the way to deal with DNS of reacting flows to understand complex turbulent and chemical phenomena in flames.

  10. A High-Order Immersed Boundary Method for Acoustic Wave Scattering and Low-Mach Number Flow-Induced Sound in Complex Geometries

    PubMed Central

    Seo, Jung Hee; Mittal, Rajat

    2010-01-01

    A new sharp-interface immersed boundary method based approach for the computation of low-Mach number flow-induced sound around complex geometries is described. The underlying approach is based on a hydrodynamic/acoustic splitting technique where the incompressible flow is first computed using a second-order accurate immersed boundary solver. This is followed by the computation of sound using the linearized perturbed compressible equations (LPCE). The primary contribution of the current work is the development of a versatile, high-order accurate immersed boundary method for solving the LPCE in complex domains. This new method applies the boundary condition on the immersed boundary to a high-order by combining the ghost-cell approach with a weighted least-squares error method based on a high-order approximating polynomial. The method is validated for canonical acoustic wave scattering and flow-induced noise problems. Applications of this technique to relatively complex cases of practical interest are also presented. PMID:21318129

  11. An algebraic variational multiscale-multigrid method for large-eddy simulation of turbulent variable-density flow at low Mach number

    NASA Astrophysics Data System (ADS)

    Gravemeier, Volker; Wall, Wolfgang A.

    2010-08-01

    An algebraic variational multiscale-multigrid method is proposed for large-eddy simulation of turbulent variable-density flow at low Mach number. Scale-separating operators generated by level-transfer operators from plain aggregation algebraic multigrid methods enable the application of modeling terms to selected scale groups (here, the smaller of the resolved scales) in a purely algebraic way. Thus, for scale separation, no additional discretization besides the basic one is required, in contrast to earlier approaches based on geometric multigrid methods. The proposed method is thoroughly validated via three numerical test cases of increasing complexity: a Rayleigh-Taylor instability, turbulent channel flow with a heated and a cooled wall, and turbulent flow past a backward-facing step with heating. Results obtained with the algebraic variational multiscale-multigrid method are compared to results obtained with residual-based variational multiscale methods as well as reference results from direct numerical simulation, experiments and LES published elsewhere. Particularly, mean and various second-order velocity and temperature results obtained for turbulent channel flow with a heated and a cooled wall indicate the higher prediction quality achievable when adding a small-scale subgrid-viscosity term within the algebraic multigrid framework instead of residual-based terms accounting for the subgrid-scale part of the non-linear convective term.

  12. Noise characteristics of jet flap type exhaust flows. [effects of Mach number, slot nozzle aspect ratio, and flap length on radiated sound power

    NASA Technical Reports Server (NTRS)

    Schrecker, G. O.; Maus, J. R.

    1974-01-01

    An experimental investigation of the aerodynamic noise and flow field characteristics of internal-flow jet-augmented flap configurations (abbreviated by the term jet flap throughout the study) is presented. The first part is a parametric study of the influence of the Mach number (subsonic range only), the slot nozzle aspect ratio and the flap length on the overall radiated sound power and the spectral composition of the jet noise, as measured in a reverberation chamber. In the second part, mean and fluctuating velocity profiles, spectra of the fluctuating velocity and space correlograms were measured in the flow field of jet flaps by means of hot-wire anemometry. Using an expression derived by Lilley, an attempt was made to estimate the overall sound power radiated by the free mixing region that originates at the orifice of the slot nozzle (primary mixing region) relative to the overall sound power generated by the free mixing region that originates at the trailing edge of the flap (secondary mixing region). It is concluded that at least as much noise is generated in the secondary mixing region as in the primary mixing region. Furthermore, the noise generation of the primary mixing region appears to be unaffected by the presence of a flap.

  13. Investigations of the Deterioration of 22 Refractory Materials in a Mach Number 2 Jet at a Stagnation Temperature of 3,800 F

    NASA Technical Reports Server (NTRS)

    Lewis, B. W.

    1961-01-01

    A limited investigation of the deterioration characteristics of 22 refractory materials was conducted by exposing them to a stagnation temperature of 3,800 F in a Mach number 2 ceramic-heated jet at the Langley Research Center. The materials tested were six materials whose major constituent was silicon carbide, five cermets whose major constituent was titanium carbide, six materials whose major constituents were metal borides, four cermets containing alumina, and one silicon nitride model. Tests consisted of obtaining weight change and appearance changes for 1/2-inch-diameter hemispherical-nose cylindrical models exposed to the air jet for 30 seconds at a time for a total of four runs or 2 minutes exposure. Curves of weight changes plotted against the number of 30-second tests in the jet were obtained. Estimates of average surface temperature near the stagnation point of the model were obtained by use of a special temperature-measuring camera. The models were examined before and after the completion of the tests for possible changes in microstructure; no significant changes were found. The data obtained were analyzed with the view that the oxidation characteristics of the materials were the main factor in deterioration of the materials under the conditions of the tests. It was concluded that only those materials which changed in weight the least could be recommended for further extensive application-oriented evaluations. The following materials fell in this category: silicon carbide - silicon, chromium - 28-percent alumina cermet, titanium boride - 5-percent boron carbide. The remainder of the materials tested had oxidation characteristics which appeared to be too severely limiting of their general applications to flight vehicles.

  14. Side forces on a tangent ogive forebody with a fineness ratio of 3.5 at high angles of attack and Mach numbers from 0.1 to 0.7

    NASA Technical Reports Server (NTRS)

    Keener, E. R.; Chapman, G. T.; Cohen, L.; Taleghani, J.

    1977-01-01

    An experimental investigation was conducted in the Ames 12-Foot Wind Tunnel to determine the subsonic aerodynamic characteristics, at high angles of attack, of a tangent ogive forebody with a fineness ratio of 3.5. The investigation included the effects of nose bluntness, nose strakes, nose booms, a simulated canopy, and boundary-layer trips. The forebody was also tested with a short afterbody attached. Static longitudinal and lateral-directional stability data were obtained at Reynolds numbers ranging from 0.3 mil. to 3.8 mil. (based on base diameter) at a Mach number of 0.25, and at a Reynolds number of 0.8 mil. at Mach numbers ranging from 0.1 to 0.7. Angle of attack was varied from 0 to 88 deg at zero sideslip, and the sideslip angle was varied from -10 to 30 deg at angles of attack of 40, 55, and 70 deg.

  15. Reynolds number effects on the aerodynamic characteristics of irregular planform wings at Mach number 0.3. [in the Ames 12 ft pressure wind tunnel

    NASA Technical Reports Server (NTRS)

    Kruse, R. L.; Lovette, G. H.; Spencer, B., Jr.

    1977-01-01

    The subsonic aerodynamic characteristics of a series of irregular planform wings were studied in wind tunnel tests conducted at M = 0.3 over a range of Reynolds numbers from 1.6 million to 26 million/m. The five basic wing planforms varied from a trapezoidal to a delta shape. Leading edge extensions, added to the basic shape, varied in approximately 5 deg increments from the wing leading edge sweep-back angle to a maximum 80 deg. Most of the tests were conducted using an NACA 0008 airfoil section with grit boundary layer trips. Tests were also conducted using an NACA 0012 airfoil section and an 8% thick wedge. In addition, the effect of free transition (no grit) was investigated. A body was used on all models.

  16. Phenomenology of Compressional Alfven Eigenmodes

    SciTech Connect

    E.D. Fredrickson; N.N. Gorelenkov; J. Menard

    2004-05-13

    Coherent oscillations with frequency 0.3 {le} {omega}/{omega}{sub ci} {le} 1, are seen in the National Spherical Torus Experiment [M. Ono, S.M. Kaye, Y-K.M. Peng, et al., Nucl. Fusion 40, 557 (2000)]. This paper presents new data and analysis comparing characteristics of the observed modes to the model of compressional Alfven eigenmodes (CAE). The toroidal mode number has been measured and is typically between 7 < n < 9. The polarization of the modes, measured using an array of four Mirnov coils, is found to be compressional. The frequency scaling of the modes agrees with the predictions of a numerical 2-D code, but the detailed structure of the spectrum is not captured with the simple model. The fast ion distribution function, as calculated with the beam deposition code in TRANSP [R.V. Budny, Nucl. Fusion 34, 1247 (1994)], is shown to be qualitatively consistent with the constraints of the Doppler-shifted cyclotron resonance drive model. This model also predicts the observed scaling of the low frequency limit for CAE.

  17. An Investigation of the Pressure Distribution on a 1/15-Scale Model of the Lockheed WS-117L Vehicle Plus Booster "B" at Mach Numbers from 1.55 to 2.35

    NASA Technical Reports Server (NTRS)

    Martin, Norman J.

    1959-01-01

    Pressure coefficients were measured over the vehicle and over the forward part of the booster at Reynolds numbers of 3.0 x 10(exp 6) per foot. Tabular results are presented for two nose shapes at Mach numbers of 1.55, 1.75, 2.00, and 2.35, at angles of attack from -4 deg to +10 deg, and at 0 deg sideslip.

  18. A Pressure-distribution Investigation of the Aerodynamic Characteristics of a Body of Revolution in the Vicinity of a Reflection Plane at Mach Numbers of 1.41 and 2.01

    NASA Technical Reports Server (NTRS)

    Gapcynski, John P; Carlson, Harry W

    1955-01-01

    The changes in the aerodynamic characteristics of a body of revolution with a fineness ratio of 8 have been determined at Mach numbers of 1.41 and 2.01, a Reynolds number, based on body length, of 4.54 x 10 to the 6th power, and angles of incidence of 0 degrees and plus or minus 3 degrees as the position of the body is varied with respect to a reflection plane. The data are compared with theoretical results.

  19. Lift and Pitching-moment Interference Between a Pointed Cylindrical Body and Triangular Wings of Various Aspect Ratios at Mach Numbers of 1.50 and 2.02

    NASA Technical Reports Server (NTRS)

    Nielsen, Jack N; Katzen, Elliott D; Tang, Kenneth K

    1956-01-01

    The lift and pitching-moment characteristics of a body alone, six triangular wings of various aspect ratios, and the combinations were measured at Mach numbers of 1.50 and 2.02 at a Reynolds number of 5.5 million (based on the body length) for angles of attack up to 5.5 degrees. The total lift and pitching-moment interference were determined and compared with theory. The agreement was found to be good.

  20. An investigation of several NACA 1 series axisymmetric inlets at Mach numbers from 0.4 to 1.29. [wind tunnel tests over range of mass-flow ratios and at angle of attack

    NASA Technical Reports Server (NTRS)

    Re, R. J.

    1974-01-01

    An investigation was conducted in the Langley 16-foot transonic tunnel to determine the performance of seven inlets having NACA 1-series contours and one inlet having an elliptical contour over a range of mass-flow ratios and at angle of attack. The inlet diameter ratio varied from 0.81 to 0.89; inlet length ratio varied from 0.75 to 1.25; and internal contraction ratio varied from 1.009 to 1.093. Reynolds number based on inlet maximum diameter varied from 3.4 million at a Mach number of 0.4 to 5.6 million at a Mach number of 1.29.

  1. Investigation of the asymmetric aerodynamic characteristics of cylindrical bodies of revolution with variations in nose geometry and rotational orientation at angles of attack to 58 degrees and Mach numbers to 2

    NASA Technical Reports Server (NTRS)

    Kruse, R. L.; Keener, E. R.; Chapman, G. T.; Claser, G.

    1979-01-01

    Wind-tunnel tests were conducted to investigate the side forces and yawing moments that can occur at high angles of attack and zero sideslip for cylindrical bodies of revolution. Two bodies having several tangent ogive forebodies with fineness ratios of 0.5, 1.5, 2.5, and 3.5 were tested. The forebodies with fineness ratios of 2.5 and 3.5 had several bluntnesses. The cylindrical afterbodies had fineness ratios of 7 and 13. The model components - tip, forebody, and afterbody - were tested in various rotational positions about their axes of symmetry. Most of the tests were conducted at a Mach number of 0.25, a Reynolds number of 0.32 x 10 to the 6th power, and with the afterbody that had a fineness ratio of 7 and with selected forebodies. The effect of Mach number was determined with the afterbody that had a fineness ratio of 13 and with selected forebodies at mach numbers from 0.25 to 2 at Reynolds number = 0.32 X 10 to the 6th power. Maximum angle of attack was 58 deg.

  2. Toroidal Alfven wave stability in ignited tokamaks

    SciTech Connect

    Cheng, C.Z.; Fu, G.Y.; Van Dam, J.W.

    1989-01-01

    The effects of fusion-product alpha particles on the stability of global-type shear Alfven waves in an ignited tokamak plasma are investigated in toroidal geometry. Finite toroidicity can lead to stabilization of the global Alfven eigenmodes, but it induces a new global shear Alfven eigenmodes, which is strongly destabilized via transit resonance with alpha particles. 8 refs., 2 figs.

  3. Formation of quasiparallel Alfven solitons

    NASA Technical Reports Server (NTRS)

    Hamilton, R. L.; Kennel, C. F.; Mjolhus, E.

    1992-01-01

    The formation of quasi-parallel Alfven solitons is investigated through the inverse scattering transformation (IST) for the derivative nonlinear Schroedinger (DNLS) equation. The DNLS has a rich complement of soliton solutions consisting of a two-parameter soliton family and a one-parameter bright/dark soliton family. In this paper, the physical roles and origins of these soliton families are inferred through an analytic study of the scattering data generated by the IST for a set of initial profiles. The DNLS equation has as limiting forms the nonlinear Schroedinger (NLS), Korteweg-de-Vries (KdV) and modified Korteweg-de-Vries (MKdV) equations. Each of these limits is briefly reviewed in the physical context of quasi-parallel Alfven waves. The existence of these limiting forms serves as a natural framework for discussing the formation of Alfven solitons.

  4. Pressure Loads Produced on a Flat-Plate Wing By Rocket Jets Exhausting in a Spanwise Direction Below the Wing and Perpendicular to a Free-Stream Flow of Mach Number 2.0

    NASA Technical Reports Server (NTRS)

    Falanga, Ralph A.; Janos, Joseph J.

    1961-01-01

    An investigation at a Reynolds number per foot of 14.4 x 10(exp 6) was made to determine the pressure loads produced on a flat-plate wing by rocket jets exhausting in a spanwise direction beneath the wing and perpendicular to a free-stream flow of Mach number 2.0. The ranges of the variables involved were (1) nozzle types - one sonic (jet Mach number of 1.00), two supersonic (jet Mach numbers of 1.74 and 3.04),. and one two-dimensional supersonic (jet Mach number of 1.71); (2) vertical nozzle positions beneath the wing of 4, 8 and 12 nozzle-throat diameters; and (3) ratios of rocket-chamber total pressure to free- stream static pressure from 0 to 130. The incremental normal force due to jet interference on the wing varied from one to two times the rocket thrust and generally decreased as the pressure ratio increased. The chordwise coordinate of the incremental-normal-force center of pressure remained upstream of the nozzle center line for the nozzle positions and pressure ratios of the investigation. The chordwise coordinate approached zero as the jet vertical distance beneath the wing increased. In the spanwise direction there was little change due to varying rocket-jet position and pressure ratio. Some boundary-layer flow separation on the wing was observed for the rocket jets close to the wing and at the higher pressure ratios. The magnitude of the chordwise and spanwise pressure distributions due to jet interference was greatest for rocket jets close to the wing and decreased as the jet was displaced farther from the wing. The design procedure for the rockets used is given in the appendix.

  5. Free-Flight Investigation of Heat Transfer to an Unswept Cylinder Subjected to an Incident Shock and Flow Interference from an Upstream Body at Mach Numbers up to 5.50

    NASA Technical Reports Server (NTRS)

    Carter, Howard S.; Carr, Robert E.

    1961-01-01

    Heat-transfer rates have been measured in free flight along the stagnation line of an unswept cylinder mounted transversely on an axial cylinder so that the shock wave from the hemispherical nose of the axial cylinder intersected the bow shock of the unswept transverse cylinder. Data were obtained at Mach numbers from 2.53 to 5.50 and at Reynolds numbers based on the transverse cylinder diameter from 1.00 x 10(exp 6) to 1.87 x 10(exp 6). Shadowgraph pictures made in a wind tunnel showed that the flow field was influenced by boundary-layer separation on the axial cylinder and by end effects on the transverse cylinder as well as by the intersecting shocks. Under these conditions, the measured heat-transfer rates had inconsistent variations both in magnitude and distribution which precluded separating the effects of these disturbances. The general magnitude of the measured heating rates at Mach numbers up to 3 was from 0.1 to 0.5 of the theoretical laminar heating rates along the stagnation line for an infinite unswept cylinder in undisturbed flow. At Mach numbers above 4 the measured heating rates were from 1.5 to 2 times the theoretical rates.

  6. Ion-neutral collision effect on an Alfven wave

    SciTech Connect

    Amagishi, Y.; Tanaka, M. Department of High Energy Engineering Science, Interdisciplinary Graduate School of Engineering Sciences, Kyushu University, Kasuga, Fukuoka 816 )

    1993-07-19

    This paper reports that ion-neutral collisions in a magnetized plasma cause a drastic change in the dispersion relation of the shear Alfven wave with poloidal mode number [ital m]=0, connecting to the branch of the [ital m]=+1 compressional Alfven wave at frequencies below the ion-cyclotron frequency. An anomaly of the dispersion then appears on the refractive index curve and a wave packet in this frequency range undergoes strong amplitude damping and profile deformation. It is confirmed that the Kramers-Kronig relation holds for the dielectric function, estimated from both the measured refractive index and damping rate.

  7. An investigation of a close-coupled canard as a direct side-force generator on a fighter model at Mach numbers from 0.40 to 0.90

    NASA Technical Reports Server (NTRS)

    Re, R. J.; Capone, F. J.

    1977-01-01

    The canard panels had 5 deg of dihedral and were deflected differentially or individually over an incidence range from 10 deg to -10 deg and a model angle-of-attack range from -4 deg to 15 deg. Significant side forces were generated in a transonic tunnel by differential and single canard-panel deflections over the Mach number and angle-of-attack ranges. The yawing moment resulting from the forward location of the generated side force would necessitate a vertical tail/rudder trim force which would augment the forebody side force and be of comparable magnitude. Incremental side forces, yawing moments, lift, and pitching moments due to single canard-panel deflections were additive; that is, their sums were essentially the same as the forces and moments produced by differential canard-panel deflections of the same magnitude. Differential and single canard-panel deflections produced negligible rolling moments over the Mach number and angle-of-attack ranges.

  8. Effect of Angle of Attack and Exit Nozzle Design on the Performance of a 16-inch Ram Jet at Mach Numbers from 1.5 to 2.0

    NASA Technical Reports Server (NTRS)

    Perchonok, Eugene; Wilcox, Fred; Pennington, Donald

    1951-01-01

    An investigation of the performance of a 16-inch ram jet engine having a single oblique-shock all-external compression inlet designed for a flight Mach number of 1.8, was conducted in the NACA Lewis 8-by 6-foot supersonic wind tunnel. Data were obtained at Mach numbers from 1.5 to 2.0 and angles of attack from 0 degrees to 10 degrees. Three exit nozzles were used; a cylindrical extension of the combustion chamber, a 4 degrees half-angle converging nozzle with a 0.71 contraction ratio, and a graphite converging-diverging nozzle having a 0.71 contraction ratio plus reexpansion to essentially major body diameter.

  9. Measurements of Local Heat Transfer and Pressure on Six 2-Inch-Diameter Blunt Bodies at a Mach Number of 4.95 and at Reynolds Numbers Per Foot up to 81 x 10(exp 6)

    NASA Technical Reports Server (NTRS)

    Cooper, Morton; Mayo, Edward E.

    1959-01-01

    Measurements of the local heat transfer and pressure distribution have been made on six 2-inch-diameter, blunt, axially symmetric bodies in the Langley gas dynamics laboratory at a Mach number of 4.95 and at Reynolds numbers per foot up to 81 x 10(exp 6). During the investigation laminar flow was observed over a hemispherical-nosed body having a surface finish from 10 to 20 microinches at the highest test Reynolds number per foot (for this configuration) of 77.4 x 10(exp 6). Though it was repeatedly possible to measure completely laminar flow at this Reynolds number for the hemisphere, it was not possible to observe completely laminar flow on the flat-nosed body for similar conditions. The significance of this phenomenon is obscured by the observation that the effects of particle impacts on the surface in causing roughness were more pronounced on the flat-nosed body. For engineering purposes, a method developed by M. Richard Dennison while employed by Lockheed Aircraft Corporation appears to be a reasonable procedure for estimating turbulent heat transfer provided transition occurs at a forward location on the body. For rearward-transition locations, the method is much poorer for the hemispherical nose than for the flat nose. The pressures measured on the hemisphere agreed very well with those of the modified Newtonian theory, whereas the pressures on all other bodies, except on the flat-nosed body, were bracketed by modified Newtonian theory both with and without centrifugal forces. For the hemisphere, the stagnation-point velocity gradient agreed very well with Newtonian theory. The stagnation-point velocity gradient for the flat- nosed model was 0.31 of the value for the hemispherical-nosed model. If a Newtonian type of flow is assumed, the ratio 0.31 will be independent of Much number and real-gas effects.

  10. Aerodynamic design and performance testing of an advanced 30 deg swept, eight bladed propeller at Mach numbers from 0.2 to 0.85

    NASA Technical Reports Server (NTRS)

    Black, D. M.; Menthe, R. W.; Wainauski, H. S.

    1978-01-01

    The increased emphasis on fuel conservation in the world has stimulated a series of studies of both conventional and unconventional propulsion systems for commercial aircraft. Preliminary results from these studies indicate that a fuel saving of from 15 to 28 percent may be realized by the use of an advanced high speed turboprop. The turboprop must be capable of high efficiency at Mach 0.8 above 10.68 km (35,000 ft) altitude if it is to compete with turbofan powered commercial aircraft. An advanced turboprop concept was wind tunnel tested. The model included such concepts as an aerodynamically integrated propeller/nacelle, blade sweep and power (disk) loadings approximately three times higher than conventional propeller designs. The aerodynamic design for the model is discussed. Test results are presented which indicate propeller net efficiencies near 80 percent were obtained at high disk loadings at Mach 0.8.

  11. Investigation of global Alfven instabilities in TFTR

    SciTech Connect

    Wong, K.L.; Paul, S.F.; Fredrickson, E.D.; Nazikian, R.; Park, H.K.; Bell, M.; Bretz, N.L.; Budny, R.; Cheng, C.Z.; Cohen, S.; Hammett, G.W.; Jobes, F.C.; Johnson, L.; Meade, D.M.; Medley, S.S.; Mueller, D.; Nagayama, Y.; Owens, D.K.; Synakowski, E.J.; Durst, R.; Fonck, R.J.; Roberts, D.R.; Sabbagh, S.

    1992-01-01

    Toroidal Alfven Eigenmodes (TAE) were excited by the energetic neutral beam ions tangentially injected into TFTR plasmas at low magnetic field such that the injection velocities were comparable to the Alfven speed. The modes were identified by measurements from Mirnov coils and beam emission spectroscopy (BES). TAE modes appear in bursts whose repetition rate increases with beam power. The neutron emission rate exhibits sawtooth-like behavior and the crashes always coincide with TAE bursts. This indicates ejection of fast ions from the plasma until these modes are stabilized. The dynamics of growth and stabilization was investigated at various plasma current and magnetic field. The results indicate that the instability can effectively clamp the number of energetic ions in the plasma. The observed instability threshold is discussed in the light of recent theories. In addition to these TAE modes, intermittent oscillations at three times the fundamental TAE frequency were observed by Mirnov coils, but no corresponding signal was found in BES. It appears that these high frequency oscillations do not have direct effect on the plasma neutron source strength.

  12. A Whitham-Theory Sonic-Boom Analysis of the TU-144 Aircraft at a Mach Number of 2.2

    NASA Technical Reports Server (NTRS)

    Mack, Robert J.

    1999-01-01

    . Therefore, an analysis of the Tu-144 was made to obtain predictions of pressure signature shape and shock strengths at cruise conditions so that the range and characteristics of the required pressure gages could be determined well in advance of the tests. Cancellation of the sonic-boom signature measurement part of the tests removed the need for these pressure gages. Since CFD methods would be used to analyze the aerodynamic performance of the Tu-144 and make similar pressure signature predictions, the relatively quick and simple Whitham-theory pressure signature predictions presented in this paper could be used for comparisons. Pressure signature predictions of sonic-boom disturbances from the Tu- 144 aircraft were obtained from geometry derived from a three-view description of the production aircraft. The geometry was used to calculate aerodynamic performance characteristics at supersonic-cruise conditions. These characteristics and Whitham/Walkden sonic-boom theory were employed to obtain F-functions and flow-field pressure signature predictions at a Mach number of 2.2, at a cruise altitude of 61000 feet, and at a cruise weight of 350000 pounds.

  13. Design and performance at a local Mach number of 6 of an inlet for an integrated Scramjet concept. [wind tunnel models - aircraft design

    NASA Technical Reports Server (NTRS)

    Trexler, C. A.; Souders, S. W.

    1975-01-01

    The development of a concept for a modular supersonic combustion ramjet which is designed to integrate with the airframe of a hypersonic vehicle is presented. The design philosophy and results of experiments at Mach 6 to evaluate the performance of the scramjet inlet are given. The inlet was designed with modest contraction ratio, fixed geometry, and three fuel injection struts which contributed to the inlet flow compression and provided a short combustor design that resulted in low internal cooling requirements. Results indicate that the inlet performance is well within the acceptable range for high engine performance.

  14. Do interplanetary Alfven waves cause auroral activity?

    NASA Technical Reports Server (NTRS)

    Roberts, D. Aaron; Goldstein, Melvyn L.

    1990-01-01

    A recent theory holds that high-intensity, long-duration, continuous auroral activity (HILDCAA) is caused by interplanetary Alfven waves propagating outward from the sun. A survey of Alfvenic intervals in over a year of ISEE 3 data shows that while Alfvenic intervals often accompany HILDCAAs, the reverse is often not true. There are many Alfvenic intervals during which auroral activity (measured by high values of the AE index) is very low, as well as times of high auroral activity that are not highly Alfvenic. This analysis supports the common conclusion that large AE values are associated with a southward interplanetary field of sufficient strength and duration. This field configuration is independent of the presence of Alfven waves (whether solar generated or not) and is expected to occur at random intervals in the large-amplitude stochastic fluctuations in the solar wind.

  15. Highly Alfvenic Slow Solar Wind

    NASA Technical Reports Server (NTRS)

    Roberts, D. Aaron

    2010-01-01

    It is commonly thought that fast solar wind tends to be highly Alfvenic, with strong correlations between velocity and magnetic fluctuations, but examples have been known for over 20 years in which slow wind is both Alfvenic and has many other properties more typically expected of fast solar wind. This paper will present a search for examples of such flows from more recent data, and will begin to characterize the general characteristics of them. A very preliminary search suggests that such intervals are more common in the rising phase of the solar cycle. These intervals are important for providing constraints on models of solar wind acceleration, and in particular the role waves might or might not play in that process.

  16. Stability Characteristics of Two Missiles of Fineness Ratios 12 and 18 with Six Rectangular Fins of Very Low Aspect Ratio Over a Mach Number Range of 1.4 to 3.2

    NASA Technical Reports Server (NTRS)

    Henning, Allen B.

    1959-01-01

    Two rocket-propelled missiles have been test flown by the Langley Pilotless Aircraft Research Division in order to study the stability characteristics of a body with six rectangular fins of very low aspect ratio. The fins, which had exposed aspect ratios of approximately o.o4 and 0.02 per fin, were mounted on bodies of fineness ratios of 12 and 18, respectively. Each body had a nose with a fineness ratio of 3.5 and a cylindrical afterbody. The body and the fin chord of the model having a fineness ratio of 12 were extended the length of 6 body diameters to produce the model with a fineness ratio of 18. The missiles were disturbed in flight by pulse rockets in order to obtain the stability data. The tests were performed over a Mach number range of 1.4 to 3.2 and a Reynolds number range of 2 x 10(exp 6) to 21 x l0(exp 6). The results of these tests indicate that these configurations with the long rectangular fins of very low aspect ratio showed little induced roll" with the missile of highest fineness ratio and longest fin chord exhibiting the least amount. Extending the body and fin chord of the shorter missile six body diameters and thereby increasing the fin area approximately 115 percent increased the lift-curve slope based on body cross-sectional area approximately 40 to 55 percent, increased the dynamic stability by a substantial amount, and increased the drag from 14 to 33 percent throughout the comparable Mach number range. The center-of-pressure location of both missiles remained constant over the Mach number range.

  17. Alfven Continuum and Alfven Eigenmodes in the National Compact Stellarator Experiment

    SciTech Connect

    Fesenyuk, O. P.; Kolesnichenko, Ya. I.; Lutsenko, V. V.; White, R. B.; Yakovenko, Yu. V.

    2004-09-17

    The Alfven continuum (AC) in the National Compact Stellarator Experiment (NCSX) is investigated with the AC code COBRA. The resonant interaction of Alfven eigenmodes and the fast ions produced by neutral beam injection is analyzed. Alfven eigenmodes residing in one of the widest gaps of the NCSX AC, the ellipticity-induced gap, are studied with the code BOA-E.

  18. Performance, Stability, and Control Investigation at Mach Numbers from 0.60 to 1.05 of a Model of the "Swallow" with Outer Wing Panels Swept 75 degree with and without Power Simulations

    NASA Technical Reports Server (NTRS)

    Schmeer, James W.; Cassetti, Marlowe D.

    1960-01-01

    An investigation of the performance, stability, and control characteristics of a variable-sweep arrow-wing model with the outer wing panels swept 75 deg. has been conducted in the Langley 16-foot transonic tunnel. Four outboard engines located above and below the wing provided propulsive thrust, and, by deflecting in the pitch direction and rotating in the lateral plane, also produced control forces. The engine nacelles incorporated swept lateral and vertical fins for aerodynamic stability and control. Jet-off data were obtained with flow-through nacelles, simulating inlet flow; jet thrust and hot-jet interference effects were obtained with faired-nose nacelles housing hydrogen peroxide gas generators. Six-component force and moment data were obtained at Mach numbers from 0.60 to 1.05 through a range of angles of attack and angles of side-slip. Control characteristics were obtained by deflecting the nacelle-fin combinations as elevators, rudders, and ailerons at several fixed angles for each control. The results indicate that the basic wing-body configuration becomes neutrally stable or unstable at a lift coefficient of 0.15; addition of nacelles with fins delayed instability to a lift coefficient of 0.30. Addition of nacelles to the wing-body configuration increased minimum drag from 0.0058 to 0.0100 at a Mach number of 0.60 and from 0.0080 to 0.0190 at a Mach number of 1.05 with corresponding reductions in maximum lift-drag ratio of 12 percent and 33 percent, respectively. The nacelle-fin combinations were ineffective as longitudinal controls but were adequate as directional and lateral controls. The model with nacelles and fins was directionally and laterally stable; the stability generally increased with increasing lift. Jet interference effects on stability and control characteristics were small but the adverse effects on drag were greater than would be expected for isolated nacelles.

  19. Nonlinear dispersive Alfven waves in dusty plasma in the transition limit, {alpha}{approx}1

    SciTech Connect

    Sah, O. P.

    2011-10-15

    Localized nonlinear structures associated with dispersive Alfven waves are investigated in dusty plasma in the transition limit, i.e., {alpha}{identical_to}({beta}/2Q){approx}1, where {beta} is the ratio of thermal to magnetic pressure and Q is electron to ion mass ratio. Sagdeev pseudopotential is obtained from the basic governing equations, which is then numerically solved to study the existence and the behaviors of the nonlinear structures. It is found that both compressive and rarefactive solitons can coexist above and below certain critical {alpha}- values determined by the wave direction cosine (K{sub Z}) and the Mach number (M); and the compressive (rarefactive) solitons are much wider than the rarefactive ones for the case MK{sub Z}). In addition, the rarefactive solitons are found to be converted into rarefactive double layers, for the case M>K{sub Z}, if the dust grains are negatively charged and their density exceeds certain critical value.

  20. Stabilizing effect of ionized background of trans-Alfvenic expansion of exploding plasmas

    SciTech Connect

    Zakharov, Yu.P.; Ponomarenko, A.G.; Dudnikova, G.I.; Vshivkov, V.A.

    1995-12-31

    Recently a lot of theoretical and numerical calculations have been performed devoted to the study of Large-Larmor-Flute Instability (LLFI). Such instability was discovered initially in laboratory and later in active experiments (AMPTE, CRRES) on expansion of a quasispherical plasma cloud in a ``vacuum`` magnetic field {rvec B}{sub 0}. In the laser-produced plasma experiments at KI-1 facility it was established for the first time, that such non-MHD instability and LHD-instability of skin-layer may effectively be suppressed by ionized background at high-Alfven Mach numbers M{sub A} {much_gt} 1 as well as in a transient regime M{sub A} {approximately} 1. In the present paper on the basis of laboratory and computer simulation the value of M{sub A} was defined more exactly and other similarity parameters characterizing the development of LLFI was founded. The laser experiments were realized in hydrogen and argon background plasmas. The computer simulations were carried out with 2D electromagnetic hybrid code. It was exposed the transition from flute increase to decrease one when M{sub A} changed from M{sub A} = 1 to M{sub A} = 3.

  1. Investigation of the NACA 4-(5)(08)-03 and NACA 4-(10)(08)-03 Two-Blade Propellers at Forward Mach Numbers to 0.725 to Determine the Effects of Camber and Compressibility on Performance

    NASA Technical Reports Server (NTRS)

    Delano, James B

    1951-01-01

    As part of a general investigation of propellers at high forward speeds, tests of two-blade propellers having the NACA 4-(5)(08)-03 and NACA 4-(10)(08)-03 blade designs were made in the Langley 8-foot high-speed tunnel through a range of blade angle from 20 degrees to 60 degrees for forward Mach numbers from 0.165 to 0.70 to determine the effect of camber and compressibility on propeller characteristics. Results previously reported for similar tests of a two-blade propeller having the NACA 4-(3)(08)-03 blade design are included for comparison.

  2. Comparison of analytical and experimental steadyand unsteady-pressure distributions at Mach number 0.78 for a high-aspect-ratio supercritical wing model with oscillating control surfaces

    NASA Technical Reports Server (NTRS)

    Mccain, W. E.

    1984-01-01

    The unsteady aerodynamic lifting surface theory, the Doublet Lattice method, with experimental steady and unsteady pressure measurements of a high aspect ratio supercritical wing model at a Mach number of 0.78 were compared. The steady pressure data comparisons were made for incremental changes in angle of attack and control surface deflection. The unsteady pressure data comparisons were made at set angle of attack positions with oscillating control surface deflections. Significant viscous and transonic effects in the experimental aerodynamics which cannot be predicted by the Doublet Lattice method are shown. This study should assist development of empirical correction methods that may be applied to improve Doublet Lattice calculations of lifting surface aerodynamics.

  3. Solitary kinetic Alfven waves in dusty plasmas

    SciTech Connect

    Li Yangfang; Wu, D. J.; Morfill, G. E.

    2008-08-15

    Solitary kinetic Alfven waves in dusty plasmas are studied by considering the dust charge variation. The effect of the dust charge-to-mass ratio on the soliton solution is discussed. The Sagdeev potential is derived analytically with constant dust charge and then calculated numerically by taking the dust charge variation into account. We show that the dust charge-to-mass ratio plays an important role in the soliton properties. The soliton solutions are comprised of two branches. One branch is sub-Alfvenic and the soliton velocity is obviously smaller than the Alfven speed. The other branch is super-Alfvenic and the soliton velocity is very close to or greater than the Alfven speed. Both compressive and rarefactive solitons can exist. For the sub-Alfvenic branch, the rarefactive soliton is bell-shaped and it is much narrower than the compressive one. However, for the super-Alfvenic branch, the compressive soliton is bell-shaped and narrower, and the rarefactive one is broadened. When the charge-to-mass ratio of the dust grains is sufficiently high, the width of the rarefactive soliton, in the super-Alfvenic branch, will broaden extremely and a electron depletion will be observed. It is also shown that the bell-shaped soliton can transition to a cusped structure when the velocity is sufficiently high.

  4. About the parametric interplay between ionic mach number, body-size, and satellite potential in determining the ion depletion in the wake of the S3-2 Satellite

    SciTech Connect

    Samir, U.; Wildman, P.J.; Rich, F.; Brinton, H.C.; Sagalyn, R.C.

    1981-12-01

    Measurements of ion current, electron temperature, and density and values of satellite potential from the U.S. Air Force Satellite S3-2 together with ion composition measurements from the Atmosphere Explorer (AE-E) satellite were used to examine the variation of the ratio ..cap alpha.. = (I/sub +/(wake))/(I/sub +/(ambient)) (where I/sub +/ is the ion current) with altitude and to examine the significance of the parametric interplay between ionic Mach number, normalized body size R/sub D/( = R0/lambda/sub D/, where R/sub 0/ is the satellite radius and lambda/sub D/ is the ambient debye length) and normalized body potenital phi/sub N/( = ephis/KT/sub e/, where phi/sub s/ is the satellite potential, T/sub e/ is the electron temperature, and e and K are constants). It was possible to separate between the influence of R/sub D/ and phi/sub N/ on ..cap alpha.. for a specific range parameters. Uncertainty, however, remains regarding the competiton between R/sub D/ and S(H/sup +/) and S(O/sup +/) are oxygen and hydrogen ionic Mach numbers, respectively) in determining the ion distribution in the nearest vicincity to the satellite surface. A brief discussion relevant to future experiments in the area of body plasma flow interactions to be conducted on board the Shuttle/Spacelab facility, is also included.

  5. Transonic Aerodynamic Characteristics of a Wing-Body Combination having a 52.5 deg Sweptback Wing of Aspect Ratio 3 with Conical Camber and Designed for a Mach Number of the Square Root of 2

    NASA Technical Reports Server (NTRS)

    Igoe, William B.; Re, Richard J.; Cassetti, Marlowe

    1961-01-01

    An investigation has been made of the effects of conical wing camber and supersonic body indentation on the aerodynamic characteristics of a wing-body configuration at transonic speeds. Wing aspect ratio was 3.0, taper ratio was 0.1, and quarter-chord line sweepback was 52.5 deg with airfoil sections of 0.03 thickness ratio. The tests were conducted in the Langley 16-foot transonic tunnel at various Mach numbers from 0.80 to 1.05 at angles of attack from -4 deg to 14 deg. The cambered-wing configuration achieved higher lift-drag ratios than a similar plane-wing configuration. The camber also reduced the effects of wing-tip flow separation on the aerodynamic characteristics. In general, no stability or trim changes below wing-tip flow separation resulted from the use of camber. The use of supersonic body indentation improved the lift-drag ratios at Mach numbers from 0.96 to 1.05.

  6. Wind-tunnel investigation at Mach numbers from 1.90 to 2.86 of a canard-controlled missile with ram-air-jet spoiler roll control. [in the Langley Unitary Plan Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Blair, A. B., Jr.

    1978-01-01

    The efficacy of using a ram-air-jet spoiler roll control device on a typical canard-controlled missile configuration was investigated. For roll control comparisons, conventional aileron controls on the tail fins were also tested. The results indicate that the roll control of the ram-air-jet spoiler tail fins at the highest free-stream Mach number compared favorably with that of the conventional 11-70 area-ratio tail fin ailerons, each deflected 10 deg. The roll control of the tail fin ailerons decreased while that of the ram-air-jet spoiler increased with free-stream Mach number. The addition of the ram-air-jet spoiler tail fins or flow-through tip chord nacelles on the tail fins resulted in only small changes in basic missile longitudinal stability. The axial force coefficient of the operating ram-air-jet spoiler is significantly larger than that of conventional ailerons and results primarily from the total pressure behind a normal shock in front of the nacelle inlets.

  7. 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.

  8. Aerodynamic pressure and heating-rate distributions in tile gaps around chine regions with pressure gradients at a Mach number of 6.6

    NASA Astrophysics Data System (ADS)

    Hunt, L. Roane; Notestine, Kristopher K.

    1990-06-01

    Surface and gap pressures and heating-rate distributions were obtained for simulated Thermal Protection System (TPS) tile arrays on the curved surface test apparatus of the Langley 8-Foot High Temperature Tunnel at Mach 6.6. The results indicated that the chine gap pressures varied inversely with gap width because larger gap widths allowed greater venting from the gap to the lower model side pressures. Lower gap pressures caused greater flow ingress from the surface and increased gap heating. Generally, gap heating was greater in the longitudinal gaps than in the circumferential gaps. Gap heating decreased with increasing gap depth. Circumferential gap heating at the mid-depth was generally less than about 10 percent of the external surface value. Gap heating was most severe at local T-gap junctions and tile-to-tile forward-facing steps that caused the greatest heating from flow impingement. The use of flow stoppers at discrete locations reduced heating from flow impingement. The use of flow stoppers at discrete locations reduced heating in most gaps but increased heating in others. Limited use of flow stoppers or gap filler in longitudinal gaps could reduce gap heating in open circumferential gaps in regions of high surface pressure gradients.

  9. Aerodynamic pressure and heating-rate distributions in tile gaps around chine regions with pressure gradients at a Mach number of 6.6

    NASA Technical Reports Server (NTRS)

    Hunt, L. Roane; Notestine, Kristopher K.

    1990-01-01

    Surface and gap pressures and heating-rate distributions were obtained for simulated Thermal Protection System (TPS) tile arrays on the curved surface test apparatus of the Langley 8-Foot High Temperature Tunnel at Mach 6.6. The results indicated that the chine gap pressures varied inversely with gap width because larger gap widths allowed greater venting from the gap to the lower model side pressures. Lower gap pressures caused greater flow ingress from the surface and increased gap heating. Generally, gap heating was greater in the longitudinal gaps than in the circumferential gaps. Gap heating decreased with increasing gap depth. Circumferential gap heating at the mid-depth was generally less than about 10 percent of the external surface value. Gap heating was most severe at local T-gap junctions and tile-to-tile forward-facing steps that caused the greatest heating from flow impingement. The use of flow stoppers at discrete locations reduced heating from flow impingement. The use of flow stoppers at discrete locations reduced heating in most gaps but increased heating in others. Limited use of flow stoppers or gap filler in longitudinal gaps could reduce gap heating in open circumferential gaps in regions of high surface pressure gradients.

  10. Compressibility and cyclotron damping in the oblique Alfven wave

    SciTech Connect

    Harmon, J.K. )

    1989-11-01

    Compressibility, magnetic compressibility, and damping rate are calculated for the obliquely propagating Alfven shear wave in high- and low-beta Vlasov plasmas. There is an overall increase in compressibility as beta is reduced from {beta} = 1 to {beta}{much lt}1. For high obliquity {theta} and low frequency ({omega} {much lt} {Omega}{sub p}) the compressibility C follows a k{sup 2} wave number dependence; for high {theta} and low {beta} the approximation C(k) {approx} k{sub n}{sup 2} {identical to} (kV{sub A}/{Omega}{sub p}){sup 2} holds for wave numbers up to the proton cyclotron resonance, where {Omega}{sub p} is the proton cyclotron frequency and V{sub A} is the Alfven velocity. Strong proton cyclotron damping sets in at k{sub n} of the order of unity; the precise k{sub n} position of the damping cutoff increases with decreasing {beta} and increasing {theta}. Hence compressibility can exceed unity near the damping cutoff for high-{theta} waves in a low-{beta} plasma. The magnetic compressibility of the oblique Alfven wave also has a k{sup 2} dependence and can reach a maximum value of the order of 10% at high wave number. It is shown that Alfven compressibility could be the dominant contributor to the near-Sun solar wind density fluctuation spectrum for k>10{sup {minus}2} km{sup {minus}1} and hence might cause some of the flattening at high wave number seen in radio scintillation measurements. This would also be consistent with the notion that the observed density spectrum inner scale is a signature of cyclotron damping.

  11. The effects of RCS jet firing on the isolated Orbiter and mated coast phases of the glide return to launch site maneuver at Mach number 6 (IA302B)

    NASA Technical Reports Server (NTRS)

    Garrett, L. V.; Buchanan, T. D.; Fryberger, P. E.

    1988-01-01

    An updated Space Shuttle aerodynamic data base was obtained in Tunnel B for two phases of the Glide Return to Launch Site (GRTLS) abort maneuver. One-and-a-quarter percent scale models of the Space Shuttle Orbiter and External Tank were used to measure the effects of various combinations of Reaction Control System (RCS) jet thrusters at Mach number 6. The angle-of-attack range for the isolated orbiter was -10 to 15 deg at sideslip angles from -5 to 10 deg during Phase 1 of testing. The angle-of-attack range for the mated orbiter and external tank was -5 to 15 deg with sideslip angles of -2 to 5 deg during Phase 2. The test was conducted at a unit Reynolds number of 0.75 million per foot.

  12. Aerodynamic Study of a Wing-fuselage Combination Employing a Wing Swept Back 63 Degrees : Characteristics at a Mach Number of 1.53 Including Effect of Small Variations of Sweep

    NASA Technical Reports Server (NTRS)

    Madden, Robert T

    1949-01-01

    Measured values of lift, drag, and pitching moment at a Mach number of 1.53 and Reynolds numbers of 0.31, 0.62, and 0.84 million are presented for a wing-fuselage combination having a wing leading-edge sweep angle of 63 degrees, an aspect ratio of 3.42, a taper ratio of 0.25, and an NACA 64A006 section in the stream direction. Data are also presented for sweep angles of 57.0 degrees, 60.4 degrees, 67.0 degrees, and 69.9 degrees. The experimentally determined characteristics were less favorable than indicated by the linear theory but the experimental and theoretical trends with sweep were in good agreement. Boundary-layer-flow tests showed that laminar boundary-layer separation was the primary cause of the differences between experiment and theory.

  13. Aerodynamic characteristics of a canard-controlled missile at Mach numbers of 0.8, 1.3, and 1.75. [in the Ames 6 by 6 foot wind tunnel

    NASA Technical Reports Server (NTRS)

    Kassner, D. L.; Wettlaufer, B.

    1977-01-01

    A typical missile model with nose-mounted canards and cruciform tail surfaces was tested in the Ames 6- by 6-Foot Wind Tunnel to determine the contributions of the component aerodynamic surfaces to the static aerodynamic characteristics at Mach numbers of 0.8, 1.3, and 1.75 and Reynolds number of 625,000 based on body diameter. Data were obtained at angles of attack ranging from 0 deg to 24 deg for various stages of model build-up (i.e., with and without canard and/or tail surfaces). In addition, two different sets of canards and tail surfaces were investigated. For the canard and tail arrangements investigated, the model was trimmable at angles of attack up to about 7 deg with canard deflections of about 10 deg. Also, the tail arrangements studied provided ample pitch stability.

  14. 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.

  15. Flight-measured base pressure coefficients for thick boundary-layer flow over an aft-facing step for Mach numbers from 0.4 to 2.5

    NASA Technical Reports Server (NTRS)

    Goecke, S. A.

    1973-01-01

    A 0.56-inch thick aft-facing step was located 52.1 feet from the leading edge of the left wing of an XB-70 airplane. A boundary-layer rake at a mirror location on the right wing was used to obtain local flow properties. Reynolds numbers were near 10 to the 8th power, resulting in a relatively thick boundary-layer. The momentum thickness ranged from slightly thinner to slightly thicker than the step height. Surface static pressures forward of the step were obtained for Mach numbers near 0.9, 1.5, 2.0, and 2.4. The data were compared with thin boundary-layer results from flight and wind-tunnel experiments and semiempirical relationships. Significant differences were found between the thick and the thin boundary-layer data.

  16. SURFACE ALFVEN WAVES IN SOLAR FLUX TUBES

    SciTech Connect

    Goossens, M.; Andries, J.; Soler, R.; Van Doorsselaere, T.; Arregui, I.; Terradas, J.

    2012-07-10

    Magnetohydrodynamic (MHD) waves are ubiquitous in the solar atmosphere. Alfven waves and magneto-sonic waves are particular classes of MHD waves. These wave modes are clearly different and have pure properties in uniform plasmas of infinite extent only. Due to plasma non-uniformity, MHD waves have mixed properties and cannot be classified as pure Alfven or magneto-sonic waves. However, vorticity is a quantity unequivocally related to Alfven waves as compression is for magneto-sonic waves. Here, we investigate MHD waves superimposed on a one-dimensional non-uniform straight cylinder with constant magnetic field. For a piecewise constant density profile, we find that the fundamental radial modes of the non-axisymmetric waves have the same properties as surface Alfven waves at a true discontinuity in density. Contrary to the classic Alfven waves in a uniform plasma of infinite extent, vorticity is zero everywhere except at the cylinder boundary. If the discontinuity in density is replaced with a continuous variation of density, vorticity is spread out over the whole interval with non-uniform density. The fundamental radial modes of the non-axisymmetric waves do not need compression to exist unlike the radial overtones. In thin magnetic cylinders, the fundamental radial modes of the non-axisymmetric waves with phase velocities between the internal and the external Alfven velocities can be considered as surface Alfven waves. On the contrary, the radial overtones can be related to fast-like magneto-sonic modes.

  17. Nonlinear Alfven waves in high-speed solar wind streams

    NASA Technical Reports Server (NTRS)

    Abraham-Shrauner, B.; Feldman, W. C.

    1977-01-01

    A nonlinear proton distribution function that is an exact stationary solution of the nonlinear Vlasov equation and Maxwell's equations and which supports a single nonlinear transverse Alfven (ion cyclotron) wave that is circularly polarized and nondispersive is proposed for most of the observations during high-speed solar wind streams. This nonlinear distribution removes the strong Alfven wave instability, inconsistent with the persistence of the observed proton distribution functions in high-speed streams, found by the linear stability analysis. Model temperature anisotropies and drift velocities of the two spatially inhomogeneous bi-Maxwellian components are consistent with typical proton velocity distributions measured in high-speed streams at 1 AU. Two derived relations for each of the wave number and the phase velocity of the wave are obeyed within experimental uncertainties by two typical proton measurements. Our model also predicts that the alpha particle bulk flow velocity exceeds the proton particle bulk flow velocity, as is observed.

  18. Measurements of Heat Transfer and Boundary-Layer Transition on an 8-Inch-Diameter Hemisphere-Cylinder in Free Flight for a Mach Number Range of 2.00 to 3.88

    NASA Technical Reports Server (NTRS)

    Garland, Benjamine J.; Chauvin, Leo T.

    1957-01-01

    Measurements of aerodynamic heat transfer have been made along the hemisphere and cylinder of a hemisphere-cylinder rocket-propelled model in free flight up to a Mach number of 3.88. The test Reynolds number based on free-stream condition and diameter of model covered a range from 2.69 x l0(exp 6) to 11.70 x 10(exp 6). Laminar, transitional, and turbulent heat-transfer coefficients were obtained. The laminar data along the body agreed with laminar theory for blunt bodies whereas the turbulent data along the cylinder were consistently lower than that predicted by the turbulent theory for a flat plate. Measurements of heat transfer at the stagnation point were, in general, lower than the theory for stagnation-point heat transfer. When the Reynolds number to the junction of the hemisphere-cylinder was greater than 6 x l0(exp 6), the transitional Reynolds number varied from 0.8 x l0(exp 6) to 3.0 x 10(exp 6); however, than 6 x l(exp 6) when the Reynolds number to the junction was less, than the transitional Reynolds number varied from 7.0 x l0(exp 6) to 24.7 x 10(exp 6).

  19. Space Shuttle Orbiter trimmed center-of-gravity extension study. Volume 8: Effects of configuration modifications on the aerodynamic characteristics of the 140 A/B Orbiter at a Mach number of 5.97

    NASA Technical Reports Server (NTRS)

    Phillips, W. P.

    1984-01-01

    Aerodynamic characteristics at M=5.97 for the 140 A/B Space Shuttle Orbiter configuration and for the configuration modified by geometric changes in the wing planform fillet region and the fuselage forebody are presented. The modifications, designed to extend the orbiter's longitudinal trim capability to more forward center of gravity locations, include reshaping the baseline wing fillet, changing the fuselage forebody camber, and adding canards. The Langley 20 inch Mach 6 Tunnel at a Reynolds number of approximately 6 million based on fuselage reference length was used. The angle of attack range of the investigation varied from about 15 deg to 35 deg at 0 deg and -5 deg sideslip angles. Data are obtained with the elevators and body flap deflected at appropriate negative and positive conditions to assess the trim limits.

  20. Tabulations of static pressure coefficients on the surfaces of 3 pylon-mounted axisymmetric flow-through nacelles at Mach numbers from 0.40 to 0.98

    NASA Technical Reports Server (NTRS)

    Re, R. J.; Peddrew, K. H.

    1982-01-01

    Three flow through nacelles mounted on an 82 deg swept pylon (10 percent thickness-to-chord ratio) were tested in the Langley 16 foot Transonic Tunnel. The long uncambered pylon was supported from a small body of revolution so that pressure measurements on the nacelle and pylon represent a pylon nacelle flow field without a wing present. Two nacelles had NACA 1-85-100 inlets and different circular arc afterbodies. The third nacelle had an NACA 1-70-100 inlet with a circular arc afterbody having the same external shape as one of the other nacelles. Nacelle length to maximum diameter ratio was 3.5. Data were obtained at angles of attack from 2 deg to 8 deg at selected Mach numbers.

  1. Longitudinal and Lateral Stability, Control Characteristics, and Vertical-Tail-Load Measurements for 0.03-Scale Model of the Avro CF-105 Airplane at Mach Number 1.41

    NASA Technical Reports Server (NTRS)

    Spearman, M. Leroy; Robinson, Ross B.; Driver, Cornelius

    1956-01-01

    An investigation has been made in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 1.41 to determine the aerodynamic characteristics of an 0.03-scale model of the Avro CF-105 airplane. The investigation included the determination of the static longitudinal and lateral stability, the control and the hinge-moment characteristics of the elevator, the aileron, and the rudder, as well as the vertical-tail-load characteristics. The results indicated a minimum drag coefficient of about 0.0270, and a maximum trimmed lift-drag ratio of about 4.25 which occurs at a lift coefficient of 0.16. The directional stability decreased with increasing angle of attack until a region of static instability occurred above an angle of attack of about 9 deg.

  2. Longitudinal and Lateral Stability and Control Characteristics and Vertical-Tail-Load Measurements for a 0.03-Scale Model of the Avro CF-105 Airplane at Mach Numbers of 1.60, 1.80, and 2.00

    NASA Technical Reports Server (NTRS)

    Silvers, H. Norman; Fournier, Roger H.; Wills, Jane S.

    1958-01-01

    An investigation has been made in the Langley Unitary Plan wind tunnel at Mach numbers of 1.60, 1.80, and 2.00 to determine the aerodynamic characteristics of a 0.03-scale model of the Avro CF-105 airplane. The investigation included the determination of the static longitudinal and lateral stability, the control and the hinge-moment characteristics of the elevator, rudder, and aileron, as well as the vertical-tail-load characteristics. Although the data are presented without analysis, a limited inspection of the longitudinal control results indicates a loss in maximum lift-drag ratio due to trimming of about 1.8 because of the large static margin. A reduction in static margin would be expected to improve the trim lift-drag ratio but would also reduce the directional stability. With the existing static margin, the configuration is directionally unstable at angles of attack above about 6 deg or 8 deg.

  3. Dynamic response of a forward-swept-wing model at angles of attack up to 15 deg at a Mach number of 0.8. [Langley transonic dynamics tunnel tests

    NASA Technical Reports Server (NTRS)

    Doggett, R. V., Jr.; Ricketts, R. H.

    1980-01-01

    Root mean square (rms) bending moments for a dynamically scaled, aeroelastic wing of a proposed forward swept wing, flight demonstrator airplane are presented for angles of attack up to 15 deg at a Mach number of 0.8 The 0.6 size semispan model had a leading edge forward sweep of 44 deg and was constructed of composite material. In addition to broad band responses, individual rms responses and total damping ratios are presented for the first two natural modes. The results show that the rms response increases with angle of attack and has a peak value at an angle of attack near 13 deg. In general, the response was characteristic of buffeting and similar to results often observed for aft swept wings. At an angle of attack near 13 deg, however, the response had characteristics associated with approaching a dynamic instability, although no instability was observed over the range of parameters investigated.

  4. Stellar winds driven by Alfven waves

    NASA Technical Reports Server (NTRS)

    Belcher, J. W.; Olbert, S.

    1973-01-01

    Models of stellar winds were considered in which the dynamic expansion of a corona is driven by Alfven waves propagating outward along radial magnetic field lines. In the presence of Alfven waves, a coronal expansion can exist for a broad range of reference conditions which would, in the absence of waves, lead to static configurations. Wind models in which the acceleration mechanism is due to Alfven waves alone and exhibit lower mass fluxes and higher energies per particle are compared to wind models in which the acceleration is due to thermal processes. For example, winds driven by Alfven waves exhibit streaming velocities at infinity which may vary between the escape velocity at the coronal base and the geometrical mean of the escape velocity and the speed of light. Upper and lower limits were derived for the allowed energy fluxes and mass fluxes associated with these winds.

  5. Flight Test of a 40-Foot Nominal Diameter Disk-Gap-Band Parachute Deployed at a Mach Number of 2.72 and a Dynamic Pressure of 9.7 Pounds per Square Foot

    NASA Technical Reports Server (NTRS)

    Eckstrom, Clinton V.; Preisser, John S.

    1968-01-01

    A 40-foot-nominal-diameter (12.2 meter) disk-gap-band parachute was flight tested as part of the NASA Supersonic Planetary Entry Decelerator (SPED-I) Program. The test parachute was deployed from an instrumented payload by means of a deployment mortar when the payload was at an altitude of 158,500 feet (48.2 kilometers), a Mach number of 2.72, and a free-stream dynamic pressure of 9.7 pounds per foot(exp 2) (465 newtons per meter(exp 2)). Suspension line stretch occurred 0.46 second after mortar firing and the resulting snatch force loading was -8.lg. The maximum acceleration experienced by the payload due to parachute opening was -27.2g at 0.50 second after the snatch force peak for a total elapsed time from mortar firing of 0.96 second. Canopy-shape variations occurred during the higher Mach number portion of the flight test (M greater than 1.4) and the payload was subjected to large amplitude oscillatory loads. A calculated average nominal axial-force coefficient ranged from about 0.25 immediately after the first canopy opening to about 0.50 as the canopy attained a steady inflated shape. One gore of the test parachute was damaged when the deployment bag with mortar lid passed through it from behind approximately 2 seconds after deployment was initiated. Although the canopy damage caused by the deployment bag penetration had no apparent effect on the functional capability of the test parachute, it may have affected parachute performance since the average effective drag coefficient of 0.48 was 9 percent less than that of a previously tested parachute of the same configuration.

  6. Alfven wave. DOE Critical Review Series

    SciTech Connect

    Hasegawa, A.; Uberoi, C.

    1982-01-01

    This monograph deals with the properties of Alfven waves and with their application to fusion. The book is divided into 7 chapters dealing with linear properties in homogeneous and inhomogeneous plasmas. Absorption is treated by means of kinetic theory. Instabilities and nonlinear processes are treated in Chapters 1 to 6, and the closing chapter is devoted to theory and experiments in plasma heating by Alfven waves. (MOW)

  7. Beam distribution modification by Alfven modes

    SciTech Connect

    White, R. B.; Gorelenkov, N.; Heidbrink, W. W.; Van Zeeland, M. A.

    2010-05-15

    Modification of a deuterium beam distribution in the presence of low amplitude toroidal Alfven eigenmodes and reversed shear Alfven eigenmodes in a toroidal magnetic confinement device is examined. Comparison to experimental data shows that multiple low amplitude modes can account for significant modification of high energy beam particle distributions. It is found that there is a stochastic threshold for beam transport, and that the experimental amplitudes are only slightly above this threshold. The modes produce a substantial central flattening of the beam distribution.

  8. Pressure Distributions and Wave Drag Due to Two-Dimensional Fabrication-Type Surface Roughness on an Ogive Cylinder at Mach Numbers of 1.61 and 2.01

    NASA Technical Reports Server (NTRS)

    Czarnecki, K. R.; Monta, William J.

    1961-01-01

    An investigation has been made at Mach numbers of 1.61 and 2.01 and over a range of free-stream Reynolds number per foot from about 1.2 x 10(exp 6) to 8.3 x 10(exp 6) to determine the pressure distributions and wave drags due to two-dimensional fabrication-type surface roughness. Ten types of surface roughness, including step, wave, crease, and swept configurations were investigated. The tests were made on an ogive cylinder of fineness ratio 12.2, the roughness elements covering the cylindrical portion of the model. The results indicate that wave drag is the major component of the drag due to roughness at supersonic speeds. The pressure distributions over the roughness elements were generally found to be in good agreement with linearized two-dimensional theory except for regions of the elements affected by boundary-layer separation and shock detachment. There was little or no effect of Reynolds number except on the pressures within the regions influenced by separation or shock detachment. Inasmuch as most of the roughness configurations were affected by flow separation and shock detachment, there was generally an effect of Reynolds number on the roughness wave drag. This wave drag decreased as the free-stream Reynolds number was decreased.

  9. Nonlinear Landau damping and Alfven wave dissipation

    NASA Technical Reports Server (NTRS)

    Vinas, Adolfo F.; Miller, James A.

    1995-01-01

    Nonlinear Landau damping has been often suggested to be the cause of the dissipation of Alfven waves in the solar wind as well as the mechanism for ion heating and selective preacceleration in solar flares. We discuss the viability of these processes in light of our theoretical and numerical results. We present one-dimensional hybrid plasma simulations of the nonlinear Landau damping of parallel Alfven waves. In this scenario, two Alfven waves nonresonantly combine to create second-order magnetic field pressure gradients, which then drive density fluctuations, which in turn drive a second-order longitudinal electric field. Under certain conditions, this electric field strongly interacts with the ambient ions via the Landau resonance which leads to a rapid dissipation of the Alfven wave energy. While there is a net flux of energy from the waves to the ions, one of the Alfven waves will grow if both have the same polarization. We compare damping and growth rates from plasma simulations with those predicted by Lee and Volk (1973), and also discuss the evolution of the ambient ion distribution. We then consider this nonlinear interaction in the presence of a spectrum of Alfven waves, and discuss the spectrum's influence on the growth or damping of a single wave. We also discuss the implications for wave dissipation and ion heating in the solar wind.

  10. Experimental Determination of the Recovery Factor and Analytical Solution of the Conical Flow Field for a 20 deg Included Angle Cone at Mach Numbers of 4.6 and 6.0 and Stagnation Temperatures to 2600 degree R

    NASA Technical Reports Server (NTRS)

    Pfyl, Frank A.; Presley, Leroy L.

    1961-01-01

    The local recovery factor was determined experimentally along the surface of a thin-walled 20 deg included angle cone for Mach numbers near 6.0 at stagnation temperatures between 1200 deg R and 2600 deg R. In addition, a similar cone configuration was tested at Mach numbers near 4.5 at stagnation temperatures of approximately 612 deg R. The local Reynolds number based on flow properties at the edge of the boundary layer ranged between 0.1 x 10(exp 4) and 3.5 x 10(exp 4) for tests at temperatures above 1200 deg R and between 6 x 10(exp 4) and 25 x 10(exp 4) for tests at temperatures near 612 deg R. The results indicated, generally, that the recovery factor can be predicted satisfactorily using the square root of the Prandtl number. No conclusion could be made as to the necessity of evaluating the Prandtl number at a reference temperature given by an empirical equation, as opposed to evaluating the Prandtl number at the wall temperature or static temperature of the gas at the cone surface. For the tests at temperatures above 1200 deg R (indicated herein as the tests conducted in the slip-flow region), two definite trends in the recovery data were observed - one of increasing recovery factor with decreasing stagnation pressure, which was associated with slip-flow effects and one of decreasing recovery factor with increasing temperature. The true cause of the latter trend could not be ascertained, but it was shown that this trend was not appreciably altered by the sources of error of the magnitude considered herein. The real-gas equations of state were used to determine accurately the local stream properties at the outer edge of the boundary layer of the cone. Included in the report, therefore, is a general solution for the conical flow of a real gas using the Beattie-Bridgeman equation of state. The largest effect of temperature was seen to be in the terms which were dependent upon the internal energy of the gas. The pressure and hence the pressure drag terms were

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

    NASA Technical Reports Server (NTRS)

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

    1957-01-01

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

  12. Study on Mach stems induced by interaction of planar shock waves on two intersecting wedges

    NASA Astrophysics Data System (ADS)

    Xiang, Gaoxiang; Wang, Chun; Teng, Honghui; Yang, Yang; Jiang, Zonglin

    2016-06-01

    The properties of Mach stems in hypersonic corner flow induced by Mach interaction over 3D intersecting wedges were studied theoretically and numerically. A new method called "spatial dimension reduction" was used to analyze theoretically the location and Mach number behind Mach stems. By using this approach, the problem of 3D steady shock/shock interaction over 3D intersecting wedges was transformed into a 2D moving one on cross sections, which can be solved by shock-polar theory and shock dynamics theory. The properties of Mach interaction over 3D intersecting wedges can be analyzed with the new method, including pressure, temperature, density in the vicinity of triple points, location, and Mach number behind Mach stems. Theoretical results were compared with numerical results, and good agreement was obtained. Also, the influence of Mach number and wedge angle on the properties of a 3D Mach stem was studied.

  13. Kepler and Mach's Principle

    NASA Astrophysics Data System (ADS)

    Barbour, Julian

    The definitive ideas that led to the creation of general relativity crystallized in Einstein's thinking during 1912 while he was in Prague. At the centenary meeting held there to mark the breakthrough, I was asked to talk about earlier great work of relevance to dynamics done at Prague, above all by Kepler and Mach. The main topics covered in this chapter are: some little known but basic facts about the planetary motions; the conceptual framework and most important discoveries of Ptolemy and Copernicus; the complete change of concepts that Kepler introduced and their role in his discoveries; the significance of them in Newton's work; Mach's realization that Kepler's conceptual revolution needed further development to free Newton's conceptual world of the last vestiges of the purely geometrical Ptolemaic world view; and the precise formulation of Mach's principle required to place GR correctly in the line of conceptual and technical evolution that began with the ancient Greek astronomers.

  14. Static Longitudinal and Lateral Stability Characteristics of an 0.065-Scale Model of the Chance Vought XRSSM-N-9a (REGULUS II) Missile at Mach Numbers from 1.6 to 2.0 (TED No. NACA AD 3122)

    NASA Technical Reports Server (NTRS)

    Hofstetter, William R.

    1957-01-01

    The static longitudinal and lateral stability charaetefistics of an 0 .065-scale model of the XRSSM-N-9a (REGULUS II) Missile at Mach number range of 1.6 to 2.0 at a Reynolds number per foot of 2.0(exp 8)

  15. Surface-Pressure and Flow-Visualization Data at Mach Number of 1.60 for Three 65 deg Delta Wings Varying in Leading-Edge Radius and Camber

    NASA Technical Reports Server (NTRS)

    McMillin, S. Naomi; Bryd, James E.; Parmar, Devendra S.; Bezos-OConnor, Gaudy M.; Forrest, Dana K.; Bowen, Susan

    1996-01-01

    An experimental investigation of the effect of leading-edge radius, camber, Reynolds number, and boundary-layer state on the incipient separation of a delta wing at supersonic speeds was conducted at the Langley Unitary Plan Wind Tunnel at Mach number of 1.60 over a free-stream Reynolds number range of 1 x 106 to 5 x 106 ft-1. The three delta wing models examined had a 65 deg swept leading edge and varied in cross-sectional shape: a sharp wedge, a 20:1 ellipse, and a 20:1 ellipse with a -9.750 circular camber imposed across the span. The wings were tested with and without transition grit applied. Surface-pressure coefficient data and flow-visualization data indicated that by rounding the wing leading edge or cambering the wing in the spanwise direction, the onset of leading-edge separation on a delta wing can be raised to a higher angle of attack than that observed on a sharp-edged delta wing. The data also showed that the onset of leading-edge separation can be raised to a higher angle of attack by forcing boundary-layer transition to occur closer to the wing leading edge by the application of grit or the increase in free-stream Reynolds number.

  16. Ernst Mach - a deeper look. Documents and new perspectives.

    NASA Astrophysics Data System (ADS)

    Blackmore, J.

    The book gives new information on Mach as a philosopher, historian, scientist and person, containing a number of biographical and philosophical manuscripts published for the first time, along with correspondence and other matters published for the first time in English. The volume also provides English translations of Mach's controversies with leading physicists and psychologists, such as Max Planck and Carl Stumpf, and offers basic evidence for resolving Mach's position on atomism and Einstein's theory of relativity.

  17. Experimental wake survey behind a 140 deg-included-angle cone at angles of attack of 0 deg and 5 deg, Mach numbers from 1.60 to 3.95, and longitudinal stations varying from 1.0 to 8.39 body diameters

    NASA Technical Reports Server (NTRS)

    Brown, C. A., Jr.; Campbell, J. F.

    1971-01-01

    The flow properties in the wake of a 140 deg-included-angle cone at Mach numbers from 1.60 to 3.95 and at angles of attack of 0 deg and 5 deg are discussed. The wake flow properties are calculated from total and static pressures measured with a pressure rake at longitudinal stations varying from 1.0 to 8.39 body diameters and at lateral stations varying from -0.42 to 3.0 body diameters. These measurements show a consistent trend throughout the range of Mach number and longitudinal distance and an increase in dynamic pressure with increasing longitudinal station.

  18. Aerothermal tests of a 12.5 percent cone at Mach 6.7 for various Reynolds numbers, angles of attack and nose shapes. [conducted in Langley 8-foot high temperature tunnel

    NASA Technical Reports Server (NTRS)

    Nowak, R. J.; Albertson, C. W.; Hunt, L. R.

    1984-01-01

    The effects of free-stream unit Reynolds number, angle of attack, and nose shape on the aerothermal environment of a 3-ft basediameter, 12.5 deg half-angle cone were investigated in the Langley 8-foot high temperature tunnel at Mach 6.7. The average total temperature was 3300 R, the freestream unit Reynolds number ranged from 400,000 to 1,400,000 per foot, and the angle of attack ranged from 0 deg to 10 deg. Three nose configurations were tested on the cone: a 3-in-radius tip, a 1-in-radius tip on an ogive frustum, and a sharp tip on an ogive frustum. Surface-pressure and cold-wall heating-rate distributions were obtained for laminar, transitional temperature in the shock layer were obtained. The location of the start of transition moved forward both on windward and leeward sides with increasing free-stream Reynolds numbers, increasing angle of attack, and decreasing nose bluntness.

  19. Experimental Investigation of the Effects of Viscosity on the Drag and Base Pressure of Bodies of Revolution at a Mach Number 1.5

    NASA Technical Reports Server (NTRS)

    Chapman, Dean R; Perkins, Edward W

    1951-01-01

    Models were tested to evaluate effects of Reynolds number for both laminar and turbulent boundary layers. Principal geometric variables investigated were afterbody shape and length-diameter ratio. Force tests and base-pressure measurements were made. Schlieren photographs were used to analyze the effects of viscosity on flow separation and shock-wave configuration and to verify the condition of the boundary layer as deduced from the force tests. The results are discussed and compared with theoretical calculations.

  20. Alfvenic waves in solar spicules

    NASA Astrophysics Data System (ADS)

    Ebadi, Hossein

    2016-07-01

    We analyzed O VI (1031.93 A) and O VI (1037.61 A line profiles from the time series of SOHO/SUMER data. The wavelet analysis is used to determine the fundamental mode and its first harmonic periods and their ratio. The period ratio, P_1/P_2 is obtained as 2.1 based on our calculations. To model the spicule oscillations, we consider an equilibrium configuration in the form of an expanding straight magnetic flux tube with varying density along tube. We used cylindrical coordinates r, phi, and z with the z-axis along tube axis. Standing Alfvenic waves with steady flows are studied. More realistic background magnetic field, plasma density, and spicule radios inferred from the actual magnetoseismology of observations are used. It is found that the oscillation periods and their ratio are shifted because of the steady flows. The observational values are reached in P_1/P_2, when the steady flows are 0.2-0.3, the values which are reported for classical spicules.

  1. Effects of ion-neutral collisions on Alfven waves: The presence of forbidden zone and heavy damping zone

    SciTech Connect

    Weng, C. J.; Lee, L. C.; Kuo, C. L.; Wang, C. B.

    2013-03-15

    Alfven waves are low-frequency transverse waves propagating in a magnetized plasma. We define the Alfven frequency {omega}{sub 0} as {omega}{sub 0}=kV{sub A}cos{theta}, where k is the wave number, V{sub A} is the Alfven speed, and {theta} is the angle between the wave vector and the ambient magnetic field. There are partially ionized plasmas in laboratory, space, and astrophysical plasma systems, such as in the solar chromosphere, interstellar clouds, and the earth ionosphere. The presence of neutral particles may modify the wave frequency and cause damping of Alfven waves. The effects on Alfven waves depend on two parameters: (1) {alpha}=n{sub n}/n{sub i}, the ratio of neutral density (n{sub n}), and ion density (n{sub i}); (2) {beta}={nu}{sub ni}/{omega}{sub 0}, the ratio of neutral collisional frequency by ions {nu}{sub ni} to the Alfven frequency {omega}{sub 0}. Most of the previous studies examined only the limiting case with a relatively large neutral collisional frequency or {beta} Much-Greater-Than 1. In the present paper, the dispersion relation for Alfven waves is solved for all values of {alpha} and {beta}. Approximate solutions in the limit {beta} Much-Greater-Than 1 as well as {beta} Much-Less-Than 1 are obtained. It is found for the first time that there is a 'forbidden zone (FZ)' in the {alpha}-{beta} parameter space, where the real frequency of Alfven waves becomes zero. We also solve the wavenumber k from the dispersion equation for a fixed frequency and find the existence of a 'heavy damping zone (HDZ).' We then examine the presence of FZ and HDZ for Alfven waves in the ionosphere and in the solar chromosphere.

  2. Nonlinear effects associated with the dispersive Alfven waves in space plasmas

    SciTech Connect

    Kumar, Sanjay; Sharma, R. P.

    2010-03-15

    This paper presents the model equations governing the nonlinear dynamics of the dispersive Alfven wave (DAW) in the low-beta plasmas (beta<Alfven waves) applicable to solar corona and intermediate-beta plasmas (m{sub e}/m{sub i}<Alfven waves) applicable to solar wind in Earth's magnetosphere. The pump DAW is perturbed by a low-frequency fast wave (FW). When the ponderomotive nonlinearities are incorporated in the DAW and FW dynamics, the model equations of DAW and FW turn out to be the modified Zakharov system of equations. Growth rate and threshold field for modulational (filamentation) instability have been calculated. The dependence of the growth rate on the perturbation wave number and the pump wave parameters (such as perpendicular wave number) has also been presented.

  3. Alfvenicity of Fluctuations Associated with Kelvin-Helmholtz Instability in Plume-Interplume Region

    NASA Astrophysics Data System (ADS)

    Parhi, S.; Suess, S.; Sulkanen, M.

    1999-05-01

    We study the velocity shear between plumes and the interplume flow in coronal holes. We model these plumes as jets (or, strictly speaking, wakes). Weak and strong magnetic fields are considered both inside and outside the jet for a shear Mach number 6. The shear can be unstable and evolve into a new less sheared pattern. As the instability sets in, the jet first develops a cocoon of intermediate speed flow and slowly a bridge develops between upstream and downstream flows. This marks the onset of jet disruption via what appears to be mass entrainment and fluid instability. This could also be induced by the jet's passage through the accompanying fast shock formation. The jet bends upon crossing the oblique shocks because all streamlines bend away from the shock normal. In a short time the downstream flow just ahead of the bending suffers a change in speed but still maintains or reestablishes supersonic conditions somehow. The transverse velocity here is very low because the instability generated in the disturbed region reduces the shear ahead. The shear ultimately must dissipate. The generation of this instability depends both temporally and spatially on the amount of shear and the time needed for nonlinear growth. To analyse the fluctuations quantitatively we perform a time series analysis at various points inside and adjacent to the jet. Specifically we consider points either in the center of the jet or just outside the transition layer- the initial location of the shear layer. We find the fully developed nonlinear fluctuations are more Alfvenic than magnetosonic in the high beta case than in low beta case.

  4. Aerodynamic Loads at Mach Numbers from 0.70 to 2.22 on a Airplane Model Having a Wing and Canard of Triangular Plan Form and Either Single or Twin Vertical Tails

    NASA Technical Reports Server (NTRS)

    Peterson, Victor L.; Menees, Gene P.

    1961-01-01

    Results of an investigation of the aerodynamic loads on a canard airplane model are presented without detailed analysis for the Mach number range of 0.70 t o 2.22. The model consisted of a triangular wing and canard of aspect ratio 2 mounted on a Sears-Haack body of fineness ratio 12.5 and either a single body-mounted vertical tail or twin wing mounted vertical tails of low aspect ratio and sweptback plan form. The body, right wing panel, single vertical tail, and left twin vertical tail were instrumented for measuring pressures. Data were obtained for angles of attack ranging from -4 degrees to +16 degrees, nominal canard deflection angles of 0 degrees and 10 degrees, and angles of sideslip of 0 degrees and 5.3 degrees. The Reynolds number was 2.9 x 10(exp 6) based on the wing mean aerodynamic chord. Selected portions of the data are presented in graphical form and attention is directed to some of the results of the investigation. All of the experimental results have been tabulated in the form of pressure coefficients and integrations of the pressure coefficients and are available as supplements to this paper. A brief summary of the contents of the tabular material is given.

  5. Effects of canard location on the aerodynamic characteristics of a blunt-nosed missile at Mach numbers of 1.5 and 2.0. [in the Ames 6x6 wind tunnel

    NASA Technical Reports Server (NTRS)

    Kassner, D. L.; Wettlaufer, B.

    1977-01-01

    A blunt-nosed missile model with nose-mounted canards and cruciform tail surfaces was tested in the Ames 6 by 6-Foot Wind Tunnel to determine the contributions of the component aerodynamic surfaces to the static aerodynamic characteristics at Mach numbers of 1.5 and 2.0 and Reynolds number of 1 million based on body diameter. Data were obtained at angles of attack ranging from -3 deg to 12 deg and canard-deflection angles from -3 deg to 15 deg for various stages of model build-up (i.e., with and without canard and/or tail surfaces). Results were obtained with the canards at two different nose locations. For the canard and tail arrangements investigated, the model was trimmable at angles of attack up to about 4 deg or 5 deg with canard deflections of 9 deg. For this blunt-nosed model, there was little effect of canard location on trim angle of attack. The tail arrangements studied provided ample pitch stability.

  6. Water-Film Cooling of an 80 deg Total-Angle Cone at a Mach Number of 2 for Airstream Total Temperatures up to 3,000 deg R

    NASA Technical Reports Server (NTRS)

    Carter, Howard S.

    1959-01-01

    Film-cooling tests, with water as the coolant, were made on an 80 deg total-angle cone in a Mach number 2 free jet at sea-level pressure. The tests were made at free-stream total temperatures from 1,500 deg R to 3,000 deg R and at free-stream Reynolds numbers per foot from 8 x 10(exp 6) to 3 x 10(exp 6). The tests showed that the downstream end of the model became very hot if the coolant rate was too small to cover the complete model with a water film. This water film was fairly symmetrical when the model was at zero angle of attack but was very asymmetrical when the model was at an angle of attack of 5 deg. A comparison with results of a previous transpiration-cooling test showed that, with water as the coolant, transpiration cooling was at least 2.5 times as efficient as the film cooling of the present tests.

  7. Transonic Aerodynamic Loading Characteristics of a Wing-Body-Tail Combination Having a 52.5 deg. Sweptback Wing of Aspect Ratio 3 With Conical Wing Camber and Body Indentation for a Design Mach Number of Square Root of 2

    NASA Technical Reports Server (NTRS)

    Cassetti, Marlowe D.; Re, Richard J.; Igoe, William B.

    1961-01-01

    An investigation has been made of the effects of conical wing camber and body indentation according to the supersonic area rule on the aerodynamic wing loading characteristics of a wing-body-tail configuration at transonic speeds. The wing aspect ratio was 3, taper ratio was 0.1, and quarter-chord-line sweepback was 52.5 deg. with 3-percent-thick airfoil sections. The tests were conducted in the Langley 16-foot transonic tunnel at Mach numbers from 0.80 to 1.05 and at angles of attack from 0 deg. to 14 deg., with Reynolds numbers based on mean aerodynamic chord varying from 7 x 10(exp 6) to 8 x 10(exp 6). Conical camber delayed wing-tip stall and reduced the severity of the accompanying longitudinal instability but did not appreciably affect the spanwise load distribution at angles of attack below tip stall. Body indentation reduced the transonic chordwise center-of-pressure travel from about 8 percent to 5 percent of the mean aerodynamic chord.

  8. Shock interference heat transfer to tank configurations mated to a straight-wing space shuttle orbiter at Mach number 10.3. [investigated in a Langley hypersonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Crawford, D. H.

    1976-01-01

    Heat transfer was measured on a space shuttle-tank configuration with no mated orbiter in place and with the orbiter in 10 different mated positions. The orbiter-tank combination was tested at angles of attack of 0 deg and 5 deg, at a Mach number of 10.3, and at a free-stream Reynolds number of one million based on the length of the tank. Comparison of interference heat transfer with no-interference heat transfer shows that shock interference can increase the heat transfer to the tank by two orders of magnitude along the ray adjacent to the orbiter and can cause high temperature gradients along the tank skin. The relative axial location of the two mated vehicles determined the location of the sharp peaks of extreme heating as well as their magnitude. The other control variables (the angle of attack, the gap, and the cross-section shape) had significant effects that were not as consistent or as extreme.

  9. Alpha particle destabilization of the toroidicity-induced Alfven eigenmodes

    SciTech Connect

    Cheng, C.Z.

    1990-10-01

    The high frequency, low mode number toroidicity-induced Alfven eigenmodes (TAE) are shown to be driven unstable by the circulating and/or trapped {alpha}-particles through the wave-particle resonances. Satisfying the resonance condition requires that the {alpha}-particle birth speed v{sub {alpha}} {ge} v{sub A}/2{vert bar}m-nq{vert bar}, where v{sub A} is the Alfven speed, m is the poloidal model number, and n is the toroidal mode number. To destabilize the TAE modes, the inverse Landau damping associated with the {alpha}-particle pressure gradient free energy must overcome the velocity space Landau damping due to both the {alpha}-particles and the core electrons and ions. The growth rate was studied analytically with a perturbative formula derived from the quadratic dispersion relation, and numerically with the aid of the NOVA-K code. Stability criteria in terms of the {alpha}-particle beta {beta}{sub {alpha}}, {alpha}-particle pressure gradient parameter ({omega}{sub {asterisk}}/{omega}{sub A}) ({omega}{sub {asterisk}} is the {alpha}-particle diamagnetic drift frequency), and (v{sub {alpha}}/v{sub A}) parameters will be presented for TFTR, CIT, and ITER tokamaks. The volume averaged {alpha}-particle beta threshold for TAE instability also depends sensitively on the core electron and ion temperature. Typically the volume averaged {alpha}-particle beta threshold is in the order of 10{sup {minus}4}. Typical growth rates of the n=1 TAE mode can be in the order of 10{sup {minus}2}{omega}{sub A}, where {omega}{sub A}=v{sub A}/qR. Other types of global Alfven waves are stable in D-T tokamaks due to toroidal coupling effects.

  10. Spatial nonlinear absorption of Alfven waves by dissipative plasma taking account bremsstrahlung

    NASA Astrophysics Data System (ADS)

    Taiurskii, A. A.; Gavrikov, M. B.

    2016-10-01

    We study numerically the nonlinear absorption of a plane Alfven wave falling on the stationary boundary of dissipative plasma. This absorption is caused by such factors as the magnetic viscosity, hydrodynamic viscosity, and thermal conductivity of electrons and ions, bremsstrahlung and energy exchange between plasma components. The relevance of this investigation is due to some works, published in 2011, with regard to the heating mechanism of the solar corona and solar wind generation as a result of the absorption of plasma Alfven waves generated in the lower significantly colder layers of the Sun. Numerical analysis shows that the absorption of Alfven waves occurs at wavelengths of the order of skin depth, in which case the classical MHD equations are inapplicable. Therefore, our research is based on equations of two-fluid magnetohydrodynamics that take into account the inertia of the electrons. The implicit difference scheme proposed here for calculating plane-parallel flows of two-fluid plasma reveals a number of important patterns of absorption and thus allows us to study the dependence of the absorption on the Alfven wave frequency and the electron thermal conductivity and viscosity, as well as to evaluate the depth and the velocity of plasma heating during the penetration of Alfven waves interacting with dissipative plasma.

  11. Theoretical Studies of Drift-Alfven and Energetic Particle Physics

    SciTech Connect

    CHEN, L.

    2014-05-14

    The research program supported by this DOE grant has been rather successful and productive in terms of both scientific investigations as well as human resources development; as demonstrated by the large number (60) of journal articles, 6 doctoral degrees, and 3 postdocs. This PI is particularly grateful to the generous support and flexible management of the DOE–SC-OFES Program. He has received three award/prize (APS Excellence in Plasma Physics Research Award, 2004; EPS Alfven Prize, 2008; APS Maxwell Prize, 2012) as the results of research accomplishments supported by this grant.

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

    NASA Technical Reports Server (NTRS)

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

    1954-01-01

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

  13. Beam Distribution Modification by Alfven Modes

    SciTech Connect

    White, R. B.; Gorelenkov, N.; Heidbrink, W. W.; Van Zeeland, M. A.

    2010-04-03

    Modification of a deuterium beam distribution in the presence of low amplitude Toroidal Alfven (TAE) eigenmodes and Reversed Shear Alfven (RSAE) eigenmodes in a toroidal magnetic confinement device is examined. Comparison with experimental data shows that multiple low amplitude modes can account for significant modification of high energy beam particle distributions. It is found that there is a stochastic threshold for beam transport, and that the experimental amplitudes are only slightly above this threshold. The modes produce a substantial central flattening of the beam distribution.

  14. Beam Distribution Modification By Alfven Modes

    SciTech Connect

    White, R. B.; Gorelenkov, N.; Heidbrink, W. W.; Van Zeeland, M. A.

    2010-01-25

    Modification of a deuterium beam distribution in the presence of low amplitude Toroidal Alfven (TAE) eigenmodes and Reversed Shear Alfven (RSAE) eigenmodes in a toroidal magnetic confinement device is examined. Comparison with experimental data shows that multiple low amplitude modes can account for significant modification of high energy beam particle distributions. It is found that there is a stochastic threshold for beam transport, and that the experimental amplitudes are only slightly above this threshold. The modes produce a substantial central flattening of the beam distribution.

  15. Shear-Alfven Waves in Gyrokinetic Plasmas

    SciTech Connect

    W.W.Lee; J.L.V.Lewandowski; T.S. Hahm; Z. Lin

    2000-10-18

    It is found that the thermal fluctuation level of the shear-Alfven waves in a gyrokinetic plasma decreases with plasma b(* cs2/uA2), where cs is the ion acoustic speed and uA is the Alfven velocity. This unique thermodynamic property based on the fluctuation-dissipation theorem is verified in this paper using a new gyrokinetic particle simulation scheme, which splits the particle distribution function into the equilibrium part as well as the adiabatic and nonadiabatic parts.

  16. Solar Coronal Heating via Alfven Wave Turbulence

    SciTech Connect

    Bigot, B.; Galtier, S.; Politano, H.

    2010-03-25

    A short review is given about the self-consistent MHD model of solar coronal heating recently proposed by Bigot et al.(2008) in which the dynamical effect of the background magnetic field along a coronal structure is taken into account through exact results from Alfven wave turbulence. The main properties of the model are given as well as the heating rate and the microturbulent velocity obtained in the case of coronal loops. The conclusion is that Alfven wave turbulence may produce an efficient background heating for the solar corona.

  17. Macroscale particle simulation of kinetic Alfven waves

    NASA Technical Reports Server (NTRS)

    Tanaka, Motohiko; Sato, Tetsuya; Hasegawa, Akira

    1987-01-01

    Two types of simulations of the kinetic Alfven wave are presented using a macroscale particle simulation code (Tanaka and Sato, 1986) which enables individual particle dynamics to be followed in the MHD scales. In this code, low frequency electromagnetic fields are solved by eliminating high frequency oscillations such as the light modes, and the scalar potential electric field is solved by eliminating Lagrangian oscillations. The dependences of the frequency and the Landau damping on the perpendicular wavenumber were studied, and good agreement was found between simulation and theoretical predictions. Some fundamental nonlinear interactions of the kinetic Alfven wave with the particles (parallel acceleration of the electrons) were also noted.

  18. Sawtooth Stabilization and Onset of Alfvenic Instabilities

    NASA Astrophysics Data System (ADS)

    Nishimura, Y.; Cheng, C. Z.

    2011-10-01

    Tokamak sawtooth instabilities can be stabilized by high energy particles as a consequence of conservation of the third adiabatic invariant.On the other hand, termination of the stabilized period is reported due to the onset of Alfvenic instabilities (and thus the absence of the stabilizing mechanism). In this work, employing a kinetic-fluid model, the interaction of m=1 resistive kink mode and high energy particles is investigated. The onset of Alfvenic instabilities is examined as a function of the inversion radius location. D.J. Campbell et al., Phys. Rev. Lett. 60, 2148 (1988); F. Porcelli, Plasma Phys. Controlled Fusion 33, 1601 (1991).

  19. Performance, Stability, and Control Investigation at Mach Numbers from 0.4 to 0.9 of a Model of the "Swallow" with Outer Wing Panels Swept 25 degree with and without Power Simulation

    NASA Technical Reports Server (NTRS)

    Runckel, Jack F.; Schmeer, James W.; Cassetti, Marlowe D.

    1960-01-01

    An investigation of the performance, stability, and control characteristics of a variable-sweep arrow-wing model (the "Swallow") with the outer wing panels swept 25 deg has been conducted in the Langley 16-foot transonic tunnel. The wing was uncambered and untwisted and had RAE 102 airfoil sections with a thickness-to-chord ratio of 0.14 normal to the leading edge. Four outboard engines located above and below the wing provided propulsive thrust, and, by deflecting in the pitch direction and rotating in the lateral plane, also produced control forces. A pair of swept lateral fins and a single vertical fin were mounted on each engine nacelle to provide aerodynamic stability and control. Jets-off data were obtained with flow-through nacelles, stimulating the effects of inlet flow; jet thrust and hot-jet interference effects were obtained with faired-nose nacelles housing hydrogen peroxide gas generators. Six-component force and moment data were obtained through a Mach number range of 0.40 to 0.90 at angles of attack and angles of sideslip from 0 deg to 15 deg. Longitudinal, directional, and lateral control were obtained by deflecting the nacelle-fin combinations as elevators, rudders, and ailerons at several fixed angles for each control.

  20. Analytical Investigation of a Flicker-Type Roll Control for a Mach Number 6 Missile with Aerodynamic Controls Over An Altitude Range of 82,000 to 282,000 feet

    NASA Technical Reports Server (NTRS)

    Lundstrom, Reginald R.; Whitman, Ruth I.

    1959-01-01

    An analytical investigation has been carried out to determine the responses of a flicker-type roll control incorporated in a missile which traverses a range of Mach number of 6.3 at an altitude of 82,000 feet to 5.26 at an altitude of 282,000 feet. The missile has 80 deg delta wings in a cruciform arrangement with aerodynamic controls attached to the fuselage near the wing trailing edge and indexed 450 to the wings. Most of the investigation was carried out on an analog computer. Results showed that roll stabilization that may be adequate for many cases can be obtained over the altitude range considered, providing that the rate factor can be changed with altitude. The response would be improved if the control deflection were made larger at the higher altitudes. lag times less than 0.04 second improve the response appreciably. Asymmetries that produce steady rolling moments can be very detrimental to the response in some cases. The wing damping made a negligible contribution to the response.

  1. Effect of nozzle lateral spacing on afterbody drag and performance of twin-jet afterbody models with convergent-divergent nozzles at Mach numbers up to 2.2

    NASA Technical Reports Server (NTRS)

    Pendergraft, O. C., Jr.; Schmeer, J. W.

    1972-01-01

    Twin-jet afterbody models were investigated by using two balances to measure the thrust-minus-total drag and the afterbody drag, separately, at static conditions and at Mach numbers up to 2.2 for an angle of attack of 0 deg. Hinged-flap convergent-divergent nozzles were tested at subsonic-cruise- and maximum-afterburning-power settings with a high-pressure air system used to provide jet-total-pressure ratios up to 20. Two nozzle lateral spacings were studied, using afterbodies with similar interfairing shapes but with different longitudinal cross-sectional area distributions. Alternate, blunter, interfairings with different shapes for the two spacings, which produced afterbodies having identical cross-sectional area progressions corresponding to an axisymmetric minimum wave-drag configuration, were also tested. The results indicate that the wide-spaced configurations improved the flow field around the nozzles, thereby reducing drag on the cruise nozzles; however, the increased surface and projected cross-sectional areas caused an increase in afterbody drag. Except for a slight advantage with cruise nozzles at subsonic speeds, the wide-spaced configurations had the higher total drag at all other test conditions.

  2. An Investigation of Single- and Dual-Rotation Propellers at Positive and Negative Thrust, and in Combination with an NACA 1-series D-Type Cowling at Mach Numbers up to 0.84

    NASA Technical Reports Server (NTRS)

    Reynolds, Robert M; Samonds, Robert I; Walker, John H

    1957-01-01

    An investigation has been made to determine the aerodynamic characteristics of the NACA 4-(5)(05)-041 four-blade, single-relation propeller and the NACA 4-(5)(05)-037 six- and eight-blade, dual-rotation propellers in combination with various spinners and NACA d-type spinner-cowling combinations at Mach numbers up to 0.84. Propeller force characteristics, local velocity distributions in the propeller planes, inlet pressure recoveries, and static-pressure distributions on the cowling surfaces were measured for a wide range of blade angles, advance ratios, and inlet-velocity ratios. Included are data showing: (a) the effect of extended cylindrical spinners on the characteristics of the single-rotation propeller, (b) the effect of variation of the difference in blade angle setting between the front and rear components of the dual-rotation propellers, (c) the negative- and static-thrust characteristics of the propellers with 1 series spinners, and (d) the effects of ideal- and platform-type propeller-spinner junctures on the pressure-recovery characteristics of the single-rotation propeller-spinner-cowling combination.

  3. Flight Test of a 40-Foot Nominal-Diameter Disk-Gap-Band Parachute Deployed at a Mach Number of 1.91 and a Dynamic Pressure of 11.6 Pounds per Square Foot

    NASA Technical Reports Server (NTRS)

    Eckstrom, Clinton V.; Preisser, John S.

    1968-01-01

    A 40-foot (12.2 meter) nominal-diameter disk-gap-band parachute was flight tested as part of the NASA Supersonic Planetary Entry Decelerator Program (SPED-I). The test parachute was ejected by a deployment mortar from an instrumented payload at an altitude of 140,000 feet (42.5 kilometers). The payload was at a Mach number of 1.91 and the dynamic pressure was 11.6 pounds per square foot (555 newtons per square meter) at the time the parachute deployment mortar was fired. The parachute reached suspension line stretch in 0.43 second with a resultant snatch force loading of 1990 pounds (8850 newtons). The maximum parachute opening load of 6500 pounds (28,910 newtons) came 0.61 second later at a total elapsed time from mortar firing of 1.04 seconds. The first full inflation occurred at 1.12 seconds and stable inflation was achieved at approximately 1.60 seconds. The parachute had an average axial-force coefficient of 0.53 during the deceleration period. During the steady-state descent portion of the flight test, the average effective drag coefficient was also 0.53 and pitch-yaw oscillations of the canopy averaged less than 10 degrees in the altitude region above 100,000 feet (30.5 meters).

  4. Flight Test of a 40-Foot Nominal Diameter Disk-Gap-Band Parachute Deployed at a Mach Number of 3.31 and a Dynamic Pressure of 10.6 Pounds per Square Foot

    NASA Technical Reports Server (NTRS)

    Eckstrom, Clinton V.

    1969-01-01

    A 40-foot-nominal-diameter (12.2 meter) disk-gap-band parachute was flight tested as part of the NASA supersonic high altitude parachute experiment (SHAPE) program. The test parachute (which included an experimental energy absorber in the attachment riser) was deployed from an instrumented payload by means of a deployment mortar when the payload was at a Mach number of 3.31 and a free-stream dynamic pressure of 10.6 pounds per square foot (508 newtons per square meter). The parachute deployed properly, the canopy inflating to a full-open condition at 1.03 seconds after mortar firing. The first full inflation of the canopy was immediately followed by a partial collapse with subsequent oscillations of the frontal area from about 30 to 75 percent of the full-open frontal area. After 1.07 seconds of operation, a large tear appeared in the cloth near the canopy apex. This tear was followed by two additional tears shortly thereafter. It was later determined that a section of the canopy cloth was severely weakened by the effects of aerodynamic heating. As a result of the damage to the disk area of the canopy, the parachute performance was significantly reduced; however, the parachute remained operationally intact throughout the flight test and the instrumented payload was recovered undamaged.

  5. Flight Test of a 30-Foot Nominal Diameter Cross Parachute Deployed at a Mach Number of 1.57 and a Dynamic Pressure of 9.7 Pounds per Square Foot

    NASA Technical Reports Server (NTRS)

    Eckstrom, Clinton V.; Preisser, John S.

    1968-01-01

    A 30-foot (9.1-meter) nominal-diameter cross-type parachute with a cloth area (reference area) of 709 square feet (65.9 square meters) was flight tested in the rocket-launched portion of the NASA Planetary Entry Parachute Program (PEPP). The test parachute was ejected from an instrumented payload by means of a mortar when the system was at a Mach number of 1.57 and a dynamic pressure of 9.7 psf. The parachute deployed to suspension-line stretch in 0.44 second with a resulting snatch-force loading of 1100 pounds (4900 newtons), Canopy inflation began at 0.58 second and a first full inflation was achieved at approximately 0.77 second. The maximum opening load occurred at 0.81 second and was 4255 pounds (18,930 newtons). Thereafter, the test item exhibited a canopy-shape instability in that the four panel arms experienced fluctuations, a "scissoring" type of motion predominating throughout the test period. Calculated values of axial-force coefficient during the deceleration portion of the test varied between 0.35 and 1.05, with an average value of 0.69. During descent, canopy-shape variations had reduced to small amplitudes and resultant pitch-yaw angles of the payload with respect to the local vertical averaged less than 10 degrees. The effective drag coefficient, based on the vertical components of velocity and acceleration during system descent, was 0.78.

  6. Aerodynamic Force Characteristics of a Series of Lifting Cone and Cone-Cylinder Configurations at a Mach Number of 6.83 and Angles of Attack up to 130 Deg

    NASA Technical Reports Server (NTRS)

    Penland, Jim A.

    1961-01-01

    Force tests of a series of right circular cones having semivertex angles ranging from 5 deg to 45 deg and a series of right circular cone-cylinder configurations having semivertex angles ranging from 5 deg to 20 deg and an afterbody fineness ratio of 6 have been made in the Langley 11-inch hypersonic tunnel at a Mach number of 6.83, a Reynolds number of 0.24 x 10.6 per inch, and angles of attack up to 130 deg. An analysis of the results made use of the Newtonian and modified Newtonian theories and the exact theory. A comparison of the experimental data of both cone and cone-cylinder configurations with theoretical calculations shows that the Newtonian concept gives excellent predictions of trends of the force characteristics and the locations with respect to angle of attack of the points of maximum lift, maximum drag, and maximum lift-drag ratio. Both the Newtonian a.nd exact theories give excellent predictions of the sign and value of the initial lift-curve slope. The maximum lift coefficient for conical bodies is nearly constant at a value of 0.5 based on planform area for semivertex angles up to 30 deg. The maximum lift-drag ratio for conical bodies can be expected to be not greater than about 3.5, and this value might be expected only for slender cones having semivertex angles of less than 5 deg. The increments of angle of attack and lift coefficient between the maximum lift-drag ratio and the maximum lift coefficient for conical bodies decrease rapidly with increasing semivertex angles as predicted by the modified Newtonian theory.

  7. Electron acceleration by inertial Alfven waves

    SciTech Connect

    Thompson, B.J.; Lysak, R.L.

    1996-03-01

    Alfven waves reflected by the ionosphere and by inhomogeneities in the Alfven speed can develop an oscillating parallel electric field when electron inertial effects are included. These waves, which have wavelengths of the order of an Earth radius, can develop a coherent structure spanning distances of several Earth radii along geomagnetic field lines. This system has characteristic frequencies in the range of 1 Hz and can exhibit electric fields capable of accelerating electrons in several senses: via Landua resonance, bounce or transit time resonance as discussed by Andre and Eliasson or through the effective potential drop which appears when the transit time of the electrons is much smaller than the wave period, so that the electric fields appear effectively static. A time-dependent model of wave propagation is developed which represents inertial Alfven wave propagation along auroral field lines. The disturbance is modeled as it travels earthward, experiences partial reflections in regions of rapid variation, and finally reflects off a conducting ionosphere to continue propagating antiearthward. The wave experiences partial trapping by the ionospheric and the Alfven speed peaks discussed earlier by Polyakov and Rapoport and Trakhtengerts and Feldstein and later by Lysak. Results of the wave simulation and an accompanying test particle simulation are presented, which indicate that inertial Alfven waves are a possible mechanism for generating electron conic distributions and field-aligned particle precipitation. The model incorporates conservation of energy by allowing electrons to affect the wave via Landau damping, which appears to enhance the effect of the interactions which heat electron populations. 22 refs., 14 figs.

  8. The many faces of shear Alfven waves

    SciTech Connect

    Gekelman, W.; Vincena, S.; Van Compernolle, B.; Morales, G. J.; Maggs, J. E.; Pribyl, P.; Carter, T. A.

    2011-05-15

    One of the fundamental waves in magnetized plasmas is the shear Alfven wave. This wave is responsible for rearranging current systems and, in fact all low frequency currents in magnetized plasmas are shear waves. It has become apparent that Alfven waves are important in a wide variety of physical environments. Shear waves of various forms have been a topic of experimental research for more than fifteen years in the large plasma device (LAPD) at UCLA. The waves were first studied in both the kinetic and inertial regimes when excited by fluctuating currents with transverse dimension on the order of the collisionless skin depth. Theory and experiment on wave propagation in these regimes is presented, and the morphology of the wave is illustrated to be dependent on the generation mechanism. Three-dimensional currents associated with the waves have been mapped. The ion motion, which closes the current across the magnetic field, has been studied using laser induced fluorescence. The wave propagation in inhomogeneous magnetic fields and density gradients is presented as well as effects of collisions and reflections from boundaries. Reflections may result in Alfvenic field line resonances and in the right conditions maser action. The waves occur spontaneously on temperature and density gradients as hybrids with drift waves. These have been seen to affect cross-field heat and plasma transport. Although the waves are easily launched with antennas, they may also be generated by secondary processes, such as Cherenkov radiation. This is the case when intense shear Alfven waves in a background magnetoplasma are produced by an exploding laser-produced plasma. Time varying magnetic flux ropes can be considered to be low frequency shear waves. Studies of the interaction of multiple ropes and the link between magnetic field line reconnection and rope dynamics are revealed. This manuscript gives us an overview of the major results from these experiments and provides a modern

  9. Heat-Transfer and Pressure Measurements from a Flight Test of the Third 1/18-Scale Model of the Titan Intercontinental Ballistic Missile up to a Mach Number of 3.86 and Reynolds Number per Foot of 23.5 x 10(exp 6) and a Comparison with Heat Transfer

    NASA Technical Reports Server (NTRS)

    Graham, John B., Jr.

    1958-01-01

    Heat-transfer and pressure measurements were obtained from a flight test of a 1/18-scale model of the Titan intercontinental ballistic missile up to a Mach number of 3.86 and Reynolds number per foot of 23.5 x 10(exp 6) and are compared with the data of two previously tested 1/18-scale models. Boundary-layer transition was observed on the nose of the model. Van Driest's theory predicted heat-transfer coefficients reasonably well for the fully laminar flow but predictions made by Van Driest's theory for turbulent flow were considerably higher than the measurements when the skin was being heated. Comparison with the flight test of two similar models shows fair repeatability of the measurements for fully laminar or turbulent flow.

  10. Dynamic Investigation of Release Characteristics of a Streamlined Internal Store from a Simulated Bomb Bay of the Republic F-105 Airplane at Mach Numbers of 0.8, 1.4, and 1.98, Coord. No. AF-222

    NASA Technical Reports Server (NTRS)

    Lee, John B.

    1956-01-01

    An investigation has been conducted in the 27- by 27-inch preflight jet of the Langley Pilotless Aircraft Research Station at Wallops Island, Va., of the release characteristics of a dynamically scaled streamlined-type internally carried store from a simulated bomb bay at Mach numbers M(sub o) of 0.8, 1.4, and 1.98. A l/17-scale model of the Republic F-105 half-fuselage and bomb-bay configuration was used with a streamlined store shape of a fineness ratio of 6.00. Simulated altitudes were 3,400 feet at M(sub o) = 0.8, 3,400, and 29,000 feet at M(sub o) = 1.4, and 29,000 feet at M(sub o) = 1.98. At supersonic speeds, high pitching moments are induced on the store in the vicinity of the bomb bay at high dynamic pressures. Successful ejections could not be made with the original configuration at supersonic speeds at near sea-level conditions. The pitching moments caused by unsymmetrical pressures on the store in a disturbed flow field were overcome by replacing the high-aspect-ratio fin with a low-aspect-ratio fin that had a 30-percent area increase which was less subject to aeroelastic effects. Release characteristics of the store were improved by orienting the fins so that they were in a more uniform flow field at the point of store release. The store pitching moments were shown to be reduced by increasing the simulated altitude. Favorable ejections were made at subsonic speeds at near sea-level conditions.

  11. Single-Mach and double-Mach reflection - Its representation in Ernst Mach's historical soot method

    NASA Astrophysics Data System (ADS)

    Krehl, P.

    In 1875 Ernst Mach discovered the effect of irregular interaction of shock waves, the so-called single Mach reflection (SMR), which for symmetric geometry is characterized by two triple points. He recorded their two trajectories on a soot-covered glass plate. Appearing as two mirror-symmetric V-branches, they form the well-known Mach soot funnel. Combining this soot method with the schlieren technique facilitates the interpretation of soot-recorded interaction phenomena as well as allows to resolve the soot removal mechanism in time. Increasing the dynamic recording range of the soot layer in terms of reflected shock pressures even renders visualization of double-Mach reflection (DMR) which, in the case of symmetric shock interaction, is characterized by a second concentric, external 'double-Mach funnel'. At transition of DMR to SMR it merges into the ordinary 'single-Mach funnel'.

  12. Experimental aspects of effects of high-energy particles on Alfven modes

    SciTech Connect

    Heidbrink, W.W.

    1994-10-01

    Global Alfven modes are observed in a number of tokamaks, including DIII-D and TFTR. Instabilities occur during neutral-beam injection and during fast-wave ICRF heating, and may recently have been observed during alpha-particle heating. Identification of toroidicity-induced Alfven eigenmodes (TAE) is based primarily on the scaling of the real frequency of the mode. Other modes, including the beta-induced Alfven eigenmode (BAE), are also observed. The stability threshold of TAE modes agree (to within a factor of two) with theoretical predictions. Toroidal mode numbers of n = 2-6 are usually most unstable, as theoretically expected. Measurements of the poloidal and radial mode structure are consistent with theoretical predictions, but the uncertainties are large. Both TAE and BAE modes can cause large, concentrated losses of fast ions. Phenomenologically, beam-driven Alfven modes usually {open_quotes}saturate{close_quotes} through bursts that expel beam ions, while modes observed during ICPF heating approach a steady saturation amplitude.

  13. Design of Mach-4 and Mach-6 Nozzles for the NASA LaRC 8-Ft High Temperature Tunnel

    NASA Technical Reports Server (NTRS)

    Gaffney, Richard L., Jr.

    2009-01-01

    The aerodynamic contours for two new nozzles have been designed for the NASA Langley Research Center 8-Foot High Temperature Tunnel. The new Mach-4 and Mach-6 contours have 54.5-inch exit-diameters allowing for testing at high dynamic pressures. The Mach-4 nozzle will extend the test capability of the facility and allow turbine-based combined-cycle propulsion systems to be tested at conditions appropriate for the transition from the turbine to the scramjet flowpath. The Mach-6 nozzle will serve a dual purpose; to provide a Mach-6 test capability at high dynamic pressure and to be used in conjunction with an existing mixer section for testing at lower enthalpy conditions. This second use will extend the life of the existing Mach-7 nozzle which has been used for this purpose. The two new nozzles, in conjunction with existing nozzles, will allow for testing at Mach numbers of 3, 4, 5 and 6 at high dynamic pressures, and Mach 4, 5 and 7 at lower dynamic pressures but larger scales.

  14. Wind tunnel and analytical investigation of over-the-wing propulsion/air frame interferences for a short-haul aircraft at Mach numbers from 0.6 to 0.78. [conducted in the Lewis 8 by 6 foot tunnel

    NASA Technical Reports Server (NTRS)

    Wells, O. D.; Lopez, M. L.; Welge, H. R.; Henne, P. A.; Sewell, A. E.

    1977-01-01

    Results of analytical calculations and wind tunnel tests at cruise speeds of a representative four engine short haul aircraft employing upper surface blowing (USB) with a supercritical wing are discussed. Wind tunnel tests covered a range of Mach number M from 0.6 to 0.78. Tests explored the use of three USB nozzle configurations. Results are shown for the isolated wing body and for each of the three nozzle types installed. Experimental results indicate that a low angle nacelle and streamline contoured nacelle yielded the same interference drag at the design Mach number. A high angle powered lift nacelle had higher interference drag primarily because of nacelle boattail low pressures and flow separation. Results of varying the spacing between the nacelles and the use of trailing edge flap deflections, wing upper surface contouring, and a convergent-divergent nozzle to reduce potential adverse jet effects were also discussed. Analytical comparisons with experimental data, made for selected cases, indicate favorable agreement.

  15. Space Shuttle Orbiter trimmed center-of-gravity extension study. Volume 9: Effects of configuration modifications on the aerodynamic characteristics of the 140 A/B Orbiter at Mach numbers of 1.5, 2.0, and 2.5. [wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Phillips, W. P.; Fournier, R. H.

    1985-01-01

    Wind-tunnel tests were conducted at Mach 1.5 to 2.5 to determine the effect of modifications designed to extend the forward center-of-gravity trim capability on the static longitudal and lateral directional characteristics of a Space shuttle 140 A/B orbiter model (0.01 scale). The modifications consisted of a forward-extended wing fillet, a flat plate canard, and a blended canard. The investigation was conducted in the low Mach number test section of the Langley unitary plan wind tunnel at a Reynolds number of approximately 2.15 million based on the fuselage reference length. The test angle of attack range was -1 deg to 32 deg and the sideslip angles were 0 deg and 5 deg.

  16. Low-n shear Alfven spectra in axisymmetric toroidal plasmas

    SciTech Connect

    Cheng, C.Z.; Chance, M.S.

    1985-11-01

    In toroidal plasmas, the toroidal magnetic field is nonuniform over a magnetic surface and causes coupling of different poloidal harmonics. It is shown both analytically and numerically that the toroidicity not only breaks up the shear Alfven continuous spectrum, but also creates new, discrete, toroidicity-induced shear Alfven eigenmodes with frequencies inside the continuum gaps. Potential applications of the low-n toroidicity-induced shear Alfven eigenmodes on plasma heating and instabilities are addressed. 17 refs., 4 figs.

  17. Multiplicity of low-shear toroidal Alfven eigenmodes

    SciTech Connect

    Candy, J.; Breizman, B.N. |; Van Dam, J.W.; Ozeki, T.

    1996-01-01

    An enlarged spectrum of ideal toroidal Alfven eigenmodes is demonstrated to exist within a toroidicity-induced Alfven gap when the inverse aspect ratio is comparable to or larger than the value of the magnetic shear. This limit is appropriate for the low-shear region in most tokamaks, especially those with low aspect ratio. The new modes may be destabilized by fusion-product alpha particles more easily than the standard toroidal Alfven eigenmodes.

  18. Characteristics of Short Wavelength Compressional Alfven Eigenmodes

    SciTech Connect

    Fredrickson, E D; Podesta, M; Bortolon, A; Crocker, N A; Gerhardt, S P; Bell, R E; Diallo, A; LeBlanc, B; Levinton, F M

    2012-12-19

    Most Alfvenic activity in the frequency range between Toroidal Alfven Eigenmodes and roughly one half of the ion cyclotron frequency on NSTX [M. Ono, et al., Nucl. Fusion 40 (2000) 557], that is, approximately 0.3 MHz up to ≈ 1.2 MHz, are modes propagating counter to the neutral beam ions. These have been modeled as Compressional and Global Alfven Eigenmodes (CAE and GAE) and are excited through a Doppler-shifted cyclotron resonance with the beam ions. There is also a class of co-propagating modes at higher frequency than the counter-propagating CAE and GAE. These modes have been identified as CAE, and are seen mostly in the company of a low frequency, n=1 kink-like mode. In this paper we present measurements of the spectrum of these high frequency CAE (hfCAE), and their mode structure. We compare those measurements to a simple model of CAE and present evidence of a curious non-linear coupling of the hfCAE and the low frequency kink-like mode.

  19. Mach and the Philosophy of Science.

    ERIC Educational Resources Information Center

    Bradley, J.

    1991-01-01

    The relationship between sense-perception and science studied by Ernst Mach is described. The author's view is that the distinction Mach makes between two different kinds of metrical concept is Mach's greatest contribution to science. (KR)

  20. Alfven continuum and Alfven eigenmodes in the National Compact Stellarator Experiment

    SciTech Connect

    Fesenyuk, O.P.; Kolesnichenko, Ya.I.; Lutsenko, V.V.; White, R.B.; Yakovenko, Yu.V.

    2004-12-01

    The Alfven continuum (AC) in the National Compact Stellarator Experiment (NCSX) [G. H. Neilson et al., in Fusion Energy 2002, 19th Conference Proceedings, Lyon, 2002 (International Atomic Energy Agency, Vienna, 2003), Report IAEA-CN-94/IC-1] is investigated with the AC code COBRA [Ya. I. Kolesnichenko et al., Phys. Plasmas 8, 491 (2001)]. The resonant interaction of Alfven eigenmodes and the fast ions produced by neutral beam injection is analyzed. Alfven eigenmodes residing in one of the widest gap of the NCSX AC, the ellipticity-induced gap, are studied with the code BOA-E [V. V. Lutsenko et al., in Fusion Energy 2002, 19th Conference Proceedings, Lyon, 2002 (International Atomic Energy Agency, Vienna, 2003), Report IAEA-CN-94-TH/P3-16].

  1. Generation of kinetic Alfven waves by beam-plasma interaction in non-uniform plasma

    SciTech Connect

    Hong, M. H.; Lin, Y.; Wang, X. Y.

    2012-07-15

    This work reports a novel mechanism of the generation of kinetic Alfven waves (KAWs) using a two-dimensional hybrid simulation: the KAWs are generated by ion beam-plasma interaction in a non-uniform plasma boundary layer, in which the bulk velocity of the ion beam is assumed to be parallel to the ambient magnetic field. As a result of the beam-plasma interaction, strong shear Alfven waves as well as fast mode compressional waves are first generated on the side of the boundary layer with a high density and thus a low Alfven speed, propagating along the background magnetic field. Later, Alfven waves also form inside the boundary layer with a continuous spectrum. As the perpendicular wave number k{sub Up-Tack} of these unstably excited waves increases with time, large-amplitude, short wavelength KAWs with k{sub Up-Tack } Much-Greater-Than k{sub ||} clearly form in the boundary layer. The physics for the generation of KAWs is discussed.

  2. Generation of shear Alfven waves by a rotating magnetic field source: Three-dimensional simulations

    SciTech Connect

    Karavaev, A. V.; Gumerov, N. A.; Papadopoulos, K.; Shao, Xi; Sharma, A. S.; Gekelman, W.; Wang, Y.; Van Compernolle, B.; Pribyl, P.; Vincena, S.

    2011-03-15

    The paper discusses the generation of polarized shear Alfven waves radiated from a rotating magnetic field source created via a phased orthogonal two-loop antenna. A semianalytical three-dimensional cold two-fluid magnetohydrodynamics model was developed and compared with recent experiments in the University of California, Los Angeles large plasma device. Comparison of the simulation results with the experimental measurements and the linear shear Alfven wave properties, namely, spatiotemporal wave structure, a dispersion relation with nonzero transverse wave number, the magnitude of the wave dependences on the wave frequency, show good agreement. From the simulations it was found that the energy of the Alfven wave generated by the rotating magnetic field source is distributed between the kinetic energy of ions and electrons and the electromagnetic energy of the wave as: {approx}1/2 is the energy of the electromagnetic field, {approx}1/2 is the kinetic energy of the ion fluid, and {approx}2.5% is the kinetic energy of electron fluid for the experiment. The wave magnetic field power calculated from the experimental data and using a fluid model differ by {approx}1% and is {approx}250 W for the experimental parameters. In both the experiment and the three-dimensional two-fluid magnetohydrodynamics simulations the rotating magnetic field source was found to be very efficient for generating shear Alfven waves.

  3. Riemann solvers and Alfven waves in black hole magnetospheres

    NASA Astrophysics Data System (ADS)

    Punsly, Brian; Balsara, Dinshaw; Kim, Jinho; Garain, Sudip

    2016-09-01

    In the magnetosphere of a rotating black hole, an inner Alfven critical surface (IACS) must be crossed by inflowing plasma. Inside the IACS, Alfven waves are inward directed toward the black hole. The majority of the proper volume of the active region of spacetime (the ergosphere) is inside of the IACS. The charge and the totally transverse momentum flux (the momentum flux transverse to both the wave normal and the unperturbed magnetic field) are both determined exclusively by the Alfven polarization. Thus, it is important for numerical simulations of black hole magnetospheres to minimize the dissipation of Alfven waves. Elements of the dissipated wave emerge in adjacent cells regardless of the IACS, there is no mechanism to prevent Alfvenic information from crossing outward. Thus, numerical dissipation can affect how simulated magnetospheres attain the substantial Goldreich-Julian charge density associated with the rotating magnetic field. In order to help minimize dissipation of Alfven waves in relativistic numerical simulations we have formulated a one-dimensional Riemann solver, called HLLI, which incorporates the Alfven discontinuity and the contact discontinuity. We have also formulated a multidimensional Riemann solver, called MuSIC, that enables low dissipation propagation of Alfven waves in multiple dimensions. The importance of higher order schemes in lowering the numerical dissipation of Alfven waves is also catalogued.

  4. Effect of Dust Grains on Solitary Kinetic Alfven Wave

    SciTech Connect

    Li Yangfang; Wu, D. J.; Morfill, G. E.

    2008-09-07

    Solitary kinetic Alfven wave has been studied in dusty plasmas. The effect of the dust charge-to-mass ratio is considered. We derive the Sagdeev potential for the soliton solutions based on the hydrodynamic equations. A singularity in the Sagdeev potential is found and this singularity results in a bell-shaped soliton. The soliton solutions comprise two branches. One branch is sub-Alfvenic and the soliton velocities are much smaller than the Alfven speed. The other branch is super-Alfvenic and the soliton velocities are very close to or greater than the Alfven speed. Both compressive and rarefactive solitons can exist in each branch. For the sub-Alfvenic branch, the rarefactive soliton is a bell shape curve which is much narrower than the compressive one. In the super-Alfvenic branch, however, the compressive soliton is bell-shaped and the rarefactive one is broadened. We also found that the super-Alfvenic solitons can develop to other structures. When the charge-to-mass ratio of the dust grains is sufficiently high, the width of the rarefactive soliton will increase extremely and an electron density depletion will be observed. When the velocity is much higher than the Alfven speed, the bell-shaped soliton will transit to a cusped structure.

  5. Heating of coronal holes by the resonant absorption and dissipation of Alfven waves and its relation to solar wind acceleration

    NASA Technical Reports Server (NTRS)

    Ofman, L.; Davila, J. M.

    1995-01-01

    Coronal hole regions are well known sources of high-speed solar wind, however to account for the observed properties of the solar wind a source of momentum and heat must be included. Alfven waves were suggested as the possible source of heating that accelerates the solar wind. We investigate the propagation of the Alfven waves in coronal holes via numerical solution of the linearized 2-D resistive MHD equations in slab geometry. The Alfven waves are driven at the lower boundary of the coronal hole and propagate into the corona. The waves are reflected at the coronal hole boundary and part of the wave energy leaks out of the coronal hole. We compare the calculated wavelengths and the attenuation rate of the fast mode Alfven waves in the leaky waveguide formed by the coronal hole with the analytical ideal MHD solutions. The formation of resonance heating layers is found to occur when shear Alfven waves propagate in an inhomogeneous coronal hole. The heating is enhanced when fast mode waves couple to the shear Alfven waves. The narrow heating layers are formed near the location of the ideal resonance, which might occur near the coronal hole boundary for a nearly constant density coronal hole, surrounded by a higher density plasma. We investigate the dependence of the heating on the driver frequency, the Lundquist number, and on the heliocentric distance. and find that the low frequency Alfven waves can be an efficient source of heating at large distances from the Sun. We discuss the relation of our results to the observed properties of high-speed solar wind and coronal holes.

  6. About the parametric interplay between ionic Mach number, body-size, and satellite potential in determining the ion depletion in the wake of the S3-2 satellite

    NASA Technical Reports Server (NTRS)

    Samir, U.; Wildman, P. J.; Sagalyn, R. C.; Rich, F.; Brinton, H. C.

    1981-01-01

    The variation of ion current depletion in the wake of the U.S. Air Force satellite S3-2 is quantitatively determined, taking into account altitudes in the range from 300 to 1100 km. The considered investigation has the objective to present results which besides of being of scientific interest per se are useful to the planning of future experiments of body-plasma electrodynamic interactions in a supersonic and sub-Alfvenic flow regime to be conducted on board the Shuttle/Spacelab. More specifically, it is expected that the outcome of investigations of the kind presented will be useful in the planning of instrument location on ejectable ensembles of probes and on the Orbiter itself in future Shuttle/Spacelab missions.

  7. Numerical Study of Pressure Fluctuations due to a Mach 6 Turbulent Boundary Layer

    NASA Technical Reports Server (NTRS)

    Duan, Lian; Choudhari, Meelan M.

    2013-01-01

    Direct numerical simulations (DNS) are used to examine the pressure fluctuations generated by a Mach 6 turbulent boundary layer with nominal freestream Mach number of 6 and Reynolds number of Re(sub t) approx. =. 464. The emphasis is on comparing the primarily vortical pressure signal at the wall with the acoustic freestream signal under higher Mach number conditions. Moreover, the Mach-number dependence of pressure signals is demonstrated by comparing the current results with those of a supersonic boundary layer at Mach 2.5 and Re(sub t) approx. = 510. It is found that the freestream pressure intensity exhibits a strong Mach number dependence, irrespective of whether it is normalized by the mean wall shear stress or by the mean pressure, with the normalized fluctuation amplitude being significantly larger for the Mach 6 case. Spectral analysis shows that both the wall and freestream pressure fluctuations of the Mach 6 boundary layer have enhanced energy content at high frequencies, with the peak of the premultiplied frequency spectrum of freestream pressure fluctuations being at a frequency of omega(delta)/U(sub infinity) approx. = 3.1, which is more than twice the corresponding frequency in the Mach 2.5 case. The space-time correlations indicate that the pressure-carrying eddies for the higher Mach number case are of smaller size, less elongated in the spanwise direction, and convect with higher convection speeds relative to the Mach 2.5 case. The demonstrated Mach-number dependence of the pressure field, including radiation intensity, directionality, and convection speed, is consistent with the trend exhibited in experimental data and can be qualitatively explained by the notion of "eddy Mach wave" radiation.

  8. Gamma-ray bursts from sheared Alfven waves

    NASA Technical Reports Server (NTRS)

    Melia, Fulvio; Fatuzzo, Marco

    1991-01-01

    The physical process by which sheared Alfven waves can accelerate electrons to a Lorentz factor of 10,000 to 100,000 within 5 km of the stellar surface is applied to a study of gamma-ray bursts, taking both resonant and nonresonant scattering into account. Several very encouraging features of the model are discussed. Although the field is oscillatory, virtually all the charges are ejected from the system, resulting in very little backheating of the stellar surface. The particle number density is accounted for naturally in terms of BA0 and m, which in principle are known from the physical manifestation of the agent causing the crustal disturbance. The resulting gamma-ray spectrum compares very favorably with the observation. The model restricts the geometry of the emission region, in the sense that only the Compton upscattering of soft photons from a warm polar cap can produce the correct spectral shape.

  9. Mach, methodology, hysteresis and economics

    NASA Astrophysics Data System (ADS)

    Cross, R.

    2008-11-01

    This methodological note examines the epistemological foundations of hysteresis with particular reference to applications to economic systems. The economy principles of Ernst Mach are advocated and used in this assessment.

  10. Ulysses Observations of Alfven and Magnetosonic Waves at High Latitude

    NASA Technical Reports Server (NTRS)

    Smith, Edward J.

    1997-01-01

    Ulysses observations provide a unique opportunity to study diverse problems related to Alfven and magnetosonic waves. The large amplitude of the Alfven waves influences the distribution functions of the spiral angle, the azimuthal field component and, possibly, the radial component such that their averages are not equal to their most probable values.

  11. Nonlinear standing Alfven wave current system at Io - Theory

    NASA Astrophysics Data System (ADS)

    Neubauer, F. M.

    1980-03-01

    A nonlinear analytical model is presented of the Alfven current tubes continuing the currents through Io generated by the unipolar inductor effect due to Io's motion relative to the magnetospheric plasma. It was shown that: (1) the portion of the currents needing Io is aligned with the Alfven characteristics at a specific angle to the magnetic field for the special case of perpendicular flow; (2) the Alfven tubes act like an external conductance; (3) the Alfven tubes may be reflected from the torus boundary or the Jovian atmosphere; and (4) from the point of view of the electrodynamic interaction, Io is unique among the Jovian satellites because of its ionosphere arising from ionized volcanic gases and a high external Alfvenic conductance.

  12. Drift-Alfven eigenmodes in inhomogeneous plasma

    SciTech Connect

    Vranjes, J.; Poedts, S.

    2006-03-15

    A set of three nonlinear equations describing drift-Alfven waves in a nonuniform magnetized plasma is derived and discussed both in linear and nonlinear limits. In the case of a cylindric radially bounded plasma with a Gaussian density distribution in the radial direction the linearized equations are solved exactly yielding general solutions for modes with quantized frequencies and with radially dependent amplitudes. The full set of nonlinear equations is also solved yielding particular solutions in the form of rotating radially limited structures. The results should be applicable to the description of electromagnetic perturbations in solar magnetic structures and in astrophysical column-like objects including cosmic tornados.

  13. Interplanetary Alfven waves and auroral (substorm) activity: IMP 8

    SciTech Connect

    Tsurutani, B.T.; Gould, T.; Goldstein, B.E. ); Gonzalez, W.D. ); Sugiura, Masahisa )

    1990-03-01

    Almost year of IMP 8 interplanetary magnetic field and plasma data (Days 1-312, 1979) have been examined to determine the interplanetary causes of geomagnetic AE activity. The nature of the interplanetary medium (Alfvenic or non-Alfvenic) and the B{sub 2} correlation with AE were examined over 12-hour increments throughout the study. It is found that Alfvenic wave intervals (defined as V{sub x}-B{sub x} cross-correlation coefficients of >0.6) are present over 60% of the time and the southward component of the Alfven waves is well correlated with AE (average peak correlation coefficient 0.62), with a median lag of 43 min. The most probable delay of AE from B{sub s} is considerably shorter, about 20-25 min. Southward magnetic fields during non-Alfvenic intervals (V{sub x}-B{sub x} cross-correlation coefficients of < 0.4) are equally effective in producing geomagnetic activity. Peak correlation coefficients and lags are similar to those of Alfvenic intervals. From this statistical study, no major differences in the magnetospheric response to Alfvenic and non-Alfvenic intervals were obvious. The high-intensity long-duration continuous AE activity (HILDCAA) events discussed previously by Tsurutani and Gonzalez (1987) are demosntrated to be caused by the southward components of the Alfven waves, presumably through the process of magnetic reconnection. The lag times of AE from B{sub s} were variable from event to event (and at different times within the Alfven wave train), ranging from 45 min to as little as 0 min. The cause of this variable delay is somewhat surprising and is not presently well understood.

  14. Mach Probe Wakes are Important in Weakly Magnetized, Collisional Plasmas

    NASA Astrophysics Data System (ADS)

    Gosselin, Jordan James; Thakur, Saikat; Sears, Stephanie; McKee, John; Scime, Earl; Tynan, George

    2015-11-01

    Mach probes are often used as the diagnostic for flow in the scrape off layer (SOL) of tokamaks and in linear devices because of their low cost and ease of construction. However, proper interpretation of the Mach number has been debated, and interpretation methods use different calibration factors for different plasma parameters. The Controlled Shear Decorrelation eXperiment (CSDX) operates in an intermediate magnetization regime. To validate theories in this regime, measurements of the parallel ion velocity were made with Mach probes and laser induced fluorescence (LIF) at magnetic fields from 400 to 1600 gauss. We find that Mach probe measurements indicate higher velocities than LIF at fields above 400 gauss. Reduced downstream plasma density due to probe shadowing is a strong candidate for the cause of the discrepancy. An advective-diffusive model for the geometric shadowing and downstream plasma density is presented. When the model for the density drop is included, the Mach probe results agree with the LIF data. This result should be included by groups using Mach probes to measure parallel velocities in plasmas where the ion-neutral mean free path is shorter than the probe shadow length, Lps = a2Cs /Dperp in linear devices, the SOL, or divertor region of tokamaks. This material is based upon work supported by the U.S. Department of Energy, Office of Science, under Awards Number DE-FG02-07ER54912.

  15. Summary of Free-Flight Zero-Lift Drag Results from Tests of 1/5-Scale Models of the Convair YF-102 and F-102A Airplanes and Several Related Small Equivalent Bodies at Mach Numbers from 0.70 to 1.46

    NASA Technical Reports Server (NTRS)

    Wallskog, Harvey A.

    1954-01-01

    One-fifth-scale rocket-propelled models of the Convair YF-102 and F-102A airplanes were tested to determine free-flight zero-lift drag coefficients through the transonic speed range at Reynolds numbers near those to be encountered by the full-scale airplane. Trim and duct characteristics were obtained along with measurements of total-, internal-, and base-drag coefficients. Additional zero-lift drag tests involved a series of small equivalent-body-of-revolution models which were launched to low supersonic speeds by means of a helium gun. The several small models tested corresponded to the following full-scale airplanes: basic, YF-102, 2-foot (full-scale) fuselage extension, F-102A, F-102A (relocated inlets), F-102A (faired nose), and F-102A (parabolic nose) . Equivalent-body models corresponding to the normal area distribution (derived for Mach number 1.0) of each of these airplane shapes were flown and, in addition, equivalent-body models designed to represent the YF-102 and F-102A airplanes at Mach number 1.2 were tested. External-drag coefficients obtained from the 115-scale tests ranged from 0.0094 to 0.0273 for the YF-102 model and from 0.0100 to 0.0255 for the F-102A model. Forebody external-pressure-drag coefficients (drag rise) at Mach number 1.05 of 0.0183 and 0.0134 were obtained from the 115-scale models of the YF-102 and F-102A, respectively, a 16-percent reduction for the F-102A model. Values of drag rise at Mach number 1.05 from the small equivalent-body tests were nearly the same for the basic, YF-102, and 2-foot-fuselage-extension airplane shapes. Equivalent-body tests of the YF-102 and F-102A shapes showed the latter to have about 25 percent less drag rise as compared with a 16-percent reduction illustrated by the 1/5-scale tests. Additional equivalent-body tests illustrating effects of modifications to the F-102A airplane shape shared that relocating the inlets on the fuselage or altering the nose shape to provide a smoother cross-sectional area

  16. Coupling of global toroidal Alfven eigenmodes and reversed shear Alfven eigenmodes in DIII-D

    SciTech Connect

    Van Zeeland, M. A.; Turnbull, A. D.; Austin, M. E.; Gorelenkov, N. N.; Kramer, G. J.; Nazikian, R.; Heidbrink, W. W.; Ruskov, E.; Makowski, M. A.; McKee, G. R.

    2007-05-15

    Reversed shear Alfven eigenmodes (RSAEs) are typically thought of as being localized near the minima in the magnetic safety factor profile, however, their spatial coupling to global toroidal Alfven eigenmodes (TAEs) has been observed in DIII-D discharges. For a decreasing minimum magnetic safety factor, the RSAE frequency chirps up through that of stable and unstable TAEs. Coupling creates a small gap at the frequency degeneracy point forming two distinct global modes. The core-localized RSAE mode structure changes and becomes temporarily global. Similarly, near the mode frequency crossing point, the global TAE extends deeper into the plasma core. The frequency splitting and spatial structure of the two modes throughout the various coupling stages, as measured by an array of internal fluctuation diagnostics, are in close agreement with linear ideal MHD calculations using the NOVA code. The implications of this coupling for eigenmode stability is also investigated and marked changes are noted throughout the coupling process.

  17. On apparent temperature in low-frequency Alfvenic turbulence

    SciTech Connect

    Nariyuki, Yasuhiro

    2012-08-15

    Low-frequency, parallel propagating Alfvenic turbulence in collisionless plasmas is theoretically studied. Alfvenic turbulence is derived as an equilibrium state (Beltrami field) in the magnetohydrodynamic equations with the pressure anisotropy and multi-species of ions. It is shown that the conservation of the total 'apparent temperature' corresponds to the Bernoulli law. A simple model of the radially expanding solar wind including Alfvenic turbulence is also discussed. The conversion of the wave energy in the 'apparent temperature' into the 'real temperature' is facilitated with increasing radial distance.

  18. Cusp Dynamics-Particle Acceleration by Alfven Waves

    NASA Technical Reports Server (NTRS)

    Ergun, Robert E.; Parker, Scott A.

    2005-01-01

    Successful results were obtained from this research project. This investigation answered and/or made progresses on each of the four important questions that were proposed: (1) How do Alfven waves propagate on dayside open field lines? (2) How are precipitating electrons influenced by propagating Alfven waves? (3) How are various cusp electron distributions generated? (4) How are Alfven waves modified by electrons? During the first year of this investigation, the input parameters, such as density and temperature altitude profiles, of the gyrofluid code on the cusp field lines were constructed based on 3-point satellite observations. The initial gyrofluid result was presented at the GEM meeting by Dr. Samuel Jones.

  19. Upper wing surface boundary layer measurements and static aerodynamic data obtained on a 0.015-scale model (42-0) or the SSV orbiter configuration 140A/B in the LTV HSWT at a Mach number of 4.6 (LA58)

    NASA Technical Reports Server (NTRS)

    Ball, J. W.; Lindahl, R. H.

    1976-01-01

    The purpose of the test was to investigate the nature of the Orbiter boundary layer characteristics at angles of attack from -4 to 32 degrees at a Mach number of 4.6. The effect of large grit, employed as transition strips, on both the nature of the boundary layer and the force and moment characteristics were investigated along with the effects of large negative elevon deflection on lee side separation. In addition, laminar and turbulent boundary layer separation phenomena which could cause asymmetric flow separation were investigated.

  20. Effects of alpha beam on the parametric decay of a parallel propagating circularly polarized Alfven wave: Hybrid simulations

    SciTech Connect

    Gao, Xinliang; Lu, Quanming; Tao, Xin; Hao, Yufei; Wang, Shui

    2013-09-15

    Alfven waves with a finite amplitude are found to be unstable to a parametric decay in low beta plasmas. In this paper, the parametric decay of a circularly polarized Alfven wave in a proton-electron-alpha plasma system is investigated with one-dimensional (1-D) hybrid simulations. In cases without alpha particles, with the increase of the wave number of the pump Alfven wave, the growth rate of the decay instability increases and the saturation amplitude of the density fluctuations slightly decrease. However, when alpha particles with a sufficiently large bulk velocity along the ambient magnetic field are included, at a definite range of the wave numbers of the pump wave, both the growth rate and the saturation amplitude of the parametric decay become much smaller and the parametric decay is heavily suppressed. At these wave numbers, the resonant condition between the alpha particles and the daughter Alfven waves is satisfied, therefore, their resonant interactions might play an important role in the suppression of the parametric decay instability.

  1. A PARALLEL-PROPAGATING ALFVENIC ION-BEAM INSTABILITY IN THE HIGH-BETA SOLAR WIND

    SciTech Connect

    Verscharen, Daniel; Bourouaine, Sofiane; Chandran, Benjamin D. G.; Maruca, Bennett A. E-mail: s.bourouaine@unh.edu E-mail: bmaruca@ssl.berkeley.edu

    2013-08-10

    We investigate the conditions under which parallel-propagating Alfven/ion-cyclotron waves are driven unstable by an isotropic (T{sub {alpha}} = T{sub Parallel-To {alpha}}) population of alpha particles drifting parallel to the magnetic field at an average speed U{sub {alpha}} with respect to the protons. We derive an approximate analytic condition for the minimum value of U{sub {alpha}} needed to excite this instability and refine this result using numerical solutions to the hot-plasma dispersion relation. When the alpha-particle number density is {approx_equal} 5% of the proton number density and the two species have similar thermal speeds, the instability requires that {beta}{sub p} {approx}> 1, where {beta}{sub p} is the ratio of the proton pressure to the magnetic pressure. For 1 {approx}< {beta}{sub p} {approx}< 12, the minimum U{sub {alpha}} needed to excite this instability ranges from 0.7v{sub A} to 0.9v{sub A}, where v{sub A} is the Alfven speed. This threshold is smaller than the threshold of {approx_equal} 1.2v{sub A} for the parallel magnetosonic instability, which was previously thought to have the lowest threshold of the alpha-particle beam instabilities at {beta}{sub p} {approx}> 0.5. We discuss the role of the parallel Alfvenic drift instability for the evolution of the alpha-particle drift speed in the solar wind. We also analyze measurements from the Wind spacecraft's Faraday cups and show that the U{sub {alpha}} values measured in solar-wind streams with T{sub {alpha}} Almost-Equal-To T{sub Parallel-To {alpha}} are approximately bounded from above by the threshold of the parallel Alfvenic instability.

  2. Ground observations of kinetic Alfven waves

    SciTech Connect

    Kloecker, N.; Luehr, H.; Robert, P.; Korth, A.

    1985-01-01

    Ground-based observations with the EISCAT magnetometer of locally confined intense drifting current systems and Geos-2 measurements during four events in November and December 1982 are examined. In the ground-based measurements near the Harang discontinuity, the events are characterized by strong pulsations with amplitudes in the horizontal component up to 1000 nT and periods of about 300 s and longer. They occur in the evening hours adjacent to the poleward side of the discontinuity with the onset of a substorm; at the same time, the inner edge of the plasma sheet passes the Geos-2 position, magnetically conjugate to ground stations. It is shown that the events can be explained in terms of kinetic Alfven waves. 8 references.

  3. Nonlinear, dispersive, elliptically polarized Alfven wavaes

    NASA Technical Reports Server (NTRS)

    Kennel, C. F.; Buti, B.; Hada, T.; Pellat, R.

    1988-01-01

    The derivative nonlinear Schroedinger (DNLS) equation is derived by an efficient means that employs Lagrangian variables. An expression for the stationary wave solutions of the DNLS that contains vanishing and nonvanishing and modulated and nonmodulated boundary conditions as subcases is then obtained. The solitary wave solutions for elliptically polarized quasiparallel Alfven waves in the magnetohydrodynamic limit (nonvanishing, unmodulated boundary conditions) are obtained. These converge to the Korteweg-de Vries and the modified Korteweg-de Vries solitons obtained previously for oblique propagation, but are more general. It is shown that there are no envelope solitary waves if the point at infinity is unstable to the modulational instability. The periodic solutions of the DNLS are characterized.

  4. Experimental Investigation of Effects of Primary Jet Flow and Secondary Flow Through a Zero-length Ejector on Base and Boattail Pressures of a Body of Revolution at Free-stream Mach Numbers of 1.62, 1.93, and 2.41

    NASA Technical Reports Server (NTRS)

    O'Donnell, Robert M; Mcdearmon, Russell W

    1954-01-01

    An investigation was made at free-stream Mach numbers of 1.62, 1.93, and 2.41 to determine the effects of a primary jet and secondary air flow on the base pressure and pressures acting over the boattailsurface of a body of revolution for two secondary discharge areas. The Mach numbers of the primary nozzles were 1 and 3.23 with the secondary mass flow being varied from 0 to 10 percent of the primary mass flow. The ratio of jet stagnation temperature to tunnel stagnation temperature was about 0.96. The Reynolds number range of the investigation was from 2.1 x 10(6) to 2.9 x 10(6)based on body length. All testing was conducted with a turbulent boundary layer along the model. This report presents results obtained with zero-length ejector and covers jet static-pressure ratios from the jet-off condition to a maximum of about 128 for the sonic nozzle and to a maximum of about 9 for the supersonic nozzle.

  5. Particle simulation of Alfven waves excited at a boundary

    SciTech Connect

    Tsung, F.S.; Tonge, J.W.; Morales, G.J.

    2005-01-01

    A particle-in-cell (PIC) code has been developed that is capable of describing the propagation of compressional and shear Alfven waves excited from a boundary. The code is used to elucidate the properties of Alfven wave cones radiated from sources having transverse scale comparable to the electron skin depth. Good agreement between theoretical predictions and simulation results is found over a wide range of frequencies. An investigation has been undertaken of the effect of hot ions on the Alfven wave cones. The PIC simulations demonstrate that as the ion temperature is increased there is a reversal in the cone angle. The reversal implies that there is a cross-field focusing of the shear Alfven waves. This is a feature which is presently being considered in studies of field-line resonances in the earth's magnetic field. The PIC results also illustrate the damping of shear modes due to the Doppler-shifted cyclotron resonance with hot ions.

  6. The Source of Alfven Waves That Heat the Solar Corona

    NASA Technical Reports Server (NTRS)

    Ruzmaikin, A.; Berger, M. A.

    1998-01-01

    We suggest a source for high-frequency Alfven waves invoked in coronal heating and acceleration of the solar wind. The source is associated with small-scale magnetic loops in the chromospheric network.

  7. Ducted kinetic Alfven waves in plasma with steep density gradients

    SciTech Connect

    Houshmandyar, Saeid; Scime, Earl E.

    2011-11-15

    Given their high plasma density (n {approx} 10{sup 13} cm{sup -3}), it is theoretically possible to excite Alfven waves in a conventional, moderate length (L {approx} 2 m) helicon plasma source. However, helicon plasmas are decidedly inhomogeneous, having a steep radial density gradient, and typically have a significant background neutral pressure. The inhomogeneity introduces regions of kinetic and inertial Alfven wave propagation. Ion-neutral and electron-neutral collisions alter the Alfven wave dispersion characteristics. Here, we present the measurements of propagating kinetic Alfven waves in helium helicon plasma. The measured wave dispersion is well fit with a kinetic model that includes the effects of ion-neutral damping and that assumes the high density plasma core defines the radial extent of the wave propagation region. The measured wave amplitude versus plasma radius is consistent with the pile up of wave magnetic energy at the boundary between the kinetic and inertial regime regions.

  8. Mach, Ernst (1838-1916)

    NASA Astrophysics Data System (ADS)

    Murdin, P.

    2000-11-01

    Moravian/Austrian physicist, scientist/philosopher, the basis of whose natural philosophy was that all knowledge is a matter of experiments, indeed sensations, so that `laws of nature' are summaries of experience provided by fallible senses. EINSTEIN, who had been taught by Mach, was very influenced by this perspective and developed the theory of relativity, expressing much the same thing by inco...

  9. Emission of radiation induced by pervading Alfven waves

    SciTech Connect

    Zhao, G. Q.; Wu, C. S.

    2013-03-15

    It is shown that under certain conditions, propagating Alfven waves can energize electrons so that consequently a new cyclotron maser instability is born. The necessary condition is that the plasma frequency is lower than electron gyrofrequency. This condition implies high Alfven speed, which can pitch-angle scatter electrons effectively and therefore the electrons are able to acquire free energy which are needed for the instability.

  10. Theory of semicollisional kinetic Alfven modes in sheared magnetic fields

    SciTech Connect

    Hahm, T.S.; Chen, L.

    1985-02-01

    The spectra of the semicollisional kinetic Alfven modes in a sheared slab geometry are investigated, including the effects of finite ion Larmor radius and diamagnetic drift frequencies. The eigenfrequencies of the damped modes are derived analytically via asymptotic analyses. In particular, as one reduces the resistivity, we find that, due to finite ion Larmor radius effects, the damped mode frequencies asymptotically approach finite real values corresponding to the end points of the kinetic Alfven continuum.

  11. The pseudospectrum of the resistive magnetohydrodynamics operator: Resolving the resistive Alfven paradox

    SciTech Connect

    Borba, D. ); Riedel, K.S. ); Kerner, W.; Huysmans, G.T.A.; Ottaviani, M. ); Schmid, P.J. )

    1994-10-01

    The Alfven paradox'' is that as resistivity decreases, the discrete eigenmodes do not converge to the generalized eigenmodes of the ideal Alfven continuum. To resolve the paradox, the [epsilon]-pseudospectrum of the resistive magnetohydrodynamic (RMHD) operator is considered. It is proven that for any [epsilon], the [epsilon]-pseudospectrum contains the Alfven continuum for sufficiently small resistivity. Formal [epsilon]-[ital pseudoeigenmodes] are constructed using the formal Wentzel--Kramers--Brillouin--Jeffreys solutions, and it is shown that the entire stable half-annulus of complex frequencies with [rho][vert bar][omega][vert bar][sup 2]=[vert bar][bold k][center dot][bold B]([ital x])[vert bar][sup 2] is resonant to order [epsilon], i.e., belongs to the [epsilon]-[ital pseudospectrum]. The resistive eigenmodes are exponentially ill-conditioned as a basis and the condition number is proportional to exp([ital R][sup 1/2][sub [ital M

  12. Effect of Wing Height and Dihedral on the Lateral Stability Characteristics at Low Lift of a 45 Deg Swept-Wing Airplane Configuration as Obtained from Time-Vector Analyses of Rocket-Propelled-Model Flights at Mach Numbers from 0.7 to 1.3

    NASA Technical Reports Server (NTRS)

    Gillis, Clarence L.; Chapman, Rowe, Jr.

    1956-01-01

    Lateral-stability flight tests were made over the Mach number range from 0.7 to 1.3 of models of three airplane configurations having 45deg sweptback wings. One model had a high wing; one, a low wing; and one, a high wing with cathedral. The models were otherwise identical. The lateral oscillations of the models resulting from intermittent yawing disturbances were interpreted in terms of full-scale airplane flying qualities and were further analyzed by the time-vector method to obtain values of the lateral stability derivatives. The effects of changes i n wing height on the static sideslip derivatives were fairly constant in the speed range investigated and agreed well with estimated values based on subsonic wind-tunnel tests. Effects of geometric dihedral on the rolling moment due to sideslip agreed well with theoretical and other experimental results and with a theoretical relation involving the damping in roll. The damping in roll, when compared with theoretical and other experimental results, shared good agreement at supersonic speeds but was somewhat higher at a Mach number of 1.0 and at subsonic speeds. The damping in yaw shared no large changes in the transonic region.

  13. Generation of Alfvenic Waves and Turbulence in Magnetic Reconnection Jets

    NASA Astrophysics Data System (ADS)

    Hoshino, M.

    2014-12-01

    The magneto-hydro-dynamic (MHD) linear stability for the plasma sheet with a localized bulk plasma flow parallel to the neutral sheet is investigated. We find three different unstable modes propagating parallel to the anti-parallel magnetic field line, and we call them as "streaming tearing'', "streaming sausage'', and "streaming kink'' mode. The streaming tearing and sausage modes have the tearing mode-like structure with symmetric density fluctuation to the neutral sheet, and the streaming kink mode has the asymmetric fluctuation. The growth rate of the streaming tearing mode decreases with increasing the magnetic Reynolds number, while those of the streaming sausage and kink modes do not strongly depend on the Reynolds number. The wavelengths of these unstable modes are of the order of the thickness of plasma sheet, which behavior is almost same as the standard tearing mode with no bulk flow. Roughly speaking the growth rates of three modes become faster than the standard tearing mode. The situation of the plasma sheet with the bulk flow can be realized in the reconnection exhaust with the Alfvenic reconnection jet, and the unstable modes may be regarded as one of the generation processes of Alfvenic turbulence in the plasma sheet during magnetic reconnection.

  14. On reflection of Alfven waves in the solar wind

    NASA Technical Reports Server (NTRS)

    Krogulec, M.; Musielak, Z. E.; Suess, S. T.; Moore, R. L.; Nerney, S. F.

    1993-01-01

    We have revisited the problem of propagation of toroidal and linear Alfven waves formulated by Heinemann and Olbert (1980) to compare WKB and non-WKB waves and their effects on the solar wind. They considered two solar wind models and showed that reflection is important for Alfven waves with periods of the order of one day and longer, and that non-WKB Alfven waves are no more effective in accelerating the solar wind than WKB waves. There are several recently published papers which seem to indicate that Alfven waves with periods of the order of several minutes should be treated as non-WKB waves and that these non-WKB waves exert a stronger acceleration force than WKB waves. The purpose of this paper is to study the origin of these discrepancies by performing parametric studies of the behavior of the waves under a variety of different conditions. In addition, we want to investigate two problems that have not been addressed by Heinemann and Olbert, namely, calculate the efficiency of Alfven wave reflection by using the reflection coefficient and identify the region of strongest wave reflection in different wind models. To achieve these goals, we investigated the influence of temperature, electron density distribution, wind velocity and magnetic field strength on the waves. The obtained results clearly demonstrate that Alfven wave reflection is strongly model dependent and that the strongest reflection can be expected in models with the base temperatures higher than 10(exp 6) K and with the base densities lower than 7 x 10(exp 7) cm(exp -3). In these models as well as in the models with lower temperatures and higher densities, Alfven waves with periods as short as several minutes have negligible reflection so that they can be treated as WKB waves; however, for Alfven waves with periods of the order of one hour or longer reflection is significant, requiring a non-WKB treatment. We also show that non-WKB, linear Alfven waves are always less effective in accelerating the

  15. Reflection of Alfven waves in the solar wind

    NASA Technical Reports Server (NTRS)

    Krogulec, M.; Musielak, Z. E.; Suess, S. T.; Nerney, S. F.; Moore, R. L.

    1994-01-01

    We have revisited the problem of propagation of toroidal and linear Alfven waves formulated by Heinemann and Olbert (1980) to compare Wentzel-Kramers-Brillouin (WKB) and non-WKB waves and their effects on the solar wind. They considered two solar wind models and showed that reflection is important for Alfven waves with periods of the order of one day and longer and that non-WKB Alfven waves are no more effective in accelerating the solar wind than in WKB waves. There are several recently published papers that seem to indicate that Alfven waves with periods of the order of several minutes should be treated as non-WKB waves and that these non-WKB waves exert a stronger acceleration force than WKB waves. The purposse of this paper is to study the origin of these discrepancies by performing parametric studies of the behavior of the waves under a variety of different conditions. In addition, we want to investigate two problems that have not been addressed by Heinimann and Olbert, namely, calculate the efficieny of Alfven wave reflection by using the reflection coefficient and identfy the region of strongest wave reflection in different wind models. To achieve these goals, we investigate the influence of temperature, electron desity distribution, wind velocity, and magnetic field strength on te waves. The obtained results clearly demonstrate that Alfven wave reflection is strongly model dependent and that the strongest reflection can be expected in models with the base temperatures higher than 10(exp 6) K and with the base densities lower than 7 x 10(exp 7)/cu cm. In these models as well as in the models with lower temperatures and higher densities Alfven waves with periods as short as several minutes have negligible reflection so that they can be treated as WKB waves; however, for Alfven waves with periods of the order of one hour or longer reflection is significant, requiring a non-WKB treatment. We also show that non-WKB, linear Alfven waves are always less effective

  16. The underexpanded jet Mach disk and its associated shear layer

    NASA Astrophysics Data System (ADS)

    Edgington-Mitchell, Daniel; Honnery, Damon R.; Soria, Julio

    2014-09-01

    High resolution planar particle image velocimetry is used to measure turbulent quantities in the region downstream of the Mach disk in an axisymmetric underexpanded jet issuing from a convergent nozzle. The internal annular shear layer generated by the slip line emanating from the triple point is shown to persist across multiple shock cells downstream. A triple decomposition based on Proper Orthogonal Decomposition shows that the external helical structure associated with the screech tone generated by the jet exerts a strong influence on velocity fluctuations in the initial region of the annular shear layer. This influence manifests as the external vortices producing oscillatory motion of the Mach disk, and thus a forcing of the internal annular shear layer. The internal shear layer is characterized by a number of azimuthal modes of varying wavenumber and type, including both helical and axisymmetric modes. Finally, the possibility of a previously hypothesized recirculation region behind the Mach disk is investigated, with no evidence found to support its existence.

  17. Radial Localization of Toroidal Alfven Eigenmode in Tokamak Plasmas

    NASA Astrophysics Data System (ADS)

    Wang, Zhixuan; Lin, Zhihong; Heidbrink, William; Tobias, Benjamin; van Zeeland, Michael

    2013-10-01

    Toroidal Alfven eigenmode (TAE) with radially extended structures can be driven unstable by pressure gradients of energetic particles (EP). These unstable Alfveneigenmodes (AE) have been routinely observed in fusion experiments to induce a large EP transport, whichcould degrade overall plasma confinement and damagefusion devices.In the well-accepted paradigm, the growth rate of the AEs can be calculated from a perturbative EP contribution to a fixedmode structure and real frequency given by magnetohydrodynamic (MHD) properties of thermal plasmas. However, linear and nonlinear kinetic effects of both EP and thermal plasmasare important and should be treated on the same footing. The gyrokinetic simulation has thus emerged as anecessary and powerful tool for studying the linear andnonlinear dynamics of AEs. In the current work, the gyrokinetic toroidal code(GTC) linear simulation of the tokamakexperiment finds a radial localization of the TAE dueto the non-perturbative EP contribution. The EP-drivenTAE has a radial mode width much smaller than thatpredicted by the MHD theory. The TAE radial positionpeaks at and moves with the location of the strongest EPpressure gradients. Experimental data confirms that the eigenfunction drifts quicklyoutward radially. The non-perturbativeEP contribution also breaks the radial symmetry of the mode structure and induces a TAE frequency dependence on the toroidal mode number, in excellent agreement with the experimental measurements.

  18. Alfven wave filamentation and dispersive phase mixing

    SciTech Connect

    Sulem, P. L.; Passot, T.; Laveder, D.; Borgogno, D.

    2009-11-10

    The formation of three-dimensional magnetic structures from quasi-monochromatic left-hand polarized dispersive Alfven waves, under the effect of transverse collapse and/or the lensing effect of density channels aligned with the ambient magnetic field is discussed, both in the context of the usual Hall-MHD and using a fluid model retaining linear Landau damping and finite Larmor radius corrections. It is in particular shown that in a small-{beta} plasma (that is stable relatively to the filamentation instability in the absence of inhomogeneities), a moderate density enhancement leads the wave energy to concentrate into a filament whose transverse size is prescribed by the dimension of the channel, while for a strong density perturbation, this structure later on evolves to thin helical ribbons where the strong gradients permit dissipation processes to become efficient and heat the plasma. The outcome of this 'dispersive phase mixing' that leads to small-scale formation on relatively extended regions contrasts with the more localized oblique shocks formed in the absence of dispersion. Preliminary results on the effect of weak collisions that lead to an increase of the transverse ion temperature are also briefly mentioned.

  19. Kinetic Alfven eigenmodes in JET and DIII-D

    SciTech Connect

    Jaun, A.; Hellsten, T.; Heidbrink, W.W.; Carolipio, E.

    1996-12-31

    Kinetic effects are studied for global Alfven eigenmodes in realistic tokamak equilibria with finite aspect ratio and plasmas, comparing calculations from the full wave code PENN with experimental measurements. The kinetic plasma model is based on a Larmor radius expansion in toroidal geometry and takes into account the gradients in the equilibrium density and temperatures. It allows for a consistent description of the mode conversion to the kinetic Alfven wave (KAW) and the effect of diamagnetic drifts on electromagnetic waves. Comparisons axe first carried out for a JET discharge, showing that multiple peeks measured in the low frequency Alfven spectrum are the signature of kinetic Alfven eigenmodes (KAE) induced through coupling between a global ellipticity Alfven eigenmode (EAE) and the KAW. In general, series of modes appear in the proximity of global fluid modes, some with a regular spacing in frequency and a very weak Landau damping of {vert_bar}{gamma}/{omega}{vert_bar} {approx_equal} 0.0007. A kinetic analysis of a DIII-D discharge shows that TAE mode wavefields reach the plasma core through electromagnetic drift waves which propagate because of finite temperature gradients in the regions of small k{sub {parallel}}. They can lead to particle diffusion and may explain the large losses of beam ions observed during the TAE instabilities. Comparisons of frequency and eigenmode structure axe carried out for resistive and kinetic models, between the theoretical calculations using the PENN code and the experimental measurements from magnetic probes.

  20. On the existence of finite amplitude, transverse Alfven waves in the interplanetary magnetic field

    NASA Technical Reports Server (NTRS)

    Sari, J. W.

    1977-01-01

    Interplanetary magnetic field data from the Mariner 10 spacecraft were examined for evidence of small and finite amplitude transverse Alfven waves, general finite amplitude Alfven waves, and magnetosonic waves. No evidence for transverse Alfven waves was found. Instead, the field fluctuations were found to be dominated by the general finite amplitude Alfven wave. Such wave modes correspond to non-plane-wave solutions of the nonlinear magnetohydrodynamic equations.

  1. Mach-like capillary-gravity wakes.

    PubMed

    Moisy, Frédéric; Rabaud, Marc

    2014-08-01

    We determine experimentally the angle α of maximum wave amplitude in the far-field wake behind a vertical surface-piercing cylinder translated at constant velocity U for Bond numbers Bo(D)=D/λ(c) ranging between 0.1 and 4.2, where D is the cylinder diameter and λ(c) the capillary length. In all cases the wake angle is found to follow a Mach-like law at large velocity, α∼U(-1), but with different prefactors depending on the value of Bo(D). For small Bo(D) (large capillary effects), the wake angle approximately follows the law α≃c(g,min)/U, where c(g,min) is the minimum group velocity of capillary-gravity waves. For larger Bo(D) (weak capillary effects), we recover a law α∼√[gD]/U similar to that found for ship wakes at large velocity [Rabaud and Moisy, Phys. Rev. Lett. 110, 214503 (2013)]. Using the general property of dispersive waves that the characteristic wavelength of the wave packet emitted by a disturbance is of order of the disturbance size, we propose a simple model that describes the transition between these two Mach-like regimes as the Bond number is varied. We show that the new capillary law α≃c(g,min)/U originates from the presence of a capillary cusp angle (distinct from the usual gravity cusp angle), along which the energy radiated by the disturbance accumulates for Bond numbers of order of unity. This model, complemented by numerical simulations of the surface elevation induced by a moving Gaussian pressure disturbance, is in qualitative agreement with experimental measurements.

  2. Mach-like capillary-gravity wakes.

    PubMed

    Moisy, Frédéric; Rabaud, Marc

    2014-08-01

    We determine experimentally the angle α of maximum wave amplitude in the far-field wake behind a vertical surface-piercing cylinder translated at constant velocity U for Bond numbers Bo(D)=D/λ(c) ranging between 0.1 and 4.2, where D is the cylinder diameter and λ(c) the capillary length. In all cases the wake angle is found to follow a Mach-like law at large velocity, α∼U(-1), but with different prefactors depending on the value of Bo(D). For small Bo(D) (large capillary effects), the wake angle approximately follows the law α≃c(g,min)/U, where c(g,min) is the minimum group velocity of capillary-gravity waves. For larger Bo(D) (weak capillary effects), we recover a law α∼√[gD]/U similar to that found for ship wakes at large velocity [Rabaud and Moisy, Phys. Rev. Lett. 110, 214503 (2013)]. Using the general property of dispersive waves that the characteristic wavelength of the wave packet emitted by a disturbance is of order of the disturbance size, we propose a simple model that describes the transition between these two Mach-like regimes as the Bond number is varied. We show that the new capillary law α≃c(g,min)/U originates from the presence of a capillary cusp angle (distinct from the usual gravity cusp angle), along which the energy radiated by the disturbance accumulates for Bond numbers of order of unity. This model, complemented by numerical simulations of the surface elevation induced by a moving Gaussian pressure disturbance, is in qualitative agreement with experimental measurements. PMID:25215822

  3. A new way to convert Alfven waves into heat in solar coronal holes - Intermittent magnetic levitation

    NASA Technical Reports Server (NTRS)

    Moore, R. L.; Hammer, R.; Musielak, Z. E.; Suess, S. T.; An, C.-H.

    1992-01-01

    In our recent analysis of Alfven wave reflection in solar coronal holes, we found evidence that coronal holes are heated by reflected Alfven waves. This result suggests that the reflection is inherent to the process that dissipates these Alfven waves into heat. We propose a novel dissipation process that is driven by the reflection, and that plausibly dominates the heating in coronal holes.

  4. Magnetospheric filter effect for Pc 3 Alfven mode waves

    NASA Technical Reports Server (NTRS)

    Zhang, X.; Comfort, R. H.; Gallagher, D. L.; Green, J. L.; Musielak, Z. E.; Moore, T. E.

    1994-01-01

    We present a ray-tracing study of the propagation of Pc 3 Alfven mode waves originating at the dayside magnetopause. This study reveals interesting features of a magnetospheric filter effect for these waves. Pc 3 Alfven mode waves cannot penetrate to low Earth altitudes unless the wave frequency is below approximately 30 mHz. Configurations of the dispersion curves and the refractive index show that the gyroresonance and pseudo-cutoff introduced by the heavy ion O(+) block the waves. When the O(+) concentration is removed from the plasma composition, the barriers caused by the O(+) no longer exist, and waves with much higher frequencies than 30 mHz can penetrate to low altitudes. The result that the 30 mHz or lower frequency Alfven waves can be guided to low altitudes agrees with ground-based power spectrum observations at high latitudes.

  5. Magnetospheric filter effect for Pc 3 Alfven mode waves

    NASA Technical Reports Server (NTRS)

    Zhang, X.; Comfort, R. H.; Gallagher, D. L.; Green, J. L.; Musielak, Z. E.; Moore, T. E.

    1995-01-01

    We present a ray-tracing study of the propagation of Pc 3 Alfven mode waves originating at the dayside magnetopause. This study reveals interesting features of magnetospheric filter effect for these waves. Pc 3 Alfven mode waves cannot penetrate to low Earth altitudes unless the wave frequency is below approximately 30 mHz. Configurations of the dispersion curves and the refractive index show that the gyroresonance and pseudo-cutoff introduced by the heavy ion O(+) block the waves. When the O(+) concentration is removed from the plasma composition, the barriers caused by the O(+) no longer exist, and waves with much higher frequencies than 30 mHz can penetrate to low altitudes. The result that the 30 mHz or lower frequency Alfven waves can be guided to low altitudes agrees with ground-based power spectrum observation at high altitudes.

  6. Alfvenically driven slow shocks in the solar chromosphere and corona

    NASA Technical Reports Server (NTRS)

    Hollweg, Joseph V.

    1992-01-01

    The nonlinear evolution of an Alfvenic impulse launched from the photosphere and its dynamical effects on the chromosphere, transition region (TR), and corona are investigated using a simple 1D model. It is found that the leading edge of the torsional pulse can steepen into a fast shock in the chromosphere if the pulse is of sufficiently large amplitude and short duration. A slow shock which develops behind the Alfvenic pulse can reflect downgoing Alfven waves back up to the corona. The upgoing reflected wave can induce a significant upward ejection of the TR. Nonlinear dynamics are found to lead to very impulsive behavior at later times. It is suggested that impulsive events occurring in the TR or corona need not be interpreted in terms of reconnection-driven microflares. It is also found that B(0) in the chromosphere can be amplified when the TR and chromosphere fall.

  7. Resonant wave-particle interactions modified by intrinsic Alfvenic turbulence

    SciTech Connect

    Wu, C. S.; Lee, K. H.; Wang, C. B.; Wu, D. J.

    2012-08-15

    The concept of wave-particle interactions via resonance is well discussed in plasma physics. This paper shows that intrinsic Alfven waves can qualitatively modify the physics discussed in conventional linear plasma kinetic theories. It turns out that preexisting Alfven waves can affect particle motion along the ambient magnetic field and, moreover, the ensuing force field is periodic in time. As a result, the meaning of the usual Landau and cyclotron resonance conditions becomes questionable. It turns out that this effect leads us to find a new electromagnetic instability. In such a process intrinsic Alfven waves not only modify the unperturbed distribution function but also result in a different type of cyclotron resonance which is affected by the level of turbulence. This instability might enable us to better our understanding of the observed radio emission processes in the solar atmosphere.

  8. Radiation from accelerated Alfven solitons in inhomogeneous plasmas

    NASA Technical Reports Server (NTRS)

    Lakhina, G. S.; Buti, B.; Tsintsadze, N. L.

    1990-01-01

    In a weakly inhomogeneous plasma, the large-amplitude Alfven waves propagating parallel to the ambient magnetic field are shown to evolve into accelerated Alfven solitons. Nonlinear interaction of the accelerated Alfven solitons with the Langmuir waves results in the emission of coherent radiations. Analytical expression for the power radiated per unit solid angle from a soliton is derived for two inhomogeneity profiles, namely the linear profile and the parabolic profile. For the case of uniform plasmas, the emission occurs via a decay-type process or resonant modes. In the presence of inhomogeneity, nonresonant modes provide a new channel for the emission of radiation. The power radiated per unit solid angle is computed for the parameters relevant to Comet Halley's plasma environment. For the nonresonant modes it is found to be several orders of magnitude higher than that for the case of resonant modes.

  9. THE ROLE OF TORSIONAL ALFVEN WAVES IN CORONAL HEATING

    SciTech Connect

    Antolin, P.; Shibata, K. E-mail: shibata@kwasan.kyoto-u.ac.j

    2010-03-20

    In the context of coronal heating, among the zoo of magnetohydrodynamic (MHD) waves that exist in the solar atmosphere, Alfven waves receive special attention. Indeed, these waves constitute an attractive heating agent due to their ability to carry over the many different layers of the solar atmosphere sufficient energy to heat and maintain a corona. However, due to their incompressible nature these waves need a mechanism such as mode conversion (leading to shock heating), phase mixing, resonant absorption, or turbulent cascade in order to heat the plasma. Furthermore, their incompressibility makes their detection in the solar atmosphere very difficult. New observations with polarimetric, spectroscopic, and imaging instruments such as those on board the Japanese satellite Hinode, or the Crisp spectropolarimeter of the Swedish Solar Telescope or the Coronal Multi-channel Polarimeter, are bringing strong evidence for the existence of energetic Alfven waves in the solar corona. In order to assess the role of Alfven waves in coronal heating, in this work we model a magnetic flux tube being subject to Alfven wave heating through the mode conversion mechanism. Using a 1.5 dimensional MHD code, we carry out a parameter survey varying the magnetic flux tube geometry (length and expansion), the photospheric magnetic field, the photospheric velocity amplitudes, and the nature of the waves (monochromatic or white-noise spectrum). The regimes under which Alfven wave heating produces hot and stable coronae are found to be rather narrow. Independently of the photospheric wave amplitude and magnetic field, a corona can be produced and maintained only for long (>80 Mm) and thick (area ratio between the photosphere and corona >500) loops. Above a critical value of the photospheric velocity amplitude (generally a few km s{sup -1}) the corona can no longer be maintained over extended periods of time and collapses due to the large momentum of the waves. These results establish several

  10. Finite Pressure Effects on Reversed Shear Alfven Eigenmodes

    SciTech Connect

    G.J. Kramer; N.N. Gorelenkov; R. Nazikian; C.Z. Cheng

    2004-09-03

    The inclusion of finite pressure in ideal-magnetohydrodynamic (MHD) theory can explain the Reversed magnetic Shear Alfven Eigenmodes (RSAE) (or Alfven cascades) that have been observed in several large tokamaks without the need to invoke the energetic particle mechanism for the existence of these modes. The chirping of the RSAEs is cased by changes in the minimum of the magnetic safety factor, q(sub)min, while finite pressure effects explains the observed non-zero minimum frequency of the RSAE when qmin has a rational value. Finite pressure effects also play a dominant role in the existence of the downward chirping RSAE branch.

  11. Resonant Alfven wave instabilities driven by streaming fast particles

    SciTech Connect

    Zachary, A.

    1987-05-08

    A plasma simulation code is used to study the resonant interactions between streaming ions and Alfven waves. The medium which supports the Alfven waves is treated as a single, one-dimensional, ideal MHD fluid, while the ions are treated as kinetic particles. The code is used to study three ion distributions: a cold beam; a monoenergetic shell; and a drifting distribution with a power-law dependence on momentum. These distributions represent: the field-aligned beams upstream of the earth's bow shock; the diffuse ions upstream of the bow shock; and the cosmic ray distribution function near a supernova remnant shock. 92 refs., 31 figs., 12 tabs.

  12. Analysis and gyrokinetic simulation of MHD Alfven wave interactions

    NASA Astrophysics Data System (ADS)

    Nielson, Kevin Derek

    The study of low-frequency turbulence in magnetized plasmas is a difficult problem due to both the enormous range of scales involved and the variety of physics encompassed over this range. Much of the progress that has been made in turbulence theory is based upon a result from incompressible magnetohydrodynamics (MHD), in which energy is only transferred from large scales to small via the collision of Alfven waves propagating oppositely along the mean magnetic field. Improvements in laboratory devices and satellite measurements have demonstrated that, while theories based on this premise are useful over inertial ranges, describing turbulence at scales that approach particle gyroscales requires new theory. In this thesis, we examine the limits of incompressible MHD theory in describing collisions between pairs of Alfven waves. This interaction represents the fundamental unit of plasma turbulence. To study this interaction, we develop an analytic theory describing the nonlinear evolution of interacting Alfven waves and compare this theory to simulations performed using the gyrokinetic code AstroGK. Gyrokinetics captures a much richer set of physics than that described by incompressible MHD, and is well-suited to describing Alfvenic turbulence around the ion gyroscale. We demonstrate that AstroGK is well suited to the study of physical Alfven waves by reproducing laboratory Alfven dispersion data collected using the LAPD. Additionally, we have developed an initialization alogrithm for use with AstroGK that allows exact Alfven eigenmodes to be initialized with user specified amplitudes and phases. We demonstrate that our analytic theory based upon incompressible MHD gives excellent agreement with gyrokinetic simulations for weakly turbulent collisions in the limit that k⊥rho i << 1. In this limit, agreement is observed in the time evolution of nonlinear products, and in the strength of nonlinear interaction with respect to polarization and scale. We also examine the

  13. Ion temperature in plasmas with intrinsic Alfven waves

    SciTech Connect

    Wu, C. S.; Yoon, P. H.; Wang, C. B.

    2014-10-15

    This Brief Communication clarifies the physics of non-resonant heating of protons by low-frequency Alfvenic turbulence. On the basis of general definition for wave energy density in plasmas, it is shown that the wave magnetic field energy is equivalent to the kinetic energy density of the ions, whose motion is induced by the wave magnetic field, thus providing a self-consistent description of the non-resonant heating by Alfvenic turbulence. Although the study is motivated by the research on the solar corona, the present discussion is only concerned with the plasma physics of the heating process.

  14. First Results of PIC Modeling of Kinetic Alfven Wave Dissipation

    NASA Technical Reports Server (NTRS)

    Chulaki, Anna; Hesse, Michael; Zenitani, Seiji

    2007-01-01

    We present first results of an investigation of the kinetic damping of Alfven wave turbulence. The methodology is based on a fully electromagnetic, three-dimensional, particle in cell code. The calculation is initialized by an Alfven wave spectrum. Subsequently, a cascade develops, and damping by coupling to both ions and electrons is observed. We discuss results of these calculations, and present first estimates of damping rates and of the effects of energy transfer on ion and electron distributions. The results pertain to solar wind heating and acceleration.

  15. The transmission of Alfven waves through the Io plasma torus

    NASA Astrophysics Data System (ADS)

    Wright, A. N.; Schwartz, S. J.

    1989-04-01

    The nature of Alfven wave propagation through the Io plasma torus was investigated using a one-dimensional model with uniform magnetic field and an exponential density decrease to a constant value. The solution was interpreted in terms of a wave that is incident upon the torus, a reflected wave, and a wave that is transmitted through the torus. The results obtained indicate that Io's Alfven waves may not propagate completely through the plasma torus, and, thus, the WKB theory and ray tracing may not provide meaningful estimates of the energy transport.

  16. De-anthropomorphizing energy and energy conservation: The case of Max Planck and Ernst Mach

    NASA Astrophysics Data System (ADS)

    Wegener, Daan

    Discussions on the relation between Mach and Planck usually focus on their famous controversy, a conflict between 'instrumentalist' and realist philosophies of science that revolved around the specific issue of the existence of atoms. This article approaches their relation from a different perspective, comparing their analyses of energy and energy conservation. It is argued that this reveals a number of striking similarities and differences. Both Mach and Planck agreed that the law was valid, and they sought to purge energy of its anthropomorphic elements. They did so in radically different ways, however, illustrating the differences between Mach's 'historical' and Planck's 'rationalistic' accounts of knowledge. Planck's attempt to de-anthropomorphize energy was part of his attempt to demarcate theoretical physics from other disciplines. Mach's attempt to de-anthropomorphize energy is placed in the context of fin-de-siècle Vienna. By doing so, this article also proposes a new interpretation of Mach as a philosopher, historian and sociologist of science.

  17. Revitalizing Ernst Mach's Popular Scientific Lectures

    NASA Astrophysics Data System (ADS)

    Euler, Manfred

    2007-06-01

    Compared to Ernst Mach's influence on the conceptual development of physics, his efforts to popularize science and his reflections on science literacy are known to a much lesser degree. The approach and the impact of Mach's popular scientific lectures are discussed in view of today's problems of understanding science. The key issues of Mach's popular scientific lectures, reconsidered in the light of contemporary science, still hold a high potential in fascinating a general audience. Moreover, Mach's grand theme, the relation of the physical to the psychical, is suited to contribute to a dialogue between different knowledge cultures, e.g. science and humanities.

  18. Tolerance of Mach 2.50 axisymmetric mixed-compression inlets to upstream flow variations

    NASA Technical Reports Server (NTRS)

    Choby, D. A.

    1972-01-01

    An investigation of the tolerances of two Mach 2.50 axisymmetric mixed-compression inlets to upstream flow variations was conducted. Tolerances of each inlet to angle of attack as a function of decreasing free-stream Mach number were obtained. A local region of overcompression was formed on the leeward side of the inlet at maximum angle of attack before unstart. This region of overcompression corresponded to local subsonic flow conditions ahead of the geometric throat. A uniform Mach number gradient of 0.10 at the cowl lip plane did not affect the inlet's pressure recovery, mass flow ratio, or diffuser exit total-pressure distortion.

  19. Experiments on Guderley Mach reflection

    NASA Astrophysics Data System (ADS)

    Skews, Beric William; Li, Gavin; Paton, Randall

    2009-06-01

    Experiments have been conducted in a large shock tube to examine the four-wave shock reflection pattern, now known as Guderley reflection (GR). The fourth wave, an expansion, is clearly identified, as is the supersonic patch behind the reflected wave. A shocklet terminating the supersonic patch behind the reflected wave is identified, which forms a second triple point further down the Mach stem. Evidence is presented showing the presence of more than one expansion wave and more than one shocklet, thus indicating the existence of more than one supersonic patch. In order to distinguish between cases with a single patch without the shocklet as originally proposed by Guderley and found in some computations, and the indications of a multi-patch geometry found here, and also in other computations, this latter case is designated Guderley Mach reflection (GMR). Multi-exposure images of the shock propagation superimposed on a single image frame enable estimates to be made of the strength of the major waves, and it is shown that the reflected wave is very weak.

  20. Evaluations of Mach Probe Characteristics in a Subsonic and Supersonic Plasma Flow

    NASA Astrophysics Data System (ADS)

    Ando, Akira; Watanabe, Takashi; Makita, Takahiro; Tobari, Hiroyuki; Hattori, Kunihiko; Inutake, Masaaki

    2004-11-01

    Mach probe characteristics are evaluated by using a directional Langmuir probe (DLP) in a fast flowing plasma produced by a Magneto-Plasma-Dynamic arcjet. The obtained data are compared with Hutchinson's simulation results which are calculated using a particle-in-cell (PIC) code in an unmagnetized plasma[1]. The ion acoustic Mach number Mi is obtained from spectroscopic measurement of plasma flow velocity and ion temperature and is compared to the DLP data in various conditions of plasma flow. The obtained data are in good agreement with the simulation results in the wide range of Mi (0.4 < Mi < 1.7). Two types of the Mach probe, where the probe tips are facing to up-down directions and perp-para ones, are compared with the model equations and the simulation results. Good agreements are obtained among these data and calibration factor to determine the Mach number are evaluated experimentally. Since the conventional Mach probe has only two probe tips, it can only determine the Mi in one direction. We have found that a multi-tip Mach probe could detect flow direction and total flow Mach number. [1] I.H. Hutchinson, Plasma Phys. Control. Fusion, 44,1953(2002).