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At least 109 records · Page 6

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NASA NTRS1972

Analysis of face deformation effects on gas film seal performance

Analyses are presented for compressible fluid flow across shaft face seals with face deformation. The solutions are obtained from an approximate integral analysis. The models, used in this analysis, can predict gas film seal behavior operating at subsonic or choked flow conditions. The flow regime can either be laminar or turbulent. Entrance losses can also be accounted for. When fluid inertia effects are negligible and the sealing faces are slightly deformed, the following results are found for both laminar and turbulent flows: (1) The pressure profiles are independent of fluid properties; and (2) the parallel film leakage equation can be used, provided a characteristic film thickness is used. Pressure profiles are presented for both divergent and convergent seal faces under choked flow conditions.

Zuk, J.↗

NASA NTRS1977

Operational experience in the Langley expansion tube with various test gases

The Langley Expansion Tube is an operational facility capable of producing good quality, highly repeatable, quasi-steady flow for test times sufficient to establish flow about blunt axisymmetric and two-dimensional models. Due to the capability of testing with arbitrary test gases, a wide range of real-gas, hypersonic-hypervelocity flow conditions may be generated. However, for a given test gas, the range of operating conditions producing useful flow is shown to be rather limited; hence, the facility yields a given flow condition for a given test gas, and variation in flow conditions comes about by using different test gases. Data are presented for operations using air and carbon dioxide as test gases. The driver gas was unheated helium at a nominal pressure of 5000 psi (34.5 MN sq m).

Miller, C. G.↗

NASA NTRS2003

Mean Flow Boundary Conditions for Computational Aeroacoustics

In this work, a new type of boundary condition for time-accurate Computational Aeroacoustics solvers is described. This boundary condition is designed to complement the existing nonreflective boundary conditions while ensuring that the correct mean flow conditions are maintained throughout the flow calculation. Results are shown for a loaded 2D cascade, started with various initial conditions.

Hixon, R.↗

NASA NTRS1993

Dynamic stall of an oscillating wing. Part 1: Evaluation of turbulence models

Unsteady flowfields of a two-dimensional oscillating wing are calculated using an implicit, finite-difference, Navier-Stokes numerical scheme using five widely used turbulence models. The objective of this study is to identify an appropriate turbulence model for accurate simulation of three-dimensional dynamic stall. Three unsteady flow conditions corresponding to attached flow, light-stall, and deep-stall of an oscillating wing experiment were chosen as test cases for computations. Results of unsteady airload hysteresis curves, harmonics of unsteady pressures, and instantaneous flow pictures are presented. Comparison of unsteady airloads with experiment show that all models are deficient in some sense and not a single model predicts all airloads consistently and in agreement with experiment for all flow conditions. For the attached flow condition, the Renormalization Group Theory (RNG), the Johnoson-King (J-K), and the Spalart-Allmaras (S-A) models have better performance. The Baldwin-Lomax (B-L) and the Baldwin-Barth (B-B) models fair poorly. At the light-stall condition, the results for the RNG, the J-K, and S-A models are in agreement with experiment for the upstroke but they all over predict the separation shown by the experiment and therefore have bigger hysteresis loops than experimental results. The B-B model results are also in good agreement for upstroke but have poor lift hysteresis for downstroke. It has superior drag and pitching-moment predictions. For deep-stall conditions, the airloads for the RNG, the B -B, and the S-A models have fair agreement with experiment, but the B-B model performed better at the extreme deep-stall condition. Overall, the RNG model provides significant improvement over the B-L model in all flow regimes with no additional computational cost. The Baldwin-Barth model is the most expensive of the models considered here, costing about 2.5 times that of the Baldwin-Lomax model. Finally, a brief discussion of the effects of grid density, time-step size, and numerical dissipation on the unsteady solutions are also presented.

Srinivasan, G. R.↗

NASA NTRS2004

Unsteady Validation of a Mean Flow Boundary Condition for Computational Aeroacoustics

In this work, a previously developed mean flow boundary condition will be validated for unsteady flows. The test cases will be several reference benchmark flows consisting of vortical gusts convecting in a uniform mean flow, as well as the more realistic case of a vortical gust impinging on a loaded 2D cascade. The results will verify that the mean flow boundary condition both imposes the desired mean flow as well as having little or no effect on the instantaneous unsteady solution.

Hixon, R.↗

NASA NTRS1993

Noise Trends of a 0.5 M (20 In.) Diameter Supersonic Throughflow Fan as Measured in an Unmodified Compressor Aerodynamic Test Facility

The tone noise levels of a supersonic throughflow fan were measured at subsonic and supersonic axial duct Mach numbers. The noise in the inlet plenum showed no blade passing and harmonic tones at subsonic or supersonic axial flow conditions. At subsonic axial flow conditions, the supersonic throughflow fan showed no inlet plenum tones at fan operating conditions where tone noise had been previously measured for a subsonic fan design. This lower inlet-quadrant noise level for the supersonic throughflow fan was the result of high subsonic inlet velocities acting to reduce the noise propagating out the inlet. The fan noise, which was prevented from propagating upstream by the high subsonic inlet velocities, appeared to increase the noise in the exhaust duct at subsonic throughflow conditions. The exhaust duct noise decreased at supersonic axial throughflow Mach numbers, with the lowest blade passing and harmonic tones levels being observed at the design axial Mach number of 2.0. Multiple pure tone noise was observed in the inlet duct at subsonic axial flow Mach numbers but was seen only in the exhaust duct at supersonic axial flow conditions.

Dittmar, James H.↗

NASA NTRS1986

Vortex shedding from solid rocket propellant inhibitors

Vortex shedding frequency caused by the protrusion of inhibitors into the flow field of a solid rocket motor is investigated by experimental and mathematical models. The time dependent Navier-Stokes equations are solved using a finite difference technique assuming incompressible, two-dimensional flow under both laminar and turbulent flow conditions. For laminar flow, explicit solutions are obtained using a vorticity-transport equation in place of the Navier-Stokes equations. For turbulent flow, a two-equation (k-epsilon) model is used for turbulent modeling and the SIMPLE algorithm is employed as the computational scheme. Cold flow tests were conducted to confirm the basic flow structure and to determine the vortex shedding frequency under both laminar and turbulent flow conditions. The vortex shedding frequencies were determined using a stroboscope to measure the oscillating frequency of yarn tufts which were fastened to one inhibitor in the models. A hot-film anemometer established the velocity history behind the inhibitor. Good agreement between the theoretical results and measurements of the vortex shedding frequencies is demonstrated.

Shu, P. H.↗

NASA NTRS1996

Analysis of Two-Phase Flow in Damper Seals for Cryogenic Turbopumps

Cryogenic damper seals operating close to the liquid-vapor region (near the critical point or slightly su-cooled) are likely to present two-phase flow conditions. Under single phase flow conditions the mechanical energy conveyed to the fluid increases its temperature and causes a phase change when the fluid temperature reaches the saturation value. A bulk-flow analysis for the prediction of the dynamic force response of damper seals operating under two-phase conditions is presented as: all-liquid, liquid-vapor, and all-vapor, i.e. a 'continuous vaporization' model. The two phase region is considered as a homogeneous saturated mixture in thermodynamic equilibrium. Th flow in each region is described by continuity, momentum and energy transport equations. The interdependency of fluid temperatures and pressure in the two-phase region (saturated mixture) does not allow the use of an energy equation in terms of fluid temperature. Instead, the energy transport is expressed in terms of fluid enthalpy. Temperature in the single phase regions, or mixture composition in the two phase region are determined based on the fluid enthalpy. The flow is also regarded as adiabatic since the large axial velocities typical of the seal application determine small levels of heat conduction to the walls as compared to the heat carried by fluid advection. Static and dynamic force characteristics for the seal are obtained from a perturbation analysis of the governing equations. The solution expressed in terms of zeroth and first order fields provide the static (leakage, torque, velocity, pressure, temperature, and mixture composition fields) and dynamic (rotordynamic force coefficients) seal parameters. Theoretical predictions show good agreement with experimental leakage pressure profiles, available from a Nitrogen at cryogenic temperatures. Force coefficient predictions for two phase flow conditions show significant fluid compressibility effects, particularly for mixtures with low mass content of vapor. Under these conditions, an increase on direct stiffness and reduction of whirl frequency ratio are shown to occur. Prediction of such important effects will motivate experimental studies as well as a more judicious selection of the operating conditions for seals used in cryogenic turbomachinery.

Arauz, Grigory L.↗

NASA NTRS1987

Comparison of solutions to bi-Maxwellian and Maxwellian transport equations for subsonic flows

Conditions corresponding to the steady state subsonic flow of a fully ionized electron-proton plasma in the terrestrial ionosphere are presently characterized by systematically comparing the solutions to the bi-Maxwellian-based 16-moment and Maxwellian-based 13-moment transport equations. The former can account for large temperature anisotropies and the flow of both parallel and perpendicular thermal energy, while the latter account for small temperature anisotropies and only a total heat flow. The comparison is conducted for 2000-10,000 K lower boundary temperatures and 1-4-K/km temperature gradients, over the 1500-13,000-km altitude range.

Demars, H. G.↗

NASA NTRS2012

Error Estimates of the Ares I Computed Turbulent Ascent Longitudinal Aerodynamic Analysis

Numerical predictions of the longitudinal aerodynamic characteristics for the Ares I class of vehicles, along with the associated error estimate derived from an iterative convergence grid refinement, are presented. Computational results are based on an unstructured grid, Reynolds-averaged Navier-Stokes analysis. The validity of the approach to compute the associated error estimates, derived from a base grid to an extrapolated infinite-size grid, was first demonstrated on a sub-scaled wind tunnel model at representative ascent flow conditions for which the experimental data existed. Such analysis at the transonic flow conditions revealed a maximum deviation of about 23% between the computed longitudinal aerodynamic coefficients with the base grid and the measured data across the entire roll angles. This maximum deviation from the wind tunnel data was associated with the computed normal force coefficient at the transonic flow condition and was reduced to approximately 16% based on the infinite-size grid. However, all the computed aerodynamic coefficients with the base grid at the supersonic flow conditions showed a maximum deviation of only about 8% with that level being improved to approximately 5% for the infinite-size grid. The results and the error estimates based on the established procedure are also presented for the flight flow conditions.

Abdol-Hamid, Khaled S.↗

NASA NTRS2016

Data Comparisons and Summary of the Second Aeroelastic Prediction Workshop

This paper presents the computational results generated by participating teams of the second Aeroelastic Prediction Workshop and compare them with experimental data. Aeroelastic and rigid configurations of the Benchmark Supercritical Wing (BSCW) wind tunnel model served as the focus for the workshop. The comparison data sets include unforced ("steady") system responses, forced pitch oscillations and coupled fluid-structure responses. Integrated coefficients, frequency response functions, and flutter onset conditions are compared. The flow conditions studied were in the transonic range, including both attached and separated flow conditions. Some of the technical discussions that took place at the workshop are summarized.

Heeg, Jennifer↗

NASA NTRS1983

Low-speed investigation of the maneuver capability of supersonic fighter wings

A low-speed investigation of wings suitable for supersonic fighter application was undertaken to understand the flow regions developed by these wings through flow visualization and to evaluate areas in which low-speed testing could provide information for extrapolation to higher Mach numbers and Reynolds numbers. This study involved cranked wings of highly-swept inboard panels and outboard panel of various sweeps with leading- and trailing-edge devices. Both theoretical and experimental results with flow visualization are presented. Flow visualization results show that a mixed flow condition (vortex flow on the highly-swept inboard section and attached flow on the cranked outboard section) is possible for wings with low sweep (20 deg) outboard cranked sections. Lift and drag measurements on the planar (no flap deflection) wings show good agreement with theoretical results for leading-edge vortex flow. There is little or no change in planar wing performance at maneuver conditions due to outboard sweep effects. Results also show that for cambered wings (flap deflected) only small improvements in performance over flat wings were obtained at high-lift maneuver conditions.

Hom, K. W.↗

NASA NTRS1993

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

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.

Calleja, John↗

NASA NTRS1980

Wavelength selection and total growth of Gortler vortices

This paper investigates a wavelength selection mechanism under ideal and nonideal flow conditions. The conditions of the total streamwise growth of Gortler vortices generated under ideal and nonideal flow conditions and its effect on the laminar-turbulent transition process is also examined. The stabilizing effects of suction on the total growth of these vortices are analyzed. It is demonstrated that the wavelength selection mechanism and the total growth of the vortices are affected by the departures from ideal flow conditions. It is shown that under ideal flow conditions, a wavelength selection mechanism can be based on the maximum amplification rate of the disturbances. However, this mechanism is easily affected by the departures from ideal flow. The case of streamwise vorticity in the basic flow is analyzed and is shown to affect disturbances through a subharmonic resonance.

Floryan, J. M.↗

NASA NTRS1991

Three-dimensional flow calculations inside SSME GGGT first stage blade rows

A numerical analysis of the first stage of the Space Shuttle Main Engine (SSME) GGGT was conducted using a 3-D Reynolds averaged Navier-Stokes flow solver. This turbine stage was designed to improve both aerodynamic efficiency and durability. The blade has an unconventional shape with a large blade thickness. No experimental data is available to verify the computational results. The objective of the current study is to analyze this turbine blade stage with a well established Navier-Stokes computational method in order to determine if the turbine is operating in the subsonic flow regime and if these are any significant separated flow regions. The stage was analyzed in a steady state flow condition. The inlet vane was analyzed with the flow conditions from the axisymmetric entire stage solution. The viscous flow solution of the first vane is used as the inlet flow condition for the rotor.

Hah, Chunill↗

NASA NTRS2006

Phased-Array Study of Dual-Flow Jet Noise: Effect of Nozzles and Mixers

A 16-microphone linear phased-array installed parallel to the jet axis and a 32-microphone azimuthal phased-array installed in the nozzle exit plane have been applied to identify the noise source distributions of nozzle exhaust systems with various internal mixers (lobed and axisymmetric) and nozzles (three different lengths). Measurements of velocity were also obtained using cross-stream stereo particle image velocimetry (PIV). Among the three nozzle lengths tested, the medium length nozzle was the quietest for all mixers at high frequency on the highest speed flow condition. Large differences in source strength distributions between nozzles and mixers occurred at or near the nozzle exit for this flow condition. The beamforming analyses from the azimuthal array for the 12-lobed mixer on the highest flow condition showed that the core flow and the lobe area were strong noise sources for the long and short nozzles. The 12 noisy spots associated with the lobe locations of the 12-lobed mixer with the long nozzle were very well detected for the frequencies 5 KHz and higher. Meanwhile, maps of the source strength of the axisymmetric splitter show that the outer shear layer was the most important noise source at most flow conditions. In general, there was a good correlation between the high turbulence regions from the PIV tests and the high noise source regions from the phased-array measurements.

Soo Lee, Sang↗

NASA NTRS2003

Demonstration of Separation Control Using Glow-Discharge Plasma Actuators

Active flow control of boundary-layer separation using glow-discharge plasma actuators is studied experimentally. Separation is induced on a flat plate installed in a closed-circuit wind tunnel by a shaped insert on the opposite wall. The flow conditions represent flow over the suction surface of a modem low-pressure-turbine airfoil. The Reynolds number, based on wetted plate length and nominal exit velocity, is varied from 50,000 to 300,000, covering cruise to takeoff conditions. Low (0.2%) and high (2.5%) free-stream turbulence intensities are set using passive grids. A spanwise-oriented phased-plasma-array actuator, fabricated on a printed circuit board, is surface-flush-mounted upstream of the separation point and can provide forcing in a wide frequency range. Static surface pressure measurements and hot-wire anemometry of the base and controlled flows are performed and indicate that the glow-discharge plasma actuator is an effective device for separation control.

Hultgren, Lennart S.↗

NASA NTRS2004

Demonstration of Separation Delay with Glow-Discharge Plasma Actuators

Active flow control of boundary-layer separation using glow-discharge plasma actuators is studied experimentally. Separation is induced on a flat plate installed in a closed-circuit wind tunnel by a shaped insert on the opposite wall. The flow conditions represent flow over the suction surface of a modern low-pressure-turbine airfoil. The Reynolds number, based on wetted plate length and nominal exit velocity, is varied from 50,000 to 300,000, covering cruise to takeoff conditions. Low (0.2 percent) and high (2.5 percent) free-stream turbulence intensities are set using passive grids. A spanwise-oriented phased-plasma-array actuator, fabricated on a printed circuit board, is surface-flush-mounted upstream of the separation point and can provide forcing in a wide frequency range. Static surface pressure measurements and hot-wire anemometry of the base and controlled flows are performed and indicate that the glow-discharge plasma actuator is an effective device for separation control.

Hultgren, Lennart S.↗