Results for “Flow Conditions” — Page 27

NASA NTRS1997

The NASA Lewis Research Center Water Tunnel Facility

A water tunnel facility specifically designed to investigate internal fluid duct flows has been built at the NASA Research Center. It is built in a modular fashion so that a variety of internal flow test hardware can be installed in the facility with minimal facility reconfiguration. The facility and test hardware interfaces are discussed along with design constraints for future test hardware. The inlet chamber flow conditioning approach is also detailed. Instrumentation and data acquisition capabilities are discussed. The incoming flow quality has been documented for about one quarter of the current facility operating range. At that range, there is some scatter in the data in the turbulent boundary layer which approaches 10 percent of the duct radius leading to a uniform core.

Wasserbauer, Charles A.↗

NASA NTRS1994

Numerical Generation of Asymmetric Flows about Slender Bodies of Revolution at Extreme Incidence

It is well known that slender bodies of revolution will develop an asymmetric, unsteady flow pattern in experimental tests, if the angle of incidence to the oncoming stream is above a critical value. It has been suggested that the origin of these asymmetric flows may stem from geometric imperfections of the model being tested, or from disturbances in the oncoming stream. In numerical simulations, it is possible to generate bodies of revolution which are perfectly symmetric about their longitudinal axis, and to impose uniform flow conditions which are free from disturbances. The current work presents numerical simulations of the flow about an ogive-cylinder configuration at 40 and 60 degree angle of incidence. These simulations. were performed using numerical algorithms which are also symmetric about the lateral plane of the cylinder body. The flowfields at 40 degree angle of attack were seen to remain symmetric to the round-off accuracy of the computer. At 60 degree angle of attack, a lateral force coefficient developed of O(1) which progressed to an alternate vortex shedding in time. The nature of this lateral force generation and vortex shedding was dependent on the choice of numerical algorithm. The origin of the asymmetries observed in the 60 degree angle of attack computations were traced to round-off errors in the implicit block-matrix inverter. A means of inverting the implicit operator matrices, which maintains the symmetry of the overall numerical algorithm was implemented.

Murman, Scott M.↗

NASA NTRS2005

On the Application of Contour Bumps for Transonic Drag Reduction(Invited)

The effect of discrete contour bumps on reducing the transonic drag at off-design conditions on an airfoil have been examined. The research focused on fully-turbulent flow conditions, at a realistic flight chord Reynolds number of 30 million. State-of-the-art computational fluid dynamics methods were used to design a new baseline airfoil, and a family of fixed contour bumps. The new configurations were experimentally evaluated in the 0.3-m Transonic Cryogenic Tunnel at the NASA Langley Research center, which utilizes an adaptive wall test section to minimize wall interference. The computational study showed that transonic drag reduction, on the order of 12% - 15%, was possible using a surface contour bump to spread a normal shock wave. The computational study also indicated that the divergence drag Mach number was increased for the contour bump applications. Preliminary analysis of the experimental data showed a similar contour bump effect, but this data needed to be further analyzed for residual wall interference corrections.

Milholen, William E., II↗

NASA NTRS2005

On the Application of Contour Bumps for Transonic Drag Reduction(Invited)

The effect of discrete contour bumps on reducing the transonic drag at off-design conditions on an airfoil have been examined. The research focused on fully-turbulent flow conditions, at a realistic flight chord Reynolds number of 30 million. State-of-the-art computational fluid dynamics methods were used to design a new baseline airfoil, and a family of fixed contour bumps. The new configurations were experimentally evaluated in the 0.3-m Transonic Cryogenic Tunnel at the NASA Langley Research center, which utilizes an adaptive wall test section to minimize wall interference. The computational study showed that transonic drag reduction, on the order of 12% - 15%, was possible using a surface contour bump to spread a normal shock wave. The computational study also indicated that the divergence drag Mach number was increased for the contour bump applications. Preliminary analysis of the experimental data showed a similar contour bump effect, but this data needed to be further analyzed for residual wall interference corrections.

Milholen, William E., II↗

NASA NTRSDate not supplied

Simultaneous Ice Water Content Measurements at Multiple Locations on the NASA DC-8 Aircraft during the 2018 HIWC RADAR Flight Campaign

Ice water content measurements were made simultaneously at three locations on the NASA DC-8 in natural, glaciated conditions during the 2018 High Ice Water Content RADAR flight campaign. The purpose of these measurements was to further evaluate efficiency factors of hot-wire total water content probes in glaciated conditions and investigate the enhancement of ice crystal concentrations near fuselage surfaces due to flow field inertial effects, and ice crystals impacting the nose, breaking up, and flowing downstream. The total water content measurements were made using either Science Engineering Associates Ice Crystal Detectors or Robust Probes. Three common sensors were mounted on an underwing canister considered to be in near free-flow conditions, a standoff from a fuselage window, and the nose of the fuselage near the pitot probes. The Ice Crystal Detector total water content sensor and Robust Probe sensor collection and retention efficiencies were evaluated through comparisons with the underwing Ice Crystal Detector and Robust Probe measurements to the IKP2 isokinetic evaporator probe, which provided the reference ice water content measurement. Local ice water content at the nose position was evaluated by comparing ratios of the nose and underwing ice crystal detectors to the IKP2. Local ice water content at the window-standoff location was also evaluated by comparing total water content sensor measurements from the window ice crystal detector to the measurements made with the underwing and nose Ice Crystal Detectors The key findings were: 1) the Ice Crystal Detector total water content sensor ice water content efficiency factor was similar to previous estimates, but reduced with increased ice crystal median mass diameter; 2) the ice water content at the fuselage nose location near the DC-8 pitot probes was approximately 2.5 times the freestream values – although this estimate is affected by a higher probe efficiency factor due to smaller particles in the debris cloud from impacts upstream of the probe; and 3) the ice water content at the 17” standoff from the port window varied from about 50% to 3 times freestream values in a complicated manner. Similar measurement locations are not uncommon on cloud research aircraft, where ice particle measurements may be subject to similar uncertainties.

Aircraft Icing↗

NASA NTRSDate not supplied

Simultaneous Ice Water Content Measurements at Multiple Locations on the NASA DC-8 Aircraft during the 2018 HIWC RADAR Flight Campaign

Ice water content measurements were made simultaneously at three locations on the NASA DC-8 in natural, glaciated conditions during the 2018 High Ice Water Content RADAR flight campaign. The purpose of these measurements was to further evaluate efficiency factors of hot-wire total water content probes in glaciated conditions and investigate the enhancement of ice crystal concentrations near fuselage surfaces due to flow field inertial effects, and ice crystals impacting the nose, breaking up, and flowing downstream. The total water content measurements were made using either Science Engineering Associates Ice Crystal Detectors or Robust Probes. Three common sensors were mounted on an underwing canister considered to be in near free-flow conditions, a standoff from a fuselage window, and the nose of the fuselage near the pitot probes. The Ice Crystal Detector total water content sensor and Robust Probe sensor collection and retention efficiencies were evaluated through comparisons with the underwing Ice Crystal Detector and Robust Probe measurements to the IKP2 isokinetic evaporator probe, which provided the reference ice water content measurement. Local ice water content at the nose position was evaluated by comparing ratios of the nose and underwing ice crystal detectors to the IKP2. Local ice water content at the window-standoff location was also evaluated by comparing total water content sensor measurements from the window ice crystal detector to the measurements made with the underwing and nose Ice Crystal Detectors The key findings were: 1) the Ice Crystal Detector total water content sensor ice water content efficiency factor was similar to previous estimates, but reduced with increased ice crystal median mass diameter; 2) the ice water content at the fuselage nose location near the DC-8 pitot probes was approximately 2.5 times the freestream values – although this estimate is affected by a higher probe efficiency factor due to smaller particles in the debris cloud from impacts upstream of the probe; and 3) the ice water content at the 17” standoff from the port window varied from about 50% to 3 times freestream values in a complicated manner. Similar measurement locations are not uncommon on cloud research aircraft, where ice particle measurements may be subject to similar uncertainties.

o Aircraft Icing↗

NASA NTRS2021

Simultaneous Ice Water Content Measurements at Multiple Locations on the NASA DC-8 Aircraft during the 2018 HIWC RADAR Flight Campaign

Ice water content measurements were made simultaneously at three locations on the NASA DC-8 in natural, glaciated conditions during the 2018 High Ice Water Content RADAR flight campaign. The purpose of these measurements was to further evaluate efficiency factors of hot-wire total water content probes in glaciated conditions and investigate the enhancement of ice crystal concentrations near fuselage surfaces due to flow field inertial effects, and ice crystals impacting the nose, breaking up, and flowing downstream. The total water content measurements were made using either Science Engineering Associates Ice Crystal Detectors or Robust Probes. Three common sensors were mounted on an underwing canister considered to be in near free-flow conditions, a standoff from a fuselage window, and the nose of the fuselage near the pitot probes. The Ice Crystal Detector concave total water content sensor and Robust Probe sensor collection and retention efficiencies were evaluated through comparisons with the underwing Ice Crystal Detector and Robust Probe measurements to the Isokinetic Probe version 2 (IKP2), which provided the reference ice water content measurement. Local ice water content at the nose position was evaluated by comparing ratios of the nose and underwing ice crystal detectors to the IKP2. Local ice water content at the window-standoff location was also evaluated by comparing total water content sensor measurements from the window probes to the measurements made with the underwing and nose probes. The key findings were: (1) the Ice Crystal Detector concave water content sensor efficiency factor to glaciated conditions was similar to previous estimates, but reduced with increased ice crystal median mass diameter; (2) the ice water content at the fuselage nose location near the DC-8 pitot probes was approximately 2.5 times the freestream values—although this estimate is affected by a higher probe efficiency factor due to smaller particles in the debris cloud from impacts upstream of the probe; and (3) the ice water content at the 17 in. standoff from the port window varied from about 50 percent to nearly three times freestream values in a complicated manner. Similar measurement locations are not uncommon on cloud research aircraft, where ice particle measurements may be subject to similar uncertainties.

Aircraft Icing↗

NASA NTRSDate not supplied

Preliminary Design of a Distributed Facesheet Acoustic Liner for Broadband Acoustic Attenuation

The purpose of this study is to investigate the acoustic performance of a liner with a distributed facesheet and a uniform depth core for broadband attenuation. The distributed facesheet is comprised of a cluster of three distinct cell resonators of varying hole diameter and porosity replicated over the active liner treatment area. A target frequency range of 1000 to 2000 Hz and an attenuation metric of 10 dB are chosen. An optimizer is used to determine the optimal facesheet designs for flow conditions of Mach 0.0 and Mach 0.3. The samples are tested in the Grazing Flow Impedance Tube at NASA Langley Research Center. The two liners are shown to achieve at least 10 dB attenuation over frequency ranges of 700 and 400 Hz, respectively, when tested at their respective designed flow speeds. This study demonstrate that a distributed facesheet with a uniform depth core can be successfully used to achieve broadband sound absorption.

acoustic liners↗

NASA NTRSDate not supplied

Evaluation of Wall-Modeled Les for Flow Over A Multi-Element Airfoil

We perform wall-modeled large-eddy simulations (WMLES) of turbulent flow over a 30P30N multielement airfoil at various angles of attack to assess its ability to accurately predict lift using an unstructured node-based finite-volume methodology. The flow conditions are based on the experiments reported in Klausmeyer & Lin [NASA /TM-112858]. The oncoming flow Mach number is 0.2, and the stowed chord-based Reynolds number (𝑅𝑒𝑐) is 9 million. The angle of attack (𝛼) is varied between 8◦ and 23◦ with the expected stall angle of around 21◦. The simulated geometry is periodic in the span, thus any three-dimensional effects present in the experiment are ignored. We first analyze the predictions with baseline isotropic and anisotropic grids containing over 16 grid points per boundary-layer thickness. The near-wall streamwise/nominal spacing ratio is 0.4 at the leading/trailing edges of the slat, main element and flap, and 1 elsewhere, and the first wall-normal/nominal spacing ratio is 0.375 over most of the airfoil for the anisotropic grid, while these are unity for the isotropic grid. It is found that the predicted lift and wall pressures, near-wall velocity profiles, and the location of onset of resolved turbulence are sensitive to the near-wall grid anisotropy. Further sensitivities were explored at lower angles of attack by independently varying the streamwise and wall-normal grid anisotropies. The effects of both off-wall and near-wall grid refinement were also quantified. For the grid resolutions used here, our results suggest that the first wall-normal spacing, which coincides with the WMLES exchange location, has a dominant effect on the predictions for this flow configuration.

Computational Fluid Dynamics↗

NASA NTRS1992

Practical design optimization of wing/body configurations using the Euler equations

The development of a practical method for the aerodynamic design of isolated wing and wing/body configurations is achieved through the coupling of existing computational fluid dynamics (CFD) analysis codes and a quasi-Newton numerical optimization method. The direct design method is generalized to treat three-dimensional aerodynamic optimization problems subject to inviscid, rotational, compressible flow conditions imposed by the Euler equations. The method couples either the FLO57 or the TEAM flow solver with a modified version of the QNMDIF numerical optimization algorithm. The method is applied, but is not limited, to supersonic design problems. A case study is presented illustrating the method's effectiveness in maximizing the lift-to-drag ratio, subject to a variety of constraints, of selected supersonic configurations at cruise conditions.

Reuther, J.↗

NASA NTRS2019

Laser Rayleigh Scattering for Flow Density and Condensation Measurements in the Supercooled Mach-14 Free-Stream at AEDC Wind Tunnel No. 9

Laser Rayleigh scattering was used to investigate clusters in the free-stream flow at Arnold Engineering Development Center’s Tunnel 9 (T9). The facility was run at Mach-14, with a pure-N2 flow medium, and at several total pressures and temperatures. Using an excimer laser operating at 248 nm, the Rayleigh instrument imaged scattering from the focused laser beam in the free-stream. As a wind-tunnel flow is accelerated, it cools and approaches the condensation boundary. As a precursor to condensation, small clusters of molecules are first formed, but the individual clusters are too small to be spatially resolved in typical images of the beam. Thus clusters effectively add a spatially smooth background signal to the pure diatomic-molecule Rayleigh signal. The main result of the present work is that clustering was not significant. After correcting for interference by small particles imbedded in the T9 flow, cluster scattering was unobservable or smaller than one standard deviation (1-sigma) of the uncertainties for almost all tunnel runs. The total light scattering level was measured to be 1.05 +/- 0.15 (1-sigma) of the expected diatomic scattering, when averaged over the entire usable data set. This result included flow conditions that were supercooled to temperatures of ~ 20 K, about 25 K below the condensation limit of ~ 45 K. Thus the Mach-14 nozzle flow is essentially cluster-free for many supercooled conditions that might be used to extend the facility operating range to larger Reynolds numbers.

Herring, Gregory C.↗

NASA NTRS1993

Near wake structure for a generic ASTV configuration

Results of a numerical study are presented for hypersonic low-density flow about a 70-deg blunt cone using the direct simulation Monte Carlo method. Particular emphasis is given to the near wake flow and its sensitivity to rarefaction and other parametric variations. The flow conditions simulated are attainable in existing low-density hypersonic wind tunnels; that is, Mach 20 nitrogen flow encompassing freestream Knudsen numbers of 0.03 to 0.001. A stable vortex forms in the near wake at and below a freestream Knudsen number of 0.01 and the size of the vortex increases with decreasing freestream Knudsen number. The base region of the flow remains in thermal nonequilibrium for all cases. There is no formation of a lip separation shock or a distinct wake shock at these rarefied conditions.

Dogra, Virendra K.↗

NASA NTRS1972

Performance of tandem-bladed transonic compressor rotor with tip speed of 1375 feet per second

The design and experimental performance of a 20-inch diameter tandem-bladed axial-flow transonic compressor rotor is presented. Radial surveys were made of the flow conditions. At design speed the peak efficiency was 0.88 and occurred at an equivalent weight flow of 63 pounds per second. At peak efficiency the total pressure and total temperature ratios were 1.77 and 1.20, respectively. The stall margin at design speed was 10 percent based on weight flows and total pressure ratios at peak efficiency and near stall.

Urasek, D. C.↗

NASA NTRS1981

Investigation of the effects of inlet shapes on fan noise radiation

The effect of inlet shape on forward radiated fan tone noise directivities was investigated under experimentally simplified zero flow conditions. Simulated fan tone noise was radiated to the far field through various shaped zero flow inlets. Baseline data were collected for the simplest baffled and unbaffled straight pipe inlets. These data compared well with prediction. The more general inlet shapes tested were the conical, circular, and exponential surfaces of revolution and an asymmetric inlet achieved by cutting a straight pipe inlet at an acute angle. Approximate theories were developed for these general shapes and some comparisons with data are presented. The conical and exponential shapes produced directivities that differed considerably from the baseline data while the circular shape produced directivities similar to the baseline data. The asymmetric inlet produced asymmetric directivities with significant reductions over the straight pipe data for some angles.

Clark, T. L.↗

NASA NTRS1982

The modeling and prediction of multiple jet VTOL aircraft flow fields in ground effect

An engineering methodology based on an empirical data base and analytical fluid dynamic models was developed for the prediction of propulsive lift system induced aerodynamic effects for multiple lift jet VTOL aircraft operating in the hover mode in and out of ground effect. The effects of aircraft geometry, aircraft orientation (pitch, roll) as well as height above ground are considered. Lift jet vector and splay directions fit the airframe, lift jet exit flow conditions, and both axisymmetric and rectangular nozzle exit geometry are also accommodated. The induced suckdown flows are computed from the potential flowfield induced by the turbulent entrainment of both the free jets and wall jets in ground effect and from the free jets alone out of ground effect. The methodology emphasized geometric considerations, computation of stagnation lines and fountain upwash inclination, fountain upwash formation and development, and fountain impingement on the airframe.

Kotansky, D. R.↗

NASA NTRS1982

Measurement and evaluation of swirl-type flow in labyrinth seals of conventional turbine stages

The effects of load factor and rotor eccentricity were determined on flow conditions in test series for two stages. The results indicate that swirl-type entry flow follows the rules of potential swirl. Within the labyrinth cavities two spatial separated flow areas are considered. A dominating flow in periphera direction nearly fills the space between the sealing strips and the ceiling of the cavity. Below this flow, an area of axial mass transport is situated, with a slight peripheral component, limited on the nearest surroundings of the seals gap and the rotor surface. Between both flows, an exchange of energy takes place. Within the gaps, flow direction depends on axial velocity and therefore on variable flow contraction. A balance of energy within the seal and the cavities interprets the results of lateral force measurements as an influence of friction at the sealing strips surface and the rotating shaft surface. Stages with their blades put together in buckets by means of shrouding segments are particularly influenced by the rotating speed of the shrouding.

Hauck, L.↗

NASA NTRS1984

Nonlinear aerodynamic effects on bodies in supersonic flow

The supersonic flow about generic bodies was analyzed to identify the elments of the nonlinear flow and to determine the influence of geometry and flow conditions on the magnitude of these nonlinearities. The nonlinear effects were attributed to separated-flow nonlinearities and attached-flow nonlinearities. The nonlinear attached-flow contribution was further broken down into large-disturbance effects and entropy effects. Conical, attached-flow bundaries were developed to illustrate the flow regimes where the nonlinear effects are significant, and the use of these boundaries for angle of attack and three-dimensional geometries was indicated. Normal-force and pressure comparisons showed that the large-disturbance and separated-flow effects were the dominant nonlinear effects at low supersonic Mach numbers and that the entropy effects were dominant for high supersonic Mach number flow. The magnitude of all the nonlinear effects increased with increasing angle of attack. A full-potential method, NCOREL, which includes an approximate entropy correction, was shown to provide accurate attached-flow pressure estimates from Mach 1.6 through 4.6.

Pittman, J. L.↗

NASA NTRS1985

Conical similarity of shock/boundary-layer interactions generated by swept and unswept fins

A parametric experimental investigation has been made of the class of three-dimensional shock wave/turbulent boundary layer interactions generated by swept and unswept leading-edge fins. The fin sweepback angles were 0-65 deg at 5, 9, and 15 deg angles of attack. Two equilibrium two-dimensional turbulent boundary layers with a freestream Mach number of 2.95 and a Reynolds number of 6.3 x 10 to the 7th/m were used as incoming flow conditions. All of the resulting interactions were found to possess conical symmetry of the surface flow patterns and pressures outside of an initial inception zone. Further, these interactions were found to obey a simple conical similarity rule based on inviscid shock wave strength, irrespective of fin sweepback or angle of attack. This is one of the first demonstrations of similarity among three-dimensional interactions produced by geometrically dissimilar shock generators.

Settles, G. S.↗