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NASA NTRS1991

Navier-Stokes calculation of transonic flow past the NTF 65 deg delta wing

Models of four delta wings were built and tested in the 8 x 8 ft Transonic Wind Tunnel at LaRC. The wings are identical in planform shape with a swept-back angle of 65 degrees, but bear different leading edge profiles. The models were tested under pressurized and cryogenic conditions to simulate true flight Reynolds numbers. Data on the aerodynamic forces and pressure distributions at various locations on the wings were taken at various flight Mach numbers and angles of attack. Effects of high Reynolds number and leading edge radius on the aerodynamic characteristics of the wings are being accessed. To thoroughly understand the turbulent, vortical flow around the wings, an effort to perform computational aerodynamic analysis is being made. The objective of the analysis is to supplement and validate the experimental data and explain the high Reynolds number and leading edge effects. GRIDGEN, a software developed by General Dynamics, is being used to generate the grid topology for the flow field around the wings. The flow solver to be used is CFL3D, a computation fluid dynamics code developed at LaRC. Based on the geometric description of the wings, a FORTRAN program called WINGSURF was written to generate the databases defining the surface geometry of the wings. To match the true geometry of the models in the wind tunnel for realistic comparison with experimental results, databases defining the sting support for the wing models were also created by two other FORTRAN programs, STINJOIN and STINREAR. Listing of the programs are attached and the geometry of a typical wing model with the sting support is shown. Other aspects of the investigation are discussed.

Wu, Chivey C.↗

DOE OSTI2021

Systematic investigations on iron cycling in phosphorus/siderophore systems: Synergism or antagonism?

Synergisms between microbial exudates on Fe (hydr)oxide dissolution as an effective Fe acquisition pathway have been recently addressed and vigorously debated. However, Fe liberation mechanisms and where siderophores and phosphorus (P) coexist received little attentions. Current study systematically investigated ferrihydrite dissolution in the presence of desferrioxamine B (DFOB) (a kind of fungally-derived siderophores) and inorganic/organic phosphorus (orthophosphate, Pi; myo-inositol hexaphosphate, IHP), as a function of solute pH, reaction time and reagent content. Reacted solids were characterized by N 2 -BET adsorption, zeta (ζ) potential analysis, sequential extraction analysis (SEDEX), field emission scanning electron microscopy (FESEM), high-resolution transmission electron microscopy (HRTEM), X-ray diffraction (XRD), X-ray photoelectron spectroscopy (XPS), micro Raman spectroscopy and attenuated total reflectance-Fourier transform infrared spectroscopy (ATR-FTIR). Our results indicate that upon reaction with P-only or (DFOB + P) systems, interfacial complexation partially switched from monolayer bidentate-binuclear surface complexes to ternary complexes, or laterally transformed into amorphous Fe–P precipitates. The Fe-Pi complex precipitated more readily under acidic conditions, and Fe-IHP complex preferentially nucleated in neutral-alkaline environments. Phosphorus slightly promoted Fe release from minerals and fixation to the leached layer or interfacial liquid zone initially, but subsequently prevented further attacks from protons and DFOB. The co-effects of P and DFOB likely correspond to two successive scenarios: 1) DFOB is preferentially attracted toward ferrihydrite surfaces by negative electrical fields induced by adsorbed phosphorus and can act synergistically with labile P–Fe complexes, resulting in intensive temporal dissolution of Fh; 2) subsequent Fe shuttling to DFOB can be prohibited by stabilized, passive P/Fe–P layers. Finally, our results emphasize the antagonism between P compounds and siderophores (i.e., DFOB here) on ferrihydrite dissolution to improve understanding of the biologically-mediated Fe cycling in natural systems.

37 INORGANIC, ORGANIC, PHYSICAL, AND ANALYTICAL CH↗

NASA NTRSDate not supplied

High Reynolds Number Testing of the NATO AVT-298 SWiFT Configuration at the National Transonic Facility

A high Reynolds number wind tunnel test of the NATO AVT-298 Swept Wing Flow Test (SWiFT) configuration was conducted in the National Transonic Facility at the NASA Langley Research Center during the summer of 2023. The SWiFT research model geometry has relevance to both blended/hybrid wing body (BWB/HWB) and unmanned combat aerial vehicle (UCAV) configurations, and the test campaign was the culmination of an international collaboration under the NATO AVT-298 research task group. The main objectives of the test were to investigate Reynolds number scaling effects on low-speed stability & control characteristics and to examine the onset and progression of flow separation on the wings particularly near the wing crank. Force & moment and surface pressure data were acquired at Mach numbers from 0.2 to 0.8, Reynolds numbers from 2.5 to 34 million, angles of attack from -3 to 20 degrees, and sideslip angles from -10 to 10 degrees. Boundary layer transition detection techniques utilizing static pressure taps, unsteady pressure transducers, and sublimating chemicals were used on the model in a free/natural transition state or a forced transition state using trip dots. Pressure sensitive paint was used to obtain a global surface pressure profile on the model and an advanced laser velocimetry technique was used to obtain velocity measurements in the wake downstream of the wing crank. The results from the test showed clear Reynolds number scaling effects on the maximum lift coefficient and pitching moment coefficient at Mach 0.2, while also capturing significant hysteresis effects. The data acquired from the test will ultimately help improve computational aerodynamic analysis and design tools for application to future BWB-type vehicle configurations.

NATO AVT-298↗

NASA NTRSDate not supplied

High Reynolds Number Testing of the NATO AVT-298 SWiFT Configuration at the National Transonic Facility

BWB/HWB↗

NASA NTRS1980

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

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

Arnaiz, H. H.↗

NASA NTRS1986

Method for the prediction of the installation aerodynamics of a propfan at subsonic speeds: User manual

This document is the user's guide for the method developed earlier for predicting the slipstream wing interaction at subsonic speeds. The analysis involves a subsonic panel code (HESS code) modified to handle the propeller onset flow. The propfan slipstream effects are superimposed on the normal flow boundary condition and are applied over the surface washed by the slipstream. The effects of the propeller slipstream are to increase the axial induced velocity, tangential velocity, and a total pressure rise in the wake of the propeller. Principles based on blade performance theory, momentum theory, and vortex theory were used to evaluate the slipstream effects. The code can be applied to any arbitrary three dimensional geometry, expressed in the form of HESS input format. The code can handle a propeller alone configuration or a propeller/nacelle/airframe configuration, operating up to high subcritical Mach numbers over a range of angles of attack. Inclusion of a viscous modelling is briefly outlined. Wind tunnel results/theory comparisons are included as examples for the application of the code to a generic supercritical wing/overwing Nacelle with a powered propfan. A sample input/output listing is provided.

Chandrasekaran, B.↗

NASA NTRS1977

Numerical solution of the viscous hypersonic flow past blunted cones at angle of attack

Hypersonic viscous flow over spherically blunted cones of large half angle is computed at small angles of attack in the plane of symmetry of the flow field. Time-dependent viscous shock-layer equations in body-oriented coordinates are used to describe the flow field. The shock wave is treated as a discontinuity, across which the Rankine-Hugoniot relations are used to compute the flow conditions behind the shock. A time-marching second-order finite-difference method is used to solve the equations for a perfect gas. The local CFL (Courant-Friedrich-Lewy) time increment is used to advance the solution in time at each grid point. A fourth-order damping is used to damp the oscillations in the flow quantities. The numerical results of the present analysis for quantities such as shock standoff distance, surface-pressure distribution, and heating rates compare well with the existing theoretical and experimental results.

Kumar, A.↗

NASA NTRS1993

Ice Accretion Prediction for a Typical Commercial Transport Aircraft

Ice accretion calculations were made for a modern commercial transport using the NASA Lewis LEWICE3D ice accretion code. The ice accretion calculations were made for the wing and horizontal tail using both isolated flow models and flow models incorporating the entire airplane. The isolated flow model calculations were made to assess the validity of using these simplified models in lieu of the entire model in the ice accretion analysis of full aircraft. Ice shapes typifying a rime and a mixed ice shape were generated for a 30 minute hold condition. In general, the calculated ice shapes looked reasonable and appeared representative of a rime and a mixed ice conditions. The isolated flow model simplification was good for the main wing except at the root where it overpredicted the amount of accreted ice relative to the full aircraft flow model. For the horizontal tail the size and amount of predicted ice compared well for the two flow models, but the position of the accretions were more towards the upper surface for the aircraft flow model relative to the isolated flow model. This was attributed to downwash from the main wing which resulted in a lower effective angle of attack for the aircraft tail.

Bidwell, C. S.↗

NASA NTRS1993

Ice accretion prediction for a typical commercial transport aircraft

Ice accretion calculations were made for a modern commercial transport using the NASA Lewis LEWICE3D ice accretion code. The ice accretion calculations were made for the wing and horizonal tail using both isolated flow models and flow models incorporating the entire airplane. The isolated flow model calculations were made to assess the validity of using these simplified models in lieu of the entire model in the ice accretion analysis for full aircraft. Ice shapes typifying a rime and a mixed ice shape were generated for a 30 minute hold condition. In general, the calculated ice shapes looked reasonable and appeared representative of a rime and a mixed ice conditions. The isolated flow model simplification was good for the main wing except at the root where it overpredicted the amount of accreted ice relative to the full aircraft flow model. For the horizontal tail the size and amount of predicted ice compared well for the two flow models, but the position of the accretions were more towards the upper surface for the aircraft flow model relative to the isolated flow model. This was attributed to downwash from the main wing which resulted in a lower effective angle of attack for the aircraft tail.

Bidwell, C. S.↗

NASA NTRS2011

C-Band Airport Surface Communications System Engineering-Initial High-Level Safety Risk Assessment and Mitigation

This document is being provided as part of ITT's NASA Glenn Research Center Aerospace Communication Systems Technical Support (ACSTS) contract: "New ATM Requirements--Future Communications, C-Band and L-Band Communications Standard Development." ITT has completed a safety hazard analysis providing a preliminary safety assessment for the proposed C-band (5091- to 5150-MHz) airport surface communication system. The assessment was performed following the guidelines outlined in the Federal Aviation Administration Safety Risk Management Guidance for System Acquisitions document. The safety analysis did not identify any hazards with an unacceptable risk, though a number of hazards with a medium risk were documented. This effort represents an initial high-level safety hazard analysis and notes the triggers for risk reassessment. A detailed safety hazards analysis is recommended as a follow-on activity to assess particular components of the C-band communication system after the profile is finalized and system rollout timing is determined. A security risk assessment has been performed by NASA as a parallel activity. While safety analysis is concerned with a prevention of accidental errors and failures, the security threat analysis focuses on deliberate attacks. Both processes identify the events that affect operation of the system; and from a safety perspective the security threats may present safety risks.

Zelkin, Natalie↗

NASA NTRS2023

HLPW-4: Wall-Modeled LES and Lattice-Boltzmann Technology Focus Group Workshop Summary

A summary of the nine submissions to the Wall-Modeled LES and Lattice-Boltzmann(WMLESLB) Technical Focus Group (TFG) at the 4th High lift Prediction Workshop is provided. The focus of this TFG was to assess the current capabilities of WMLES and LB methods on a complex high-lift configuration across a wide range of angles-of-attack. Analysis of the submitted data suggests that>250𝑀spatial degrees of freedom are needed to accurately predict pitching moments at high angles-of-attack due to large pressure gradients present on the outboard slat and main element for 𝛼 >17◦(corrected for free-air). While some scatter is reported in pitching moment coefficient at the low-angles of attack (𝛼 <11◦) - excellent agreement is observed between submissions near the 𝐶𝐿,max state. Objective superiority of WMLES methods over RANS can be seen in terms of lack of excess outboard separation; a majority of the WMLES and LB submissions predict wedge-shaped separation patterns consistent with the experimental oil flow. The in-tunnel simulations show excellent agreement with the experiment in terms of a) integrated loads, b) surface flow-topology, and c) mechanism for the onset of inboard stall. Further evidence is provided to demonstrate both qualitative and quantitative superiority of the WMLES submissions over RANS.

TTT↗

NASA NTRS1993

Analysis of a high speed civil transport configuration at subsonic flow conditions using a Navier-Stokes solver

Computations of three dimensional vortical flows over a generic High Speed Civil Transport (HSCT) configuration with an aspect ratio of 3.04 are performed using a thin-layer Navier-Stokes solver. The HSCT cruise configuration is modeled without leading or trailing edge flap deflections and without engine nacelles. The flow conditions, which correspond to tests done in the NASA Langley 8-Foot Transonic Pressure Tunnel (TPT), are a subsonic Mach number of 0.3 and Reynolds number of 4.4 million for a range-of-attack (-.23 deg to 17.78 deg). The effects of the farfield boundary location with respect to the body are investigated. The boundary layer is assumed turbulent and simulated using an algebraic turbulence model. The key features of the vortices and their interactions are captured. Grid distribution in the vortex regions is critical for predicting the correct induced lift. Computed forces and surface pressures compare reasonably well with the experimental TPT data.

Lessard, Victor R.↗

NASA NTRS1973

An improved method for the aerodynamic analysis of wing-body-tail configurations in subsonic and supersonic flow. Part 1: Theory and application

A new method has been developed for calculating the pressure distribution and aerodynamic characteristics of wing-body-tail combinations in subsonic and supersonic potential flow. A computer program has been developed to perform the numerical calculations. The configuration surface is subdivided into a large number of panels, each of which contains an aerodynamic singularity distribution. A constant source distribution is used on the body panels, and a vortex distribution having a linear variation in the streamwise direction is used on the wing and tail panels. The normal components of velocity induced at specified control points by each singularity distribution are calculated and make up the coefficients of a system of linear equations relating the strengths of the singularities to the magnitude of the normal velocities. The singularity strengths which satisfy the boundary condition of tangential flow at the control points for a given Mach number and angle of attack are determined by solving this system of equations using an interactive procedure. Once the singularity strengths are known, the pressure coefficients are calculated, and the forces and moments acting on the configuration determined by numerical integration.

Woodward, F. A.↗

NASA NTRS1985

Unsteady analysis of rotor blade tip flow

The development of the VSAERO-TS and VSAERO-H computer programs for calculating the unsteady aerodynamic characteristics of arbitrarily shaped wings oscillating in pitch is presented. The effect of several wake parameters on chordwise pressure distribution in VSAERO-TS is given and the convergence characteristics of both programs are discussed. In the program, the influence coefficient for each panel is formulated for a planar surface and so a skewed panel is represented by a projected flat quadrilateral lying in the mean plane. Since panels in the extreme roll-up region of the tip vortex are highly skewed the program was modified to treat each highly skewed panel as a pair of triangles. The programs are validated by comparing the chordwise pressure distribution of several blade tip planforms with experimental data. The comparison, for the most part, is good. The triangular panel representation improved the chordwise pressure distribution near the tip region for higher mean angle of attack.

Maskey, B.↗

NASA NTRS1985

Wide-field shadowgraph flow visualization of tip vortices generated by a helicopter rotor

The vortex trajectory and vortex wake generated by helicopter rotors were visualized using a wide-field shadowgraph technique. Use of a retro-reflective Scotchlite screen made it possible to investigate the flow field generated by full-scale tail rotors. Tip vortex trajectories were visible in shadowgraphs for a range for tip Mach number of 0.38 - 0.60. The effect of the angle of attack was substantial. At an angle of attack greater than 8 degrees, the visibility of the vortex core was significant even at relatively low tip Mach numbers. The theoretical analysis of the visibility was carried out for a rotating blade. This analysis demonstrated that the visibility decreases with increasing dimensionless core radius (ro/c) and increases with increasing tip Mach number. Based on this investigation, it is concluded that the wide-field shadowgraph flow visualization technique should be feasible to study the flow field generated by a large main rotor in a wind tunnel and in an outdoor full-scale test stand. Of note is that the shadowgraph technique could easily be used with other on-going tests such as aerodynamic performance study, noise measurements, velocity measurements with LDV or hot-wire anemometer, local pressure measurement on the rotor surface, etc.

Parthasarathy, S. P.↗

NASA NTRS1991

Applications of hot-film anemometers in hypersonic shear layers

A wind tunnel test was conducted on a flat plate at zero angle of attack with a rearward facing 2D cooling film injector nozzle. The freestream Mach number was 8 and the injector Mach number was 3. The freestream Reynolds number varied from 0.43 to 3.3 million per ft during the test, and the injector flow rate was such that the jet exit and freestream static pressures were matched. The analysis reported herein will focus on data obtained at a freestream Reynolds number of 0.85 million per ft. The data consists of heat-transfer measurements obtained upstream and downstream of the injector nozzle and flowfield surveys obtained downstream of the injector nozzle with a pitot, total temperature, hot-film anemometer and hot-wire anemometer probes. The flowfield surveys were made at stations 0.1 to 9 in. downstream of the injector nozzle from near the model surface to approximately 2 in above the model surface. The hot-film anemometer was used to define the fluctuations in the shear layer separating the flows. The hot-film results are integrated with conventional measurement techniques to obtain a more complete description of the complicated shear layer separating hypersonic and supersonic flows.

Grubb, J. P.↗

NASA NTRS2020

Wind Tunnel Investigation of the Supersonic Stage Separation Aerodynamics of a Generic 0.0175-Scale Bimese Two-Stage-to-Orbit Reusable Launch Vehicle Configuration

A wind tunnel investigation was conducted of the supersonic stage separation aerodynamics of a generic two-stage-to-orbit bimese wingbody configuration in the NASA Langley Research Center Unitary Plan Wind Tunnel. Proximity and isolated model testing was conducted at Mach numbers of 2.3, 3.0, and 4.5 and a unit Reynolds number of 2.0 million per foot using 0.0175-scale models of the Langley Glide-Back Booster concept designated as the orbiter and booster in belly-to-belly and back-to-belly configurations. Longitudinal forces and moments were obtained on both models and surface static pressure measurements were obtained on the orbiter model at 328 relative proximity locations and at relative angles of attack of 0 degrees and 5 degrees. The test results supported a larger effort to develop and validate experimental and computational tools applicable to the design and simulation of stage separation and abort procedures for reusable launch vehicles composed of multiple bodies, including winged bodies. An initial proof-of-concept experiment featuring low-cost uninstrumented models was conducted to verify an emerging automated model control system and new support system hardware, and to identify potential model and support system blockage and unsteady aerodynamics/model dynamics prior to committing to higher-fidelity instrumented models. This investigation led to upgrades in the facility stage separation hardware, calibration and testing techniques and capabilities, and data analysis and documentation methodologies that have been extended to the more recent NASA Constellation and Space Launch System crew and cargo launch vehicle programs. A virtual diagnostics interface methodology was used to facilitate the design of the stage separation support hardware, to position the models in the test section, and to define the experimental test space. Advances in the facility automated model positioning system established a foundation for the development of a continuous-sweep data acquisition technique that is responsible for significant productivity improvements to the current NASA Space Launch System test program. The automated model positioning capability was leveraged to conduct a companion statistically-designed stage separation experiment requiring randomization of the relative proximity positions of the orbiter and booster models. The respective zones of influence and interference effects of the orbiter and booster were identified from three-dimensional scatter plots, contour and influence maps, and two-dimensional plotting methods. The highly-nonlinear, shock-dominated aerodynamic characteristics of the orbiter and booster in the Unitary Plan Wind Tunnel exhibited good agreement with independent test data obtained in a NASA Marshall Space Flight Center wind tunnel and with computational fluid dynamics predictions using a compressible, three-dimensional flow solver and an inviscid, unstructured Cartesian method.

Erickson, Gary E.↗

NASA NTRS1999

Low Speed Analysis of Mission Adaptive Flaps on a High Speed Civil Transport Configuration

Thin-layer Navier-Stokes analyses were done on a high speed civil transport configuration with mission adaptive leading-edge flaps. The flow conditions simulated were Mach = 0.22 and Reynolds number of 4.27 million for angles-of-attack ranging from 0 to 18 degrees. Two turbulence closure models were used. Analyses were done exclusively with the Baldwin-Lomax turbulence model at low angle-of-attack conditions. At high angles-of-attack where considerable flow separation and vortices occurred the Spalart-Allmaras turbulence model was also considered. The effects of flow transition were studied. Predicted aerodynamic forces, moment, and pressure are compared to experimental data obtained in the 14- by 22-Foot Subsonic Tunnel at NASA Langley. The forces and moments correlated well with experimental data in terms of trends. Drag and pitching moment were consistently underpredicted. Predicted surface pressures compared well with experiment at low angles-of-attack. Above 10 angle-of-attack the pressure comparisons were not as favorable. The two turbulent models affected the pressures on the flap considerably and neither produced correct results at the high angles-of-attack.

Lessard, Victor R.↗