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FUN3D Analyses of the 5th AIAA Propulsion Aerodynamics Workshop Inlet Test Case

This work presents the results of FUN3D analyses that were performed for the 5th AIAA Propulsion Aerodynamics Workshop. The workshop was separated into two sections that focused on inlet and nozzle flows. This paper focuses on the inlet section of the workshop, which considered an axisymmetric inlet in close-proximity to the ground and subjected to a crosswind. Four levels of grid refinement were provided by the workshop committee. Participants were asked to perform simulations for three crosswind velocities on a minimum of three levels of grid refinement, for a total of nine required cases. This work employed an updated set of grids provided by the committee after completion of the workshop based on feedback from participants. The updated grids featured a larger computational domain and increased refinement near the ground vortex region. This work employed the three coarsest grid levels from the updated grid set. Note that an additional extension downstream of the inlet was added to the provided grids for this work to mitigate observed convergence issues. Additionally, a limited turbulence model study was performed for the highest value of crosswind velocity and the finest grid. The results show that grid convergence was only achieved for the case with the lowest value of crosswind velocity. However, the results were generally observed to improve, relative to the experimental data, with increasing grid refinement. The FUN3D predictions were observed to agree qualitatively with the provided experimental data. However, FUN3D exhibited a tendency to underpredict the minimum value of the radially averaged total pressure at the Aerodynamic Interface Plane, which is consistent with the general findings from the workshop. Overall, FUN3D was found to perform well relative to its peers for the PAW5 workshop inlet test case.

PAI↗

NASA NTRS1998

A Summary of Validation Results for LEWICE 2.0

A research project is underway at NASA Lewis to produce a computer code which can accurately predict ice growth under any meteorological conditions for any aircraft surface. This report will present results from version 2.0 of this code, which is called LEWICE. This version differs from previous releases due to its robustness and its ability to reproduce results accurately for different point spacing, and time step criteria across general computing platforms. It also differs in the extensive amount of effort undertaken to compare the results in a quantifiable manner against the database of ice shapes which have been generated in the NASA Lewis Icing, Research Tunnel (IRT), The complete set of data used for this comparison is available in a recent contractor report . The result of this comparison shows that the difference between the predicted ice shape from LEWICE 2.0 and the average of the experimental data is 7.2% while the variability of the experimental data is 2.5%.

Wright, William B.↗

NASA NTRS1998

Estimation of Aircraft Nonlinear Unsteady Parameters From Wind Tunnel Data

Aerodynamic equations were formulated for an aircraft in one-degree-of-freedom large amplitude motion about each of its body axes. The model formulation based on indicial functions separated the resulting aerodynamic forces and moments into static terms, purely rotary terms and unsteady terms. Model identification from experimental data combined stepwise regression and maximum likelihood estimation in a two-stage optimization algorithm that can identify the unsteady term and rotary term if necessary. The identification scheme was applied to oscillatory data in two examples. The model identified from experimental data fit the data well, however, some parameters were estimated with limited accuracy. The resulting model was a good predictor for oscillatory and ramp input data.

Klein, Vladislav↗

NASA NTRS1996

Overview of HATP Experimental Aerodynamics Data for the Baseline F/A-18 Configuration

Determining the baseline aerodynamics of the F/A-18 was one of the major objectives of the High-Angle-of-Attack Technology Program (HATP). This paper will review the key data bases that have contributed to our knowledge of the baseline aerodynamics and the improvements in test techniques that have resulted from the experimental program. Photographs are given highlighting the forebody and leading-edge-extension (LEX) vortices. Other data representing the impact of Mach and Reynolds numbers on the forebody and LEX vortices will also be detailed. The level of agreement between different tunnels and between tunnels and flight will be illustrated using pressures, forces, and moments measured on a 0.06-scale model tested in the Langley 7- by 10-Foot High Speed Tunnel, a 0.16-scale model in the Langley 30- by 60-Foot Tunnel, a full-scale vehicle in the Ames 80- by 120-Foot Wind Tunnel, and the flight F/A-18 High Alpha Research Vehicle (HARV). Next, creative use of wind tunnel resources that accelerated the validation of the computational fluid dynamics (CFD) codes will be described. Lastly, lessons learned, deliverables, and program conclusions are presented.

Hall, Robert M.↗

NASA NTRS1991

Viscous three-dimensional calculations of transonic fan performance

A 3-D flow analysis code was used to compute the design speed operating line of a transonic fan rotor, and the results were compared with experimental data. The code is an explicit finite difference code with an algebraic turbulence model. The transonic fan, called rotor 67, was tested experimentally at NASA-Lewis with conventional aerodynamic probes and with user anemometry and was included as one of the AGARD test cases for the computation of internal flows. The experimental data are described. Maps of total pressure ratio and adiabatic efficiency versus mass flow were computed and are compared with the experimental maps, with good agreement. Detailed comparisons between calculations and experiment are made at two operating points, one near peak efficiency and the other near stall. Blade-to-blade contour plots are used to show the shock structure. Comparisons of circumferentially integrated flow quantities downstream of the rotor show spanwise distributions of several aerodynamic parameters. Calculated Mach number distributions are compared with laser anemometer data within the blade row and the wake to quantify the accuracy of the calculations. Particle traces are used to show the nature of secondary flow.

Chima, Rodrick V.↗

NASA NTRS1992

Viscous three-dimensional calculations of transonic fan performance

A 3-D flow analysis code was used to compute the design speed operating line of a transonic fan rotor, and the results were compared with experimental data. The code is an explicit finite difference code with an algebraic turbulence model. The transonic fan, called Rotor 67, was tested experimentally at NASA Lewis conventional aerodynamic probes and with user anemometry and was included as one of the AGARD test cases for the computation of internal flows. The experimental data are described. Maps of total pressure ratio and adiabatic efficiency vs mass flow were computed and are compared with the experimental maps, with good agreement. Detailed comparisons between calculations and experiment are made at two operating points, one near peak efficiency and the other near stall. Blade-to-blade contour plots are used to show the shock structure. Comparisons of circumferentially integrated flow quantities downstream of the rotor show spanwise distributions of several aerodynamic parameters. Calculated Mach number distributions are compared with laser anemometer data within the blade row and the wake to quantify the accuracy of the calculations. Particle traces are used to show the nature of secondary flow.

Chima, Rodrick V.↗

NASA NTRS1972

An experimental investigation of a cold jet emitting from a body of revolution into a subsonic free stream

An experimental program was undertaken to determine the pressure distribution induced on aerodynamic bodies by a subsonic cold jet exhausting normal to the body surface and into a subsonic free stream. The investigation was limited to two bodies with single exhaust jets a flat plate at zero angle of attack with respect to the free-stream flow and a cylinder, fitted with a conical nose, with the longitudinal axis alined with the free-stream flow. Experimental data were obtained for free-stream velocity to jet velocity ratios between 0.3 and 0.5. The experimental data are presented in tabular form with appropriate graphs to indicate pressure coefficient contours, pressure coefficient decay, pitching-moment characteristics, and lift characteristics.

Ousterhout, D. S.↗

NASA NTRS1985

A review of dynamic inflow and its effect on experimental correlations

A review is given of the relationship between experimental data and the development of modern dynamic-inflow theory. Some of the most interesting data, first presented 10 years ago at the Dynamic Specialist's Meeting, is now reviewed in light of the newer theories. These pure blade-flapping data correlate very well with analyses that include the new dynamic inflow theory, thus verifying the theory. Experimental data are also presented for damping with coupled inplane and body motions. Although inclusion of dynamic inflow is often required to correlate this coupled data, the data cannot be used to verify any particular dynamic inflow theory due to the uncertainties in modeling the inplane degree of freedom. For verification, pure flapping is required. However, the coupled data do show that inflow is often important in such computations.

Gaonkar, G. H.↗

NASA NTRS1986

Application of a computational model for vortex generators in subsonic internal flows

A model for the analysis of vortex generators in a fully viscous subsonic internal flow is evaluated. A vorticity source term is used in a modified form of the Parabolized Navier-Stokes equations to model the shed vortex. Computed results are compared with idealized flow vortex paths, and with experimental data for vortex generators embedded in a thick turbulent boundary layer. The analysis is also compared with experimental data for a separated diffusing S-duct and for a diffusing S-duct with vortex generators. Quantitative comparisons are shown for the latter three cases. Emphasis is placed on verifying the ability of the model to predict global distortions in the flow field.

Kunik, W. G.↗

NASA NTRS1986

Application of a computational model for vortex generators in subsonic internal flows

Kunik, W. G.↗

NASA NTRS1988

Numerical simulation of helicopter multi-bladed rotor flow

A numerical method for solving the Euler equations for multiblade rotors has been developed. The computed results, including one nonlifting forward-flight and several lifting hovering cases, are compared with experimental data. For forward-flight case, the unsteady growth and decay of the shock agree well with experimental results. For hovering case, at low and moderate transonic tip speed, the results show good agreement with experimental data in the tip region. The near-wake effects can be captured without wake modeling.

Chen, C. L.↗

NASA NTRS1993

A preferential vibration dissociation coupling model for nonequilibrium flowfields

A preferential vibration-dissociation coupling model is incorporated into a radiatively coupled viscous shock laver code that also includes chemical, radiative, and thermal nonequilibrium. Stagnation point flow profiles are obtained for various Fire 2 flight conditions and for a typical 14 km/sec AOTV case, and comparisons are made with Fire 2 experimental data. Adjustments in molecular absorption coefficients are also made for several diatomic species. Based on comparisons with experimental data, very little preferential dissociation behavior is present in the Fire 2 flight conditions.

Mcgough, David E.↗

NASA NTRS1994

Integrated active and passive control design methodology for the LaRC CSI evolutionary model

A general design methodology to integrate active control with passive damping was demonstrated on the NASA LaRC CSI Evolutionary Model (CEM), a ground testbed for future large, flexible spacecraft. Vibration suppression controllers designed for Line-of Sight (LOS) minimization were successfully implemented on the CEM. A frequency-shaped H2 methodology was developed, allowing the designer to specify the roll-off of the MIMO compensator. A closed loop bandwidth of 4 Hz, including the six rigid body modes and the first three dominant elastic modes of the CEM was achieved. Good agreement was demonstrated between experimental data and analytical predictions for the closed loop frequency response and random tests. Using the Modal Strain Energy (MSE) method, a passive damping treatment consisting of 60 viscoelastically damped struts was designed, fabricated and implemented on the CEM. Damping levels for the targeted modes were more than an order of magnitude larger than for the undamped structure. Using measured loss and stiffness data for the individual damped struts, analytical predictions of the damping levels were very close to the experimental values in the (1-10) Hz frequency range where the open loop model matched the experimental data. An integrated active/passive controller was successfully implemented on the CEM and was evaluated against an active-only controller. A two-fold increase in the effective control bandwidth and further reductions of 30 percent to 50 percent in the LOS RMS outputs were achieved compared to an active-only controller. Superior performance was also obtained compared to a High-Authority/Low-Authority (HAC/LAC) controller.

Voth, Christopher T.↗

NASA NTRS2003

Blade Row Interaction Effects on the Performance of a Moderately Loaded NASA Transonic Compressor Stage

Blade row interaction effects on loss generation in compressors have received increased attention as compressor work-per-stage and blade loading have increased. Two dimensional Laser Doppler Velocimeter measurements of the velocity field in a NASA transonic compressor stage show the magnitude of interactions in the velocity field at the peak efficiency and near stall operating conditions. The experimental data are presented along with an assessment of the velocity field interactions. In the present study the experimental data are used to confirm the fidelity of a three-dimensional, time-accurate, Navier Stokes calculation of the stage using the MSU-TURBO code at the peak efficiency and near stall operating conditions. The simulations are used to quantify the loss generation associated with interaction phenomena. At the design point the stator pressure field has minimal effect on the rotor performance. The rotor wakes do have an impact on loss production in the stator passage at both operating conditions. A method for determining the potential importance of blade row interactions on performance is presented.

VanZante, Dale E.↗

NASA NTRS2016

Experimental Evaluation of Inlet Distortion on an Ejector Powered Hybrid Wing Body at Take-off and Landing Conditions

As part of the NASA Environmentally Responsible Aircraft project, an ultra high bypass ratio engine integration on a hybrid wing body demonstration was planned. The goal was to include engine and airframe integration concepts that reduced fuel consumption by at least 50% while still reducing noise 42 db cumulative on the ground. Since the engines would be mounted on the upper surface of the aft body of the aircraft, the inlets may be susceptible to vortex ingestion from the wing leading edge at high angles of attack and sideslip, and separated wing/body flow. Consequently, experimental and computational studies were conducted to collect flow surveys useful for characterizing engine operability. The wind tunnel tests were conducted at two NASA facilities, the 14- by 22-foot at NASA Langley and the 40- by 80-foot at NASA Ames Research Center. The test results included in this paper show that the distortion and pressure recovery levels were acceptable for engine operability. The CFD studies conducted to compare to experimental data showed excellent agreement for the angle of attacks examined, although failed to match the low speed experimental data at high sideslip angles.

Carter, Melissa B.↗

NASA NTRS1995

Development of a two-equation turbulence model for hypersonic flows: Evaluation of a low Reynolds number correction to the Kappa - epsilon two equation compressible turbulence model - Volume 1

The objective of the current research is the development of an improved k-epsilon two-equation compressible turbulence model for turbulent boundary layer flows experiencing strong viscous-inviscid interactions. The development of an improved model is important in the design of hypersonic vehicles such as the National Aerospace Plane (NASP) and the High Speed Civil Transport (HSCT). Improvements have been made to the low Reynolds number functions in the eddy viscosity and dissipation of solenoidal dissipation of the k-epsilon turbulence mode. These corrections offer easily applicable modifications that may be utilized for more complex geometries. The low Reynolds number corrections are functions of the turbulent Reynolds number and are therefore independent of the coordinate system. The proposed model offers advantages over some current models which are based upon the physical distance from the wall, that modify the constants of the standard model, or that make more corrections than are necessary to the governing equations. The code has been developed to solve the Favre averaged, boundary layer equations for mass, momentum, energy, turbulence kinetic energy, and dissipation of solenoidal dissipation using Keller's box scheme and the Newton spatial marching method. The code has been validated by removing the turbulent terms and comparing the solution with the Blasius solution, and by comparing the turbulent solution with an existing k-epsilon model code using wall function boundary conditions. Excellent agreement is seen between the computed solution and the Blasius solution, and between the two codes. The model has been tested for both subsonic and supersonic flat-plate turbulent boundary layer flow by comparing the computed skin friction with the Van Driest II theory and the experimental data of Weighardt; by comparing the transformed velocity profile with the data of Weighardt, and the Law of the Wall and the Law of the Wake; and by comparing the computed results of an adverse pressure gradient with the experimental data of Fernando and Smits. Good agreement is obtained with the experimental correlations for all flow conditions.

Knight, Doyle D.↗

NASA NTRS2011

Waterhammer Transient Simulation and Model Anchoring for the Robotic Lunar Lander Propulsion System

Waterhammer transients have the potential to adversely impact propulsion system design if not properly addressed. Waterhammer can potentially lead to system plumbing, and component damage. Multi-thruster propulsion systems also develop constructive/destructive wave interference which becomes difficult to predict without detailed models. Therefore, it is important to sufficiently characterize propulsion system waterhammer in order to develop a robust design with minimal impact to other systems. A risk reduction activity was performed at Marshall Space Flight Center to develop a tool for estimating waterhammer through the use of anchored simulation for the Robotic Lunar Lander (RLL) propulsion system design. Testing was performed to simulate waterhammer surges due to rapid valve closure and consisted of twenty-two series of waterhammer tests, resulting in more than 300 valve actuations. These tests were performed using different valve actuation schemes and three system pressures. Data from the valve characterization tests were used to anchor the models that employed MSCSoftware.EASY5 v.2010 to model transient fluid phenomena by using transient forms of mass and energy conservation. The anchoring process was performed by comparing initial model results to experimental data and then iterating the model input to match the simulation results with the experimental data. The models provide good correlation with experimental results, supporting the use of EASY5 as a tool to model fluid transients and provide a baseline for future RLL system modeling. This paper addresses tasks performed during the waterhammer risk reduction activity for the RLL propulsion system. The problem of waterhammer simulation anchoring as applied to the RLL system is discussed with results from the corresponding experimental valve tests. Important factors for waterhammer mitigation are discussed along with potential design impacts to the RLL propulsion system.

Stein, William B.↗

NASA NTRS1991

On the quasi-conical flowfield structure of the swept shock wave-turbulent boundary layer interaction

The swept oblique shock-wave/turbulent-boundary-layer interaction generated by a 20-deg sharp fin at Mach 4 and Reynolds number 21,000 is investigated via a series of computations using both conical and three-dimensional Reynolds-averaged Navier-Stokes equations with turbulence incorporated through the algebraic turbulent eddy viscosity model of Baldwin-Lomax. Results are compared with known experimental data, and it is concluded that the computed three-dimensional flowfield is quasi-conical (in agreement with the experimental data), the computed three-dimensional and conical surface pressure and surface flow direction are in good agreement with the experiment, and the three-dimensional and conical flows significantly underpredict the peak experimental skin friction. It is pointed out that most of the features of the conical flowfield model in the experiment are observed in the conical computation which also describes the complete conical streamline pattern not included in the model of the experiment.

Knight, Doyle D.↗