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Alaa Elmiligui

Publications and source records attributed to Alaa Elmiligui.

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2023

Evaluation of Limiter Functions for Supersonic Applications

Limiters commonly used in the simulations of flows with discontinuities are compared with the new limiter function proposed by Nishikawa using idealized test cases in two dimensions as well as complex three-dimensional problems. The Nishikawa limiter is observed to be consistently the least dissipative in idealized test cases as well as complex practical problems, for both steady and time-dependent problems. In the case of steady simulations, its iterative convergence characteristics are either similar or better than other limiter functions. The nearfield sonic boom signature of a low-boom demonstrator is computed to demonstrate the utility of the Nishikawa limiter function for realistic supersonic aircraft configurations.

Computational fluid dynamics↗
NASA NTRS2022

USM3D Simulations for the Third AIAA Sonic Boom Prediction Workshop

The NASA USM3D flow solver was used to compute test cases for the Third AIAA Sonic Boom Prediction Workshop (SBPW3). The test cases included a nearfield biconvex shock-plume interaction wind tunnel model and the C608 Low Boom Flight Demonstrator. Numerical simulations were conducted on the mixed-element and tetrahedral grids provided by the workshop committee, as well as a family of grids generated by an in-house approach known as BoomGrid. The nearfield pressure signatures were extracted and propagated to the ground, and the perceived loudness levels on the ground were computed. The USM3D nearfield pressure signatures, corresponding ground signatures, and loudness levels on the ground were compared with that of mean values from other workshop participants. The effect of three flux-limiters on the accuracy of nearfield pressure signature prediction was investigated and results were compared with that of mean values from other workshop participants. The effect of using wall-function models for predicting nearfield pressure signatures was also evaluated. Results showed that the overpressure signatures extracted from the wall-modeled simulations and the wall-resolved simulations are in good agreement. The use of wall functions (wall-modeled simulations) allowed for approximately 23% savings in the time to solution and yielded comparable perceived loudness levels on the ground to the wall-resolved Reynolds-averaged Navier Stokes simulations.

Supersonics↗
NASA NTRS2020

Artificial Thickening of a Transonic Boundary Layer in the Presence of a Pressure Gradient Associated with a Boundary Layer Ingestion Concept

Boundary layer ingestion is an aeropropulsive concept associated with a propulsion airframe integration technique that integrates viscous aerodynamics into a propulsion system to achieve more efficient flight. The effectiveness of this concept is heavily dependent on how much of the boundary layer is being ingested into the propulsor. Scaling this concept for transonic wind tunnel testing is often plagued with blockage issues and requires a modification to the model dimensions such as fuselage length and diameter and wingspan. To simulate the boundary layer for these modified models requires a manipulation of the boundary layer height to achieve the appropriate ratio of boundary layer to propulsor inlet diameter or height. This paper will focus on 15 different transonic boundary layer manipulators to achieve varying turbulent boundary layer heights for a representative single-aisle transport utilizing Boundary Layer Ingestion.

Gregory S Jones↗
NASA NTRS2007

PAB3D Simulations for the CAWAPI F-16XL

Numerical simulations of the flow around F-16XL are performed as a contribution to the Cranked Arrow Wing Aerodynamic Project International (CAWAPI) using the PAB3D CFD code. Two turbulence models are used in the calculations: a standard k-! model, and the Shih-Zhu-Lumley (SZL) algebraic stress model. Seven flight conditions are simulated for the flow around the F-16XL where the free stream Mach number varies from 0.242 to 0.97. The range of angles of attack varies from 0° to 20°. Computational results, surface static pressure, boundary layer velocity profiles, and skin friction are presented and compared with flight data. Numerical results are generally in good agreement with flight data, considering that only one grid resolution is utilized for the different flight conditions simulated in this study. The ASM results are closer to the flight data than the k-ε model results. The ASM predicted a stronger primary vortex, however, the origin of the vortex and footprint is approximately the same as in the k-ε predictions.

Flight Data↗
NASA NTRSDate not supplied

USM3D Simulations for Third Sonic Boom Workshop

The NASA USM3D flow solver was used to compute test cases for the Third AIAA Sonic Boom Prediction Workshop (SBPW3). The test cases include an axisymmetric equivalent area body, a near field biconvex shock-plume interaction wind tunnel model, and the C608 Low Boom Flight Demonstrator. Numerical simulations were conducted on the mixed element grids and the tetrahedral grids provided by the workshop committee, as well as a family of grids generated by an in-house approach for sonic boom analyses known as Boom Grid. The near-field pressure signatures were extracted, propagated to the ground and the perceived loudness levels on the ground was computed. The USM3D near-field pressure signatures,corresponding ground signatures, and loudness levels on the ground are compared with mean values from other workshop participants.

Supersonics↗
NASA NTRSDate not supplied

USM3D Simulations for Third Sonic Boom Workshop

The NASA USM3D flow solver was used to compute test cases for the Third AIAA Sonic Boom Prediction Workshop (SBPW3). The test cases include a nearfield biconvex shock-plume interaction wind tunnel model, and the C608 Low Boom Flight Demonstrator. Numerical simulations were conducted on the mixed-element and tetrahedral grids provided by the workshop committee, as well as a family of grids generated by an in-house approach for sonic boom analyses known as BoomGrid. The near-field pressure signatures were extracted, propagated to the ground and the perceived loudness levels on the ground were computed. The USM3D near-field pressure signatures, corresponding ground signatures, and loudness levels on the ground are compared with that of mean values from other workshop participants. The effect of three flux-limiters on the accuracy of near-field pressure signatures prediction was investigated and results are compared with that of mean values from other workshop participants. The effect of using wall function grids in accurately predicting near-field pressure signatures was also evaluated.

Supersonics↗
NASA NTRSDate not supplied

USM3D Simulations for the Third Sonic Boom Prediction Workshop (SBPW-3)

The NASA USM3D flow solver was used to compute test cases for the Third AIAA Sonic Boom Prediction Workshop (SBPW3). The test cases include a nearfield biconvex shock-plume interaction wind tunnel model, and the C608 Low Boom Flight Demonstrator. Numerical simulations were conducted on the mixed-element and tetrahedral grids provided by the workshop committee, as well as a family of grids generated by an in-house approach for sonic boom analyses known as BoomGrid. The near-field pressure signatures were extracted, propagated to the ground and the perceived loudness levels on the ground were computed. The USM3D near-field pressure signatures, corresponding ground signatures, and loudness levels on the ground are compared with that of mean values from other workshop participants. The effect of three flux-limiters on the accuracy of near-field pressure signatures prediction was investigated and results are compared with that of mean values from other workshop participants. The effect of using wall function grids in accurately predicting near-field pressure signatures was also evaluated.

Supersonics↗
NASA NTRSDate not supplied

Numerical Simulation of the High-Speed Leg of the National Transonic Facility

Numerical simulations for the flow inside the high-speed leg of the National Transonic Facility was conducted. The NASA Tetrahedral Unstructured Software System (TetrUSS) with its USM3D_ME solver was used to perform the numerical simulations. USM3D_ME is developed and maintained by NASA Langley Research Center. Simulations were conducted for three configurations: empty tunnel, body of revolution installed, and NASA Common Research Model installed in the test section. Simulations were performed for a test section Mach number of 0.7 and 0.85 and a corresponding Reynolds number of 8 million per ft. A controller was developed that automated dynamic outflow boundary and streamlined the process of running multiple simulations. The use of a dynamic outflow boundary was the key parameter to drive simulation to the desired tunnel conditions. The numerical simulations captured expected flow features in NTF test section and in the tunnel plenum. The simulations depicted that separation of the flow inside the diffuser appears to be asymmetric and more extensive toward the top and bottom walls. Analysis of the flow field were conducted, and the computed drag coefficient compared to wind tunnel data.

NTF↗
NASA NTRSDate not supplied

Development of Neural Network for X-59 Air Data Probe Calibration

Two wind tunnel tests were conducted in the NASA Glenn Research Center (GRC) 8- by 6-Foot Supersonic Wind Tunnel to calibrate the nose-mounted air data probe for the X-59. The probe will be the primary instrument for the flight test of the X-59 for determining angle of attack, angle of sideslip, airspeed, and pressure altitude of the aircraft in flight. This paper documents the development of several neural net algorithms that can be used to relate air data probe port pressures from those tests to relevant X-59 flight parameters, such as angle of attack, angle of sideslip, mach number, and static and dynamic pressure. These calibration algorithms address accuracy effects for significant real-world conditions, such as initial sensor measurement uncertainties, assumed parameter constraints on NN modelling errors, and malfunctioning or plugged ports. Resultant integrated uncertainties are provided for each algorithm; these uncertainties are then used to estimate the effects on relevant X-59 mission parameters.

Kurtis Roy Long↗
NASA NTRSDate not supplied

X-59 Air Data Probe Calibration Wind Tunnel Test

NASA's aeronautical innovators are working with Lockheed Martin to design and build a supersonic X-plane, the X-59, under the the Low Boom Flight Demonstrator (LBFD) Project, that produces a gentle thump rather than a loud sonic boom. The plane will be used to collect human response data over select U.S. communities. The data from the X-59 test flights will be provided to U.S. and international regulators to potentially allow supersonic flight over land, drastically reducing travel time within the U.S. and around the world. Two wind tunnel tests were conducted in the NASA Glenn Research Center (GRC) 8- by 6-Foot Supersonic Wind Tunnel to calibrate the nose probe for the X-59. The probe was successfully tested and calibrated at 19 Mach numbers between Mach 0.25 and Mach 1.7. Most data were collected using continuous roll angle sweeps at a set Mach number and pitch angle. About 725 high quality data runs were recorded during testing to be used in the calibration of the probe. A backup nose probe was also run through the same test matrix as the primary probe. The probe will be the primary instrument for the flight test of the X-59 for determining angle of attack, angle of sideslip, airspeed, and pressure altitude of the aircraft in flight. Knowing the aircraft's speed and attitude is critical for flight safety and is also critical data for the Sonic Boom mission. Areas of interest for the test were sea level takeoff and landing at Mach 0.2, subsonic cruise at Mach 0.9, and supersonic cruise at Mach 1.4. During Phase I flight testing, the air data nose probe and air data system as a whole will be further calibrated in flight to account for additional installed effects of the airframe on the air data system.

X-59↗
NASA NTRSDate not supplied

Evaluation of Limiter Functions for Supersonic Applications

Limiters commonly used in the simulations of flows with discontinuities are compared with the new limiter function proposed by Nishikawa using idealized test cases in two dimensions as well as complex three-dimensional problems. The Nishikawa limiter is observed to be consistently the least dissipative in idealized test cases as well as complex practical problems, for both steady and time-dependent problems. In the case of steady simulations, its iterative convergence characteristics are either similar or better than other limiter functions. The nearfield sonic boom signature of a low-boom demonstrator is computed to demonstrate the utility of the Nishikawa limiter function for realistic supersonic aircraft configurations.

Computational fluid dynamics↗
NASA NTRSDate not supplied

X-59 Air Data Probe Calibration Wind Tunnel Test

NASA's aeronautical innovators are working with Lockheed Martin to design and build a supersonic X-plane, the X-59, under the the Low Boom Flight Demonstrator (LBFD) Project, that produces a gentle thump rather than a loud sonic boom. The plane will be used to collect human response data over select U.S. communities. The data from the X-59 test flights will be provided to U.S. and international regulators to potentially allow supersonic flight over land, drastically reducing travel time within the U.S. and around the world. Two wind tunnel tests were conducted in the NASA Glenn Research Center (GRC) 8- by 6-Foot Supersonic Wind Tunnel to calibrate the nose probe for the X-59. The probe was successfully tested and calibrated at 19 Mach numbers between Mach 0.25 and Mach 1.7. Most data were collected using continuous roll angle sweeps at a set Mach number and pitch angle. About 725 high quality data runs were recorded during testing to be used in the calibration of the probe. A backup nose probe was also run through the same test matrix as the primary probe. The probe will be the primary instrument for the flight test of the X-59 for determining angle of attack, angle of sideslip, airspeed, and pressure altitude of the aircraft in flight. Knowing the aircraft's speed and attitude is critical for flight safety and is also critical data for the Sonic Boom mission. Areas of interest for the test were sea level takeoff and landing at Mach 0.2, subsonic cruise at Mach 0.9, and supersonic cruise at Mach 1.4. During Phase I flight testing, the air data nose probe and air data system as a whole will be further calibrated in flight to account for additional installed effects of the airframe on the air data system.

X-59↗
NASA NTRSDate not supplied

Measurement Accuracy and Uncertainty Analysis of the X-59 Air Data Probe Calibration Test Entry One

The X-59 is being developed to demonstrate quiet sonic boom technology. Data was obtained via two test entries in the NASA Glenn Research Center’s 8- by 6-Foot Supersonic Wind Tunnel to calibrate the X-59’s nose air data probe. The first entry tested two theoretically identical probes with the intention of one of the probes becoming the primary X-59 nose air data probe flight hardware and the other becoming a backup. A measurement accuracy and uncertainty analysis was performed on data obtained from the first test entry. The analysis showed that the uncertainties in the probe pressures were nearly identical for the two probes, with an average difference of 2.51 x 10-5 for the non-dimensional total pressure and 2.88 x 10-5 for the non-dimensional static pressures. The analysis also showed that the uncertainty in the probe yaw and pitch angles were 0.00034° and 0.01416°, respectively. This gives confidence in the X-59’s flight air data system.

Air Data Probe↗
NASA NTRSDate not supplied

USM3D-ME Solutions for RANS Test Suite of High-Fidelity CFD Verification Workshop 2024

USM3D-ME solutions for the High-Fidelity CFD Verification Workshop 2024 are presented. Solutions are computed for the Reynolds-Averaged Navier-Stokes equations using the Spalart-Allmaras one equation turbulence model, SA-neg-QCR2000-R, that is enhanced with a nonlinear correction to the turbulence stresses and a correction for flow rotation. USM3D-ME solutions have been computed for the three verification cases, namely, a Joukowski airfoil, a subsonic three-dimensional flow around an extruded NACA 0012 wing in tunnel, and subsonic flow around a high-lift wing-body configuration. Iterative convergence of USM3D-ME solutions on several grids within various grid families is demonstrated. Grid convergence of integrated forces, pitching moment and sectional variations of surface pressure and skin friction are also presented.

CFD↗