Ken Gee

Generation of Aerodynamic Data using a Design of Experiment and Data Fusion Approach

‡As one component of an expert system to generate aerodynamic data using Computational Fluid Dynamics (CFD) tools, a new approach utilizing Design of Experiment (DOE) and data fusion is outlined in the following paper. The goal of combining data fusion (merging of various fidelity solutions into a single, coherent database) with an adaptive DOE design is to improve the efficiency of the data generation process. A comparison between databases created using this novel approach and a more conventional full-factorial design shows that the new process can dramatically reduce the computational times required to generate data.

Unsteady Simulation of Viscous Flowfield Around F-18 Aircraft at Large Incidence

This article describes the numerical simulation of the unsteady viscous flow around the F-18 aircraft at high angles of attack. A generalized overset zonal grid scheme is used to decompose the computational space around the complete aircraft, included deflected control surfaces. The grids around various components of the aircraft are created numerically using a three-dimensional hyperbolic grid generation procedure. The Reynolds-averaged Navier-Stokes equations are integrated using a time-accurate, implicit procedure. Results for the turbulent flow around the F-18 aircraft at 30 deg angle of attack show the details of the flowfleld structure, including the unsteadiness created by the vortex burst and the resulting fluctuating airloads exerted on the vertical tail. The computed results agree fairly well with flight data for surface pressure, surface flow pattern, vortex burst location, and the dominant frequency for tail load fluctuations.

Numerical prediction of the unsteady flowfield around the F-18 aircraft at large incidence

This paper describes a numerical method capable of solving the steady and unsteady viscous flow around complete aircraft configurations at high angles of attack. This method is used to simulate the external flow around the F-18 aircraft, including deflected control surfaces. The current technique employs a generalized overset zonal grid scheme to decompose the computational space around the aircraft. The grid around various components of the aircraft are created numerically using a three-dimensional hyperbolic grid generation procedure. The Reynolds-averaged Navier-Stokes equations are integrated using a time-accurate, implicit procedure. Results for the turbulent flow around the F-18 aircraft at 30 degrees angle of attack show the details of the flowfield structure, including the unsteadiness created by the vortex burst and the resulting fluctuating airloads exerted on the vertical tail. The computed results agree fairly well with flight data for surface pressure, surface flow pattern, vortex burst location, and the dominant frequency for tail load fluctuations.

A multi-level parallelization concept for high-fidelity multi-block solvers

The integration of high-fidelity Computational Fluid Dynamics (CFD) analysis tools with the industrial design process benefits greatly from the robust implementations that are transportable across a wide range of computer architectures. In the present work, a hybrid domain-decomposition and parallelization concept was developed and implemented into the widely-used NASA multi-block Computational Fluid Dynamics (CFD) solvers employed in ENSAERO and OVERFLOW advanced flow analysis packages. These advanced engineering and scientific analysis packages include more than 300,000 lines of code written in FORTRAN 77 language in more than 1300 individual subprograms. The new parallel solver concept, PENS (Parallel Euler Navier-Stokes Solver), employs both fine and coarse granularity with data partitioning as well as data coalescing to obtain the desired load-balance characteristics on the available computer platforms for these legacy packages. This multi-level parallelism implementation itself introduces no changes to the numerical results, hence the original fidelity of the packages are identically preserved. The present implementation uses the Message Passing Interface (MPI) library for interprocessor message passing and memory accessing. By choosing an appropriate combination of the available partitioning and coalescing possibilities only during the execution stage, the PENS solver is used on different computer architectures from shared-memory to distributed-memory platforms with varying degrees of parallelism. Improvements in computational load-balance and speeds are extremely crucial on the realistic problems in the design of aerospace vehicles. The PENS implementation on the IBM SP2 distributed memory environment at the NASA Ames Research Center obtains 85 percent scalable parallel performance using fine-grain partitioning of single-block CFD domains using up to 128 wide computational nodes. Multi-block CFD simulations of complete aircraft geometries achieve 85 percent perfect load-balanced executions using data coalescing and the two levels of parallelism. SGI PowerChallenge, SGI Onyx2, and Cray T3E are the other platforms where the robustness, performance behavior, and the parallel scalability of the implementation are tested and fine-tuned for actual production run environments.

Numerical investigation of tail buffet on F-18 aircraft

Numerical investigation of vortex induced tail buffet is conducted on the F-18 aircraft at high angles of attack. The Reynolds-averaged Navier-Stokes equations are integrated using a time-accurate, implicit procedure. A generalized overset zonal grid scheme is used to decompose the computational space around the complete aircraft with faired-over inlet. A weak coupling between the aerodynamics and structures is assumed to compute the structural oscillation of the flexible vertical tail. Time-accurate computations of the turbulent flow around the F-18 aircraft at 30 degrees angle of attack show the surface and off-surface flowfield details, including the unsteadiness created by the vortex burst and its interaction with the vertical twin tail which causes the tail buffet. The effect of installing a LEX fence on modifying the vortex structure upstream of the tail is also examined.

Numerical simulation of the viscous flow around a simplified F/A-18 at high angles of attack

A numerical method developed for solving the viscous flow around three-dimensional complex configurations is presently used to simulate the flow around a simplified F/A-18 configuration encompassing forebody, wing, leading-edge extension, faired-over inlet, and deflected wing leading-edge flaps, at Mach 0.243 and 30.3 deg angle of attack. The computational results show the details of the flowfield structure, including primary, secondary, and tertiary separation lines, the development of forebody and leading-edge extension vortex, and the burst of this vortex. A grid-refinement study is conducted to assess the effect of grid characteristics on solution accuracy. Substantial agreement is obtained between these results and flight data.

Simulation assisted risk assessment applied to launch vehicle conceptual design

A simulation-based risk assessment approach is presented and is applied to the analysis of abort during the ascent phase of a space exploration mission. The approach utilizes groupings of launch vehicle failures, referred to as failure bins, which are mapped to corresponding failure environments. Physical models are used to characterize the failure environments in terms of the risk due to blast overpressure, resulting debris field, and the thermal radiation due to a fireball. The resulting risk to the crew is dynamically modeled by combining the likelihood of each failure, the severity of the failure environments as a function of initiator and time of the failure, the robustness of the crew module, and the warning time available due to early detection. The approach is shown to support the launch vehicle design process by characterizing the risk drivers and identifying regions where failure detection would significantly reduce the risk to the crew.

Analysis of a pneumatic forebody flow control concept about a full aircraft geometry

A full aircraft geometry is used to computationally analyze the effectiveness of a pneumatic forebody flow control concept. An overset grid technique is employed to model the aircraft and slot geometry. Steady-state solutions for both isolated forebody and full aircraft configurations are carried out using a thin-layer Navier-Stokes flow solver. A solution obtained using the full aircraft geometry and a flight sideslip condition investigates the effect of sideslip on the leading edge extention vortex burst point. A no-sideslip blowing solution using the isolated forebody at full-scale wind tunnel test conditions is compared with experimental data to determine the accuracy of the numerical method. A solution employing the full geometry and slot blowing at flight conditions is obtained.

Simulation-Assisted Risk Assessment

A probabilistic risk assessment (PRA) approach has been developed and applied to the risk analysis of capsule abort during ascent. The PRA is used to assist in the identification of modeling and simulation applications that can significantly impact the understanding of crew risk during this potentially dangerous maneuver. The PRA approach is also being used to identify the appropriate level of fidelity for the modeling of those critical failure modes. The Apollo launch escape system (LES) was chosen as a test problem for application of this approach. Failure modes that have been modeled and/or simulated to date include explosive overpressure-based failure, explosive fragment-based failure, land landing failures (range limits exceeded either near launch or Mode III trajectories ending on the African continent), capsule-booster re-contact during separation, and failure due to plume-induced instability. These failure modes have been investigated using analysis tools in a variety of technical disciplines at various levels of fidelity. The current paper focuses on the roles and impacts of the higher-fidelity methods on this process and, by association, the roles and impacts of the high performance computing resources of the Columbia supercomputer system at NASA Ames Research Center.

Computational investigation of tangential slot blowing on a generic chined forebody

The effect of tangential slot blowing on the flowfield about a generic chined forebody at high angles of attack is investigated numerically using solutions of the thin-layer, Reynolds-averaged, Navier-Stokes equations. The effects of jet mass now ratios, angle of attack, and blowing slot location in the axial and circumferential directions are studied. The computed results compare well with available wind-tunnel experimental data. Computational results show that for a given mass now rate, the yawing moments generated by slot blowing increase as the body angle of attack increases. It is observed that greater changes in the yawing moments are produced by a slot located closest to the lip of the nose. Also, computational solutions show that inboard blowing across the top surface is more effective at generating yawing moments than blowing outboard from the bottom surface.