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In the event of a launch vehicle failure during ascent, a manned space launch system requires an ascent launch abort system that will be able to separate the crew module from the launch vehicle and return the crew safely to earth. One measure of the effectiveness and reliability of the launch abort system is its ability to reduce the risk of loss of crew from the failure environments, such as blast overpressure and debris strikes from an exploding launch vehicle, resulting from the launch vehicle failure scenario. Physics-based models are used to assess the evolution of, and risks presented by, the failure environments. In the case of debris strikes, the probability of the crew module being hit by at least one piece of debris is computed by modeling the evolution of the debris field over time and determining its relative position to the crew module. The characteristics of the debris field, including the number of pieces, the mass and reference area of each piece, the imparted velocity magnitude and direction and the ballistic coefficient, are defined by a debris catalog. A model has been created to generate a debris catalog using a combination of empirical- and physics-based models. The debris catalog model accounts for design features of the launch vehicle and the failure mechanisms involved in determining the number of pieces and imparted velocity. The model results are compared with a published catalog for the Space Shuttle external tank. The sensitivity of the risk prediction to the number of pieces and the imparted velocities are studied. The debris catalog generation model provides an additional tool in the risk assessment of ascent aborts for manned launch systems.

The generation of significant side forces and yawing moments on an F/A-18 fuselage through tangential slot blowing is analyzed using computational fluid dynamics. The effects of freestream Mach number, jet exit conditions, jet length, and jet location are studied. The effects of over- and underblowing on force and moment production are analyzed. Non-time-accurate solutions are obtained to determine the steady-state side forces, yawing moments, and surface pressure distributions generated by tangential slot blowing. Time-accurate solutions are obtained to study the force onset time lag of tangential slot blowing. Comparison with available experimental data from full-scale wind-tunnel and subscale wind-tunnel tests are made. This computational analysis complements the experimental results and provides a detailed understanding of the effects of tangential slot blowing on the flowfield about the isolated F/A-18 forebody. Additionally, it extends the slot-blowing database to transonic maneuvering Mach numbers.

The effect of forebody tangential slot blowing on the flowfield about an F/A-18 aircraft is investigated numerically using solutions of the Navier-Stokes equations. Computed solutions are obtained for an aircraft geometry which includes the fuselage, a wing with deflected leading-edge flap, empennage, and a faired-over engine inlet. The computational slot geometry corresponds to that used in full-scale wind-tunnel tests. Solutions are computed using flight test conditions and jet mass flow ratios equivalent to wind-tunnel test conditions. The effect of slot location is analyzed by computing two nontime-accurate solutions with a 16-in. slot located 3 in. and 11 in. aft of the nose of the aircraft. These computations resolve the trends observed in the full-scale wind-tunnel test data. The flow aft of the leading-edge extension vortex burst is unsteady. A time-accurate solution is obtained to investigate the flow characteristics aft of the vortex burst, including the effect of blowing on tail buffet.

The effectiveness of a tangential slot blowing concept for generating lateral control forces on an aircraft forebody is analyzed using computational fluid dynamics. The flow about a fighter forebody is computed using a multiple-zone, thin-layer Navier-Stokes code. Tangential slot blowing is modeled by the use of an actuator plane. The effects of slot location and slot length on the efficiency of the system are analyzed. Results of the study indicate that placement of the slot near the nose of the aircraft greatly enhances the efficiency of the system, while the length and circumferential location of the slot are of secondary importance. Efficiency is defined by the amount of side force or yawing moment obtained per unit blowing coefficient. The effect of sideslip on the system is also analyzed. The system is able to generate incremental changes in forces and moments in flows with sideslip angles up to 10 deg comparable to those obtained at zero sideslip. These results are used to determine a baseline configuration for an experimental study of the tangential slot blowing concept.

A framework for representing intermediate failure states in an assessment of abort risk during the ascent phase of a crewed space launch vehicle mission is presented. The framework refines a previously established, simulation-based risk assessment approach (Ref. 1) by improving the characterization of vehicle failure prior to the onset of final, potentially catastrophic “loss of crew” (LOC) hazards through the introduction of system-level “loss of mission” (LOM) failure states, or “LOM environments.” The intermediate failure state framework is found to improve the risk analysis with respect to both risk model fidelity and risk model management. While LOM environments primarily provide an incremental point of reference for failure evolution modeling and refinement, they also serve to consolidate the risk analysis in late failure evolution and thereby reduce overall analysis effort. Ultimately, the logical boundary in failure event sequences formed by LOM environments is found to effectively delineate areas of failure analysis responsibility between the teams that provide inputs to the risk assessment, and furthermore, define the data interface between those teams. The merits of this framework are illustrated in a case study concerning the treatment of upper stage liquid fuel engine failures.

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.

In the event of a space launch vehicle explosion during ascent, the debris field generated by the explosion poses a risk to the crew. To evaluate this risk, a model of the debris environment was created and used to determine the probability of a debris strike on the crew module. The model uses experimental data to determine the initial debris field due to a launch vehicle explosion and computes the trajectory of each piece of debris. The trajectory of the crew module after the abort is also computed. The relative position of the debris field and the crew module is determined as a function of time after abort and explosion. A debris flux about the crew module is computed based on this information. The debris flux is used to compute the probability of a debris strike on the crew module using the Poisson distribution. The effect of system and model parameters - such as warning time, the number of debris pieces and abort system thrust- on the debris strike probability is assessed.

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.

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.

‡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.

The Tool Agent Framework—Java (TAF-J) software was developed to provide a set of services to access, manage, and integrate aerospace analysis tools with the Information Power Grid (IPG), a distributed set of heterogeneous compute resources. Analysis tools in several disciplines were incorporated into the framework. The framework also included a number of helper utilities created to link the various tools together to perform multidisciplinary analyses. Analysis and tool selection, job creation, submission and monitoring, solution monitoring, and data storage and retrieval were managed using prototype software agents. These agents were designed to perform their functions with minimal user input. The framework was used to compute ascent abort trajectories for a reusable launch vehicle design. Several design of experiments (DOE) methods within the framework were tested for computational speed and accuracy. The data generated were then used in the trajectory analysis. The trajectory analysis agent automatically selected an appropriate emergency landing site based upon the abort parameters.

Methods to combine a number of analysis tools into an agent-based integrated design system (IDS) for aerospace vehicles were developed. Computer-aided Design (CAD) and Finite Element Method (FEM) thermal and structural analysis packages were integrated with low-fidelity aero- and thermodynamic methods and high-fidelity CFD flow solvers. Analyses were characterized by the generation of an FEM unstructured grid from the CAD data, generation of a structured grid from the FEM grid, computation of the aerodynamic solution, interpolation of the flow data onto the FEM grid, and solution of the thermal and/or structural response of the geometry to the aerodynamic loading. A level of autonomy was provided by an expert system that managed the input data, selected the appropriate flow solver, and built the required input files. The ability to dynamically link available tools together to solve a given problem was also included. The framework integrated commercial tools to enable the use of distributed, heterogeneous computing systems. The framework was used to analyze the effect of wedge angle and nose radius on the heating characteristics of the nose region of a reusable launch vehicle.

A numerical study is performed to investigate the effects of tangential slot blowing on a generic chined forebody. The Reynolds-averaged, thin-layer, Navier-Stokes equations are solved to obtain the high-angle-of-attack viscous flow field about a generic chined forebody. Tangential slot blowing is investigated as a means of forebody flow control to generate side force and yawing moment on the forebody. The effects of jet mass flow ratios, angle of attack, and blowing slot location in the axial and circumferential directions are studied. The computed results are compared with available wind tunnel experimental data. The solutions with and without blowing are also analyzed using helicity density contours, surface flow patterns, and off-surface instantaneous streamlines. The results of this analysis provide details of the flow field about the generic chined forebody, as well as show that tangential slot blowing can be used as a means of forebody flow control to generate side force and yawing moment.

The effect of forebody tangential slot blowing on the flowfield about an F/A-18 aircraft is investigated numerically using solutions of the Navier-Stokes equations. Computed solutions are obtained for a full aircraft geometry, including the fuselage, wing with deflected leading-edge flap, empennage, and a faired-over engine inlet. The computational slot geometry corresponds to that used in full-scale wind tunnel tests. Solutions are computed using flight test conditions and jet mass flow ratios equivalent to wind tunnel test conditions. The effect of slot location is analyzed by computing two non-time-accurate solutions with a 16 in. slot located 3 in. and 11 in. aft of the nose of the aircraft. These computations resolve the trends observed in the full-scale wind tunnel test data. The flow aft of the leading edge extension (LEX) vortex burst is unsteady. A time-accurate solution is obtained to investigate the flow characteristics aft of the vortex burst, including the effect of blowing on tail buffet.

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.

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.

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 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.

The F3D thin-layer Navier-Stokes code presently used to numerically investigate the three-dimensional separated flow about a prolate spheroid at high incidence analyzes the effect of different turbulence models on the flowfield solution and the characteristics of the predicted flow. The Johnson-King (1984) model is applied in order to evaluate the importance of modeling nonequilibrium effects in predicting flow about a slender body at high incidence; the computations in question are for steady-state, fully turbulent flow. Insight is gained into the effects of turbulence models on flow characteristics, and model effects on the accurate prediction of highly separated and vortical flows about a slender body are demonstrated.