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LaCETS baseline flight study include: 29 high-quality nearfield shock structure probings at three Mach numbers; Shocks in exhaust plume measured; ! CFD study of simplified nozzle shows similar plume structures as flight data; ! Phase II flights scheduled for October 2008; and ! US Industry and Academia invited to participate in analysis, review, and assessment of LaNCETS data.
A method of automated computational fluid dynamics (CFD) has been invented for the generation of performance tables for an object subject to fluid flow. The method is applicable to the generation of tables that summarize the effects of two-dimensional flows about airfoils and that are in a format known in the art as C81. (A C81 airfoil performance table is a text file that lists coefficients of lift, drag, and pitching moment of an airfoil as functions of angle of attack for a range of Mach numbers.) The method makes it possible to efficiently generate and tabulate data from simulations of flows for parameter values spanning all operational ranges of actual or potential interest. In so doing, the method also enables filling of gaps and resolution of inconsistencies in C81 tables generated previously from incomplete experimental data or from theoretical calculations that involved questionable assumptions.
A new framework is presented for shape optimization using analytical shape functions and high-fidelity computational fluid dynamics (CFD) via Cart3D. The focus of the paper is the system-level integration of several key enabling analysis tools and automation methods to perform shape optimization and reduce sonic boom footprint. A boom mitigation case study subject to performance, stability and geometrical requirements is presented to demonstrate a subset of the capabilities of the framework. Lastly, a design space exploration is carried out to assess the key parameters and constraints driving the design.
New ideas are forthcoming to break existing bottlenecks in using CFD during design. CAD-based automated grid generation. Multi-disciplinary use of embedded, overset grids to eliminate complex gridding problems. Use of time-averaged detached-eddy simulations as norm instead of "steady" RANS to include effects of self-excited unsteadiness. Combined GPU/Core parallel computing to provide over an order of magnitude increase in performance/price ratio. Gas-turbine applications are shown here but these ideas can be used for other Air Force, Navy, and NASA applications.
A novel Computational Fluid Dynamics (CFD) coupling framework using a conventional Reynolds-Averaged Navier-Stokes (BANS) solver to resolve the near-body flow field and a Particle-based Vorticity Transport Method (PVTM) to predict the evolution of the far field wake is developed, refined, and evaluated for fixed and rotary wing cases. For the rotary wing case, the RANS/PVTM modules are loosely coupled to a Computational Structural Dynamics (CSD) module that provides blade motion and vehicle trim information. The PVTM module is refined by the addition of vortex diffusion, stretching, and reorientation models as well as an efficient memory model. Results from the coupled framework are compared with several experimental data sets (a fixed-wing wind tunnel test and a rotary-wing hover test).
Several computational fluid dynamics (CFD) codes were used to analyze the Versatile Integrated Inlet Propulsion Aerodynamics Rig (VIIPAR) located at NASA Glenn Research Center. The rig consists of a serpentine inlet, a rake assembly, inlet guide vanes, a 12-in. diameter tip-turbine driven fan stage, exit rakes or probes, and an exhaust nozzle with a translating centerbody. The analyses were done to develop computational capabilities for modeling inlet/fan interaction and to help interpret experimental data. Three-dimensional Reynolds averaged Navier-Stokes (RANS) calculations of the fan stage were used to predict the operating line of the stage, the effects of leakage from the turbine stream, and the effects of inlet guide vane (IGV) setting angle. Coupled axisymmetric calculations of a bellmouth, fan, and nozzle were used to develop techniques for coupling codes together and to investigate possible effects of the nozzle on the fan. RANS calculations of the serpentine inlet were coupled to Euler calculations of the fan to investigate the complete inlet/fan system. Computed wall static pressures along the inlet centerline agreed reasonably well with experimental data but computed total pressures at the aerodynamic interface plane (AIP) showed significant differences from the data. Inlet distortion was shown to reduce the fan corrected flow and pressure ratio, and was not completely eliminated by passage through the fan
A combined computational and experimental effort has been undertaken to study fuselage drag reduction on a generic, non-proprietary rotorcraft fuselage by the application of active ow control. Fuselage drag reduction is an area of research interest to both the United States and France and this area is being worked collaboratively as a task under the United States/France Memorandum of Agreement on Helicopter Aeromechanics. In the first half of this task, emphasis is placed on the US generic fuselage, the ROBIN-mod7, with the experimental work being conducted on the US side and complementary US and French CFD analysis of the baseline and controlled cases. Fuselage simulations were made using Reynolds-averaged Navier-Stokes ow solvers and with multiple turbulence models. Comparisons were made to experimental data for numerical simulations of the isolated fuselage and for the fuselage as installed in the tunnel, which includes modeling of the tunnel contraction, walls, and support fairing. The numerical simulations show that comparisons to the experimental data are in good agreement when the tunnel and model support are included. The isolated fuselage simulations compare well to each other, however, there is a positive shift in the centerline pressure when compared to the experiment. The computed flow separation locations on the rear ramp region had only slight differences with and without the tunnel walls and model support. For the simulations, the flow control slots were placed at several locations around the flow separation lines as a series of eight slots that formed a nearly continuous U-shape. Results from the numerical simulations resulted in an estimated 35% fuselage drag reduction from a steady blowing flow control configuration and a 26% drag reduction for unsteady zero-net-mass flow control configuration. Simulations with steady blowing show a delayed flow separation at the rear ramp of the fuselage that increases the surface pressure acting on the ramp, thus decreasing the overall fuselage pressure drag.
The focus here is on the early development (mid 1970's-1980's) at NASA Ames Research Center of implicit methods in Computational Fluid Dynamics (CFD). A class of implicit finite difference schemes of the Beam and Warming approximate factorization type will be addressed. The emphasis will be on the Euler equations. A review of material pertinent to the solution of the Euler equations within the framework of implicit methods will be presented. The eigensystem of the equations will be used extensively in developing a framework for various methods applied to the Euler equations. The development and analysis of various aspects of this class of schemes will be given along with the motivations behind many of the choices. Various acceleration and efficiency modifications such as matrix reduction, diagonalization and flux split schemes will be presented.
This viewgraph presentation describes two nozzle fluid flow interactions. They include: 1) Cold flow nozzle tests with fluid-structure interaction at nozzle separated flow; and 2) CFD analysis for nozzle flow and side loads of nozzle extensions with various out-of-round cases.
The objectives of this presentation are: (1) Share lessons learned from a recent exercise in CFD validation of protuberance heating (2) Impact of experimental data reduction assumptions and techniques on validation activity (3) Advanced data reduction techniques may provide useful data from non-typical test methods (4) Significance of the recovery factor for high-speed flows (5) Show typical results of the Lag turbulence model on protuberances (6) Introduce and inform the listener of a protuberance heating dataset which will soon be available for comparison
Data from the recent UH-60A Airloads Test in the National Full-Scale Aerodynamics Complex 40- by 80- Foot Wind Tunnel at NASA Ames Research Center are presented and compared to predictions computed by a loosely coupled Computational Fluid Dynamics (CFD)/Comprehensive analysis. Primary calculations model the rotor in free-air, but initial calculations are presented including a model of the tunnel test section. The conditions studied include a speed sweep at constant lift up to an advance ratio of 0.4 and a thrust sweep at constant speed into deep stall. Predictions show reasonable agreement with measurement for integrated performance indicators such as power and propulsive but occasionally deviate significantly. Detailed analysis of sectional airloads reveals good correlation in overall trends for normal force and pitching moment but pitching moment mean often differs. Chord force is frequently plagued by mean shifts and an overprediction of drag on the advancing side. Locations of significant aerodynamic phenomena are predicted accurately although the magnitude of individual events is often missed.
A common open volume is created by the stacking of the Orion vehicle onto the Ares I Upper Stage. Called the Shared Volume, both vehicles contribute to its gas, fluid, and thermal environment. One of these environments is related to hazardous hydrogen gas. While both vehicles use inert purge gas to mitigate any hazardous gas buildup, there are concerns that hydrogen gas may still accumulate and that the Ares I Hazardous Gas Detection System will not be sufficient for monitoring the integrated volume. This Computational Fluid Dynamics (CFD) analysis has been performed to examine these topics. Results of the analysis conclude that the Ares I Hazardous Gas Detection System will be able to sample the vent effluent containing the highest hydrogen concentrations. A second conclusion is that hydrogen does not accumulate under the Orion Service Module (SM) avionics ring as diffusion and purge flow mixing sufficiently dilute the hydrogen to safe concentrations. Finally the hydrogen concentrations within the Orion SM engine nozzle may slightly exceed the 1 percent volume fraction when the entire worse case maximum full leak is directed vertically into the engine nozzle.
Demanding aerodynamic modeling requirements for military and civilian aircraft have provided impetus for researchers to improve computational and experimental techniques. Model validation is a key component for these research endeavors so this study is an initial effort to extend conventional time history comparisons by comparing model parameter estimates and their standard errors using system identification methods. An aerodynamic model of an aircraft performing one-degree-of-freedom roll oscillatory motion about its body axes is developed. The model includes linear aerodynamics and deficiency function parameters characterizing an unsteady effect. For estimation of unknown parameters two techniques, harmonic analysis and two-step linear regression, were applied to roll-oscillatory wind tunnel data and to computational fluid dynamics (CFD) simulated data. The model used for this study is a highly swept wing unmanned aerial combat vehicle. Differences in response prediction, parameters estimates, and standard errors are compared and discussed
The goal of the study is to assess the impacts of free water propagation in the Waste and Hygiene Compartment (WHC). Free water can be generated inside the WHC in small quantities due to crew hygiene activity. To mitigate potential impact of free water in Node 3 cabin the WHC doorway is enclosed by a waterproof bump-out, Kabin, with openings at the top and bottom. At the overhead side of the rack, there is a screen that prevents large drops of water from exiting. However, as the avionics fan in the WHC causes airflow toward the deck side of the rack, small quantities of free water may exit at the bottom of the Kabin. A Computational Fluid Dynamics (CFD) analysis of Node 3 cabin airflow made possible to identify the paths of water transport. The Node 3 airflow was computed for several ventilation scenarios. To simulate the droplet transport the Lagrangian discrete phase approach was used. Various initial droplet distributions were considered in the study. The droplet diameter was varied in the range of 2-20 mm. The results of the computations showed that most of the drops fall to the rack surface not far from the WHC curtain. The probability of the droplet transport to the adjacent rack surface with electronic equipment was predicted.
At NASA/MSFC, Structural Dynamics personnel continue to perform advanced analysis for the turbomachinery in the J2X Rocket Engine, which is under consideration for the new Space Launch System. One of the most challenging analyses in the program is predicting turbine blade structural capability. Resonance was predicted by modal analysis, so comprehensive forced response analyses using high fidelity cyclic symmetric finite element models were initiated as required. Analysis methodologies up to this point have assumed the flow field could be fully described by a sector, so the loading on every blade would be identical as it travelled through it. However, in the J2X the CFD flow field varied over the 360 deg of a revolution because of the flow speeds and tortuous axial path. MSFC therefore developed a complex procedure using Nastran Dmap's and Matlab scripts to apply this circumferentially varying loading onto the cyclically symmetric structural models to produce accurate dynamic stresses for every blade on the disk. This procedure is coupled with static, spin, and thermal loading to produce high cycle fatigue safety factors resulting in much more accurate analytical assessments of the blades.
This paper describes the use of targeted Loci/CHEM CFD simulations to evaluate the effects of a dual-engine first-stage hot-fire test on an evolving integrated launch pad/test article design. This effort was undertaken as a part of the NESC Independent Assessment of the Taurus II Stage Test Series. The underlying conceptual model included development of a series of computational models and simulations to analyze the plume induced environments on the pad, facility structures and test article. A pathfinder simulation was first developed, capable of providing quick-turn around evaluation of plume impingement pressures on the flame deflector. Results from this simulation were available in time to provide data for an ongoing structural assessment of the deflector. The resulting recommendation was available in a timely manner and was incorporated into construction schedule for the new launch stand under construction at Wallops Flight Facility. A series of Reynolds-Averaged Navier-Stokes (RANS) quasi-steady simulations representative of various key elements of the test profile was performed to identify potential concerns with the test configuration and test profile. As required, unsteady Hybrid-RANS/LES simulations were performed, to provide additional insight into critical aspects of the test sequence. Modifications to the test-specific hardware and facility structures thermal protection as well as modifications to the planned hot-fire test profile were implemented based on these simulation results.
In an attempt to explain the high loss of panels from the south face of the Vehicle Assembly Building (VAB) during Hurricane Frances, a three-dimensional computational fluid dynamics (3-D CFD) model was developed to simulate local velocity and pressure distributions resulting from such a storm. A preconditioned compressible Navier-Stokes flow solver 1 was used to compute the flow field around the VAB complex, including the Launch Control Center, the Low and High Bays of the VAB, and several outbuildings in the immediate LC-39 area. The mapping of the forces and velocities on and along the affected faces of the VAB correlated surprisingly well with the extensive damage areas realized on both on the south face and on the southeast section of the roof. The model results were also consistent with the minimal damage seen on the east, north, and west faces of the structure.
A solution has been developed to the challenges of computation of derivatives with respect to geometry, which is not straightforward because these are not typically direct inputs to the computational fluid dynamics (CFD) solver. To overcome these issues, a procedure has been devised that can be used without having access to the mesh generator, while still being applicable to all types of meshes. The basic approach is inspired by the mesh motion algorithms used to deform the interior mesh nodes in a smooth manner when the surface nodes, for example, are in a fluid structure interaction problem. The general idea is to model the mesh edges and nodes as constituting a spring-mass system. Changes to boundary node locations are propagated to interior nodes by allowing them to assume their new equilibrium positions, for instance, one where the forces on each node are in balance. The main advantage of the technique is that it is independent of the volumetric mesh generator, and can be applied to structured, unstructured, single- and multi-block meshes. It essentially reduces the problem down to defining the surface mesh node derivatives with respect to the geometry parameters of interest. For analytical geometries, this is quite straightforward. In the more general case, one would need to be able to interrogate the underlying parametric CAD (computer aided design) model and to evaluate the derivatives either analytically, or by a finite difference technique. Because the technique is based on a partial differential equation (PDE), it is applicable not only to forward mode problems (where derivatives of all the output quantities are computed with respect to a single input), but it could also be extended to the adjoint problem, either by using an analytical adjoint of the PDE or a discrete analog.