ARTLAB

Current and Recent Research Projects:

This research effort proposes to develop a variable-fidelity computational tool for the design of optimal aeroshell configurations for hypersonic entry at Mars, providing NASA with trend and trade-study information on slender and blunt bodies. Hypersonic configurations which maximize landed capability by maximizing drag-area, while satisfying a navigation or deceleration imposed constraint on lift-to-drag ratio will be identified. System-level trades between aerodynamic performance, aerothermodynamic considerations, volumetric efficiency, center-of-mass position, structural design, and aeroshell mass fraction will be explored. By enabling high mass Mars entry systems for future robotic and human exploration missions, this proposed effort is directly relevant to the NASA Strategic Plan, sub-goals 3E.3 and 3C.

To date, a scaled variant of the Viking 70◦ sphere-cone aeroshell has been employed on every Mars landing mission due to its relatively high hypersonic CD (zero angle-of-attack hypersonic CD of approximately 1.68) and the existence of a broad set of aerodynamic performance data on this shape. This aeroshell configuration has been flown successfully along different entry trajectories and at angles-of-attack between 0◦ and 11◦. In this respect, capsules and other blunt shapes tend to compare favorably to slender-body designs that offer lift and higher L/D at the expense of drag (and therefore altitude and timeline).

In this investigation, alternative aeroshell configurations suitable for high mass Mars entry systems will be numerically determined. Vehicle design issues associated with transition to supersonic terminal descent will be assessed in addition to hypersonic aerodynamics, flight dynamics, aeroheating, and aeroshell mass metrics. All disciplinary analyses will be integrated around a numerical optimization framework in which design robustness can be assessed. Variable-fidelity analysis and both direct and indirect optimization techniques will be examined. This variable-fidelity approach will include a high-fidelity Navier-Stokes method based on FUN3D with validation from well-established codes such as DPLR or LAURA. A mid-fidelity approach based on a Cartesian grid analysis coupled to an integral boundary layer approach and a low-fidelity method based on modified Newtonian theory will also be developed. This aeroshell design tool will provide a unique and valuable capability in the analysis of future, high mass Mars entry systems.

Identification and Assessment of Aeroshell Shapes that Maximize Landed Mass Capability for Future Robotic and Human Mars Missions

Research Projects

The computational grid generation process remains a troublesome bottleneck in development of aerodynamic databases for complex geometries and multi-body configurations. The ability of Cartesian-grid based methods to perform automatic mesh generation and dynamic mesh adaptation allows for substantially reduced manpower, time and cost. The advantages of octree-based Cartesian-mesh methods have been already demonstrated for inviscid flow simulations. The goal of this STTR project is to expand the Cartesian-mesh CFD methods for viscous flow simulations. During Phase I, we will explore several innovative methodologies including binary-tree based anisotropic mesh refinement near embedded boundaries, gas-kinetic scheme with embedded boundaries, a normal ray refinement technique and hybrid Cartesian/mesh-free methods. These approaches will be evaluated and ranked for efficient treatment of viscous flow effects near complex boundaries. The ease of implementation, accuracy, efficiency and generality will be investigated for selected test cases. Our framework will provide capabilities for direct simulations up to high Reynolds numbers with support of different turbulence models. The experience of the CFDRC team with Cartesian solvers will be enhanced by the Georgia Tech team to demonstrate the proposed innovations. During Phase II, we will fully implement, demonstrate and validate the prototype algorithms selected in Phase I with particular emphasis to Navy systems.

Viscous Cartesian Flow Solver with AMR Capabilities for Automated Flow Simulation

The radiative heating environment encountered during a high-speed reentry poses a unique problem for the design of thermal protection systems (TPS). The radiation heat flux reaching a body is the result of a complex interplay of many different systems. The emission of radiation is a strong function of the temperature and the species composition present in the shock layer, and in turn, the energy exchange due to the radiation affects the fluid mechanics and chemistry. Ablation and spallation of the surface of the body further influence these factors.

It is proposed to extend the capability of the NASA Ames flow solver Data-Parallel Line Relaxation (DPLR) to model the effects of material response (ablation and spallation) on the radiative heating environment. Ablation-induced turbulence is a well-understood phenomenon and has been implemented by others. However, the effects of radiation caused by the increased diffusion of carbonaceous species into the shock later will be studied. Additionally, the effect of spalled particles in the shock layer on radiative heating will be evaluated. This effort is an extension of previous work with DPLR-NEQAIR coupling, and the methods and capabilities developed will be used here.

Investigation of Ablation and Spallation-Enhanced Radiative Heating Predictions using DPLR

The Boeing Company is under a Phase I study contact to DARPA for the Rapid Eye program.  Boeing will be performing trade studies to determine the optimal concepts for a Rapid Eye Operational System (REOS) and Rapid Eye Demonstration System (REDS).  As part of this activity Boeing will evaluating various concepts that utilize Inflatable Aerodynamic Decelerator (IAD) technology.

As part of the Boeing inflatable aerodynamic decelerator development effort, Georgia Tech is performing systems analysis efforts which focus on 1) A spreadsheet based IAD mass estimation tool, 2) a technology maturation plan, and 3) aerodynamic and structural analysis through the hypersonic phase of flight.

Inflatable Aerodynamic Decelerator Systems Analysis, Technology Maturation, and Aerodynamics and Structural Analysis for the Boeing Rapid Eye program

Issues related to airframe propulsion integration and design simplicity have led naturally to rectangular configurations as the most examined type of scramjet device.  However, this type of cross-section inlet yields some non-desirable structures along several cross-flow planes. In the present work, three injector configurations are examined for circular supersonic combustors.  These arrangements, detailed later, are determined based on results with the rectangular configuration.  Each simulation is analyzed to understand the phenomenology and guide further investigation.  A systematic study is performed to examine both non-reacting and reacting conditions, as well as the effect of inflow profile (uniform versus distorted).

High-fidelity simulations are conducted to perform a parametric analysis of injector placement and orientation in circular cross-section supersonic combustors using a hydrocarbon-air mixture.  Three configurations are compared under conditions where the inflow is uniform as well as when it is coupled to a “Jaws” inward-turning inlet.  The integrated scramjet (i.e., inlet and combustor combination) performance is compared at on-design conditions of Mach 6 and 19.8kPa (1500 psf).  The flame-holder cavity is constructed to be the axisymmetric equivalent of the more traditional rectangular type.  The injector configurations were chosen based on the most effective arrangements observed in previous rectangular cross-section combustors.  Although inward turning inlets enhance inlet performance, flow field irregularities and distortions resulted in non-axis-symmetric combustion downstream and lower efficiency. Several injector configurations derived from rectangular approaches are explored for the circular cross-section cavity flameholder based design.  Parametric simulations delineate the effect of inflow distortion and reactions are described in terms of penetration and mixing.  The results suggest that the penetration depth and fuel-air proportions in the cavity are strongly impacted by inlet distortion and thus influence optimum injector configuration.

Numerical Study of Innovative Scramjet Inlets Coupled to Combustors Using Hydrocarbon-Air Mixture

The cost and time to certify or qualify a rotorcraft for flight in forecast icing has been a major impediment to the development of ice protection systems for helicopter rotors. Development and flight test programs for those aircraft that have achieved certification or qualification for flight in icing conditions have taken many years, and the costs have been very high. NASA, Sikorsky, and others have conducted research into alternatives to natural in flight icing tests to provide information for the development of ice protection systems and to substantiate the airworthiness of a rotor ice protection system. Much progress has been made both in ice accretion modeling and in coupling the ice accretion models to rotary wing aerodynamic analyses. Many of these approaches however were developed for 2-D airfoil sections and 3-D wings with ice formation modeled in a strip theory fashion. Additional work is needed on extending these approaches to 3-D ice accretion models for rotors, for coupling the ice accretion models to the aeroelastic analysis of rotors. This proposal focuses on the coupling process. Physics-based tools are proposed for adapting existing ice accretion models for helicopter rotor applications, and for assessing the attendant loss in performance and increased vibratory loads in forward flight due to icing. A 3-D Navier-Stokes analysis, coupled with multi-body dynamics tool is used to model the aeroelastic behavior of the rotor and trim the vehicle. This analysis is coupled to an unstructured Cartesian grid based analysis that represents the irregular 3-D ice shapes with high fidelity. The growth of ice and the resulting changes to the airloads are modeled by a time-marching scheme. Although the proposed research employs current generation CFD tools and ice accretion tools, it is not limited to these tools. The framework used in this study and employs open interfaces for coupling the analyses, permitting replacement of current generation tools with next generation CFD analyses and ice accretion models. Validation of the methodology and identification of areas that require future work is done using NASA and Sikorsky data for airfoils, wings, model rotors, and full scale flight tests.

 

Physics-Based Modeling, Simulation, and Validation of Ice-Accretion Phenomena for Rotary Wing Aircraft

Over the past several years, rotorcraft designers have begun to make steadily increasing use of flow control devices (usually suction, blowing, or zero mass jets) for improving rotorcraft performance. Other devices such as micro-flaps, plasma jets, and compliant surfaces have also been proposed. These devices, when employed on airframes are effective in reducing downloads and vehicle parasite drag. Compliant wall devices are known to postpone transition. There is also computational and experimental evidence that flow control may be used to modify and alleviate dynamic stall and BVI phenomena. With the advent of improved CFD methods, some of the research in flow control has been directed at modeling flow control. These studies can eliminate much of the trial-and error approach that would otherwise be needed for the proper placement of actuators and sensors.The objective of this study is to develop a multi-scale modeling approach that efficiently models the details of the flow at the device level, properly exchanges information between the outer flow and the device flow field, and solves the outer flow equations. A secondary objective of this approach is to validate this concept through application to a number of flow control problems that have been documented in literature.

The novel approach of this task is inspired by multi-scale models which are used in several fields of engineering. In fluid mechanics, Prandtl’s boundary layer theory and its variants (e.g. triple deck theory) may be viewed as multi-scale models. RANS approach using turbulence models may also be visualized as a multi-scale approach that uses two different set of equations (mean flow equations and transport equations for turbulence properties).

Multi-Scale Modeling of Flow Control Concepts and Devices

The design of future hypersonic vehicles requires detailed understanding of flow regimes ranging from rarefied to continuum. Moreover, hypervelocity flows are characterized by high temperatures, excitation of vibrational level of molecules, nonequilibrium dissociation and ionization.  The goal of this project is to develop unified kinetic/continuum solution methods with proper domain decomposition for a wide range of Air Force Applications. The Unified Flow Solver (UFS) with Adaptive Mesh and Algorithm Refinement (AMAR) will be further developed and demonstrated for viscous/inviscid problems covering rarefied and continuum flow regimes. The octree based Cartesian mesh methods of UFS will be further improved by implementing immersed boundary techniques and anisotropic mesh refinement capabilities to better resolve viscous boundary layers and improve heat transfer simulations near surfaces.  Phase I work will demonstrate the feasibility of kinetic/continuum algorithms with Cartesian mesh to compute heat transfer for dissociating and ionizing hypersonic flows. During Phase II, the advanced numerical techniques will be incorporated into a user-friendly code, a general-purpose chemistry module and turbulence models will be added to the continuum solvers. The code will be validated for heat transfer simulations with Cartesian mesh and demonstrated for several benchmark cases including heat transfer prediction on a Mach 16 flow over bi-conic body.

Unified Kinetic/Continuum Flow Solver with Adaptive Cartesian Mesh for Hypersonic Flows in the Earth Atmosphere

School of Aerospace Engineering