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NASA NTRS1999

Configuration Aerodynamics: Past - Present - Future

The Configuration Aerodynamics (CA) element of the High Speed Research (HSR) program is managed by a joint NASA and Industry team, referred to as the Technology Integration Development (ITD) team. This team is responsible for the development of a broad range of technologies for improved aerodynamic performance and stability and control characteristics at subsonic to supersonic flight conditions. These objectives are pursued through the aggressive use of advanced experimental test techniques and state of the art computational methods. As the HSR program matures and transitions into the next phase the objectives of the Configuration Aerodynamics ITD are being refined to address the drag reduction needs and stability and control requirements of High Speed Civil Transport (HSCT) aircraft. In addition, the experimental and computational tools are being refined and improved to meet these challenges. The presentation will review the work performed within the Configuration Aerodynamics element in 1994 and 1995 and then discuss the plans for the 1996-1998 time period. The final portion of the presentation will review several observations of the HSR program and the design activity within Configuration Aerodynamics.

Wood, Richard M.↗

NASA NTRS2001

Characterizing Aeroelastic Systems Using Eigenanalysis, Explicitly Retaining The Aerodynamic Degrees of Freedom

Discrete time aeroelastic models with explicitly retained aerodynamic modes have been generated employing a time marching vortex lattice aerodynamic model. This paper presents analytical results from eigenanalysis of these models. The potential of these models to calculate the behavior of modes that represent damped system motion (noncritical modes) in addition to the simple harmonic modes is explored. A typical section with only structural freedom in pitch is examined. The eigenvalues are examined and compared to experimental data. Issues regarding the convergence of the solution with regard to refining the aerodynamic discretization are investigated. Eigenvector behavior is examined; the eigenvector associated with a particular eigenvalue can be viewed as the set of modal participation factors for that particular mode. For the present formulation of the equations of motion, the vorticity for each aerodynamic element appears explicitly as an element of each eigenvector in addition to the structural dynamic generalized coordinates. Thus, modal participation of the aerodynamic degrees of freedom can be assessed in M addition to participation of structural degrees of freedom.

Heeg, Jennifer↗

NASA NTRS2001

System Identification and POD Method Applied to Unsteady Aerodynamics

The representation of unsteady aerodynamic flow fields in terms of global aerodynamic modes has proven to be a useful method for reducing the size of the aerodynamic model over those representations that use local variables at discrete grid points in the flow field. Eigenmodes and Proper Orthogonal Decomposition (POD) modes have been used for this purpose with good effect. This suggests that system identification models may also be used to represent the aerodynamic flow field. Implicit in the use of a systems identification technique is the notion that a relative small state space model can be useful in describing a dynamical system. The POD model is first used to show that indeed a reduced order model can be obtained from a much larger numerical aerodynamical model (the vortex lattice method is used for illustrative purposes) and the results from the POD and the system identification methods are then compared. For the example considered, the two methods are shown to give comparable results in terms of accuracy and reduced model size. The advantages and limitations of each approach are briefly discussed. Both appear promising and complementary in their characteristics.

Tang, Deman↗

NASA NTRS2002

Supersonic Aerodynamic Characteristics of Proposed Mars '07 Smart Lander Configurations

Supersonic aerodynamic data were obtained for proposed Mars '07 Smart Lander configurations in NASA Langley Research Center's Unitary Plan Wind Tunnel. The primary objective of this test program was to assess the supersonic aerodynamic characteristics of the baseline Smart Lander configuration with and without fixed shelf/tab control surfaces. Data were obtained over a Mach number range of 2.3 to 4.5, at a free stream Reynolds Number of 1 x 10(exp 6) based on body diameter. All configurations were run at angles of attack from -5 to 20 degrees and angles of sideslip of -5 to 5 degrees. These results were complemented with computational fluid dynamic (CFD) predictions to enhance the understanding of experimentally observed aerodynamic trends. Inviscid and viscous full model CFD solutions compared well with experimental results for the baseline and 3 shelf/tab configurations. Over the range tested, Mach number effects were shown to be small on vehicle aerodynamic characteristics. Based on the results from 3 different shelf/tab configurations, a fixed control surface appears to be a feasible concept for meeting aerodynamic performance metrics necessary to satisfy mission requirements.

Murphy, Kelly J.↗

NASA NTRS2003

An Experimental and Computational Investigation of Oscillating Airfoil Unsteady Aerodynamics at Large Mean Incidence

A major challenge in the design and development of turbomachine airfoils for gas turbine engines is high cycle fatigue failures due to flutter and aerodynamically induced forced vibrations. In order to predict the aeroelastic response of gas turbine airfoils early in the design phase, accurate unsteady aerodynamic models are required. However, accurate predictions of flutter and forced vibration stress at all operating conditions have remained elusive. The overall objectives of this research program are to develop a transition model suitable for unsteady separated flow and quantify the effects of transition on airfoil steady and unsteady aerodynamics for attached and separated flow using this model. Furthermore, the capability of current state-of-the-art unsteady aerodynamic models to predict the oscillating airfoil response of compressor airfoils over a range of realistic reduced frequencies, Mach numbers, and loading levels will be evaluated through correlation with benchmark data. This comprehensive evaluation will assess the assumptions used in unsteady aerodynamic models. The results of this evaluation can be used to direct improvement of current models and the development of future models. The transition modeling effort will also make strides in improving predictions of steady flow performance of fan and compressor blades at off-design conditions. This report summarizes the progress and results obtained in the first year of this program. These include: installation and verification of the operation of the parallel version of TURBO; the grid generation and initiation of steady flow simulations of the NASA/Pratt&Whitney airfoil at a Mach number of 0.5 and chordal incidence angles of 0 and 10 deg.; and the investigation of the prediction of laminar separation bubbles on a NACA 0012 airfoil.

Capece, Vincent R.↗

NASA NTRS2005

Static Aerodynamics of the Mars Exploration Rover Entry Capsule

The static aerodynamics for the Mars Exploration Rover (MER) aeroshell are presented. This aerodynamic database was an integral part of the end-to-end simulation used in pre- entry analysis for determining the MER entry design requirements for development of the MER entry system, as well as targeting the MER landing sites. The database was constructed using the same approach used for Mars Pathfinder (MPF). However, the MER aerodynamic database is of much higher fidelity and tailored to the MER entry trajectories. This set of data includes direct simulation Monte Carlo calculations covering the transitional regime of the entry trajectory and computational fluid dynamics calculations describing the aerodynamics in the hypersonic and supersonic continuum regimes. An overview of the methodology used to generate the data is given along with comparisons to important features in the MPF aerodynamics and related heritage data. The MER and MPF comparison indicates that trajectory specific data is required to properly model the flight characteristics of a.blunt entry capsule at Mars.

Schoenenberger, Mark↗

NASA NTRS2005

Aerodynamic Parameter Estimation for the X-43A (Hyper-X) from Flight Data

Aerodynamic parameters were estimated based on flight data from the third flight of the X-43A hypersonic research vehicle, also called Hyper-X. Maneuvers were flown using multiple orthogonal phase-optimized sweep inputs applied as simultaneous control surface perturbations at Mach 8, 7, 6, 5, 4, and 3 during the vehicle descent. Aerodynamic parameters, consisting of non-dimensional longitudinal and lateral stability and control derivatives, were estimated from flight data at each Mach number. Multi-step inputs at nearly the same flight conditions were also flown to assess the prediction capability of the identified models. Prediction errors were found to be comparable in magnitude to the modeling errors, which indicates accurate modeling. Aerodynamic parameter estimates were plotted as a function of Mach number, and compared with estimates from the pre-flight aerodynamic database, which was based on wind-tunnel tests and computational fluid dynamics. Agreement between flight estimates and values computed from the aerodynamic database was excellent overall.

Morelli, Eugene A.↗

NASA NTRS2007

The DaveMLTranslator: An Interface for DAVE-ML Aerodynamic Models

It can take weeks or months to incorporate a new aerodynamic model into a vehicle simulation and validate the performance of the model. The Dynamic Aerospace Vehicle Exchange Markup Language (DAVE-ML) has been proposed as a means to reduce the time required to accomplish this task by defining a standard format for typical components of a flight dynamic model. The purpose of this paper is to describe an object-oriented C++ implementation of a class that interfaces a vehicle subsystem model specified in DAVE-ML and a vehicle simulation. Using the DaveMLTranslator class, aerodynamic or other subsystem models can be automatically imported and verified at run-time, significantly reducing the elapsed time between receipt of a DAVE-ML model and its integration into a simulation environment. The translator performs variable initializations, data table lookups, and mathematical calculations for the aerodynamic build-up, and executes any embedded static check-cases for verification. The implementation is efficient, enabling real-time execution. Simple interface code for the model inputs and outputs is the only requirement to integrate the DaveMLTranslator as a vehicle aerodynamic model. The translator makes use of existing table-lookup utilities from the Langley Standard Real-Time Simulation in C++ (LaSRS++). The design and operation of the translator class is described and comparisons with existing, conventional, C++ aerodynamic models of the same vehicle are given.

Hill, Melissa A.↗

NASA NTRS2008

NASA Perspective on Requirements for Development of Advanced Methods Predicting Unsteady Aerodynamics and Aeroelasticity

Over the past three years, the National Aeronautics and Space Administration (NASA) has initiated design, development, and testing of a new human-rated space exploration system under the Constellation Program. Initial designs within the Constellation Program are scheduled to replace the present Space Shuttle, which is slated for retirement within the next three years. The development of vehicles for the Constellation system has encountered several unsteady aerodynamics challenges that have bearing on more traditional unsteady aerodynamic and aeroelastic analysis. This paper focuses on the synergy between the present NASA challenges and the ongoing challenges that have historically been the subject of research and method development. There are specific similarities in the flows required to be analyzed for the space exploration problems and those required for some of the more nonlinear unsteady aerodynamic and aeroelastic problems encountered on aircraft. The aggressive schedule, significant technical challenge, and high-priority status of the exploration system development is forcing engineers to implement existing tools and techniques in a design and application environment that is significantly stretching the capability of their methods. While these methods afford the users with the ability to rapidly turn around designs and analyses, their aggressive implementation comes at a price. The relative immaturity of the techniques for specific flow problems and the inexperience with their broad application to them, particularly on manned spacecraft flight system, has resulted in the implementation of an extensive wind tunnel and flight test program to reduce uncertainty and improve the experience base in the application of these methods. This provides a unique opportunity for unsteady aerodynamics and aeroelastic method developers to test and evaluate new analysis techniques on problems with high potential for acquisition of test and even flight data against which they can be evaluated. However, researchers may be required to alter the geometries typically used in their analyses, the types of flows analyzed, and even the techniques by which computational tools are verified and validated. This paper discusses these issues and provides some perspective on the potential for new and innovative approaches to the development of methods to attack problems in nonlinear unsteady aerodynamics.

Schuster, David M.↗

NASA NTRS2009

Flight Dynamics of an Aeroshell Using an Attached Inflatable Aerodynamic Decelerator

An aeroelastic analysis of the behavior of an entry vehicle utilizing an attached inflatable aerodynamic decelerator during supersonic flight is presented. The analysis consists of a planar, four degree of freedom simulation. The aeroshell and the IAD are assumed to be separate, rigid bodies connected with a spring-damper at an interface point constraining the relative motion of the two bodies. Aerodynamic forces and moments are modeled using modified Newtonian aerodynamics. The analysis includes the contribution of static aerodynamic forces and moments as well as pitch damping. Two cases are considered in the analysis: constant velocity flight and planar free flight. For the constant velocity and free flight cases with neutral pitch damping, configurations with highly-stiff interfaces exhibit statically stable but dynamically unstable aeroshell angle of attack. Moderately stiff interfaces exhibit static and dynamic stability of aeroshell angle of attack due to damping induced by the pitch angle rate lag between the aeroshell and IAD. For the free-flight case, low values of both the interface stiffness and damping cause divergence of the aeroshell angle of attack due to the offset of the IAD drag force with respect to the aeroshell center of mass. The presence of dynamic aerodynamic moments was found to influence the stability characteristics of the vehicle. The effect of gravity on the aeroshell angle of attack stability characteristics was determined to be negligible for the cases investigated.

Cruz, Juan R.↗

NASA NTRS2010

Modeling the Launch Abort Vehicle's Subsonic Aerodynamics from Free Flight Testing

An investigation into the aerodynamics of the Launch Abort Vehicle for NASA's Constellation Crew Launch Vehicle in the subsonic, incompressible flow regime was conducted in the NASA Langley 20-ft Vertical Spin Tunnel. Time histories of center of mass position and Euler Angles are captured using photogrammetry. Time histories of the wind tunnel's airspeed and dynamic pressure are recorded as well. The primary objective of the investigation is to determine models for the aerodynamic yaw and pitch moments that provide insight into the static and dynamic stability of the vehicle. System IDentification Programs for AirCraft (SIDPAC) is used to determine the aerodynamic model structure and estimate model parameters. Aerodynamic models for the aerodynamic body Y and Z force coefficients, and the pitching and yawing moment coefficients were identified.

Hartman, Christopher L.↗

NASA NTRS2011

Aerodynamic Simulation of Ice Accretion on Airfoils

This report describes recent improvements in aerodynamic scaling and simulation of ice accretion on airfoils. Ice accretions were classified into four types on the basis of aerodynamic effects: roughness, horn, streamwise, and spanwise ridge. The NASA Icing Research Tunnel (IRT) was used to generate ice accretions within these four types using both subscale and full-scale models. Large-scale, pressurized windtunnel testing was performed using a 72-in.- (1.83-m-) chord, NACA 23012 airfoil model with high-fidelity, three-dimensional castings of the IRT ice accretions. Performance data were recorded over Reynolds numbers from 4.5 x 10(exp 6) to 15.9 x 10(exp 6) and Mach numbers from 0.10 to 0.28. Lower fidelity ice-accretion simulation methods were developed and tested on an 18-in.- (0.46-m-) chord NACA 23012 airfoil model in a small-scale wind tunnel at a lower Reynolds number. The aerodynamic accuracy of the lower fidelity, subscale ice simulations was validated against the full-scale results for a factor of 4 reduction in model scale and a factor of 8 reduction in Reynolds number. This research has defined the level of geometric fidelity required for artificial ice shapes to yield aerodynamic performance results to within a known level of uncertainty and has culminated in a proposed methodology for subscale iced-airfoil aerodynamic simulation.

Broeren, Andy P.↗

NASA NTRS2011

Aerodynamic and Aeroacoustic Wind Tunnel Testing of the Orion Spacecraft

The Orion aerodynamic testing team has completed more than 40 tests as part of developing the aerodynamic and loads databases for the vehicle. These databases are key to achieving good mechanical design for the vehicle and to ensure controllable flight during all potential atmospheric phases of a mission, including launch aborts. A wide variety of wind tunnels have been used by the team to document not only the aerodynamics but the aeroacoustic environment that the Orion might experience both during nominal ascents and launch aborts. During potential abort scenarios the effects of the various rocket motor plumes on the vehicle must be accurately understood. The Abort Motor (AM) is a high-thrust, short duration motor that rapidly separates Orion from its launch vehicle. The Attitude Control Motor (ACM), located in the nose of the Orion Launch Abort Vehicle, is used for control during a potential abort. The 8 plumes from the ACM interact in a nonlinear manner with the four AM plumes which required a carefully controlled test to define the interactions and their effect on the control authority provided by the ACM. Techniques for measuring dynamic stability and for simulating rocket plume aerodynamics and acoustics were improved or developed in the course of building the aerodynamic and loads databases for Orion.

Ross, James C.↗

NASA NTRS2011

Unsteady Aerodynamic Model Tuning for Precise Flutter Prediction

A simple method for an unsteady aerodynamic model tuning is proposed in this study. This method is based on the direct modification of the aerodynamic influence coefficient matrices. The aerostructures test wing 2 flight-test data is used to demonstrate the proposed model tuning method. The flutter speed margin computed using only the test validated structural dynamic model can be improved using the additional unsteady aerodynamic model tuning, and then the flutter speed margin requirement of 15 percent in military specifications can apply towards the test validated aeroelastic model. In this study, unsteady aerodynamic model tunings are performed at two time invariant flight conditions, at Mach numbers of 0.390 and 0.456. When the Mach number for the unsteady aerodynamic model tuning approaches to the measured fluttering Mach number, 0.502, at the flight altitude of 9,837 ft, the estimated flutter speed is approached to the measured flutter speed at this altitude. The minimum flutter speed difference between the estimated and measured flutter speed is -0.14 percent.

Pak, Chan-gi↗

NASA NTRS2011

Basis Function Approximation of Transonic Aerodynamic Influence Coefficient Matrix

A technique for approximating the modal aerodynamic influence coefficients matrices by using basis functions has been developed and validated. An application of the resulting approximated modal aerodynamic influence coefficients matrix for a flutter analysis in transonic speed regime has been demonstrated. This methodology can be applied to the unsteady subsonic, transonic, and supersonic aerodynamics. The method requires the unsteady aerodynamics in frequency-domain. The flutter solution can be found by the classic methods, such as rational function approximation, k, p-k, p, root-locus et cetera. The unsteady aeroelastic analysis for design optimization using unsteady transonic aerodynamic approximation is being demonstrated using the ZAERO flutter solver (ZONA Technology Incorporated, Scottsdale, Arizona). The technique presented has been shown to offer consistent flutter speed prediction on an aerostructures test wing 2 configuration with negligible loss in precision in transonic speed regime. These results may have practical significance in the analysis of aircraft aeroelastic calculation and could lead to a more efficient design optimization cycle.

Li, Wesley W.↗

NASA NTRS2012

Preliminary Aerodynamic Investigation of Fan Rotor Blade Morphing

Various new technologies currently under development may enable controlled blade shape variability, or so-called blade morphing, to be practically employed in aircraft engine fans and compressors in the foreseeable future. The current study is a relatively brief, preliminary computational fluid dynamics investigation aimed at partially demonstrating and quantifying the aerodynamic potential of fan rotor blade morphing. The investigation is intended to provide information useful for near-term planning, as well as aerodynamic solution data sets that can be subsequently analyzed using advanced acoustic diagnostic tools, for the purpose of making fan noise comparisons. Two existing fan system models serve as baselines for the investigation: the Advanced Ducted Propulsor fan with a design tip speed of 806 ft/sec and a pressure ratio of 1.294, and the Source Diagnostic Test fan with a design tip speed of 1215 ft/sec and a pressure ratio of 1.470. Both are 22-in. sub-scale, low-noise research fan/nacelle models that have undergone extensive experimental testing in the 9- by 15-foot Low Speed Wind Tunnel at the NASA Glenn Research Center. The study, restricted to fan rotor blade morphing only, involves a fairly simple blade morphing technique. Specifically, spanwise-linear variations in rotor blade-section setting angle are applied to alter the blade shape; that is, the blade is linearly retwisted from hub to tip. Aerodynamic performance comparisons are made between morphed-blade and corresponding baseline configurations on the basis of equal fan system thrust, where rotor rotational speed for the morphed-blade fan is varied to change the thrust level for that configuration. The results of the investigation confirm that rotor blade morphing could be a useful technology, with the potential to enable significant improvements in fan aerodynamic performance. Even though the study is very limited in scope and confined to simple geometric perturbations of two existing fan systems, the aerodynamic effectiveness of blade morphing is demonstrated by the configurations analyzed. In particular, for the Advanced Ducted Propulsor fan it is demonstrated that the performance levels of the original variable-pitch baseline design can be achieved using blade morphing instead of variable pitch, and for the Source Diagnostic Test fan the performance at important off-design operating points is substantially increased with blade morphing.

Tweedt, Daniel L.↗

NASA NTRS2013

Assessment of the Reconstructed Aerodynamics of the Mars Science Laboratory Entry Vehicle

On August 5, 2012, the Mars Science Laboratory entry vehicle successfully entered Mars atmosphere, flying a guided entry until parachute deploy. The Curiosity rover landed safely in Gale crater upon completion of the Entry Descent and Landing sequence. This paper compares the aerodynamics of the entry capsule extracted from onboard flight data, including Inertial Measurement Unit (IMU) accelerometer and rate gyro information, and heatshield surface pressure measurements. From the onboard data, static force and moment data has been extracted. This data is compared to preflight predictions. The information collected by MSL represents the most complete set of information collected during Mars entry to date. It allows the separation of aerodynamic performance from atmospheric conditions. The comparisons show the MSL aerodynamic characteristics have been identified and resolved to an accuracy better than the aerodynamic database uncertainties used in preflight simulations. A number of small anomalies have been identified and are discussed. This data will help revise aerodynamic databases for future missions and will guide computational fluid dynamics (CFD) development to improved prediction codes.

Schoenenberger, Mark↗

NASA NTRS2013

Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Technology Development Overview

The successful flight of the Inflatable Reentry Vehicle Experiment (IRVE)-3 has further demonstrated the potential value of Hypersonic Inflatable Aerodynamic Decelerator (HIAD) technology. This technology development effort is funded by NASA's Space Technology Mission Directorate (STMD) Game Changing Development Program (GCDP). This paper provides an overview of a multi-year HIAD technology development effort, detailing the projects completed to date and the additional testing planned for the future. The effort was divided into three areas: Flexible Systems Development (FSD), Mission Advanced Entry Concepts (AEC), and Flight Validation. FSD consists of a Flexible Thermal Protection Systems (FTPS) element, which is investigating high temperature materials, coatings, and additives for use in the bladder, insulator, and heat shield layers; and an Inflatable Structures (IS) element which includes manufacture and testing (laboratory and wind tunnel) of inflatable structures and their associated structural elements. AEC consists of the Mission Applications element developing concepts (including payload interfaces) for missions at multiple destinations for the purpose of demonstrating the benefits and need for the HIAD technology as well as the Next Generation Subsystems element. Ground test development has been pursued in parallel with the Flight Validation IRVE-3 flight test. A larger scale (6m diameter) HIAD inflatable structure was constructed and aerodynamically tested in the National Full-scale Aerodynamics Complex (NFAC) 40ft by 80ft test section along with a duplicate of the IRVE-3 3m article. Both the 6m and 3m articles were tested with instrumented aerodynamic covers which incorporated an array of pressure taps to capture surface pressure distribution to validate Computational Fluid Dynamics (CFD) model predictions of surface pressure distribution. The 3m article also had a duplicate IRVE-3 Thermal Protection System (TPS) to test in addition to testing with the Aerocover configuration. Both the Aerocovers and the TPS were populated with high contrast targets so that photogrammetric solutions of the loaded surface could be created. These solutions both refined the aerodynamic shape for CFD modeling and provided a deformed shape to validate structural Finite Element Analysis (FEA) models. Extensive aerothermal testing has been performed on the TPS candidates. This testing has been conducted in several facilities across the country. The majority of the testing has been conducted in the Boeing Large Core Arc Tunnel (LCAT). HIAD is continuing to mature testing methodology in this facility and is developing new test sample fixtures and control methodologies to improve understanding and quality of the environments to which the samples are subjected. Additional testing has been and continues to be performed in the NASA LaRC 8ft High Temperature Tunnel, where samples up to 2ft by 2ft are being tested over representative underlying structures incorporating construction features such as sewn seams and through-thickness quilting. With the successful completion to the IRVE-3 flight demonstration, mission planning efforts are ramping up on the development of the HIAD Earth Atmospheric Reenty Test (HEART) which will demonstrate a relevant scale vehicle in relevant environments via a large-scale aeroshell (approximately 8.5m) entering at orbital velocity (approximately 7km/sec) with an entry mass on the order of 4MT. Also, the Build to Print (BTP) hardware built as a risk mitigation for the IRVE-3 project to have a "spare" ready to go in the event of a launch vehicle delivery failure is now available for an additional sub-orbital flight experiment. Mission planning is underway to define a mission that can utilize this existing hardware and help the HIAD project further mature this technology.

Hughes, Stephen J.↗