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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 451 records · Page 25
All sourcesNASA NTRSDOE OSTI
NASA NTRS1982

Space Shuttle entry longitudinal aerodynamic comparisons of flight 2 with preflight predictions

An overview of the aerodynamic performance, longitudinal trim, and control surface effectiveness as developed from the second Shuttle flight is presented, along with comparisons with preflight predictions. The Orbiter enters the atmosphere at Mach 29 and descends through all the Mach speeds until touchdown. Control mechanisms for the Orbiter are outlined, noting the lessening of angle of attack as the terminal configuration is reached. Methods of developing the wind tunnel predictions are reviewed, including performance in different altitudes and speed regimes, and the impetus to decouple the flight control system from aerodynamics with respect to uncertainties. Preflight predictions agreed well with the lift/drag ratio above Mach 1, but higher L/D was exhibited than predicted below Mach 1. A deficiency in the wind tunnel environment was concluded for the inaccuracies in the predictions for hypersonic trim characteristics.

Romere, P. O.↗
NASA NTRS1990

The calculation of rotor/fuselage interaction for two-dimensional bodies

Unsteady rotor wake interactions with the empennage, tail boom, and other aerodynamic surfaces have a significant influence on the aerodynamic performance of the helicopter, ride quality, and vibration. A Computational Fluid Dynamic (CFD) method for computing the aerodynamic interaction between an interacting vortex wake and the viscous flow about arbitrary 2-D bodies was developed to address this helicopter problem. The vorticity and flow field velocities are calculated on a body-fitted computational mesh using an uncoupled iterative solution. The interacting vortex wake is represented by an array of discrete vortices which, in turn, are represented by a finite core model. The evolution of the interacting vortex wake is calculated by Lagrangian techniques. The flow around circular and elliptic cylinders in the absence of an interacting vortex wake was calculated. These results compare very well with other numerical results and with results obtained from experiment and thereby demonstrate the accuracy of the viscous solution. The interaction of a simulated rotor wake with the flow about 2-D bodies, representing cross sections of fuselage components, was calculated to address the vortex interaction problem. The vortex interaction was calculated for the flow about a circular and an elliptic cylinder at 45 and 90 degrees incidence. The results demonstrate the significant variation in lift and drag on the 2-D bodies during the vortex interaction.

Stremel, Paul M.↗
NASA NTRSDate not supplied

Free-Flight CFD Simulations and Dynamic StabilityAnalysis of the Orion Crew Module

Dynamic stability analysis of the Multi-Purpose Crew Vehicle (MPCV) is performed us-ing the US3D flow solver and a plugin library to solve rigid-body dynamics, Free-Flight CFD(FF-CFD). Previous efforts investigated the free-flight behavior of higher speed, open-backaeroshell shapes. Three primary experimental sources are used to evaluate the predictivecapability of the FF-CFD solver in the low supersonic range (Mach≈1). First, the ballisticrange results obtained at the HFFAF facility at NASA Ames is used to verify the six degree-of-freedom (6-DoF) dynamic capability of the FF-CFD solver. Next, FF-CFD simulationsare preformed using restricted motion and the resultant trajectories are post-processed toobtain pitch damping coefficient as a function of angle-of-attack. The comparison of thepitch damping results obtained with one degree-of-freedom (1-DoF) FF-CFD compare wellwith experimental fits derived from ballistic range data. Finally, the atmospheric flight ca-pability of FF-CFD is compared to data from the Ascent Abort-2 (AA-2) flight experiment.Two simulations were performed using forced and forced-free flight to investigate surfacepressure predictive capability and free-flight aerodynamic performance through a varyingatmosphere at low speed (Mach 0.6-0.2). Surface pressure predicted with FF-CFD agreeswell with experimental trends, with slight over-prediction near the end of the trajectory.The total angle-of-attack for the free-flight portion agrees well with experimental data.

FFCFD↗
NASA NTRSDate not supplied

Exploring the Impact of Sideslip Angle on Acoustic Predictions of an Open Rotor System Using Scale-Resolving Simulations

The development of innovative open-rotor propulsion systems that meet performance, fuel efficiency, and noise standards for commercial use depends significantly on the availability of reliable and computationally efficient tools for aerodynamic performance and noise prediction.This investigation demonstrates the capability of the structured curvilinear overset gridComputational Fluid Dynamics solver within the Launch, Ascent, and Vehicle Aerodynamics framework to predict the aerodynamic and aeroacoustic characteristics of a Contra-RotatingOpen Rotor propulsion system under various sideslip angles. In this publication, results obtained using a Hybrid Reynolds-Averaged Navier-Stokes/Large Eddy Simulation-based scale-resolving turbulence closure approach are validated against existing experimental measurements including aerodynamic loads, moments, and tones.

AATT↗
NASA NTRS2004

NASA's Aeronautics Vision

Six long-term technology focus areas are: 1. Environmentally Friendly, Clean Burning Engines. Focus: Develop innovative technologies to enable intelligent turbine engines that significantly reduce harmful emissions while maintaining high performance and increasing reliability. 2. New Aircraft Energy Sources and Management. Focus: Discover new energy sources and intelligent management techniques directed towards zero emissions and enable new vehicle concepts for public mobility and new science missions. 3. Quiet Aircraft for Community Friendly Service. Focus: Develop and integrate noise reduction technology to enable unrestricted air transportation service to all communities. 4. Aerodynamic Performance for Fuel Efficiency. Focus: Improve aerodynamic efficiency,structures and materials technologies, and design tools and methodologies to reduce fuel burn and minimize environmental impact and enable new vehicle concepts and capabilities for public mobility and new science missions. 5. Aircraft Weight Reduction and Community Access. Focus: Develop ultralight smart materials and structures, aerodynamic concepts, and lightweight subsystems to increase vehicle efficiency, leading to high altitude long endurance vehicles, planetary aircraft, advanced vertical and short takeoff and landing vehicles and beyond. 6. Smart Aircraft and Autonomous Control. Focus: Enable aircraft to fly with reduced or no human intervention, to optimize flight over multiple regimes, and to provide maintenance on demand towards the goal of a feeling, seeing, sensing, sentient air vehicle.

Tenney, Darrel R.↗
DOE OSTI2020

Fast Evaluation of Aircraft Icing Severity Using Machine Learning Based on XGBoost

Aircraft icing represents a serious hazard in aviation which has caused a number of fatal accidents over the years. In addition, it can lead to substantial increase in drag and weight, thus reducing the aerodynamics performance of the airplane. The process of ice accretion on a solid surface is a complex interaction of aerodynamic and environmental variables. The complex relationship makes machine learning-based methods an attractive alternative to traditional numerical simulation-based approaches. In this study, we introduce a purely data-driven approach to find the complex pattern between different flight conditions and aircraft icing severity prediction. The supervised learning algorithm Extreme Gradient Boosting (XGBoost) is applied to establish the prediction framework which makes prediction based on any set of observations. The input flight conditions for the proposed prediction framework are liquid water content, droplet diameter and exposure time. The proposed approach is demonstrated in three cases: maximum ice thickness prediction, icing area prediction and icing severity level evaluation. Performance comparison studies and error analysis are also conducted to verify the effectiveness and performance of the proposed method. Results show that the proposed method has reasonable capability in evaluating aircraft icing severity.

Li, Sibo (ORCID:000000021705844X)↗
NASA NTRS1972

Performance, static stability, and control effectiveness of a parametric space shuttle launch vehicle

This test was run as a continuation of a prior investigation of aerodynamic performance and static stability tests for a parametric space shuttle launch vehicle. The purposes of this test were: (1) to obtain a more complete set of data in the transonic flight region, (2) to investigate new H-0 tank noseshapes and tank diameters, (3) to obtain control effectiveness data for the orbiter at 0 degree incidence and with a smaller diameter H-0 tank, and (4) to determine the effects of varying solid rocket motor-to-H0 tank gap size. Experimental data were obtained for angles of attack from -10 to +10 degrees and for angles of sideslip from +10 to -10 degrees at Mach numbers ranging from .6 to 4.96.

Buchholz, R. E.↗
NASA NTRS1982

The effect of rotor blade thickness and surface finish on the performance of a small axial flow turbine

An experimental investigation was conducted to determine the effect of blade profile inaccuracies and surface finish on the aerodynamic performance of a 11.13 cm tip diameter turbine. The as-received cast rotor blades had a significantly thicker profile than the design intent and a fairly rough surface finish. Stage test results showed an increase of one point in efficieny by smoothing the surface finish and another three points by thinning the blade profiles to near the design profile. Most of the performance gain between the as-cast thick and the thinned rotor blades both with the same surface finish, was attributed to reduced trailing edge losses of the recontoured blades.

Roelke, R. J.↗
NASA NTRS1999

Aeroelastic Tailoring for Stability Augmentation and Performance Enhancements of Tiltrotor Aircraft

The requirements for increased speed and productivity for tiltrotors has spawned several investigations associated with proprotor aeroelastic stability augmentation and aerodynamic performance enhancements. Included among these investigations is a focus on passive aeroelastic tailoring concepts which exploit the anisotropic capabilities of fiber composite materials. Researchers at Langley Research Center and Bell Helicopter have devoted considerable effort to assess the potential for using these materials to obtain aeroelastic responses which are beneficial to the important stability and performance considerations of tiltrotors. Both experimental and analytical studies have been completed to examine aeroelastic tailoring concepts for the tiltrotor, applied either to the wing or to the rotor blades. This paper reviews some of the results obtained in these aeroelastic tailoring investigations and discusses the relative merits associated with these approaches.

Nixon, Mark W.↗
NASA NTRS1997

Unsteady Velocity Measurements Taken Behind a Model Helicopter Rotor Hub in Forward Flight

Drag caused by separated flow behind the hub of a helicopter has an adverse effect on aerodynamic performance of the aircraft. To determine the effect of separated flow on a configuration used extensively for helicopter aerodynamic investigations, an experiment was conducted using a laser velocimeter to measure velocities in the wake of a model helicopter hub operating at Mach-scaled conditions in forward flight. Velocity measurements were taken using a laser velocimeter with components in the vertical and downstream directions. Measurements were taken at 13 stations downstream from the rotor hub. At each station, measurements were taken in both a horizontal and vertical row of locations. These measurements were analyzed for harmonic content based on the rotor period of revolution. After accounting for these periodic velocities, the remaining unsteady velocities were treated as turbulence. Turbulence intensity distributions are presented. Average turbulent intensities ranged from approximately 2 percent of free stream to over 15 percent of free stream at specific locations and azimuths. The maximum average value of turbulence was located near the rear-facing region of the fuselage.

Berry, John D.↗
NASA NTRS2001

Toward an Efficient Icing CFD Process Using an Interactive Software Toolkit: Smagglce 2D

Two-dimensional CID analysis for iced airfoils can be a labor-intensive task. The software toolkit SmaggIce 2D is being developed to help streamline the CID process and provide the unique features needed for icing. When complete, it will include a combination of partially automated and fully interactive tools for all aspects of the tasks leading up to the flow analysis: geometry preparation, domain decomposition. block boundary demoralization. gridding, and linking with a flow solver. It also includes tools to perform ice shape characterization, an important aid in determining the relationship between ice characteristics and their effects on aerodynamic performance. Completed tools, work-in-progress, and planned features of the software toolkit are presented here.

Vickerman, Mary B.↗
NASA NTRS2002

Toward an Efficient Icing CFD Process Using an Interactive Software Toolkit--SmaggIce 2D

Two-dimensional CFD analysis for iced airfoils can be a labor-intensive task. The software toolkit SmaggIce 2D is being developed to help streamline the CFD process and provide the unique features needed for icing. When complete, it will include a combination of partially automated and fully interactive tools for all aspects of the tasks leading up to the flow analysis: geometry preparation, domain decomposition, block boundary discretization. gridding, and linking with a flow solver. It also includes tools to perform ice shape characterization, an important aid in determining the relationship between ice characteristics and their effects on aerodynamic performance. Completed tools, work-in-progress, and planned features of the software toolkit are presented here.

Vickerman, Mary B.↗
NASA NTRS1998

Langley Aerothermodynamic Facilities Complex: Enhancements and Testing Capabilities

Description, capabilities, recent upgrades, and utilization of the NASA Langley Research Center (LaRC) Aerothermodynamic Facilities Complex (AFC) are presented. The AFC consists of five hypersonic, blow-down-to-vacuum wind tunnels that collectively provide a range of Mach number from 6 to 20, unit Reynolds number from 0.04 to 22 million per foot and, most importantly for blunt configurations, normal shock density ratio from 4 to 12. These wide ranges of hypersonic simulation parameters are due, in part, to the use of three different test gases (air, helium, and tetrafluoromethane), thereby making several of the facilities unique. The Complex represents nearly three-fourths of the conventional (as opposed to impulse)-type hypersonic wind tunnels operational in this country. AFC facilities are used to assess and optimize the hypersonic aerodynamic performance and aeroheating characteristics of aerospace vehicle concepts and to provide benchmark aerodynamic/aeroheating data fr generating the flight aerodynamic databook and final design of the thermal protection system (TPS) (e.g., establishment of flight limitations not to exceed TPS design limits). Modifications and enhancements of AFC hardware components and instrumentation have been pursued to increase capability, reliability, and productivity in support of programmatic goals. Examples illustrating facility utilization in recent years to generate essentially all of the experimental hypersonic aerodynamic and aeroheating information for high-priority, fast-paced Agency programs are presented. These programs include Phase I of the Reusable Launch Vehicle (RLV) Advanced Technology Demonstrator, X-33 program, PHase II of the X-33 program, X-34 program, the Hyper-X program ( a Mach 5,7, and 10 airbreathing propulsion flight experiment), and the X-38 program (Experimental Crew Return Vehicle, X-CRV). Current upgrades/enchancements and future plans for the AFC are discussed.

Micol, J. R.↗
NASA NTRS2015

Aeroelastic Analysis of a Flexible Wing Wind Tunnel Model with Variable Camber Continuous Trailing Edge Flap Design

This paper presents data analysis of a flexible wing wind tunnel model with a variable camber continuous trailing edge flap (VCCTEF) design for drag minimization tested at the University of Washington Aeronautical Laboratory (UWAL). The wind tunnel test was designed to explore the relative merit of the VCCTEF concept for improved cruise efficiency through the use of low-cost aeroelastic model test techniques. The flexible wing model is a 10%-scale model of a typical transport wing and is constructed of woven fabric composites and foam core. The wing structural stiffness in bending is tailored to be half of the stiffness of a Boeing 757-era transport wing while the torsional stiffness is about the same. This stiffness reduction results in a wing tip deflection of about 10% of the wing semi-span. The VCCTEF is a multi-segment flap design having three chordwise camber segments and five spanwise flap sections for a total of 15 individual flap elements. The three chordwise camber segments can be positioned appropriately to create a desired trailing edge camber. Elastomeric material is used to cover the gaps in between the spanwise flap sections, thereby creating a continuous trailing edge. Wind tunnel data analysis conducted previously shows that the VCCTEF can achieve a drag reduction of up to 6.31% and an improvement in the lift-to-drag ratio (L=D) of up to 4.85%. A method for estimating the bending and torsional stiffnesses of the flexible wingUWAL wind tunnel model from static load test data is presented. The resulting estimation indicates that the stiffness of the flexible wing is significantly stiffer in torsion than in bending by as much as 9 to 1. The lift prediction for the flexible wing is computed by a coupled aerodynamic-structural model. The coupled model is developed by coupling a conceptual aerodynamic tool Vorlax with a finite-element model of the flexible wing via an automated geometry deformation tool. Based on the comparison of the lift curve slope, the lift prediction for the rigid wing is in good agreement with the estimated lift coefficients derived from the wind tunnel test data. Due to the movement of the VCCTEF during the wind tunnel test, uncertainty in the lift prediction due to the indicated variations of the VCCTEF deflection is studied. The results show a significant spread in the lift prediction which contradicts the consistency in the aerodynamic measurements, thus suggesting that the indicated variations as measured by the VICON system may not be reliable. The lift prediction of the flexible wing agrees very well with the measured lift curve for the baseline configuration. The computed bending deflection and wash-out twist of the flexible wing also match reasonably well with the aeroelastic deflection measurements. The results demonstrate the validity of the aerodynamic-structural tool for use to analyze aerodynamic performance of flexible wings.

Wind Tunnel↗
NASA NTRSDate not supplied

TPSAS-NF1676L-13013-DND

Description, capabilities and utilization of the NASA Langley Aerothermodynamics Laboratory (LAL) are presented. The LAL consists of four hypersonic, blow-down-to-vacuum wind tunnels that collectively provide a range of Mach number from 6 to 10 (18 simulation), unit Reynolds number from 0.05 to 8 million per foot and, most importantly for blunt configurations, normal shock density ratio from 4 to 12. These wide ranges of hypersonic simulation parameters are due, in part, to the use of two different test gases (air and tetrafluoromethane, CF4), thereby making the facilities unique. The complex represents half of the conventional (as opposed to impulse) -type hypersonic wind tunnels operational in this country. LAL facilities are used to assess and optimize the hypersonic aerodynamic performance and aeroheating characteristics of aerospace vehicle concepts and to provide benchmark aerodynamic/aeroheating data for generating the flight aerodynamic databook and final design of the thermal protection system (TPS). Modifications and enhancements of LAL hardware components and instrumentation have been pursued to increase capability, reliability, and productivity in support of programmatic goals. The facilities are utilized to generate essentially all of the experimental hypersonic aerodynamic and aeroheating information for high-priority, fast-paced Agency programs and were a major contributor to the Shuttle Columbia Accident Investigation, space shuttle Return-to-Flight program and Hypersonic Thermodynamic Infrared Measurements (HYTHIRM) project. The LAL has also contributed to the development of the Orion Crew Exploration Vehicle, Ares launch system, Mars Science Laboratory, Genesis Sample Return Capsule and supported DOD programs including HIFiRE, X-37, Falcon HTV-2 and X-51A.

John R Micol↗
NASA NTRS2003

Design and Analysis of Turbines for Space Applications

In order to mitigate the risk of rocket propulsion development, efficient, accurate, detailed fluid dynamics analysis of the turbomachinery is necessary. This analysis is used for component development, design parametrics, performance prediction, and environment definition. To support this requirement, a task was developed at NASAh4arshall Space Flight Center (MSFC) to improve turbine aerodynamic performance through the application of advanced design and analysis tools. There are four major objectives of this task: 1) to develop, enhance, and integrate advanced turbine aerodynamic design and analysis tools; 2) to develop the methodology for application of the analytical techniques; 3) to demonstrate the benefits of the advanced turbine design procedure through its application to a relevant turbine design point; and 4) to verify the optimized design and analysis with testing. The turbine chosen on which to demonstrate the procedure was a supersonic design suitable for a reusable launch vehicle (RLV). The hot gas path and blading were redesigned to obtain an increased efficiency. The redesign of the turbine was conducted with a consideration of system requirements, realizing that a highly efficient turbine that, for example, significantly increases engine weight, is of limited benefit. Both preliminary and detailed designs were considered. To generate an improved design, one-dimensional (1D) design and analysis tools, computational fluid dynamics (CFD), response surface methodology (RSM), and neural nets (NN) were used.

Griffin, Lisa W.↗
NASA NTRS2003

Navier-Stokes and Comprehensive Analysis Performance Predictions of the NREL Phase VI Experiment

A vortex lattice code, CAMRAD II, and a Reynolds-Averaged Navier-Stoke code, OVERFLOW-D2, were used to predict the aerodynamic performance of a two-bladed horizontal axis wind turbine. All computations were compared with experimental data that was collected at the NASA Ames Research Center 80- by 120-Foot Wind Tunnel. Computations were performed for both axial as well as yawed operating conditions. Various stall delay models and dynamics stall models were used by the CAMRAD II code. Comparisons between the experimental data and computed aerodynamic loads show that the OVERFLOW-D2 code can accurately predict the power and spanwise loading of a wind turbine rotor.

NREL (NATIONAL RENEWABLE ENERGY LABORATORY)↗
NASA NTRS2014

Comparison of Computational Approaches for Rapid Aerodynamic Assessment of Small UAVs

Computational Fluid Dynamic (CFD) methods were used to determine the basic aerodynamic, performance, and stability and control characteristics of the unmanned air vehicle (UAV), Kahu. Accurate and timely prediction of the aerodynamic characteristics of small UAVs is an essential part of military system acquisition and air-worthiness evaluations. The forces and moments of the UAV were predicted using a variety of analytical methods for a range of configurations and conditions. The methods included Navier Stokes (N-S) flow solvers (USM3D, Kestrel and Cobalt) that take days to set up and hours to converge on a single solution; potential flow methods (PMARC, LSAERO, and XFLR5) that take hours to set up and minutes to compute; empirical methods (Datcom) that involve table lookups and produce a solution quickly; and handbook calculations. A preliminary aerodynamic database can be developed very efficiently by using a combination of computational tools. The database can be generated with low-order and empirical methods in linear regions, then replacing or adjusting the data as predictions from higher order methods are obtained. A comparison of results from all the data sources as well as experimental data obtained from a wind-tunnel test will be shown and the methods will be evaluated on their utility during each portion of the flight envelope.

Shafer, Theresa C.↗