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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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NASA NTRS2005

Navier-Stokes Computations of Longitudinal Forces and Moments for a Blended Wing Body

The object of this paper is to investigate the feasibility of applying CFD methods to aerodynamic analyses for aircraft stability and control. The integrated aerodynamic parameters used in stability and control, however, are not necessarily those extensively validated in the state of the art CFD technology. Hence, an exploratory study of such applications and the comparison of the solutions to available experimental data will help to assess the validity of the current computation methods. In addition, this study will also examine issues related to wind tunnel measurements such as measurement uncertainty and support interference effects. Several sets of experimental data from the NASA Langley 14x22-Foot Subsonic Tunnel and the National Transonic Facility are presented. Two Navier-Stokes flow solvers, one using structured meshes and the other unstructured meshes, were used to compute longitudinal static stability derivatives for an advanced Blended Wing Body configuration over a wide range of angles of attack. The computations were performed for two different Reynolds numbers and the resulting forces and moments are compared with the above mentioned wind tunnel data.

Pao, S. Paul↗
NASA NTRS2011

Aerodynamic Analyses and Database Development for Lift-Off/Transition and First Stage Ascent of the Ares I A106 Vehicle

NASA Langley Research Center, in partnership with NASA Marshall Space Flight Center and NASA Ames Research Center, was involved in the aerodynamic analyses, testing, and database development for the Ares I A106 crew launch vehicle in support of the Ares Design and Analysis Cycle. This paper discusses the development of lift-off/transition and ascent databases. The lift-off/transition database was developed using data from tests on a 1.75% scale model of the A106 configuration in the NASA Langley 14x22 Subsonic Wind Tunnel. The power-off ascent database was developed using test data on a 1% A106 scale model from two different facilities, the Boeing Polysonic Wind Tunnel and the NASA Langley Unitary Plan Wind Tunnel. The ascent database was adjusted for differences in wind tunnel and flight Reynolds numbers using USM3D CFD code. The aerodynamic jet interaction effects due to first stage roll control system were modeled using USM3D and OVERFLOW CFD codes.

Pamadi, Bandu N.↗
NASA NTRS2012

The Development and Hover Test Application of a Projection Moire Interferometry Blade Displacement Measurement System

Projection moir interferometry (PMI) was employed to measure blade deflections during a hover test of a generic model-scale rotor in the NASA Langley 14x22 subsonic wind tunnel s hover facility. PMI was one of several optical measurement techniques tasked to acquire deflection and flow visualization data for a rotor at several distinct heights above a ground plane. Two of the main objectives of this test were to demonstrate that multiple optical measurement techniques can be used simultaneously to acquire data and to identify and address deficiencies in the techniques. Several PMI-specific technical challenges needed to be addressed during the test and in post-processing of the data. These challenges included developing an efficient and accurate calibration method for an extremely large (65 inch) height range; automating the analysis of the large amount of data acquired during the test; and developing a method to determinate the absolute displacement of rotor blades without a required anchor point measurement. The results indicate that the use of a single-camera/single-projector approach for the large height range reduced the accuracy of the PMI system compared to PMI systems designed for smaller height ranges. The lack of the anchor point measurement (due to a technical issue with one of the other measurement techniques) limited the ability of the PMI system to correctly measure blade displacements to only one of the three rotor heights tested. The new calibration technique reduced the data required by 80 percent while new post-processing algorithms successfully automated the process of locating rotor blades in images, determining the blade quarter chord location, and calculating the blade root and blade tip heights above the ground plane.

Sekula, Martin K.↗
NASA NTRS2014

Shielding Characteristics Using an Ultrasonic Configurable Fan Artificial Noise Source to Generate Modes - Experimental Measurements and Analytical Predictions

An Ultrasonic Configurable Fan Artificial Noise Source (UCFANS) was designed, built, and tested in support of the NASA Langley Research Center's 14x22 wind tunnel test of the Hybrid Wing Body (HWB) full 3-D 5.8% scale model. The UCFANS is a 5.8% rapid prototype scale model of a high-bypass turbofan engine that can generate the tonal signature of proposed engines using artificial sources (no flow). The purpose of the program was to provide an estimate of the acoustic shielding benefits possible from mounting an engine on the upper surface of a wing; a flat plate model was used as the shielding surface. Simple analytical simulations were used to preview the radiation patterns - Fresnel knife-edge diffraction was coupled with a dense phased array of point sources to compute shielded and unshielded sound pressure distributions for potential test geometries and excitation modes. Contour plots of sound pressure levels, and integrated power levels, from nacelle alone and shielded configurations for both the experimental measurements and the analytical predictions are presented in this paper.

Artifical Noise↗
NASA NTRS2014

Hybrid Wing Body Shielding Studies Using an Ultrasonic Configurable Fan Artificial Noise Source Generating Typical Turbofan Modes

An Ultrasonic Configurable Fan Artificial Noise Source (UCFANS) was designed, built, and tested in support of the NASA Langley Research Center's 14x22 wind tunnel test of the Hybrid Wing Body (HWB) full 3-D 5.8% scale model. The UCFANS is a 5.8% rapid prototype scale model of a high-bypass turbofan engine that can generate the tonal signature of proposed engines using artificial sources (no flow). The purpose of the test was to provide an estimate of the acoustic shielding benefits possible from mounting the engine on the upper surface of an HWB aircraft using the projected signature of the engine currently proposed for the HWB. The modal structures at the rating points were generated from inlet and exhaust nacelle configurations - a flat plate model was used as the shielding surface and vertical control surfaces with correct plan form shapes were also tested to determine their additional impact on shielding. Radiated acoustic data were acquired from a traversing linear array of 13 microphones, spanning 36 inches. Two planes perpendicular, and two planes parallel, to the axis of the nacelle were acquired from the array sweep. In each plane the linear array traversed 4 sweeps, for a total span of 168 inches acquired. The resolution of the sweep is variable, so that points closer to the model are taken at a higher resolution. Contour plots of Sound Pressure Levels, and integrated Power Levels, from nacelle alone and shielded configurations are presented in this paper; as well as the in-duct mode power levels.

Sutliff, Daniel L.↗
NASA NTRS2016

NASA Environmentally Responsible Aviation Hybrid Wing Body Flow-Through Nacelle Wind Tunnel CFD

Wind tunnel tests of a 5.75 scale model of the Boeing Hybrid Wing Body (HWB) configuration were conducted in the NASA Langley Research Center (LaRC) 14x22 and NASA Ames Research Center (ARC) 40x80 low speed wind tunnels as part of the NASA Environmentally Responsible Aviation (ERA) Project. Computational fluid dynamics (CFD) simulations of the flow-through nacelle (FTN) configuration of this model were performed before and after the testing. This paper presents a summary of the experimental and CFD results for the model in the cruise and landing configurations.

Wind Tunnel Testing↗
NASA NTRS2015

Advancing Test Capabilities at NASA Wind Tunnels

NASA maintains twelve major wind tunnels at three field centers capable of providing flows at 0.1 M 10 and unit Reynolds numbers up to 45106m. The maintenance and enhancement of these facilities is handled through a unified management structure under NASAs Aeronautics and Evaluation and Test Capability (AETC) project. The AETC facilities are; the 11x11 transonic and 9x7 supersonic wind tunnels at NASA Ames; the 10x10 and 8x6 supersonic wind tunnels, 9x15 low speed tunnel, Icing Research Tunnel, and Propulsion Simulator Laboratory, all at NASA Glenn; and the National Transonic Facility, Transonic Dynamics Tunnel, LAL aerothermodynamics laboratory, 8 High Temperature Tunnel, and 14x22 low speed tunnel, all at NASA Langley. This presentation describes the primary AETC facilities and their current capabilities, as well as improvements which are planned over the next five years. These improvements fall into three categories. The first are operations and maintenance improvements designed to increase the efficiency and reliability of the wind tunnels. These include new (possibly composite) fan blades at several facilities, new temperature control systems, and new and much more capable facility data systems. The second category of improvements are facility capability advancements. These include significant improvements to optical access in wind tunnel test sections at Ames, improvements to test section acoustics at Glenn and Langley, the development of a Supercooled Large Droplet capability for icing research, and the development of an icing capability for large engine testing. The final category of improvements consists of test technology enhancements which provide value across multiple facilities. These include projects to increase balance accuracy, provide NIST-traceable calibration characterization for wind tunnels, and to advance optical instruments for Computational Fluid Dynamics (CFD) validation. Taken as a whole, these individual projects provide significant enhancements to NASA capabilities in ground-based testing. They ensure that these wind tunnels will provide accurate and relevant experimental data for years to come, supporting both NASAs mission and the missions of our government and industry customers.

wind tunnel testing↗
NASA NTRS2015

Comparison of Computational and Experimental Microphone Array Results for an 18%-Scale Aircraft Model

An 18%-scale, semi-span model is used as a platform for examining the efficacy of microphone array processing using synthetic data from numerical simulations. Two hybrid RANS/LES codes coupled with Ffowcs Williams-Hawkings solvers are used to calculate 97 microphone signals at the locations of an array employed in the NASA LaRC 14x22 tunnel. Conventional, DAMAS, and CLEAN-SC array processing is applied in an identical fashion to the experimental and computational results for three different configurations involving deploying and retracting the main landing gear and a part span flap. Despite the short time records of the numerical signals, the beamform maps are able to isolate the noise sources, and the appearance of the DAMAS synthetic array maps is generally better than those from the experimental data. The experimental CLEAN-SC maps are similar in quality to those from the simulations indicating that CLEAN-SC may have less sensitivity to background noise. The spectrum obtained from DAMAS processing of synthetic array data is nearly identical to the spectrum of the center microphone of the array, indicating that for this problem array processing of synthetic data does not improve spectral comparisons with experiment. However, the beamform maps do provide an additional means of comparison that can reveal differences that cannot be ascertained from spectra alone.

Lockard, David P.↗
NASA NTRS2016

NASA Hybrid Wing Aircraft Aeroacoustic Test Documentation Report

This report summarizes results of the Hybrid Wing Body (HWB) N2A-EXTE model aeroacoustic test. The N2A-EXTE model was tested in the NASA Langley 14- by 22-Foot Subsonic Tunnel (14x22 Tunnel) from September 12, 2012 until January 28, 2013 and was designated as test T598. This document contains the following main sections: Section 1 - Introduction, Section 2 - Main Personnel, Section 3 - Test Equipment, Section 4 - Data Acquisition Systems, Section 5 - Instrumentation and Calibration, Section 6 - Test Matrix, Section 7 - Data Processing, and Section 8 - Summary. Due to the amount of material to be documented, this HWB test documentation report does not cover analysis of acquired data, which is to be presented separately by the principal investigators. Also, no attempt was made to include preliminary risk reduction tests (such as Broadband Engine Noise Simulator and Compact Jet Engine Simulator characterization tests, shielding measurement technique studies, and speaker calibration method studies), which were performed in support of this HWB test. Separate reports containing these preliminary tests are referenced where applicable.

Heath, Stephanie L.↗
NASA NTRS2015

NASA's Aeroacoustic Tools and Methods for Analysis of Aircraft Noise

Aircraft community noise is a significant concern due to continued growth in air traffic, increasingly stringent environmental goals, and operational limitations imposed by airport authorities. The ability to quantify aircraft noise at the source and ultimately at observers is required to develop low noise aircraft designs and flight procedures. Predicting noise at the source, accounting for scattering and propagation through the atmosphere to the observer, and assessing the perception and impact on a community requires physics-based aeroacoustics tools. Along with the analyses for aero-performance, weights and fuel burn, these tools can provide the acoustic component for aircraft MDAO (Multidisciplinary Design Analysis and Optimization). Over the last decade significant progress has been made in advancing the aeroacoustic tools such that acoustic analyses can now be performed during the design process. One major and enabling advance has been the development of the system noise framework known as Aircraft NOise Prediction Program2 (ANOPP2). ANOPP2 is NASA's aeroacoustic toolset and is designed to facilitate the combination of acoustic approaches of varying fidelity for the analysis of noise from conventional and unconventional aircraft. The toolset includes a framework that integrates noise prediction and propagation methods into a unified system for use within general aircraft analysis software. This includes acoustic analyses, signal processing and interfaces that allow for the assessment of perception of noise on a community. ANOPP2's capability to incorporate medium fidelity shielding predictions and wind tunnel experiments into a design environment is presented. An assessment of noise from a conventional and Hybrid Wing Body (HWB) aircraft using medium fidelity scattering methods combined with noise measurements from a model-scale HWB recently placed in NASA's 14x22 wind tunnel are presented. The results are in the form of community noise metrics and auralizations.

Rizzi, Stephen A.↗
NASA NTRS2016

NASA ERA Integrated CFD for Wind Tunnel Testing of Hybrid Wing-Body Configuration

NASAs Environmentally Responsible Aviation (ERA) Project explores enabling technologies to reduce aviations impact on the environment. One research challenge area for the project has been to study advanced airframe and engine integration concepts to reduce community noise and fuel burn. In order to achieve this, complex wind tunnel experiments at both the NASA Langley Research Centers (LaRC) 14x22 and the Ames Research Centers 40x80 low-speed wind tunnel facilities were conducted on a Boeing Hybrid Wing Body (HWB) configuration. These wind tunnel tests entailed various entries to evaluate the propulsion airframe interference effects including aerodynamic performance and aeroacoustics. In order to assist these tests in producing high quality data with minimal hardware interference, extensive Computational Fluid Dynamic (CFD) simulations were performed for everything from sting design and placement for both the wing body and powered ejector nacelle systems to the placement of aeroacoustic arrays to minimize its impact on the vehicles aerodynamics. This paper will provide a high level summary of the CFD simulations that NASA performed in support of the model integration hardware design as well as some simulation guideline development based on post-test aerodynamic data. In addition, the paper includes details on how multiple CFD codes (OVERFLOW, STAR-CCM+, USM3D, and FUN3D) were efficiently used to provide timely insight into the wind tunnel experimental setup and execution.

Computational Fluid Dynamics↗
NASA NTRS2014

The D8 Aircraft: An Aerodynamics Study of Boundary Layer and Wake Ingestion Benefit

Goal: Quantify benefits of BLI (Boundary Layer Ingesting) and wake ingestion for the D8. Approach: Overset CFD (Computational Fluid Dynamics) using CGT (Chimera Grid Tools software package) and Overflow-­‐2; CFD validation; NASA LaRC (Langley Research Center) 14x22 WT (wind tunnel) data for a 1:11 scale model; Quantifying the BLI and wake ingestion benefit - Direct Comparison between: Efficient conventional (podded nacelle) configuration; BLI (integrated nacelle) configuration.

D8↗
NASA NTRS2019

Wind Tunnel Testing of Active Flow Control on High-Lift Common Research Model

A 10%-scale high-lift version of the Common Research Model (CRM-HL) and an Active Flow Control (AFC) version of the model equipped with a simple-hinged flap (CRM-SHLAFC) were successfully tested. The tests were performed in the 14- by 22-Foot Subsonic Tunnel (14x22) at the NASA Langley Research Center (LaRC). The CRM-HL has a set of 37° inboard and outboard single-element Fowler flaps. The CRM-SHL-AFC has a set of 50° inboard and 55° outboard simple-hinged flaps equipped with integrated modular AFC cartridges on the flap shoulder. Both high-lift configurations share the same 30° slats and engine nacelle. Three new types of AFC devices were examined: the Double-Row Sweeping Jets (DRSWJ), the Alternating Pulsed Jets (APJ), and the High Efficiency Low Power (HELP) actuators. The DRSWJ and the APJ actuators used two rows of unsteady jets, whereas the HELP actuators used a combination of unsteady and steady jets, to overcome strong adverse pressure gradients while minimizing the mass flow usage. Nozzle pressure ratio, mass flow consumption and the power coefficient, which takes account of both supply air pressure and mass flow usage for the actuators, were used for judging the performance efficiency of the AFC devices. A prestall lift performance degradation for the CRM-HL configuration was resolved with a properly placed nacelle chine. The configuration with nacelle chine was chosen as the representative reference conventional high-lift case for comparison with the CRMSHL- AFC. The AFC-induced lift coefficient increment (DCL) was maintained for the entire lift curve over the CRM-SHL-AFC case with no AFC for almost all flow-control cases examined. The lift curve of the reference CRM-HL have a slightly steeper slope compared to those of the CRM-SHL-AFC configurations. The HELP actuation concept was extremely effective in controlling flow separation in the “linear region” of the curves comparing lift coefficient to mass flow rate. The HELP actuation achieved a targeted DCL of 0.50 using a moderate amount of mass flow and supply air pressure. The CRM-SHL-AFC configuration equipped with HELP actuation was able to match or exceed the lift performance of the reference conventional high-lift configuration (i.e., CRM-HL equipped with a nacelle chine), thus meeting the NASA Advanced Air Transport Technology (AATT) project goal.

Lin, John C.↗
NASA NTRS2019

Surface Flow Visualization of the High-Lift Common Research Model

A 10% scale version of the High-Lift Common Research Model (CRM-HL) was tested in the NASA Langley 14- by 22-Foot Subsonic Tunnel (14x22) in support of the NASA Advanced Air Transport Technology (AATT) Project. The CRM-HL experiment included various configurations such as conventional and simple-hinged flaps, with and without engine nacelle/pylon, with and without nacelle chine, different Active Flow Control (AFC) methods (sweeping jets, alternating pulsed jets, and preconditioned boundary layer blowing), and their various parameters. This particular study is focused on the surface flow visualization of the conventional CRM-HL model at landing configuration. The conventional CRM-HL model with the single-slotted Fowler flap system serves as a baseline for the AFC-enabled simplified high-lift configuration as well as a high-lift technology development platform due to its publicly open geometry. Surface flow visualizations were performed using fluorescent minitufts, which were found to be nonintrusive to the aerodynamic performance. Tuft flow visualizations are supplemented with the relevant pressure and force measurements in order to understand the flow characteristics developed on the conventional CRM- HL model. In addition, three dimensional, unsteady, compressible Computational Fluid Dynamic (CFD) simulations were performed for selective cases. The surface streamlines and transverse velocity fluctuations obtained by the CFD simulations are qualitatively compared to the tuft direction and tuft unsteadiness, respectively. Force measurements of the CRM-HL model show performance degradation at higher angles of attack. Surface flow visualizations revealed the performance loss due to the nacelle/pylon wake that grows with angle of attack and eventually promotes flow separation over the inboard wing. This performance loss was successfully recovered by placing a chine on the engine nacelle.

Koklu, Mehti↗
NASA NTRS2020

Wind Tunnel Testing of High Efficiency Low Power (HELP) Actuation for Active Flow Control

A High Efficiency Low Power (HELP) methodology has been successfully developed and tested on an active flow control (AFC) version of the 10%-scale high-lift Common Research Model (CRM-HL) at the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel (14x22). The AFC variant of the CRM-HL, designated as CRM-SHL-AFC, was integrated with modular HELP actuator cartridges on the shoulder of its highly deflected (≥50°) simplehinged flaps. A reference conventional CRM-HL configuration equipped with Fowler flaps and a nacelle chine was chosen to provide the targeted lift enhancement goals and for comparison with the results of the CRM-SHL-AFC. The current data are presented with the Transonic Wall Interference Correction System (TWICS) method applied. The HELP actuators, which use a combination of unsteady sweeping jets and steady discrete jets in tandem, were designed to overcome strong adverse pressure gradients, while minimizing the pneumatic power usage. The power coefficient (C(sub π)), which takes account of both supply air pressure and mass flow usage for the AFC actuators, is a useful parameter for judging the actuators’ performance efficiency and perhaps for scaling-up of the AFC system. Full HELP actuation coverage with a constant spanwise nozzle pressure ratio (NPR) was able to achieve the lift enhancement goals for the entire lift curve and was found to be the most effective AFC case. For the most effective case, mass flow rates greater than ~0.91 lbm/s and NPR values greater than ~1.8 (corresponding to C(sub π) = 0.18) are needed to achieve the lift enhancement goal at α = 9°, whereas mass flow rates greater than ~1.23 lbm/s and NPR values greater than ~2.3 (corresponding to C(sub π) = 0.3) are needed to achieve the lift enhancement goal at α = 17°. Surface pressure distributions indicate that the HELP actuation on the flap shoulder increased the suction pressures (and flow circulation) globally in both the streamwise and the spanwise directions, and thereby enhanced the lift over the entire high-lift system. The CRM-SHLAFC configuration equipped with HELP actuation was able to match or exceed the lift performance of the reference conventional CRM-HL, thus meeting the objective of the research.

John C Lin↗
NASA NTRSDate not supplied

Distributed Pressure Sensing for Enabling Self-Aware Autonomous Aerial Vehicles

Autonomous aerial transportation will be a fixture of future robotic societies, simultaneously requiring more stringent safety requirements and fewer resources for characterization than current commercial air transportation. More robust, adaptable, self-state estimation will be necessary to create such autonomous systems. We present a modular, scalable, distributed pressure sensing skin for aerodynamic state estimation of a large, flexible aerostructure. This skin used a network of 22 nodes that performed in-situ computation and communication of data collected from 74 pressure sensors, which were embedded into the skin panels of an ultra-lightweight 14-foot wingspan made from commutable, lattice-based subcomponents, and tested at NASA Langley Research Center's 14X22 wind tunnel. The density of the pressure sensors allowed for the use of a novel distributed algorithm to generate estimates of the wing lift contribution that were more accurate than the direct integration of the pressure distribution over the wing surface.

Daniel Cellucci↗
NASA NTRS2019

TPSAS-NF1676L-32828-DND

The 14-by 22-Foot Subsonic Tunnel has the capability to simulate propulsion and augmented lift using high pressure air supplied by a variety of air stations. During a customer test in 2018, the chosen air station did not provide the necessary air flow required for the test, so an effort was made to revive a decommissioned air station in a brief time window. This presentation summarizes the high pressure air capabilities of the 14x22 and the challenges of reactivating an air station.

Air station↗
NASA NTRSDate not supplied

Unsteady Pressure Measurement on a Simplified High-Lift Configuration of the Common Research Model using Active Flow Control

The High-Lift Common Research Model (CRM-HL) is a product of the NASA Advanced Air Transport Technology (AATT) Project intended to study various high-lift aerodynamic phenomena including enhanced lift, acoustic noise generation, and optimum placement of leading and trailing-edge devices. The current study was conducted in the NASA Langley 14-by 22-Foot Subsonic Tunnel (14x22) for the purpose of evaluating enhanced lift through the use of active flow control (AFC). The current paper documents the effect of sweeping-jet actuation combined with steady blowing on a simplified high-lift configuration using unsteady pressure measurements. The time-series parameters for the measurements are documented as well as the data acquisition system used to acquire the data. An uncertainty analysis will be presented in order to document the system performance and provide an estimate of the data quality. Repeatability of the data will be shown along with trends associated with model angle of attack, and those associated with changes in the nozzle pressure ratio (NPR) of the AFC system.

Common Research Model↗