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At least 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS2019

Measurement and Analysis of Terminal Shock Oscillation and Buffet Forcing Functions on a Launch Vehicle Payload Fairing

The buffet loads on a launch vehicle payload shroud can be impacted by the unsteadiness associated with a terminal shock at high subsonic speeds. At these conditions, flow accelerates to supersonic speeds on the nose of the payload fairing and is terminated by a normal shock on the cylindrical section downstream of the nose cone/cylinder shoulder. The location of the terminal shock and associated separated boundary layer is affected by the freestream Mach number, Reynolds number, and the pitch/yaw of the launch vehicle. Furthermore, even when the freestream conditions and vehicle attitude are constant, this terminal shock oscillates on the surface of the vehicle. The time-varying surface pressure associated with the terminal shock results in unsteady aerodynamic loads that may interact with vehicle structural dynamic modes and the guidance and control of the vehicle. Buffet testing of a 3-percent scale rigid buffet model of a launch vehicle cargo configuration with a tangent-ogive payload shroud was conducted in 2012 and in 2016. Initial buffet forcing functions (BFFs) utilized a coarse pressure sensor distribution on the vehicle surface in which a single longitudinal station with eight sensors observed the terminal shock environment at Mach 0.90. An examination of these circumferential pressures reveal large impulse-like pressure fluctuations and an asymmetry in pressure when the vehicle is at a nonzeroangle of attack that result in high BFFs. Revisions to the shock integration region were made based on computational fluid dynamics and shadowgraph video of shock motion to better represent the BFFs and reduce the high loads resulting from this environment. To more clearly understand this terminal shock environment, a second wind tunnel test was conducted with a dense distribution of 256 sensors at the terminal shock location. These sensor arrays presents a unique opportunity to observe the unsteady terminal shock environment and to characterize the impact of various integration schemes on the BFFs. This paper presents a summary of the development of BFFs for this terminal shock and a detailed analyses of shock region pressure coefficients, coherence, BFFs, shock location time histories, and power spectral density to help guide development of BFFs for other launch vehicle test and analysis programs.

Piatak, David J.↗
NASA NTRS2002

Durability Issues for the Protection of Materials from Atomic Oxygen Attack in Low Earth Orbit

Low Earth orbital atomic oxygen is capable of eroding most polymeric materials typically used on spacecraft. Solar array blankets, thermal control polymers, and carbon fiber matrix composites are readily oxidized to become thinner and less capable of supporting the loads imposed upon them. Protective coatings have been developed that are durable to atomic oxygen to prevent oxidative erosion of the underlying polymers. However, the details of the surface roughness, coating defect density, and coating configuration can play a significant role as to whether or not the coating provides long duration atomic oxygen protection. Identical coatings on different surface roughness surfaces can have drastically different durability results. Examples and analysis of the causes of resultant differences in atomic oxygen protection are presented. Implications based on in-space experiences, ground laboratory testing, and computational modeling indicate that thin film vacuum-deposited aluminum protective coatings offer much less atomic oxygen protection than sputter-deposited silicon dioxide coatings.

Banks, Bruce↗
DOE OSTI2021

Critical role of solvent-modulated hydrogen-binding strength in the catalytic hydrogenation of benzaldehyde on palladium

Protic and non-protic solvents influence the rate of benzaldehyde hydrogenation on Pd by one order of magnitude in the following sequence of increasing reactivity, dioxane < tetrahydrofuran < water < methanol. Despite these differences, the reaction pathway does not change; the majority of turnovers occurs via a Langmuir-Hinshelwood mechanism by stepwise addition of sorbed H to sorbed benzaldehyde, first to the carbonyl O and then to the C atom of the formyl group, forming benzyl alcohol. An analysis of the solvation energies shows that both ground and transition states are destabilized by the solvents compared to the same reaction at the gas-solid interface. The destabilization extent of the reacting organic substrates in both states are similar and, therefore, compensate each other, making the net kinetic effects inconsequential. In consequence, the marked reactivity differences arise solely from the differences in the surface stabilization of adsorbed hydrogen adatoms moderated by the four solvents-more weakly bound hydrogen adatoms, because of their solvation by solvent, exhibit higher rates of their hydrogenation reaction in kinetically relevant attack on the C atom.

37 INORGANIC, ORGANIC, PHYSICAL, AND ANALYTICAL CH↗
NASA NTRS1992

Improvements to the missile aerodynamic prediction code DEMON3

The computer program DEMON3 was developed for the aerodynamic analysis of nonconventional supersonic configurations comprising a body with noncircular cross section and up to two wing or fin sections. Within a wing or fin section, the lifting surfaces may be cruciform, triform, planar, or low profile layouts; the planforms of the lifting surfaces allow for breaks in sweep. The body and fin sections are modeled by triplet and constant u-velocity panels, respectively, accounting for mutual body-fin interference. Fin thickness effects are included for the use of supersonic planar source panels. One of the unique features of DEMON3 is the modeling of high angle of attack vortical effects associated with the lifting surfaces and the body. In addition, shock expansion and Newtonian pressure calculation methods can be optionally engaged. These two dimensional nonlinear methods are augmented by aerodynamic interference determined from the linear panel methods. Depending on the geometric details of the body, the DEMON3 program can be used to analyze nonconventional configurations at angles of attack up to 25 degrees for Mach numbers from 1.1 to 6. Calculative results and comparisons with experimental data demonstrate the capabilities of DEMON3. Limitations and deficiencies are listed.

Dillenius, Marnix F. E.↗
NASA NTRS1995

Evaluation of F/A-18A HARV inlet flow analysis with flight data

The F/A-18A aircraft has experienced engine stalls at high angles-of-attack and yaw flight conditions which were outside of its flight envelope. Future aircraft may be designed to operate routinely in this flight regime. Therefore, it is essential that an understanding of the inlet flow field at these flight conditions be obtained. Due to the complex interactions of the fuselage and inlet flow fields, a study of the flow within the inlet must also include external effects. Full Navier-Stokes (FNS) calculations on the F/A-18A High Alpha Research Vehicle (HARV) inlet for several angles-of-attack with sideslip and free stream Mach numbers have been obtained. The predicted forebody/fuselage surface static pressures agreed well with flight data. The surface static pressures along the inlet lip are in good agreement with the numerical predictions. The major departure in agreement is along the bottom of the lip at 30 deg and 60 deg angle-of-attack where a possible streamwise flow separation is not being predicted by the code. The circumferential pressure distributions at the engine face are in very good agreement with the numerical results. The variation in surface static pressure in the circumferential direction is very small with the exception of 60 angle-of-attack. Although the simulation does not include the effect of the engine, it appears that this omission has a second order effect on the circumferential pressure distribution. An examination of the unsteady flight test data base has shown that the secondary vortex migrates a significant distance with time. In fact, the extent of this migration increases with angle-of-attack with increasing levels of distortion. The effects of the engine on this vortex movement is unknown. This implies that the level of flow unsteadiness increases with increasing distortion. Since the computational results represent an asymptotic solution driven by steady boundary conditions, these numerical results may represent an arbitrary point in time. A comparison of the predicted total pressure contours with flight data indicates that the numerical results are within the excursion range of the unsteady data which is the best the calculations can attain unless an unsteady simulation is performed.

Smith, C. Frederic↗
NASA NTRS1999

X-33 (Rev-F) Aeroheating Results of Test 6770 in NASA Langley 20-Inch Mach 6 Air Tunnel

Aeroheating characteristics of the X-33 Rev-F configuration have been experimentally examined in the Langley 20-Inch Mach 6 Air Tunnel (Test 6770). Global surface heat transfer distributions, surface streamline patterns, and shock shapes were measured on a 0.013-scale model at Mach 6 in air. Parametric variations include angles-of-attack of 20-deg, 30-deg, and 40-deg; Reynolds numbers based on model length of 0.9 to 4.9 million; and body-flap deflections of 10-deg and 20-deg. The effects of discrete roughness elements on boundary layer transition, which included trip height, size, and location, both on and off the windward centerline, were investigated. This document is intended to serve as a quick release of preliminary data to the X-33 program; analysis is limited to observations of the experimental trends in order to expedite dissemination.

Berry, Scott A.↗
NASA NTRS1972

Experimental and theoretical aerodynamic heating and flow field analysis of a space shuttle orbiter

The results of work on the McDonnell Douglas delta wing orbiter are summarized and compared to other data and theories where applicable. The tests were conducted in the continuous hypersonic Tunnel B at Mach 8 and in the hypervelocity hotshot Tunnel F at Mach 10.5. Photographic data are used to illustrate the regions of peak heating on the lee-side and to show the vortical nature of the flow. Windward centerline data include measurements of shock angle, surface pressure, flow field pitot pressure and total temperature, and heat transfer rate coefficient distributions. Shock angles, surface pressures, and local Mach number distributions are compared with tangent cone theory over an angle of attack range from 10 to 60 deg. The measured heat transfer rate coefficient distributions are compared to both laminar and turbulent theories and boundary layer transition data are compared to the McDonnell Douglas criteria.

Matthews, R. K.↗
DOE OSTI2024

Open-air plasma-assisted deposition of organosilicon coating for corrosion protection of AZ91D Mg alloy

As a corrosion barrier layer, open-air plasma-assisted organosilicon coating was applied on AZ91D Mg alloy. Organosilicon-coated AZ91D samples, prepared from two different plasma coating processes, were used for corrosion evaluation based on multistep electrochemical and H 2 collection measurements in 3.5 wt% sodium chloride (NaCl) solution. Some coated AZ91D samples were characterized with and without corrosion exposure in NaCl solution using electron microscopies and X-ray chemical analysis techniques. Further, the results indicate that the organosilicon coatings from open-air plasma processes are effective to delay the initiation of corrosion attack and also reduce the corrosion rate for AZ91D substrate. The corrosion attack was considered to develop after permeation of NaCl solution through the coating layer.

36 MATERIALS SCIENCE↗
DOE OSTI2024

Attack Surface of Renewable Energy Technologies

This slide deck presents an overview of the threat landscape for renewable energy technologies. It highlights he growing penetration of renewable resources and potential impact of an attack. Standard architectures are shared to highlight where vulnerabilities may exist and attack paths to reach critical infrastructure. We discuss threat actors and attack paths. Several recent events impacting renewable energy assets or companies are explained.

14 SOLAR ENERGY↗
NASA NTRS2013

Analysis and Results from a Flush Airdata Sensing (FADS) System in Close Proximity to Firing Rocket Nozzles

This presentation presents information regarding the nose-cap flush airdata sensing (FADS) system on Orion's Pad Abort 1 (PA-1) vehicle. The purpose of the nose-cap FADS system was to test whether or not useful data could be obtained from a FADS system if it was placed in close proximity to firing rockets nozzles like the attitude control motor (ACM) nozzles on the PA-1 launch abort system (LAS). The nose-cap FADS systems use pressure measurements from a series of pressure ports which are arranged in a cruciform pattern and flush with the surface of the vehicle to estimate values of angle of attack, angle of side-slip, Mach number, impact pressure and free-stream static pressure.

Pad Abort 1↗
NASA NTRSDate not supplied

Separation and Transition on the ROTEX-T Cone-Flare

As part of NATO STO AVT-346 “Predicting Hypersonic Boundary-Layer Transition on Complex Geometries,” coordinated experimental and computational studies were conducted on the ROTEX-T, a cone-flare geometry used in a successful flight-test experiment. At the as-flown conditions, a separation bubble existed at the compression corner. Separation, reattachment, and the multifaceted linear instability paths leading this bubble to transition to turbulence are challenging to predict, but have significant impact on surface pressure and heat flux. High-resolution background-oriented schlieren and infrared thermography measurements were made in the AFOSR–Notre Dame Large Mach-6 Quiet Tunnel at freestream unit Reynolds numbers from 5.8×10(exp 6) to 12.2×10(exp 6) m-1and nominally zero angle of attack. High-speed self-aligned focusing schlieren, infrared thermography, and focused laser differential interferometry measurements were made in the AFRL Mach-6 Ludwieg Tube from 2.2×10(exp 6) to 24.7×10(exp 6) m-1and nominally zero degrees angle of attack. The surface heat-flux and Stanton number distributions were computed. Separation and reattachment locations, as well as the flow state at each, were determined from the combination of surface and off-wall measurements. The convective and global boundary-layer instabilities of the axisymmetric laminar flow at the experimental conditions were investigated computationally. Amplification of Mack’s first and second modes were observed to have logarithmic amplification factors between 5 to 7.5 at the separation location, depending on conditions. The flow was found to be globally unstable to stationary three-dimensional disturbances concentrated in the reattachment region. Previous analysis of the ROTEX-T flight data had not assessed reattachment location or the flow state upon reattachment. Thanks to the insights gained from the coordinated, on- and off-wall ground-test measurements, these evaluations have now been made. The separation location indicated by laminar simulations is consistently numerically predicted to be upstream of the experimentally observed location for a transitional separation bubble. The cause of this difference is understood to lie in the steady-state and axisymmetric assumptions made by both solvers employed to compute the basic states analyzed, as flow topology considerations assert that unsteady two-dimensional or axisymmetric separation bubbles are structurally unstable and will become three-dimensional. Computed laminar heating rates prior to separation agreed well with experiment; transitional heating rates after reattachment were between laminar and turbulent computations.

CFD Validation↗
NASA NTRSDate not supplied

Separation and Transition on the ROTEX-T Cone-Flare

As part of NATO STO AVT-346 “Predicting Hypersonic Boundary-Layer Transition on Complex Geometries,” coordinated experimental and computational studies were conducted on the ROTEX-T, a cone-flare geometry used in a successful flight-test experiment. At the as-flown conditions, a separation bubble existed at the compression corner. Separation, reattachment, and the multifaceted linear instability paths leading this bubble to transition to turbulence are challenging to predict, but have significant impact on surface pressure and heat flux. High-resolution background-oriented schlieren and infrared thermography measurements were made in the AFOSR–Notre Dame Large Mach-6 Quiet Tunnel at freestream unit Reynolds numbers from 5.8×10(exp 6) to 12.2×10(exp 6) m-1and nominally zero angle of attack. High-speed self-aligned focusing schlieren, infrared thermography, and focused laser differential interferometry measurements were made in the AFRL Mach-6 Ludwieg Tube from 2.2×10(exp 6) to 24.7×10(exp 6) m-1and nominally zero degrees angle of attack. The surface heat-flux and Stanton number distributions were computed. Separation and reattachment locations, as well as the flow state at each, were determined from the combination of surface and off-wall measurements. The convective and global boundary-layer instabilities of the axisymmetric laminar flow at the experimental conditions were investigated computationally. Amplification of Mack’s first and second modes were observed to have logarithmic amplification factors between 5 to 7.5 at the separation location, depending on conditions. The flow was found to be globally unstable to stationary three-dimensional disturbances concentrated in the reattachment region. Previous analysis of the ROTEX-T flight data had not assessed reattachment location or the flow state upon reattachment. Thanks to the insights gained from the coordinated, on- and off-wall ground-test measurements, these evaluations have now been made. The separation location indicated by laminar simulations is consistently numerically predicted to be upstream of the experimentally observed location for a transitional separation bubble. The cause of this difference is understood to lie in the steady-state and axisymmetric assumptions made by both solvers employed to compute the basic states analyzed, as flow topology considerations assert that unsteady two-dimensional or axisymmetric separation bubbles are structurally unstable and will become three-dimensional. Computed laminar heating rates prior to separation agreed well with experiment; transitional heating rates after reattachment were between laminar and turbulent computations.

CFD Validation↗
NASA NTRS1974

Results of tests OA26 and IA16 in the NASA/ARC 3.5-foot hypersonic wind tunnel on an 0.015-scale model (36-OTS) of the space shuttle configuration 140A/B to obtain pressures for venting analysis

Tests were conducted, from November 15 to December 4, 1973, to obtain surface pressure data on an 0.015-scale replica of the Space Shuttle Vehicle 4. Data were obtained at Mach numbers of 5.3, 7.4, and 10.3, to support the venting analysis for both launch and entry conditions. These tests were the final tests in a series covering a Mach number range from 0.6 to 10.3. The model was instrumented with pressure orifices in the vicinity of the cargo bay door hinge and parting lines, and on the side of the fuselage at the crew compartment, and below the orbital maneuvering system pods at the aft compartment. The model was tested at angles of attack and sideslip consistent with expected divergencies from the nominal trajectory.

Spangler, R. H.↗
NASA NTRS1992

Numerical investigation of the effects of icing on fixed and rotary wing aircraft

A 2-D multi-element airfoil code was modified to study the effects of icing on the aerodynamic characteristics of high lift systems. In each zone of the flow field, the solver numerically integrates the 2-D compressible Navier-Stokes equations using a time marching scheme. The surface pressure distribution is generated over a GAW 130 airfoil/flap combination for a flap setting of 25 degrees, and an angle of attack equal to 5 degrees, at a freestream Mach number equal to 0.3. A series of calculations were performed to determine the effects of small scale ice build up on the high lift characteristics of this arifoil/flap combination. The appendix summarizes this progress. Joint studies on correlation of a 3-D iced wing code with experimental data reviewed new measured laser Doppler velocimeter data in the separated region behind the leading edge ice shape. A version of the iced wing analysis using the Roe scheme was developed to evaluate the poor correlation between the computed and measured velocities in the separated region. Work on the extension of the wing-alone analysis to wing body configuration began with modifications to the 3-D iced wing analyses to accept externally generated grids and multi-block grids.

Sankar, L. N.↗
NASA NTRS2003

Active Control of F/A-18 Vertical Tail Buffeting using Piezoelectric Actuators

Vertical tail buffeting is a serious multidisciplinary problem that limits the performance of twin-tail fighter aircraft. The buffet problem occurs at high angles of attack when the vortical flow breaks down ahead of the vertical tails resulting in unsteady and unbalanced pressure loads on the vertical tails. This paper describes a multidisciplinary computational investigation for buffet load alleviation of full F/A-18 aircraft using distributed piezoelectric actuators. The inboard and outboard surfaces of the vertical tail are equipped with piezoelectric actuators to control the buffet responses in the first bending and torsion modes. The electrodynamics of the smart structure are expressed with a three-dimensional finite element model. A single-input-single-output controller is designed to drive the active piezoelectric actuators. High-fidelity multidisciplinary analysis modules for the fluid dynamics, structure dynamics, electrodynamics of the piezoelectric actuators, fluid-structure interfacing, and grid motion are integrated into a multidisciplinary computing environment that controls the temporal synchronization of the analysis modules. Peak values of the power spectral density of tail tip acceleration are reduced by as much as 22% in the first bending mode and by as much as 82% in the first torsion mode. RMS values of tip acceleration are reduced by as much as 12%.

Sheta, Essam F.↗
DOE OSTI2024

Attack Surface of Wind Energy Technologies in the United States [Slides]

This slide deck presents an overview of the threat landscape for wind energy technologies. It highlights unique cybersecurity considerations for wind, the growing penetration and potential impact of an attack. Standard architectures are shared to highlight where vulnerabilities may exist and attack paths to reach critical infrastructure. We discuss threat actors and attack paths. Several recent events impacting wind assets or wind companies are explained.

17 WIND ENERGY↗
DOE OSTI2021

Mechanism and kinetics of light alkane dehydrogenation and cracking over isolated Ga species in Ga/H-MFI

The objective of this study is to examine the mechanisms and kinetics of C2H6 dehydrogenation and n-C4H10 dehydrogenation and cracking over isolated Ga species in Ga/HMFI and to compare these results to those reported previously for C3H8 dehydrogenation and cracking. C2H6 dehydrogenation is found to be catalyzed by both [GaH]2+ and [GaH2]+ cations at similar turnover frequencies. Rate measurements over Ga/H-MFI containing predominantly [GaH2]+ cations reveal that C2H6 dehydrogenation rates exhibit a Langmuir.Hinshelwood dependence on C2H6 partial pressure at elevated temperatures (>730 K), consistent with the involvement of chemisorbed [C2H5.GaH]+ species. The reaction kinetics suggest that C2H6 dehydrogenation proceeds via heterolytic C-H cleavage of adsorbed C2H6 by [GaH2]+ cations to form H2 and [C2H5-GaH]+ species, which further decompose via β-hydride elimination to form C2H4. By contrast, C4H10 dehydrogenation and both terminal and central cracking are catalyzed exclusively by [GaH]2+ cations. All three reactions exhibit a Langmuir-Hinshelwood dependence on C4H10 partial pressure and are inhibited by H2. Ratios of dehydrogenation to cracking (total) and terminal to central cracking are approximately independent of C4H10 partial pressure consistent with the involvement of a common C4H10-derived surface intermediate. The observed reaction kinetics are consistent with an alkyl-mediated mechanism occurring over [GaH]2+, analogous to that reported previously for C3H8 dehydrogenation/cracking over Ga/H-MFI (Phadke, N. M.; et al. J. Am. Chem. Soc. 2019, 141, 1614-1627). The mechanism proceeds via facile, heterolytic dissociation of adsorbed C4H10 to form [C4H9-GaH]+-H+ cation pairs via methyl C-H-activated pathways. Dehydrogenation then proceeds via β-hydride elimination, respectively, forming butene, while terminal and central cracking proceed via C-H-activated H+ attack. Methylene activation was also considered but found to occur at a significantly lower rate. Theoretical analysis of the proposed reaction pathways leads to apparent activation enthalpies in good agreement with values extracted from the measured kinetics, thereby supporting the proposed pathways and the roles of [GaH]2+ and [GaH2]+ cations in the dehydrogenation and cracking of light alkanes on Ga/H-MFI.

Phadke, NM↗
NASA NTRS1992

Natural flow wing

The invention is a natural flow wing and a method for constructing the same. The method comprises contouring a three-dimensional upper surface and a three-dimensional lower surface of the natural flow wing independently of one another into a prescribed shape. Experimental data and theoretical analysis show that flow and pressure-loading over an upper surface of a wing tend to be conical about an apex of the wing, producing favorable and unfavorable regions of performance based on drag. The method reduces these unfavorable regions by shaping the upper surface such that the maximum thickness near a tip of the natural flow wing moves aft, thereby, contouring the wing to coincide more closely with the conical nature of the flow on the upper surface. Nearly constant compressive loading characterizes the flow field over a lower surface of the conventional wing. Magnitude of these compressive pressures on the lower surface depends on angle of attack and on a streamwise curvature of the lower surface of the wing and not on a cross-sectional spanwise curvature. The method, thereby, shapes the lower surface to create an area as large as possible with negative slopes. Any type of swept wing may be used to obtain the final, shaped geometry of the upper and lower surfaces of the natural flow wing.

Wood, Richard M.↗