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NASA NTRS1942

Study of Turning Performance of a Fighter-Type Airplane Particularly as Affected by Flaps and Increased Supercharging, Special Report

Results of a study to determine the effects on turning performance due to various assumed modifications to a typical Naval fighter airplane are presented. The modifications considered included flaps of various types, both part and full space, increased supercharging, and increased wing loading. The calculations indicated that near the low-speed end of the speed range, the turning performance, as defined by steady level turns at a given speed, would be improved to some extent by any of the flaps considered at altitudes up to about 25,000 feet. (If turning is not restricted to the conditions of no loss of speed or altitude, more rapid turning can, of course, be accomplished with the aid of flaps, regardless of altitude.) Fowler flaps and NACA slotted flaps appeared somewhat superior to split or perforated split flaps for maneuvering purposes, particularly if the flap position is not adjustable. Similarly, better turning performance should be realized with full-span than with part-span flaps. Turning performance over the lower half of the speed range would probably not be materially improved at any altitude by increased supercharging of the engine unless the propeller were redesigned to absorb the added power more effectively; with a suitable propeller the turning performance at high altitudes could probably be greatly improved with increased supercharging. A reduction in wing area with the aspect ratio held constant would result in impairment of turning performance over practically the entire speed range at all altitudes.

Wetmore, J. W.↗

NASA NTRS2009

Investigation of Orbital Debris Impacts on Shuttle Radiator Panels

This paper documents the data collected from two hypervelocity micro-meteoroid orbital debris (MMOD) impact events where the shuttle payload bay door radiator sandwich panel was completely perforated. Scanning Electron Microscope/Energy-Dispersive x-ray Spectroscopy (SEM/EDS) analysis of impact residue provided evidence to identify the source of each impact. Impact site features that indicate projectile directionality are discussed, along with hypervelocity impact testing on representative samples conducted to simulate the impact event. The paper provides results of a study of impact risks for the size of particles that caused the MMOD damage and the regions of the orbiter vehicle that would be vulnerable to an equivalent projectile

Hyde, James L.↗

NASA NTRS2009

Impedance Eduction in Ducts with Higher-Order Modes and Flow

An impedance eduction technique, previously validated for ducts with plane waves at the source and duct termination planes, has been extended to support higher-order modes at these locations. Inputs for this method are the acoustic pressures along the source and duct termination planes, and along a microphone array located in a wall either adjacent or opposite to the test liner. A second impedance eduction technique is then presented that eliminates the need for the microphone array. The integrity of both methods is tested using three sound sources, six Mach numbers, and six selected frequencies. Results are presented for both a hardwall and a test liner (with known impedance) consisting of a perforated plate bonded to a honeycomb core. The primary conclusion of the study is that the second method performs well in the presence of higher-order modes and flow. However, the first method performs poorly when most of the microphones are located near acoustic pressure nulls. The negative effects of the acoustic pressure nulls can be mitigated by a judicious choice of the mode structure in the sound source. The paper closes by using the first impedance eduction method to design a rectangular array of 32 microphones for accurate impedance eduction in the NASA LaRC Curved Duct Test Rig in the presence of expected measurement uncertainties, higher order modes, and mean flow.

Watson, Willie R.↗

NASA NTRS2009

Advanced Computational and Experimental Techniques for Nacelle Liner Performance Evaluation

The Curved Duct Test Rig (CDTR) has been developed to investigate sound propagation through a duct of size comparable to the aft bypass duct of typical aircraft engines. The axial dimension of the bypass duct is often curved and this geometric characteristic is captured in the CDTR. The semiannular bypass duct is simulated by a rectangular test section in which the height corresponds to the circumferential dimension and the width corresponds to the radial dimension. The liner samples are perforate over honeycomb core and are installed on the side walls of the test section. The top and bottom surfaces of the test section are acoustically rigid to simulate a hard wall bifurcation or pylon. A unique feature of the CDTR is the control system that generates sound incident on the liner test section in specific modes. Uniform air flow, at ambient temperature and flow speed Mach 0.275, is introduced through the duct. Experiments to investigate configuration effects such as curvature along the flow path on the acoustic performance of a sample liner are performed in the CDTR and reported in this paper. Combinations of treated and acoustically rigid side walls are investigated. The scattering of modes of the incident wave, both by the curvature and by the asymmetry of wall treatment, is demonstrated in the experimental results. The effect that mode scattering has on total acoustic effectiveness of the liner treatment is also shown. Comparisons of measured liner attenuation with numerical results predicted by an analytic model based on the parabolic approximation to the convected Helmholtz equation are reported. The spectra of attenuation produced by the analytic model are similar to experimental results for both walls treated, straight and curved flow path, with plane wave and higher order modes incident. The numerical model is used to define the optimized resistance and reactance of a liner that significantly improves liner attenuation in the frequency range 1900-2400 Hz. A liner impedance descriptor is used to determine the liner parameters that achieve the optimum impedance.

Gerhold, Carl H.↗

NASA NTRS1954

Permeability and Strength Measurements on Sintered, Porous, Hollow Turbine Blades Made by the American Electro Metal Corporation under Office of Naval Research Contract N-ONR-295 (01)

An experimental investigation was made to determine the permeability and strength characteristics of a number of sintered, porous, hollow turbine rotor blades and to determine the effectiveness of the blade fabrication method on permeability control. The test blades were fabricated by the American Electro Metal Corporation under a contract with the Office of Naval Research, Department of the Navy, and were submitted to the NACA for testing. Of the 22 test blades submitted, ten were sintered but not coined, five were sintered and coined, and seven were sintered and not coined but contained perforated reinforcements integral with the blade shells. Representative samples of each group of blades were tested. Large variations in permeability in both chordwise and spanwise directions were found. Local deviations as large as 155 to -85 percent from prescribed values were found in chordwise permeability. Only one blade, an uncoined one, had a chordwise permeability variations which reasonably approached that specified. Even for this blade, local deviations exceeded 10 percent. Spanwise permeability, specified to be held constant, varied as much as 50 percent from root to tip for both an uncoined and a coined blade. Previous NACA analyses have shown that in order to maintain proper control of blade wall temperatures, permeability variations must not exceed plus or minus 10 percent. Satisfactory control of permeability in either the chordwise or the spanwise direction was not achieved in the blades tested. Spin tests made at room temperature for six blades revealed the highest material rupture strength to be 8926 pounds per square inch. This value is about one third the strength required for rotor blades in present-day turbojet engines. The lowest value of blade strength was 1436 pounds per square inch.

Richards, Hadley T.↗

NASA NTRS2009

Performance of Whipple Shields at Impact Velocities above 9 km/s

Whipple shields were first proposed as a means of protecting spacecraft from the impact of micrometeoroids in 1947 [1] and are currently in use as micrometeoroid and orbital debris shields on modern spacecraft. In the intervening years, the function of the thin bumper used to shatter or melt threatening particles has been augmented and enhanced by the use of various types and configurations of intermediate layers of various materials. All shield designs serve to minimize the threat of a spall failure or perforation of the main wall of the spacecraft as a result of the impact of the fragments. With increasing use of Whipple shields, various ballistic limit equations (BLEs) for guiding the design and estimating the performance of shield systems have been developed. Perhaps the best known and most used are the "new" modified Cour-Palais (Christiansen) equations [2]. These equations address the three phases of impact: (1) ballistic (<3 km/s), where the projectile is moving too slowly to fragment and essentially penetrates as an intact projectile; (2) shatter (3 to 7 km/s), where the projectile fragments at impact and forms an expanding cloud of debris fragments; and (3) melt/vaporization (>7 km/s), where the projectile melts or vaporizes at impact. The performance of Whipple shields and the adequacy of the BLEs have been examined for the first two phases using the results of impact tests obtained from two-stage, light-gas gun test firings. Shield performance and the adequacy of the BLEs has not been evaluated in the melt/vaporization phase until now because of the limitations of launchers used to accelerate projectiles with controlled properties to velocities above 7.5 km/s. A three-stage, light-gas gun, developed at the University of Dayton Research Institute (UDRI) [3], is capable of launching small, aluminum spheres to velocities above 9 km/s. This launcher was used to evaluate the ballistic performance of two Whipple shield systems, various thermal protection system materials, and other spacecraft-related materials to the impact of 1.6-mm- to 2.6-mm-diameter, 2017-T4 aluminum spheres at impact velocities ranging from 8.91 km/s to 9.28 km/s. Test results, details of the shield systems, and nominal ballistic limits for the two Whipple shields are shown in Figures 1 and 2.

Christiansen, Eric L.↗

NASA NTRS2009

Ultra-Large Solar Sail

UltraSail is a next-generation ultra-large (km2 class) sail system. Analysis of the launch, deployment, stabilization, and control of these sails shows that high-payload-mass fractions for interplanetary and deep-space missions are possible. UltraSail combines propulsion and control systems developed for formation-flying microsatellites with a solar sail architecture to achieve controllable sail areas approaching 1 km2. Electrically conductive CP-1 polyimide film results in sail subsystem area densities as low as 5 g/m2. UltraSail produces thrust levels many times those of ion thrusters used for comparable deep-space missions. The primary innovation involves the near-elimination of sail-supporting structures by attaching each blade tip to a formation- flying microsatellite, which deploys the sail and then articulates the sail to provide attitude control, including spin stabilization and precession of the spin axis. These microsatellite tips are controlled by microthrusters for sail-film deployment and mission operations. UltraSail also avoids the problems inherent in folded sail film, namely stressing, yielding, or perforating, by storing the film in a roll for launch and deployment. A 5-km long by 2 micrometer thick film roll on a mandrel with a 1 m circumference (32 cm diameter) has a stored thickness of 5 cm. A 5 m-long mandrel can store a film area of 25,000 m2, and a four-blade system has an area of 0.1 sq km.

Burton, Rodney↗

NASA NTRS2010

Characteristics of Whipple Shield Performance in the Shatter Regime

Ballistic limit equations define the failure of metallic Whipple shields in three parts: low velocity, shatter, and hypervelocity. Failure limits in the shatter regime are based on a linear interpolation between the onset of projectile fragmentation, and impulsive rupture of the shield rear wall. A series of hypervelocity impact tests have been performed on aluminum alloy Whipple shields to investigate failure mechanisms and performance limits in the shatter regime. Test results demonstrated a more rapid increase in performance than predicted by the latest iteration of the JSC Whipple shield BLE following the onset of projectile fragmentation. This increase in performance was found to level out between 4.0-5.0 km/s, with a subsequent decrease in performance for velocities up to 6.0 km/s. For a detached spall failure criterion, the failure limit was found to continually decrease up to a velocity of 7.0 km/s, substantially varying from the BLE, while for perforation-based failure an increase in performance was observed. An existing phenomenological ballistic limit curve was found to provide a more accurate reproduction of shield behavior that the BLE, however a number of underlying assumptions such as the occurrence of complete projectile fragmentation and the effect on performance of incipient projectile melt were found to be inaccurate. A cratering relationship based on the largest residual fragment size has been derived for application at velocities between 3.0-4.0 km/s, and was shown to accurately reproduce the trends of the experimental data. Further investigation is required to allow a full analytical description of shatter regime performance for metallic Whipple shields.

Ryan, S.↗

NASA NTRS2010

Nondestructive Evaluation Methods for the Ares I Common Bulkhead

A large scale bonding demonstration test article was fabricated to prove out manufacturing techniques for the current design of the NASA Ares I Upper Stage common bulkhead. The common bulkhead serves as the single interface between the liquid hydrogen and liquid oxygen portions of the Upper Stage propellant tank. The bulkhead consists of spin-formed aluminum domes friction stir welded to Y-rings and bonded to a perforated phenolic honeycomb core. Nondestructive evaluation methods are being developed for assessing core integrity and the core-to-dome bond line of the common bulkhead. Detection of manufacturing defects such as delaminations between the core and face sheets as well as service life defects such as crushed or sheared core resulting from impact loading are all of interest. The focus of this work will be on the application of thermographic, shearographic, and phased array ultrasonic methods to the bonding demonstration article as well as various smaller test panels featuring design specific defect types and geometric features.

Walker, James↗

NASA NTRS2010

Experimental Investigation and Analysis of an Annular Pogo Accumulator

An experimental investigation was conducted on a scaled annular pogo accumulator for the Ares I Upper Stage. The test article was representative of the LO2 feedline and preliminary accumulator design, and included multiple designs of a perforated ring connecting the accumulator to the core feedline flow. The system was pulse tested in water over a range of pulse frequency and flow rates. Time dependent measurements of pressure at various locations in the test article were used to extract system compliance, inertance, and resistance. Results indicate a significant deviation from standard orifice flow theory and suggest a strong dependence on feedline average velocity. In addition, several CFD analyses were conducted to investigate the details of the time variant flow field. Both two-dimensional and three-dimensional simulations were performed with time varying boundary conditions used to represent system pulsing. The CFD results compared well with the sub-scale results and demonstrated the influence of feedline average velocity on the flow into and out of the accumulator

Peugeot, John↗

NASA NTRS2010

Validation of a New Procedure for Impedance Eduction in Flow

A new impedance eduction procedure is validated by comparing the educed impedance spectrum to that of an older but well-tested eduction procedure. The older procedure requires the installation of a microphone array in the liner test section but the new procedure removes this requirement. A 12.7-mm stainless steel plate and a conventional liner consisting of a perforated plate bonded to a honeycomb core are tested. Test data is acquired from a grazing flow, impedance tube facility for a range of source frequencies and mean flow Mach numbers for which only plane waves are cut on. For the stainless steel plate, the educed admittance spectrum using the new procedure shows an improvement over that of the old procedure. This improvement shows up primarily in the educed conductance spectrum. Both eduction procedures show discrepancies in educed admittance in the mid-frequency range. Indications are that this discrepancy is triggered by an inconsistency between the measured eduction data (that contains boundary layer effects) and the two eduction models (for which the boundary layer is neglected). For the conventional liner, both eduction procedures are in very good agreement with each other. Small discrepancies occur for one or two frequencies in the mid-frequency range and for frequencies beyond the cut on frequency of higher-order duct modes. This discrepancy in the midfrequency range occurs because an automated optimizer is used to educe the impedance and the objective function used by the optimizer is extremely flat and therefore sensitive to initial starting values. The discrepancies at frequencies beyond the cut on frequency of higher order duct modes are due to the assumption of only plane waves in the impedance eduction model, although higher order modes are propagating in the impedance tube facility.

Watson, W. R.↗

NASA NTRS2010

Effects of Flow Profile on Educed Acoustic Liner Impedance

This paper presents results of an investigation of the effects of shear flow profile on impedance eduction processes employed at NASA Langley. Uniform and 1-D shear-flow propagation models are used to educe the acoustic impedance of three test liners based on aeroacoustic data acquired in the Langley Grazing Flow Impedance Tube, at source levels of 130, 140 and 150 dB, and at centerline Mach numbers of 0.0, 0.3 and 0.5. A ceramic tubular, calibration liner is used to evaluate the propagation models, as this liner is expected to be insensitive to SPL, grazing flow Mach number, and flow profile effects. The propagation models are then used to investigate the effects of shear flow profile on acoustic impedances educed for two conventional perforate-over-honeycomb liners. Results achieved with the uniform-flow models follow expected trends, but those educed with the 1-D shear-flow model do not, even for the calibration liner. However, when the flow profile used with the shear-flow model is varied to increase the Mach number gradient near the wall, results computed with the shear-flow model are well matched to those achieved with the uniform-flow model. This indicates the effects of flow profile on educed acoustic liner impedance are small, but more detailed investigations of the flow field throughout the duct are needed to better understand these effects.

Jones, Michael G.↗

NASA NTRS2009

Explicit Finite Element Modeling of Multilayer Composite Fabric for Gas Turbine Engine Containment Systems, Phase II: Ballistic Impact Testing - Part 2

This report summarizes the ballistic impact testing that was conducted to provide validation data for the development of numerical models of blade-out events in fabric containment systems. The ballistic impact response of two different fiber materials - Kevlar(TradeName) 49 and Zylon(TradeName) AS (as spun) was studied by firing metal projectiles into dry woven fabric specimens using a gas gun. The shape, mass, orientation, and velocity of the projectile were varied and recorded. In most cases, the tests were designed so the projectile would perforate the specimen, allowing measurement of the energy absorbed by the fabric. The results for both Zylon and Kevlar presented here represent a useful set of data for the purposes of establishing and validating numerical models to predict the response of fabrics under conditions that simulate those of a jet engine blade-release situation. In addition, some useful empirical observations were made regarding the effects of projectile orientation and the relative performance of the different fabric materials.

Zylon↗

NASA NTRS2001

Hypervelocity Impact Testing of International Space Station Meteoroid/Orbital Debris Shielding Using an Inhibited Shaped Charge Launcher

Engineers at the NASA Johnson Space Center have conducted hypervelocity impact (HVI) performance evaluations of spacecraft meteoroid and orbital debris (M/OD) shields at velocities in excess of 7 km/s. The inhibited shaped charge launcher (ISCL), developed by the Southwest Research Institute, launches hollow, circular, cylindrical jet tips to approximately 11 km/s. Since traditional M/OD shield ballistic limit performance is defined as the diameter of sphere required to just perforate or spall a spacecraft pressure wall, engineers must decide how to compare ISCL derived data with those of the spherical impactor data set. Knowing the mass of the ISCL impactor, an equivalent sphere diameter may be calculated. This approach is conservative since ISCL jet tips are more damaging than equal mass spheres. A total of 12 tests were recently conducted at the Southwest Research Institute (SWRI) on International Space Station M/OD shields. Results of these tests are presented and compared to existing ballistic limit equations. Modification of these equations is suggested based on the results.

Kerr, Justin H.↗

NASA NTRS2010

Recent Shuttle Post Flight MMOD Inspection Highlights

Four significant micro-meteoroid orbital debris (MMOD) impact events on the shuttle are documented: two perforations of the payload bay door radiator sandwich panel along with craters in a crew module window and a wing leading edge reinforced carbon-carbon panel. Results from Scanning Electron Microscope/Energy-Dispersive x-ray Spectroscopy (SEM/EDS) analysis of impact residue are provided in an attempt to identify the source of each impact. Particle sizes that could produce the observed damage are discussed, along with hypervelocity impact testing on representative samples conducted to simulate the impact event.

Hyde, James L.↗

NASA NTRS2010

Evaluation of Sandwich Structure Bonding In Out-of-Autoclave Processing

The out-of-autoclave-vacuum-bag-only (OOA-VBO) process is low in capital expenditures compared to the traditional autoclave, however, the material challenges for OOA-VBO workable material systems are high. Presently there are few such aerospace grade prepreg materials available commercially. In this study, we evaluated processing and properties of honeycomb sandwich structure (HC/SS) panels fabricated by co-curing composite face sheet with adhesives by the OOA-VBO process in an oven. The prepreg materials were IM7/MTM 45-1 and T40-800B/5320. Adhesives studied were AF-555M, XMTA-241/PM15, FM-309-1M and FM-300K. Aluminum H/C cores with and without perforations were included. It was found that adhesives in IM7/MTM 45-1/AF-555M, T40-800B/5320/FM 309-1M and T40-800B/5320/FM-300K panels all foamed but yielded high flatwise tensile (FWT) strength values above 8,275 kPA (1,200 psi). IM7/MTM 45-1/XMTA-241/PM15 did not foam, yet yielded a low FWT strength. SEM photomicrographs revealed that the origin of this low strength was poor adhesion in the interfaces between the adhesive and face sheet composite due to poor wetting associated with the high initial viscosity of the XMTA-241/PM15 adhesive.

Hou, Tan-Hung↗

NASA NTRS2010

Modeling Neutral Densities Downstream of a Gridded Ion Thruster

The details of a model for determining the neutral density downstream of a gridded ion thruster are presented. An investigation of the possible sources of neutrals emanating from and surrounding a NEXT ion thruster determined that the most significant contributors to the downstream neutral density include discharge chamber neutrals escaping through the perforated grids, neutrals escaping from the neutralizer, and vacuum facility background neutrals. For the neutral flux through the grids, near- and far-field equations are presented for rigorously determining the neutral density downstream of a cylindrical aperture. These equations are integrated into a spherically-domed convex grid geometry with a hexagonal array of apertures for determining neutral densities downstream of the ion thruster grids. The neutrals escaping from an off-center neutralizer are also modeled assuming diffuse neutral emission from the neutralizer keeper orifice. Finally, the effect of the surrounding vacuum facility neutrals is included and assumed to be constant. The model is used to predict the neutral density downstream of a NEXT ion thruster with and without neutralizer flow and a vacuum facility background pressure. The impacts of past simplifying assumptions for predicting downstream neutral densities are also examined for a NEXT ion thruster.

Soulas, George C.↗

NASA NTRS2011

Apparatus and methods for filtering granular solid material

Apparatuses for screening granular solid particulate material include a generally planar first screen and a second screen. A plurality of apertures extends through the first screen. At least a portion of the second screen is oriented at an angle to the first screen, and apertures extend through a perforated region of the second screen. The second screen includes at least one region configured to prevent at least some particles of solid material from passing through the second screen.

Backes, Douglas J.↗