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At least 289 records · Page 16

Observations on the structure of an edge-tone flowfield

The unsteady flowfield in the region in which a rectangular jet interacts with the leading edge of a 20-deg wedge is investigated experimentally. Edge-tone flowfields at Mach 0.8 in the blowdown facility at Stanford University are recorded by phase-locked Schlieren photography using the apparatus and techniques described by Krothapalli and Horne (1984); detailed surface-pressure and velocity measurements are obtained at Mach 0.23 in the NASA Ames anechoic-chamber facility and presented graphically. It is shown that the jet-edge interaction produces a secondary vortex at the leading edge when an established shear layer or primary vortex is present, and that the primary and secondary vortices then evolve from a counterrotating vortex pair to a large single vortex which migrates along the wedge surface (in agreement with the observations of Rockwell et al., 1986). From a comparison with data on airfoil-trailing-edge flows (Satyanarayana, 1977) it is inferred that an instantaneous zero pressure differential is present at the leading edge of the wedge.

Krothapalli, A.↗

Investigation of leading-edge flap performance on delta and double-delta wings at supersonic speeds

An investigation of the aerodynamic performance of leading-edge flaps on three clipped delta and three clipped double-delta wing planforms with aspect ratios of 1.75, 2.11, and 2.50 was conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.90, and 2.16. A primary set of fullspan leading-edge flaps with similar root and tip chords were investigated on each wing, and several alternate flap planforms were investigated on the aspect-ratio-1.75 wings. All leading-edge flap geometries were effective in reducing the drag at lifting conditions over the range of wing aspect ratios and Mach numbers tested. Application of a primary flap resulted in better flap performance with the double-delta planform than with the delta planform. The primary flap geometry generally yielded better performance than the alternate flap geometries tested. Trim drag due to flap-induced pitching moments was found to reduce the leading-edge flap performance more for the delta planform than for the double-delta planform. Flow-visualization techniques showed that leading-edge flap deflection reduces crossflow shock-induced separation effects. Finally, it was found that modified linear theory consistently predicts only the effects of leading-edge flap deflection as related to pitching moment and lift trends.

Covell, Peter F.↗

Navier Stokes computation of the flow field over delta wings with spanwise leading edge blowing

The concept of spanwise leading edge blowing, a means of controlling the position and strength of leading edge vortices, is analyzed by numerical solutions of the three-dimensional Thin-Layer Navier Stokes equations. The leading edge jet is simulated by defining a permeable boundary, corresponding to the jet slot, where suitable boundary conditions are implemented. Numerical results agree favorably with experimental measurements. It is found that the use of spanwise leading edge blowing not only magnifies the size and strength of the leading edge vortices, but also moves the vortex cores outboard and upward. As a result, the increase in lift comes primarily from the greater nonlinear vortex lift. The presence of the leading edge jet stream displaces the flow outboard, thereby increasing the effective aspect ratio of the delta wing. However, blowing causes earlier vortex breakdown, thus decreasing the stall angle.

Yeh, David T.↗

The effects of ion gun beam voltage on the electrical characteristics of NbCN/PbBi edge junctions

The authors have succeeded in fabricating high-quality submicron NbCN edge junctions using a technique which is commonly used to make Nb edge junctions. A modified commercial ion gun was used to cut an edge in SiO2/NbCN films partially covered with photoresist. An insulating barrier was then formed on the exposed edge by reactive ion beam oxidation, and a counterelectrode of PbBi was deposited. The electrical quality of the resulting junctions was found to be strongly influenced by the ion beam acceleration voltages used to cut the edge and to oxidize it. For low ion beam voltages, the junction quality parameter was as high as Vm = 55 mV (measured at 3 mV), but higher ion beam voltages yielded strikingly poorer quality junctions. In light of the small coherence length of NbN, the dependence of the electrical characteristics on ion beam voltage is presumably due to mechanical damage of the NbCN surface. In contrast, for similar ion beam voltages, no such dependence was found for Nb edge junctions.

Lichtenberger, A. W.↗

Wind-tunnel investigations of wings with serrated sharp trailing edges

Exploratory wind-tunnel force measurements are presented for two wing geometries with small-scale planar and nonplanar serrated trailing-edge devices (chord-Reynolds numbers ranged from 1.0 to 3.7 million). The planar serrated trailing-edge extensions reduced the drag at conditions when trailing-edge separation occurred at low angles of attack. The introduction of serrations reduced or eliminated the drag penalty, due to the small (1-2 percent of the chord length) nonplanar trailing-edge flaps, while maintaining the effects of increase in camber. The presence of streamwise vortices immediately downstream of the serrated trailing edges is believed to have favorably affected the boundary-layer flow approaching the trailing edge and the near-wake development, resulting in reduced pressure (form) drag.

Vijgen, P. M. H. W.↗

Wind-tunnel investigations of wings with serrated sharp trailing edges

Exploratory wind-tunnel force measurements are presented for two wing geometries with small-scale planar and nonplanar serrated trailing-edge devices (chord-Reynolds numbers ranged from 1.0 - 3.7 x 10 to the 6th). The planar serrated trailing-edge extensions reduced the drag at conditions when trailing-edge separation occurred at low angles of attack. The introduction of serrations reduced or eliminated the drag penalty due to the small (1-2 percent of the chord length) nonplanar trailing-edge flaps, while maintaining the effects of increase in camber. The presence of streamwise vortices immediately downstream of the serrated trailing edges is believed to have favorably affected the boundary-layer flow approaching the trailing edge and the near-wake development, resulting in reduced pressure drag.

Vijgen, P. M. H. W.↗

Thermostructural applications of heat pipes for cooling leading edges of high-speed aerospace vehicles

Heat pipes have been considered for use on wing leading edge for over 20 years. Early concepts envisioned metal heat pipes cooling a metallic leading edge. Several superalloy/sodium heat pipes were fabricated and successfully tested for wing leading edge cooling. Results of radiant heat and aerothermal testing indicate the feasibility of using heat pipes to cool the stagnation region of shuttle-type space transportation systems. The test model withstood a total seven radiant heating tests, eight aerothermal tests, and twenty-seven supplemental radiant heating tests. Cold-wall heating rates ranged from 21 to 57 Btu/sq ft-s and maximum operating temperatures ranged from 1090 to 1520 F. Follow-on studies investigated the application of heat pipes to cool the stagnation regions of single-stage-to-orbit and advanced shuttle vehicles. Results of those studies indicate that a 'D-shaped' structural design can reduce the mass of the heat-pipe concept by over 44 percent compared to a circular heat-pipe geometry. Simple analytical models for heat-pipe startup from the frozen state (working fluid initially frozen) were adequate to approximate transient, startup, and steady-state heat-pipe performance. Improvement in analysis methods has resulted in the development of a finite-element analysis technique to predict heat-pipe startup from the frozen state. However, current requirements of light-weight design and reliability suggest that metallic heat pipes embedded in a refractory composite material should be used. This concept is the concept presently being evaluated for NASP. A refractory-composite/heat-pipe-cooled wing leading edge is currently being considered for the National Aero-Space Plane (NASP). This concept uses high-temperature refractory-metal/lithium heat pipes embedded within a refractory-composite structure and is significantly lighter than an actively cooled wing leading edge because it eliminates the need for active cooling during ascent and descent. Since the NASP vehicle uses cryogenic hydrogen to cool structural components and then burns this fuel in the combustor, hydrogen necessary for descent cooling only, when the vehicle is unpowered, is considered to be a weight penalty. Details of the design of the refractory-composite/heat-pipe-cooled wing leading edge are currently being investigated. Issues such as thermal contact resistance and thermal stress are also being investigated.

Camarda, Charles J.↗

Search for partial systemic Lyman edges in nearby quasars

IUE archival spectra of nearby quasars with redshifts in the range 0.4-2.3 were searched for changes in the continuum at the systemic Lyman edge positions. Thirty-nine percent of the sample do not show any discontinuities. Twenty-nine percent of the sample show complete Lyman edges at redshifts less than the quasar redshift. These objects have associated narrow absorption lines; hence the intervening gas responsible for the absorption is extrinsic to the quasar and its environment. Twenty-two percent of the sample show Lyman edges at the quasar redshift with associated narrow absorption lines. In these objects the gas is associated with the AGN or its host galaxy. Ten percent of the sample have Lyman edges with no known associated absorption lines. If Lyman edge discontinuities are due to optically thick, geometrically thin accretion disks, which are randomly oriented to the line of sight, these results indicate that the Lyman edge discontinuity, either in emission or absorption, has to be not more than 15 percent in most quasars.

Koratkar, A. P.↗

Effect of leading- and trailing-edge flaps on clipped delta wings with and without wing camber at supersonic speeds

An experimental investigation of the aerodynamic characteristics of thin, moderately swept fighter wings has been conducted to evaluate the effect of camber and twist on the effectiveness of leading- and trailing-edge flaps at supersonic speeds in the Langley Unitary Plan Wind Tunnel. The study geometry consisted of a generic fuselage with camber typical of advanced fighter designs without inlets, canopy, or vertical tail. The model was tested with two wing configurations an uncambered (flat) wing and a cambered and twisted wing. Each wing had an identical clipped delta planform with an inboard leading edge swept back 65 deg and an outboard leading edge swept back 50 deg. The trailing edge was swept forward 25 deg. The leading-edge flaps were deflected 4 deg to 15 deg, and the trailing-edge flaps were deflected from -30 deg to 10 deg. Longitudinal force and moment data were obtained at Mach numbers of 1.60, 1.80, 2.00, and 2.16 for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.16 x 10(exp 6) per foot and for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.0 x 10(exp 6) per foot. Vapor screen, tuft, and oil flow visualization data are also included.

Hernandez, Gloria↗

Hypersonic Engine Leading Edge Experiments in a High Heat Flux, Supersonic Flow Environment

A major concern in advancing the state-of-the-art technologies for hypersonic vehicles is the development of an aeropropulsion system capable of withstanding the sustained high thermal loads expected during hypersonic flight. Three aerothermal load related concerns are the boundary layer transition from laminar to turbulent flow, articulating panel seals in high temperature environments, and strut (or cowl) leading edges with shock-on-shock interactions. A multidisciplinary approach is required to address these technical concerns. A hydrogen/oxygen rocket engine heat source has been developed at the NASA Lewis Research Center as one element in a series of facilities at national laboratories designed to experimentally evaluate the heat transfer and structural response of the strut (or cowl) leading edge. A recent experimental program conducted in this facility is discussed and related to cooling technology capability. The specific objective of the experiment discussed is to evaluate the erosion and oxidation characteristics of a coating on a cowl leading edge (or strut leading edge) in a supersonic, high heat flux environment. Heat transfer analyses of a similar leading edge concept cooled with gaseous hydrogen is included to demonstrate the complexity of the problem resulting from plastic deformation of the structures. Macro-photographic data from a coated leading edge model show progressive degradation over several thermal cycles at aerothermal conditions representative of high Mach number flight.

Gladden, Herbert J.↗

A test of a vortex method for the computation of flap side edge noise

Upon approach to landing, a major source location of airframe noise occurs at the side edges of the part span, trailing edge flaps. In the vicinity of these flaps, a complex arrangement of spanwise flow with primary and secondary tip vortices may form. Each of these vortices is observed to become fully three-dimensional. In the present study, a numerical model is developed to investigate the noise radiated from the side edge of a flap. The inherent three-dimensionality of this flow forces us to carefully consider a numerical scheme which will be both accurate in its prediction of the flow acoustics and also computationally efficient. Vortex methods have offered a fast and efficient means of simulating many two and three-dimensional, vortex dominated flows. In vortex methods, the time development of the flow is tracked by following exclusively the vorticity containing regions. Through the Biot-Savart law, knowledge of the vorticity field enables one to obtain flow quantities at any desired location during the flow evolution. In the present study, a numerical procedure has been developed which incorporates the Lagrangian approach of vortex methods into a calculation for the noise radiated by a flow-surface interaction. In particular, the noise generated by a vortex in the presence of a flat half plane is considered. This problem serves as a basic model of flap edge flow. It also permits the direct comparison between our computed results and previous acoustic analyses performed for this problem. In our numerical simulations, the mean flow is represented by the complex potential W(z) = Aiz(exp l/2), which is obtained through conformal mapping techniques. The magnitude of the mean flow is controlled by the parameter A. This mean flow has been used in the acoustic analysis by Hardin and is considered a reasonable model of the flow field in the vicinity of the edge and away from the leading and trailing edges of the flap. To represent the primary vortex which occurs near the flap, a point vortex is introduced just below the flat half plane. Using a technique from panel methods, boundary conditions on the flap surface are satisfied by the introduction of a row of stationary point vortices along the extent of the flap. At each time step in the calculation, the strength of these vortices is chosen to eliminate the normal velocity at intermediary collocation points. The time development of the overall flow field is then tracked using standard techniques from vortex methods. Vortex trajectories obtained through this computation are in good agreement with those predicted by the analytical solution given by Hardin, thus verifying the viability of this procedure for more complex flow arrangements. For the flow acoustics, the Ffowcs Williams-Hawkings equation is numerically integrated. This equation supplies the far field acoustic pressure based upon pressures occurring along the flap surface. With our vortex method solution, surface pressures may be obtained with exceptional resolution. The Ffowcs Williams-Hawkings equation is integrated using a spatially fourth order accurate Simpson's rule. Rational function interpolation is used to obtain the surface pressures at the appropriate retarded times. Comparisons between our numerical results for the acoustic pressure and those predicted by the Hardin analysis have been made. Preliminary results indicate the need for an improved integration technique. In the future, the numerical procedure developed in this study will be applied to the case of a rectangular flap of finite thickness and ultimately modified for application to the fully three-dimensional problem.

Martin, James E.↗

Effects of Tangential Edge Constraints on the Postbuckling Behavior of Flat and Curved Panels Subjected to Thermal and Mechanical Loads

A parametric study of the effects of tangential edge constraints on the postbuckling response of flat and shallow curved panels subjected to thermal and mechanical loads is presented. The mechanical loads investigated are uniform compressive edge loads and transverse lateral pressure. The temperature fields considered are associated with spatially nonuniform heating over the panels, and a linear through-the-thickness temperature gradient. The structural model is based on a higher-order transverse-shear-deformation theory of shallow shells that incorporates the effects of geometric nonlinearities, initial geometric imperfections, and tangential edge motion constraints. Results are presented for three-layer sandwich panels made from transversely isotropic materials. Simply supported panels are considered in which the tangential motion of the unloaded edges is either unrestrained, partially restrained, or fully restrained. These results focus on the effects of the tangential edge restraint on the postbuckling response. The results of this study indicate that tangentially restraining the edges of a curved panel can make the panel insensitive to initial geometric imperfections in some cases.

Lin, W.↗

A Thermostructural Analysis of a Diboride Composite Leading Edge

In an effort to support the design of zirconium diboride composite leading edges for hypersonic vehicles, a finite element model (FEM) of a prototype leading edge was created and finite element analysis (FEA) was employed to assess its thermal and structural response to aerothermal boundary conditions. Unidirectional material properties for the structural components of the leading edge, a continuous fiber reinforced diboride composite, were computed with COSTAR. These properties agree well with those experimentally measured. To verify the analytical approach taken with COSMOS/M, an independent FEA of one of the leading edge assembly components was also done with COSTAR. Good agreement was obtained between the two codes. Both showed that a unidirectional lay-up had the best margin of safety for a simple loading case. Both located the maximum stress in the same region and ply. The magnitudes agreed within 4 percent. Trajectory based aerothermal heating was then applied to the leading edge assembly FEM created with COSMOS/M to determine steady state temperature response, displacement, stresses, and contact forces due to thermal expansion and thermal strains. Results show that the leading edge stagnation line temperature reached 4700 F. The maximum computed failure index for the laminated composite components peaks at 4.2, and is located at the bolt flange in layer 2 of the side bracket. The temperature gradient in the tip causes a compressive stress of 279 ksi along its width and substantial tensile stresses within its depth.

Kowalski, Tom↗

Turbulence Measurements on a Flap-Edge Model

Turbulence measurements have been made on a flap-edge and leading-edge slat model using hot-wire anemometry, and, later, particle image velocimetry. The properties of hot-wire anemometry were studied using facilities at NASA Ames Research Center. Hot-film probes were used because of their durability, but cross-films were limited by non-linear end effects. As a warm-up exercise, hot-film probes were used to measure velocities in the farfield wake of a cylinder with an airfoil in the near-field wake. The airfoil reduced the drag coefficient of the system by 10%. A single-wire hot-film probe was used to measure velocity profiles over the top of a NACA 63(sub 2)-215 Mod. B wing with a Fowler flap and leading,-edge slat. Results showed the size of slat wake was dependent upon the slat deflection angle. Velocity increased through the slat gap with increased deflection. The acoustically modified slat decreased the chance of separation. Measurements were taken at the flap edge with a single hot-film. Trends in the data indicate velocity and turbulence levels increase at the flap edge. The acoustically modified flap modifies the mean flow near the flap edge. Correlations were made between the hot-film signal and the unsteady pressure transducers on the wing which were published in a NASA CDTM. The principles of Particle Image Velocimetry (PIV) were studied at Florida State University. Spectral PIV was used to measure the spectra of a subsonic jet. Measured frequencies were close to the predicted frequency of jet shedding. Spectral PIV will be used to measure the spectra of the slat flow in the second 7 x lO-ft. wind tunnel test. PIV has an advantage that it can measure velocity and spectra of the entire flowfield instantaneously. However, problems arise when trying, to store this massive amount of PIV data. Support for this research has continued through a NASA Graduate Student Program Fellowship which will end in June 1999. The thesis should be completed by this time.

Moriarty, Patrick↗

Exploring the Relationship Between Reflectance Red Edge and Chlorophyll Content in Slash Pine

Chlorophyll is a key indicator of the physiological status of a forest canopy. However, its distribution may vary greatly in time and space, so that the estimation of chlorophyll content of canopies or branches by extrapolation from leaf values obtained by destructive sampling is labor intensive and potentially inaccurate. Chlorophy11 content is related positively to the point of maximum slope in vegetation reflectance spectra which occurs at wavelengths between 690-740 nm and is known as the "red edge." The red edge of needles on individual slash pine (Piniis elliottii Engelm.) branches and in whole forest canopies was measured with a spectroradiometer. Branches were measured on the ground against a spectrally flat reflectance target and canopies were measured from observation towers against a spectrally variable understory and forest floor. There was a linear relationship between red edge and chlorophyll content of branches (R(exp 2) = 0.91). Measurements of the red edge and this relationship were used to estimate the chlorophyll content of other branches with an error that was lower than that associated with the colorimetric (laboratory) method. There was no relationship between the red edge and the chlorophyll content of whole canopies. This can be explained by the overriding influence of the understory and forest floor, an influence that was illustrated by spectral mixture modeling. The results suggest that the red edge could be used to estimate the chlorophyll content in branches but it is unlikely to be of value for the estimation of chlorophyll content in canopies unless the canopy cover is high.

Curran, Paul J.↗

Test Cases for a Clipped Delta Wing with Pitching and Trailing-Edge Control Surface Oscillations

Steady and unsteady measured pressures for a Clipped Delta Wing (CDW) undergoing pitching oscillations and trailing-edge control surface oscillations have been presented . From the several hundred compiled data points, 22 static cases, 12 pitching-oscillation cases, and 12 control-surface-oscillation cases have been proposed for Computational Test Cases to illustrate the trends with Mach number, reduced frequency, and angle of attack. The planform for this wing was derived by simplifying the planform of a proposed design for a supersonic transport which is described as the Boeing 2707-300. The strake was deleted, the resulting planform was approximated by a trapezoid with an unswept trailing edge, and the twist and camber were removed. In order to facilitate pressure instrumentation, the thickness was increased to 6 percent from the typical 2.5 to 3 percent for the supersonic transport. The airfoil is thus a symmetrical circular arc section with t/c = 0.06. A wing of similar planform but with a thinner airfoil of t/c = 0.03 was used in the flutter investigations, and the buffet and stall flutter investigation . Flutter results are also reported both for the 3 per cent thick simplified wing and for a more complex SST model. One of the consequences of the increased thickness of the clipped delta wing is that transonic effects are enhanced for Mach numbers near one. They are significantly stronger than would be the case for the thinner wing. Also, with the combination of high leading edge sweep of 50.5, and the sharp leading edge, a leading edge vortex forms on the wing at relatively low angles of attack, on the order of three degrees. The Appendix discusses some of the vortex flow effects. In addition, a shock develops over the aft portion of the wing at transonic speeds such that at some angles of attack, there is both a leading edge vortex and a shock wave on the wing. Such cases are a computational challenge. Some previous applications of this data set have been for the evaluation of an aerodynamic panel method and for evaluation of a Navier-Stokes capability. Linear theory and panel method results are also presented, which demonstrated the need for inclusion of transonic effects. Flutter calculations for the related wing with t/c=O.O3 are given. In this report several Test Cases are selected to illustrate trends for a variety of different conditions with emphasis on transonic flow effects. An overview of the model and tests are given, and the standard formulary for these data is listed. For each type of data, a sample table and a sample plot of the measured pressures are presented. A complete tabulation and plotting of the Test Cases is given. Only the static pressures and the 1st harmonic real and imaginary parts of the pressures are available. All of the data for the test are included in a microfiche document in the original report and are available in electronic file form. The Test Cases are also available as separate electronic files.

Bennett, Robert M.↗

Quickbird Satellite in-orbit Modulation Transfer Function (MTF) Measurement Using Edge, Pulse and Impulse Methods for Summer 2003

The spatial characteristics of an imaging system cannot be expressed by a single number or simple statement. However, the Modulation Transfer Function (MTF) is one approach to measure the spatial quality of an imaging system. Basically, MTF is the normalized spatial frequency response of an imaging system. The frequency response of the system can be evaluated by applying an impulse input. The resulting impulse response is termed the Point Spread function (PSF). This function is a measure of the amount of blurring present in the imaging system and is itself a useful measure of spatial quality. An underlying assumption is that the imaging system is linear and shift-independent. The Fourier transform of the PSF is called the Optical Transfer Function (OTF) and the normalized magnitude of the OTF is the MTF. In addition to using an impulse input, a knife-edge in technique has also been used in this project. The sharp edge exercises an imaging system at all spatial frequencies. The profile of an edge response from an imaging system is called an Edge Spread Function (ESF). Differentiation of the ESF results in a one-dimensional version of the Point Spread Function (PSF). Finally, MTF can be calculated through use of Fourier transform of the PSF as stated previously. Every image includes noise in some degree which makes MTF of PSF estimation more difficult. To avoid the noise effects, many MTF estimation approaches use smooth numerical models. Historically, Gaussian models and Fermi functions were applied to reduce the random noise in the output profiles. The pulse-input method was used to measure the MTF of the Landsat Thematic Mapper (TM) using 8th order even functions over the San Mateo Bridge in San Francisco, California. Because the bridge width was smaller than the 30-meter ground sample distance (GSD) of the TM, the Nyquist frequency was located before the first zero-crossing point of the sinc function from the Fourier transformation of the bridge pulse. To avoid the zero-crossing points in the frequency domain from a pulse, the pulse width should be less than the width of two pixels (or 2 GSD's), but the short extent of the pulse results in a poor signal-to-noise ratio. Similarly, for a high-resolution satellite imaging system such as Quickbird, the input pulse width was critical because of the zero crossing points and noise present in the background area. It is important, therefore, that the width of the input pulse be appropriately sized. Finally, the MTF was calculated by taking ratio between Fourier transform of output and Fourier transform of input. Regardless of whether the edge, pulse and impulse target method is used, the orientation of the targets is critical in order to obtain uniformly spaced sub-pixel data points. When the orientation is incorrect, sample data points tend to be located in clusters that result in poor reconstruction of the edge or pulse profiles. Thus, a compromise orientation must be selected so that all spectral bands can be accommodated. This report continues by outlining the objectives in Section 2, procedures followed in Section 3, descriptions of the field campaigns in Section 4, results in Section 5, and a brief summary in Section 6.

Helder, Dennis↗

Nondestructive Evaluation for the Space Shuttle's Wing Leading Edge

The loss of the Space Shuttle Columbia highlighted concerns about the integrity of the Shuttle's thermal protection system, which includes Reinforced Carbon-Carbon (RCC) on the leading edge. This led NASA to investigate nondestructive evaluation (NDE) methods for certifying the integrity of the Shuttle's wing leading edge. That investigation was performed simultaneously with a large study conducted to understand the impact damage caused by errant debris. Among the many advanced NDE methods investigated for applicability to the RCC material, advanced digital radiography, high resolution computed tomography, thermography, ultrasound, acoustic emission and eddy current systems have demonstrated the maturity and success for application to the Shuttle RCC panels. For the purposes of evaluating the RCC panels while they are installed on the orbiters, thermographic detection incorporating principal component analysis (PCA) and eddy current array scanning systems demonstrated the ability to measure the RCC panels from one side only and to detect several flaw types of concern. These systems were field tested at Kennedy Space Center (KSC) and at several locations where impact testing was being conducted. Another advanced method that NASA has been investigating is an automated acoustic based detection system. Such a system would be based in part on methods developed over the years for acoustic emission testing. Impact sensing has been demonstrated through numerous impact tests on both reinforced carbon-carbon (RCC) leading edge materials as well as Shuttle tile materials on representative aluminum wing structures. A variety of impact materials and conditions have been evaluated including foam, ice, and ablator materials at ascent velocities as well as simulated hypervelocity micrometeoroid and orbital debris impacts. These tests have successfully demonstrated the capability to detect and localize impact events on Shuttle's wing structures. A first generation impact sensing system has been designed for the next Shuttle flight and is undergoing final evaluation for deployment on the Shuttle's first return to flight. This system will employ wireless accelerometer sensors that were qualified for other applications on previous Shuttle flights. These sensors will be deployed on the wing's leading edge to detect impacts on the RCC leading edge panels. The application of these methods will help to insure the continued integrity of the Shuttle wing's leading edge system as the Shuttle flights resume and until their retirement.

Eric I Madaras↗