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Support was provided for the maintenance and modifications of software for the production and detailed analysis of data from the DE-A spacecraft and new software developed for this end. Software for the analysis of the data from the Spacelab Experimental Particle Accelerator (SEPAC) was also developed.
A transient technique was used to measure heat transfer coefficients on stator airfoils in a high-temperature annular cascade at real engine conditions. The transient response of thin film thermocouples on the airfoil surface to step changes in the gas stream temperature was used to determine these coefficients. In addition, gardon gages and paired thermocouples were also utilized to measure heat flux on the airfoil pressure surface at steady state conditions. The tests were conducted at exit gas stream Reynolds numbers of one-half to 1.9 million based on true chord. The results from the transient technique show good comparison with the steady-state results in both trend and magnitude. In addition, comparison is made with the STAN5 boundary layer code and shows good comparison with the trends. However, the magnitude of the experimental data is consistently higher than the analysis.
Data are reported from simultaneous measurements of sky radiance and aerosol optical thickness using a transmissometer tuned to seven bands between 441-872 nm. The measurements were made near Washington, DC in February and November 1985 in generally cloud-free conditions. The data indicated that satellite radiance measurements in one band would be highly correlated with AEC values in other bands, suggesting that transmission and reflection of sunlight in nonabsorbing gaseous bands is highly correlated with path radiance.
The hierarchy of codes; boundary layers and turbulence models; uniqueness and artificial viscosity; efficiency of solution algorithms; data display and analysis; and cost of experimental validation are briefly examined.
Plastic deformation accompanying the wear process in fully dense and porous metals was studied in order to develop more abradable gas path seal material systems for turbine engine applications. The finite element method was used to model high-rate viscoplastic deformation in the vicinity of a moving contact. Techniques were developed for studying such deformation in both solid (fully dense) and porous metals. Results of the analysis were verified by comparison with experimental data obtained using a pendulum-type rub test apparatus and a pin-on-disk sliding wear tester. It was found that the near-surface plastic deformation in fully dense solids was different from that for porous materials because of the compressibility of the latter. Both types of materials exhibit deformation features which can be taken advantage of in the design of more abradable material systems.
Results are presented from the comparison study of two computer codes for crack growth analysis - NASCRAC and NASA/FLAGRO. The two computer codes gave compatible conservative results when the part through crack analysis solutions were analyzed versus experimental test data. Results showed good correlation between the codes for the through crack at a lug solution. For the through crack at a lug solution, NASA/FLAGRO gave the most conservative results.
The two-phase frictional pressure drop and vapor volume fraction in the vertical boiling and adiabatic flow of the refrigerant, R11, have been simultaneously measured by a liquid balancing column and differential magnetic reluctance pressure transducers. An account is given of the experimental apparatus and procedure, data acquisition and analysis, and error estimation employed. All values of two-phase multipliers evaluated on the basis of the measured frictional pressure drop data in vertical upflow fall in the range bounded by the predictions of the Chisholm correlation and the homogeneous model.
Some illustrative results obtained from work accomplished under the aerothermoelasticity work breakdown structure (WBS) element of the National Aerospace Plane (NASP) Technology Maturation Program (TMP) are presented and discussed. The objectives of the aerothermoelasticity element were to develop analytical methods applicable to aerospace plane type configurations, to conduct analytical studies to identify potential problems, to evaluate potential solutions to problems, and to provide an experimental data base to verify codes and analytical trends. Work accomplished in the three areas of experimental data base, unsteady aerodynamics, and integrated analysis methodology are described. Some of the specific topics discussed are: (1) transonic wind tunnel aeroelastic model tests of cantilever delta wing models, of an all-moveable delta-wing model, and of aileron buzz models; (2) unsteady aerodynamic theory correlation with experiment and theory improvements; and (3) integrated analysis methodology results for thermal effects on vibration, for thermal effects on flutter, and for improving aeroelastic performance by using active controls.
Locating damaged truss members in large space structures will involve a combination of sensing and diagnostic techniques. Methods developed for damage location require experimental verification prior to on-orbit applications. To this end, a series of experiments for locating damaged members using a generic, ten bay truss structure were conducted. A 'damaged' member is a member which has been removed entirely. Previously developed identification methods are used in conjunction with the experimental data to locate damage. Preliminary results to date are included, and indicate that mode selection and sensor location are important issues for location performance. A number of experimental data sets representing various damage configurations were compiled using the ten bay truss. The experimental data and the corresponding finite element analysis models are available to researchers for verification of various methods of structure identification and damage location.
A qualitative analysis is conducted to reconcile the experimental data from critical ionization velocity (CIV) studies with CIV theories. The experimental data are reviewed demonstrating that: (1) the wave frequency is variable and low; (2) the wave polarization is almost isotropic; (3) electron energization is not easily reconciled with the observed wave spectrum; and (4) ambient electron density plays a role in determining CIV triggering conditions. Analytical treatment is given to the dispersion relation of the lower hybrid wave (LWH) instability driven by the streaming of an ion beam generated by the interaction of the neutral cloud with the ambient atmosphere. By incorporating the LWH instabilities of strong turbulence and finite-size effects into theoretical CIV relationships, the observations can be interpreted. The issues raised by the experimental data are understood within the context of a hypothesis of backward propagating nonlinearly collapsing wavepackets.
Analysis of some recent experimental sonic boom data has revived the hypothesis that there is a closeness limit to the near-field separation distance from which measured wind tunnel pressure signatures can be extrapolated to the ground as though generated by a supersonic-cruise aircraft. Geometric acoustic theory is used to derive an estimate of this distance and the sample data is used to provide a preliminary indication of practical separation distance values.
Viewgraphs on the Middeck 0-gravity Dynamics Experiment (MODE): Structural Test Article (STA) are presented. Topics covered include: MODE: structural test article motivation; hardware; sensors and actuators; experimental support module; data; preliminary results; supporting analysis program; and modeling approach.
A study of transverse cracking mechanism in composite laminates is presented using a singular hybrid finite element model. The model provides the global structural response as well as the precise local crack-tip stress fields. An elasticity basis for the problem is established by employing Lekhnitskii's complex variable potentials and method of eigenfunction expansion. Stress singularities associated with the transverse crack are obtained by decomposing the deformation into the symmetric and antisymmetric modes and proper boundary conditions. A singular hybrid element is thereby formulated based on the variational principle of a modified hybrid functional to incorporate local crack singularities. Axial stiffness reduction due to transverse cracking is studied. The results are shown to be in very good agreement with the existing experimental data. Comparison with simple shear lag analysis is also given. The effects of stress intensity factors and strain energy density on the increase of crack density are analyzed. The results reveal that the parameters approach definite limits when crack densities are saturated, an evidence of the existence of characteristic damage state.
An approximate analytical solution including the effect of end leakage from the oil film of short plain bearings is presented because of the importance of endwise flow in sleeve bearings of the short lengths commonly used. The analytical approximation is supported by experimental data, resulting in charts which facilitate analysis of short plain bearings. The analytical approximation includes the endwise flow and that part of the circumferential flow which is related to surface velocity and film thickness but neglects the effect of film pressure on the circumferential flow. In practical use, this approximation applies best to bearings having a length-diameter ratio up to 1, and the effects of elastic deflection, inlet oil pressure, and changes of clearance with temperature minimize the relative importance of the neglected term. The analytical approximation was found to be an extension of a little-known pressure-distribution function originally proposed by Michell and Cardullo.
Wing rock is a high angle of attack dynamic phenomenon of limited cycle motion predominantly in roll. The wing rock is one of the limitations to combat effectiveness of the fighter aircraft. Roll Attractor is the steady state or equilibrium trim angle (phi(sub trim)) attained by the free-to-roll model, held at some angle of attack, and released form rest at a given initial roll (bank) angle (phi(sub O)). Multiple roll attractors are attained at different trim angles depending on initial roll angle. The test facility (Vigyan's low speed wind tunnel) and experimental work is presented here along with mathematical modelling of roll attractor phenomenon and analysis and comparison of predictions with experimental data.
An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 120 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6), 60 x 10(exp 6), and 120 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.
An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 120 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6) and 60 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.
An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 36 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at a Reynolds number of 6 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.