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Effects of wing-crank, leading-edge chord extensions & horizontal-tail height on longitudinal stability of swept-wing models
Wind tunnel and flight tests were conducted to determine the effects of several discontinuous drooped wing leading-edge configurations on the spinning characteristics of a light, single-engine, low-wing research airplane. Particular emphasis was placed on the identification of modifications which would improve the spinning characteristics. The spanwise length of a discontinuous outboard droop was varied and several additional inboard segments were added to determine the influence of such leading-edge configurations on the spin behavior. Results of the study indicated that the use of only the discontinuous outboard droop, over a specific spanwise area, was most effective towards improving spin and spin recovery characteristics, whereas the segmented configurations having both inboard and outboard droop exhibited a tendency to enter a flat spin.
The effects of free-jet phenomena, jetting, in a 90-deg-sharp-edge inlet tube are analyzed. Mass-limiting flow data and associated pressure profiles for tubes of 53, 64, 73 and 105 L/D with a 90 deg-sharp-edge or orifice-type inlet are compared to Borda-type inlet data to determine bounds of the free-jet phenomena. For smooth tubes the limits appear to be one-dimensional and dependent only on inlet stagnation conditions. The upper L/D boundary is related by stagnation characteristics and the lower bound appears to be saturation conditions at the inlet. It is noted that similar free-jet effects were found for fluid hydrogen indicating that fluid jetting might be common to all fluids flowing through 90 deg-sharp-edge inlet geometries.
The relationship between habitat area, spatial dynamics of the landscape, and species diversity is an important theme in population and conservation biology. Of particular interest is how populations of various species are affected by increasing habitat edges due to fragmentation. Over the last decade, assumptions regarding the effects of habitat edges on biodiversity have fluctuated wildly, from the belief that they have a positive effect to the belief that they have a clearly negative effect. This change in viewpoint has been brought about by an increasing recognition of the importance of geographic scale and a reinterpretation of natural history observations. In this preliminary report from an ongoing project, we explore the use of remote sensing technology and geographic information systems to further our understanding of how species diversity and population density are affected by habitat heterogeneity and landscape composition. A primary feature of this study is the investigation of SAR for making more rigorous investigations of habitat structure by exploiting the interaction between radar backscatter and vegetation structure and biomass. A major emphasis will be on the use of SAR data to define relative structural types based on measures of structural consolidation using the vegetation surface area to volume ratio (SA/V). Past research has shown that SAR may be sensitive to this form of structural expression which may affect biodiversity.
Existing interpretations of the trailing edge condition, addressing both theoretical and experimental works in steady, as well as unsteady flows are critically reviewed. The work of Kutta and Joukowski on the trailing edge condition in steady flow is reviewed. It is shown that for most practical airfoils and blades (as in the case of most turbomachine blades), this condition is violated due to rounded trailing edges and high frequency effects, the flow dynamics in the trailing edge region being dominated by viscous forces; therefore, any meaningful modelling must include viscous effects. The question of to what extent the trailing edge condition affects acoustic radiation from the edge is raised; it is found that violation of the trailing edge condition leads to significant sound diffraction at the tailing edge, which is related to the problem of noise generation. Finally, various trailing edge conditions in unsteady flow are discussed, with emphasis on high reduced frequencies.
Broad iron absorption structures have been observed in the X-ray spectra of both AGNs and black hole X-ray binaries (BHXBs). A correctly modeled absorption structure can reveal the physical condition of the source, help to determine the continuum spectra and thus help to estimate other spectral lifes more accurately. The absorption structures are usually thought to be caused by the reflection of X-rays by the accretion disks around the central black holes, and the broadening can be a ttributed to the ionization states of the disk and relativistic effects.
Three-dimensional sidewall compression scramjet inlets with leading edge sweeps of 30 and 70 degrees have been tested in the Langley Hypersonic CF4 Tunnel at Mach 6 and a ratio of specific heats of 1.2. The effects of cowl position, contraction ratio, and Reynolds number were investigated. The models were instrumented with 42 static pressure orifices distributed on the sidewalls, baseplate, and cowl. Schlieren movies were made of each test for flow visualization of the entrance plane and cowl region. In order to obtain an approximate characterization of the flow field, a modification to two-dimensional inviscid oblique shock theory was derived to accommodate the three-dimensional effects of leading edge sweep. This theory qualitatively predicted the reflected shock structure/sidewall impingement locations and the observed increase in spillage (flow upturning) with increasing leading edge sweep. The primary effect of moving the cowl forward is capturing the flow which would have otherwise spilled out ahead of the cowl. Increasing the contraction ratio (moving the sidewalls closer together) increases the number of internal shock reflections and hence incrementally increases the sidewall pressure distribution. Significant Reynolds number effects were noted over a small range of Reynolds number.
The effects of cylindrical leading edge sweep on surface pressure and heat transfer rate for swept shock wave interference were investigated. Experimental tests were conducted in the Calspan 48-inch Hypersonic Shock Tunnel at a nominal Mach number of 8, nominal unit Reynolds number of 1.5 x 10 to the 6th power per foot, leading edge and incident shock generator sweep angles of 0, 15, and 30 deg, and incident shock generator angle-of-attack fixed at 12.5 deg. Detailed surface pressure and heat transfer rate on the cylindircal leading edge of a swept shock wave interference model were measured at the region of the maximum surface pressure and heat transfer rate. Results show that pressure and heat transfer rate on the cylindrical leading edge of the shock wave interference model were reduced as the sweep was increased over the range of tested parameters. Peak surface pressure and heat transfer rate on the cylinder were about 10 and 30 times the undisturbed flow stagnation point value, respectively, for the 0 deg sweep test. A comparison of the 15 and 30 deg swept results with the 0 deg swept results showed that peak pressure was reduced about 13 percent and 44 percent, respectively, and peak heat transfer rate was reduced about 7 percent and 27 percent, respectively.
Two dimensional subsonic wind tunnel tests were conducted on a 20% thickness: chord ratio circulation controlled elliptic aerofoil section equipped with forward and reverse blowing slots. Overall performance measurements were made over a range of trailing edge blowing momentum coefficients from 0 to 0.04; some included the effect of leading edge blowing. A detailed investigation of the trailing edge wall jet, using split film probes, hot wire probes and total head tubes, provided measurements of mean velocity components, Reynolds normal and shear stresses, and radial static pressure. The closure of the two dimensional angular momentum and continuity equations was examined using the measured data, with and without correction, and the difficulty of obtaining a satisfactory solution illustrated. Suggestions regarding the nature of the flow field which should aid the understanding of Coanda effect and the theoretical solution of highly curved wall jet flows are presented.
The results are presented of a two-dimensional investigation conducted to determine the effect of blowing over various types of trailing-edge flaps on a wing having the NACA 0006 airfoil section and a drooped-nose flap. The position and profile of the trailing-edge flap, the nozzle height, and the location of the flap with respect to the nozzle were found to be important variables. Data from many investigations were used to make an evaluation of the effects of blowing on lift. An analysis was made of flow and power relationships for blowing systems.
The paper presents numerical results, based on Donnell shell theory, for the axial compressive buckling loads for eccentrically stringer-stiffened circular cylindrical panels, in a study of the effect of boundary conditions and panel width on the buckling load. The two cases of inside and outside stiffeners were considered. The complete cylinder buckling load was reached only for panels under classical simply supported boundary conditions. The prevention of circumferential displacement is found to be the most important in-plane boundary condition from the point of view of increasing the buckling load. Clamping is found more effective in increasing the buckling loads of panels with free circumferential edge displacement than of panels with zero edge displacement. When panel width is equal to or greater than 180 deg, the panel buckling loads are within 10% of the complete cylinder load for all cases except one simply supported panel with outside stringers. Buckling loads were higher for outside stringers, except for very narrow panels that are restrained against circumferential edge displacement. Eccentricity effects are generally similar for clamped and simply supported panels with the same in-plane boundary conditions.
The solution of von Karman's fundamental equations for large deflections of plates is presented for the case of a simply supported rectangular plate under combined edge compression and lateral loading. Numerical solutions are given for square plates and for rectangular plates with a width-span ratio of 3:1. The effective widths under edge compression are compared with effective widths according to von Karman, Bengston, Marguerre, and Cox and with experimental results by Ramberg, McPherson, and Levy. The deflections for a square plate under lateral pressure are compared with experimental and theoretical results by Kaiser. It is found that the effective widths agree closely with Marguerre's formula and with the experimentally observed values and that the deflections agree with the experimental results and with Kaiser's work.
The solution of Von Karman's fundamental equations for large deflections of plates is presented for the case of a simply supported rectangular plate under combined edge compression and lateral loading. Numerical solutions are given for square plates and for rectangular plates with a width-span ratio of 3:1. The effective widths under edge compression are compared with effective widths according to Von Karman, Bengston, Marguerre, and Cox and with experimental results by Ramberg, Mcpherson, and Levy. The deflections of a square plate under lateral pressure are compared with experimental and theoretical results by Kaiser. It is found that the effective widths agree closely with Marguerre's formula and with the experimentally observed values and that the deflections agree with the experimental results and with Kaiser's work.
Low speed wind sweep effect on parawings with equal-length leading edges & keel - aerodynamic characteristics
Several local edge detection operators were applied to a set of ERTS pictures of the Monterey, Calif. area. Gradient operators performed consistently better than laplacian operators in detecting edges. It was also found that if a grayscale normalization operation, histogram flattening, was applied to the pictures first, the edge detector outputs were greatly enhanced. The use of interpolation for more accurate location of edges on a digital picture was also briefly investigated. Curve detection operators were applied to the edge detector outputs; this had the effect of enhancing the edges while suppressing noise.
A rigorous mathematical model was obtained for the boundary-layer free-edge stress singularity in angleplied and crossplied fiber composite laminates. The solution was obtained using a method consisting of complex-variable stress function potentials and eigenfunction expansions. The required order of the boundary-layer stress singularity is determined by solving the transcendental characteristic equation obtained from the homogeneous solution of the partial differential equations. Numerical results obtained show that the boundary-layer stress singularity depends only upon material elastic constants and fiber orientation of the adjacent plies. For angleplied and crossplied laminates the order of the singularity is weak in general.
Models of four delta wings were built and tested in the 8 x 8 ft Transonic Wind Tunnel at LaRC. The wings are identical in planform shape with a swept-back angle of 65 degrees, but bear different leading edge profiles. The models were tested under pressurized and cryogenic conditions to simulate true flight Reynolds numbers. Data on the aerodynamic forces and pressure distributions at various locations on the wings were taken at various flight Mach numbers and angles of attack. Effects of high Reynolds number and leading edge radius on the aerodynamic characteristics of the wings are being accessed. To thoroughly understand the turbulent, vortical flow around the wings, an effort to perform computational aerodynamic analysis is being made. The objective of the analysis is to supplement and validate the experimental data and explain the high Reynolds number and leading edge effects. GRIDGEN, a software developed by General Dynamics, is being used to generate the grid topology for the flow field around the wings. The flow solver to be used is CFL3D, a computation fluid dynamics code developed at LaRC. Based on the geometric description of the wings, a FORTRAN program called WINGSURF was written to generate the databases defining the surface geometry of the wings. To match the true geometry of the models in the wind tunnel for realistic comparison with experimental results, databases defining the sting support for the wing models were also created by two other FORTRAN programs, STINJOIN and STINREAR. Listing of the programs are attached and the geometry of a typical wing model with the sting support is shown. Other aspects of the investigation are discussed.
The electronics module cover for the leading edge (Row D 9) experiment M0003-8 was fabricated from T300 graphite/934 epoxy unidirectional prepreg tape in a (O(sub 2), +/- 45, O(sub 2), +/- 45, 90, 0)(sub s) layup. This 11.75 in x 16.75 in panel was covered with thermal control coatings in three of the four quadrants with the fourth quadrant uncoated. The composite panel experienced different thermal cycling extremes in each quadrant due to the different optical properties of the coatings and bare composite. The panel also experienced ultraviolet (UV) and atomic oxygen (AO) attack as well as micrometeoroid and space debris impacts. An AO reactivity of 0.99 x 10(exp -24) cm(sup 3)/atom was calculated for the bare composite based on thickness loss. The white urethane thermal control coatings (A276 and BMS 1060) prevented AO attack of the composite substrate. However, the black urethane thermal control coating (Z306) was severely eroded by AO, allowing some AO attack of the composite substrate. An interesting banding pattern on the AO eroded bare composite surface was investigated and found to match the dimensions of the graphite fiber tow widths as prepregged. Also, erosion depths were greater in the darker bands. Five micrometeoroid/space debris impacts were cross sectioned to investigate possible structural damage as well as impact/AO interactions. Local crushing and delaminations were found to some extent in all of the impacts. No signs of coating undercutting were observed despite the extensive AO erosion patterns seen in the exposed composite material at the impact sites. An extensive microcrack study was performed on the panel along with modeling of the thermal environment to estimate temperature extremes and thermal shock. The white coated composite substrate displayed almost no microcracking while the black coated and bare composite showed extensive microcracking. Significant AO erosion was seen in many of the cracks in the bare composite.