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A payload retention concept that may be utilized on Space Shuttle missions is investigated. This concept appears to embody all the desirable features for the very demanding requirements for space flight. Desirable features include light weight, low cost, high reliability, excellent load distribution, versatility, and insensitivity to critical alignment.
The measurement of velocity in large wind tunnels is normally performed using mechanical probes. This practice can entail considerable complexity in connection with the structure often required for probe support. The use of laser velocimetry for the flowfield measurements in large wind tunnels is, therefore, far more desirable since it can be implemented with minimal mechanical complexity and with minimal modifications to the test section structure. A description is provided of a velocimeter which is intended for use in investigations which include V/STOL vehicle flowfield mapping for location and sizing of wakes and plumes, wing-loading distribution determination without the need for pressure taps, the aerodynamics of high-lift airfoils as well as rotating rotor blades, and vehicle wake studies. Use is made of a single-color dual-beam backscatter system which is capable of sensing two orthogonal components of velocity.
The vertical gust response of the helicopter rotor in cruising flight is studied analytically by means of the local momentum theory (LMT) and experimentally in a wind tunnel utilizing a gust generator. By introducing the unsteady aerodynamic effects into the LMT and by considering the elastic deformation of the rotor blade, the vibratory characteristics of flapping blades and of the rotor forces can be obtained. Since the LMT makes it possible to calculate the instantaneous load distribution along the rotor blade in desired azimuthal or timewise intervals, the effects of gradual penetration of the rotor into the gust can be studied and the Fourier analysis or power spectrum analysis of the rotor response applied to any kind of gust input. Some results obtained analytically are verified by the experimental tests performed in the wind tunnel which generates step, sinusoidal and random vertical gusts by the motion of a series of cascaded vanes upstream of the rotor.
A joint NASA/USAF program was conducted to accomplish the following objectives: (1) evaluate the benefits that could be achieved from the application of winglets to KC-135 aircraft; and (2) determine the ability of wind tunnel tests and analytical analysis to predict winglet characteristics. The program included wind-tunnel development of a test winglet configuration; analytical predictions of the changes to the aircraft resulting from the application of the test winglet; and finally, flight tests of the developed configuration. Pressure distribution, loads, stability and control, buffet, fuel mileage, and flutter data were obtained to fulfill the objectives of the program.
The source code for the latest production version, MARK IV, of the NASA-Langley Vortex Lattice Computer Program is presented. All viable subcritical aerodynamic features of previous versions were retained. This version extends the previously documented program capabilities to four planforms, 400 panels, and enables the user to obtain vortex-flow aerodynamics on cambered planforms, flowfield properties off the configuration in attached flow, and planform longitudinal load distributions.
The application of an experimental flight test maneuver autopilot test technique for collecting aerodynamic and structural flight research data on a highly maneuverable aircraft is described in this paper. This technique, which was developed to increase the quality and quantity of data obtained during flight test, was applied to the highly maneuverable aircraft technology (HiMAT) vehicle. A primary flight experiment was to verify the design techniques used to develop the HiMAT aerodynamics and structures. This required the performance of maneuvers for collection of large quantities of high-quality pressure distribution, loads, and wing and canard deflection data. Flight data obtained while executing these research maneuvers are presented to demonstrate the effectiveness of this new technique.
The integral conservation laws of fluid mechanics are used to assess the drag efficiency of lifting wings, both CTOL and various out-of-plane configurations. The drag-due-to-lift is separated into two major components: (1) the induced drag-due-to-lift that depends on aspect ratio but is relatively independent of Reynolds number; (2) the form drag-due-to-lift that is independent of aspect ratio but dependent on the details of the wing section design, planform and Reynolds number. For each lifting configuration there is an optimal load distribution that yields the minimum value of drag-due-to-lift. For well designed high aspect ratio CTOL wings the two drag components are independent. With modern design technology CTOL wings can be (and usually are) designed with a drag-due-to-lift efficiency close to unity. Wing tip-devices (winglets, feathers, sails, etc.) can improve drag-due-to-lift efficiency by 10 to 15% if they are designed as an integral part of the wing. As add-on devices they can be detrimental. It is estimated that 25% improvements of wing drag-due-to-lift efficiency can be obtained with joined tip configurations and vertically separated lifting elements without considering additional benefits that might be realized by improved structural efficiency. It is strongly recommended that an integrated aerodynamic/structural approach be taken in the design of (or research on) future out-of-plane configurations.
In the present treatment of the calculation of forces on a wing that is suddenly brought into motion at a constant speed, attention is given to the unsteady potential's contribution to the force balance. Total bound vorticity is produced at the initial impulse. The results obtained are independent of wing aspect ratio; as time increases, this effect on the drag force becomes smaller as the vortex emanating from the trailing edge is left behind. The second contributor to induced drag is the spanwise vorticity shedding that results from the spanwise load distribution of three-dimensional wings. This contribution grows with time as the length of the wake grows.
A full Navier-Stokes solver has been used to model transonic flow over three airfoil sections. The method uses a two-dimensional, implicit, conservative finite difference scheme for solving the compressible Navier-Stokes equations. Results are presented as prescribed for the Viscous Transonic Airfoil Workshop to be held at the AIAA 25th Aerospace Sciences Meeting. The NACA 0012, RAE 2822 and Jones airfoils have been investigated for both attached and separated transonic flows. Predictions for pressure distributions, loads, skin friction coefficients, boundary layer displacement thickness and velocity profiles are included and compared with experimental data when possible. Overall, the results are in good agreement with experimental data.
Conceptual efforts to develop a Space Station based system for docking and/or berthing the NSTS Orbiter are described. Past docking and berthing systems are reviewed, the general requirements and options for mating the Orbiter and Space Station are discussed, and the rationale for locating the system on the Station is established. One class of Station-based system is developed in several variations and evaluated with respect to weight distribution, loads, safety, reliability, viewing, and maintainability. An evolutionary presentation of the variations provides insight into the development process and the problems encountered. An overall evaluation of the Station-based variations compared to an optimized Orbiter-based system demonstrates the potential benefits of this approach as well as the issues that must be resolved to realize the benefits.
A full Navier-Stokes solver has been used to model transonic flow over three airfoil sections. The method uses a two-dimensional, implicit, conservative finite difference scheme for solving the compressible Navier-Stokes equations. Results are presented as prescribed for the Viscous Transonic Airfoil Workshop to be held at the AIAA 25th Aerospace Sciences Meeting. The NACA 0012, RAE 2822 and Jones airfoils have been investigated for both attached and separated transonic flows. Predictions for pressure distributions, loads, skin friction coefficients, boundary layer displacement thickness and velocity profiles are included and compared with experimental data when possible. Overall, the results are in good agreement with experimental data.
A simple and fast linear optimum method for trimming aircraft with a combination of three longitudinal aerodynamic or propulsive control effectors is developed, with application to preliminary design or in-flight scheduling. A Lagrangian formulation is used to determine the longitudinal load distribution which minimizes trim drag. Trim drag results obtained with the linear optimum method are similar to those obtained with a nonlinear optimum method, and significant trim drag reductions are found using either optimum method, in comparison with results obtained using various ad hoc methods.
A simple analytical method has been introduced for aerohydrodynamic load analysis of an airborne configuration during water ditching. The method employs an aerodynamic panel code, based on linear potential flow theory, to simulate the flow of air and water around an aircraft configuration. The free surface separating the air and water region is represented by doublet sheet singularities. Although all the theoretical load distributions are computed for air, provisions are made to correct the pressure coefficients obtained on the configuration wetted surfaces to account for the water density. As an analytical tool, the Vortex Separation Aerodynamic (VSAERO) code is chosen to carry out the present investigation. After assessing the validity of the method, its first application is to analyze the water ditching of the Space Shuttle configuration at a 12 degree attitude.
The nondestructive testing (NDT) on the Space Shuttle Solid Rocket Booster (SRB) filament wound case (FWC) short stact structural test articles 2 (STA-2A) during test of phases 1B-9C is described. The primary objective of this testing was to verify the structural integrity of the SRB-FWC for critical design loads. Another objective was to quantify the effect of load distributions in the aft skirt. The NDT objectives were to determine the acoustic emission characteristics of the FWC-SRB and to identify possible design deficiencies or defect growth. The results from the posttest inspection of the samples shows the depth measurements taken were accurate until exceeding .260 inches thickness. The data then show that pulse echo measurements exceeded actual part thickness by 10 to 14 percent. The mapping of forward boundaries of delaminations proved to be within the tolerance of the equipment. Using the ZIP probe, the maximum difference between the pulse echo boundary and the visual boundary was expected to be no greater than one half the diameter of the probe. The NDT performance on STA-2A shows how NDT can be used to assist design engineering in evaluating the structural integrity of composite test articles.
A computational method for prediction of external store carriage loads at transonic speeds is described. The geometric flexibility required for treatment of isolated and underwing, pylon-mounted stores is achieved by computing solutions on a five-level embedded grid arrangement. A completely automated grid generation procedure facilitates applications. Store modeling capability consists of bodies of revolution with multiple fore and aft fins. A body-conforming grid improves the accuracy of the computed store body flow field. A nonlinear finite difference relaxation scheme, developed specifically for modified transonic small disturbance flow equations, enhances numerical stability and accuracy. As a result, more accurate treatment of low aspect ratio, highly swept and tapered wing planforms is possible. A limited supersonic freestream capability is also provided. Pressure, load distribution, force and moment correlations show good agreement for several test cases.
A method is described for the analysis of the unsteady, incompressible potential flow associated with a helicopter rotor and it's wake in forward flight. This method is particularly useful in low advance ratio flight due to the major contribution, in the near field, of the deformed wake. The rotor geometry is prescribed and the unsteady wake geometry is computed from the local flow perturbation velocities. The wake is modeled as a full vortex lattice. The rotor geometry is arbitrary and several rotor blades can be represented. The unsteady airloads on the rotor blades are computed in the presence of the deformed rotor wake by a time-stepping technique. Solution for the load distribution on the blade surfaces is found by prescribing boundary conditions in a reference system which rotates with the blade tips. Transformation tensors are used to describe the contribution of the wake in the inertial system to the rotor in the rotating reference system. The effects of blade cyclic pitch variation are computed using a rotation tensor. The deformation of the wake is computed in the inertial frame. The wake is started impulsively from rest, allowing a natural convection of the wake with time.
A simple analytical method has been introduced for aerohydrodynamic load analysis of an airborne configuration during water ditching. The method employs an aerodynamic panel code, based on linear potential flow theory, to simulate the flow of air and water around an aircraft configuration. The free surface separating the air and water region is represented by doublet sheet singularities. Although all the theoretical load distributions are computed for air, provisions are made to correct the pressure coefficients obtained on the configuration wetted surfaces to account for the water density. As an analytical tool, the Vortex Separation Aerodynamic (VSAERO) code is chosen to carry out the present investigation. After assessing the validity of the method, its first application is to analyze the water ditching of the Space Shuttle configuration at a 12 degree attitude.
A computer program, the Propeller Nacelle Aerodynamic Performance Prediction Analysis (PANPER), was developed for the prediction and analysis of the performance and airflow of propeller-nacelle configurations operating over a forward speed range inclusive of high speed flight typical of recent propfan designs. A propeller lifting line, wake program was combined with a compressible, viscous center body interaction program, originally developed for diffusers, to compute the propeller-nacelle flow field, blade loading distribution, propeller performance, and the nacelle forebody pressure and viscous drag distributions. The computer analysis is applicable to single and coaxial counterrotating propellers. The blade geometries can include spanwise variations in sweep, droop, taper, thickness, and airfoil section type. In the coaxial mode of operation the analysis can treat both equal and unequal blade number and rotational speeds on the propeller disks. The nacelle portion of the analysis can treat both free air and tunnel wall configurations including wall bleed. The analysis was applied to many different sets of flight conditions using selected aerodynamic modeling options. The influence of different propeller nacelle-tunnel wall configurations was studied. Comparisons with available test data for both single and coaxial propeller configurations are presented along with a discussion of the results.