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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 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.
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 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.
Airplane designs are currently being proposed with a multitude of lifting and control devices. Because of the redundancy in ways to generate moments and forces, there are a variety of strategies for trimming each airplane. A linear optimum trim solution (LOTS) is derived using a Lagrange formulation. LOTS enables the rapid calculation of the longitudinal load distribution resulting in the minimum trim drag in level, steady-state flight for airplanes with a mixture of three or more aerodynamic surfaces and propulsive control effectors. Comparisons of the trim drags obtained using LOTS, a direct constrained optimization method, and several ad hoc methods are presented for vortex-lattice representations of a three-surface airplane and two-surface airplane with thrust vectoring. These comparisons show that LOTS accurately predicts the results obtained from the nonlinear optimization and that the optimum methods result in trim drag reductions of up to 80 percent compared to the ad hoc methods.
Multiple coordinated robot arms are modeled by considering the arms: (1) as closed kinematic chains, and (2) as a force constrained mechanical system working on the same object simultaneously. In both formulations a new dynamic control method is discussed. It is based on a feedback linearization and simultaneous output decoupling technique. Applying a nonlinear feedback and a nonlinear coordinate transformation, the complicated model of the multiple robot arms in either formulation is converted into a linear and output decoupled system. The linear system control theory and optimal control theory are used to design robust controllers in the task space. The first formulation has the advantage of automatically handling the coordination and load distribution among the robot arms. In the second formulation, by choosing a general output equation, researchers can superimpose the position and velocity error feedback with the force-torque error feedback in the task space simultaneously.
Multiple coordinated robot arms are modeled by considering the arms as closed kinematic chains and as a force-constrained mechanical system working on the same object simultaneously. In both formulations, a novel dynamic control method is discussed. It is based on feedback linearization and simultaneous output decoupling technique. By applying a nonlinear feedback and a nonlinear coordinate transformation, the complicated model of the multiple robot arms in either formulation is converted into a linear and output decoupled system. The linear system control theory and optimal control theory are used to design robust controllers in the task space. The first formulation has the advantage of automatically handling the coordination and load distribution among the robot arms. In the second formulation, it was found that by choosing a general output equation it became possible simultaneously to superimpose the position and velocity error feedback with the force-torque error feedback in the task space.
This paper deals with the dynamic response of anisotropic laminated composite plates to stationary random excitation. The theory of laminated plates used here includes shear deformation and rotary inertia effects in the same manner as Mindlin's theory for isotropic homogeneous plates. Two cases of random pressure fields are considered in this analysis. In the first case, it is modeled as a uniformly distributed load, random in time with an exponential decaying correlation function, while in the second one as a jet noise, fully correlated in one direction and homogeneous in the perpendicular one. The analysis presented herein, as well as the obtained response characteristics, expressed in terms of mean-squares and angle-dependent mean-squares, may be useful in the design of composite plates subjected to random pressure fields and in the evaluation of their reliability.
An experimental study of the postbuckling behavior of square compression-loaded graphite-epoxy plates and isotropic plates with a central circular cutout is presented. Results are presented for unidirectional (0 sub 10)s and (90 sub 10)s plates, (0/90 sub 5)s plates, and for aluminum plates. Results are also presented for (+ or - O sub 6)s angle-ply plates for values of O = 30, 46, and 60 degrees. The experimental results indicate that the change in axial stiffness of a plate at buckling is strongly dependent upon cutout size and plate orthotropy. The presence of a cutout gives rise to an internal load distribution that changes, sometimes dramtically, as a function of cutout size coupled with the plate orthotropy. In the buckled state, the role of orthotropy becomes more significant since bending in addition to membrane orthotropy is present. Most of the plates with cutouts exhibited less postbuckling stiffness than the corresponding plate without a cutout, and the postbuckling stiffness decreased with increasing cutout size. However, some of the highly orthotropic plates with cutouts exhibited more postbuckling stiffness than the corresponding plate without a cutout.