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Linear theory is used, without two dimensional or slender body assumptions, to predict flow field produced by aircraft wing, nose, inlet, and pylons. Aircraft shock wave locations are predicted, and their effect on flow field is included through transformation of aircraft geometry. Program was written in FORTRAN IV for CDC 6400 computer.
A static flexible fighter aircraft wing configuration is presently addressed by a multilevel optimization technique, based on both a full-potential concept and a rapid structural optimization program, which can be applied to such aircraft-design problems as maneuver load control, aileron reversal, and lift effectiveness. It is found that nonlinearities are important in the design of an aircraft whose flight envelope encompasses the transonic regime, and that the present structural suboptimization produces a significantly lighter wing by reducing ply thicknesses.
The performance of a small reconfigurable unmanned aerial vehicle (UAV) is evaluated, combining a multidisciplinary approach in the computational analysis of additive manufactured structures, fluid dynamics, and experiments. Reconfigurable UAVs promise cost savings and efficiency, without sacrificing performance, while demonstrating versatility to fulfill different mission profiles. The use of computational fluid dynamics (CFD) in UAV design produces higher accuracy aerodynamic data, which is particularly important for complex aircraft concepts such as blended wing bodies. To address challenges relating to anisotropic materials, the Tsai–Wu failure criterion is applied to the structural analysis, using CFD solutions as load inputs. Aerodynamic performance results show the low-speed variant attains an endurance of 1 h, 48 min, whereas its high-speed counterpart is 29 min at a 66.7% higher cruise speed. Each variant serves different aspects of small UAS deployment, with low speed envisioned for high-endurance surveying, and high speed for long-range or time-critical missions such as delivery. The experimental and simulation results suggest room for design iteration, in wing area and geometry adjustments. Structural simulations demonstrated the need for airframe improvements to the low-speed configuration. This paper highlights the potential of reconfigurable UAVs to be useful across multiple industries, advocating for further research and design improvements.
A program was conducted at Douglas Aircraft Company to develop the technology for critical structural joints of composite wing structure that meets design requirements for a 1990 commercial transport aircraft. The prime objective of the program was to demonstrate the ability to reliably predict the strength of large bolted composite joints. Ancillary testing of 180 specimens generated data on strength and load-deflection characteristics which provided input to the joint analysis. Load-sharing between fasteners in multirow bolted joints was computed by the nonlinear analysis program A4EJ. This program was used to predict strengths of 20 additional large subcomponents representing strips from a wing root chordwise splice. In most cases, the predictions were accurate to within a few percent of the test results. In some cases, the observed mode of failure was different than anticipated. The highlight of the subcomponent testing was the consistent ability to achieve gross-section failure strains close to 0.005. That represents a considerable improvement over the state of the art.
One of the powered lift systems currently being considered for use on STOL vehicles obtains additional lift by impinging the flow of a jet exhaust on the flaps and turning it downward. This direct impingement subjects large areas of the flap to the fluctuating pressure loads of the turbulent jet flow; these loads have the potential for causing high vibration levels and sonic fatigue failures. A test was conducted on a full-scale mockup of an externally blown flap system, which consisted of a TF-34 engine and a section of a wing having a triple-slotted flap. Data were obtained and evaluated on the fluctuating pressures on the flaps and it is the purpose of this paper to present some of the results describing the loads.
The effects of realistic atmospheric density variations on reentry vehicles are investigated with the aid of a random model of atmospheric density and a series of Monte Carlo reentry trajectories generated for a typical delta-wing orbiter design. The random behavior of five major performance indicators is studied, namely, maximum dynamic pressure, maximum g load, maximum stagnation heating rate, down-range distance, and cross-range distance.
Changes in angle of attack in system were sensed indirectly by gages which responded to strains induced in wing structure. Output signals were amplified, filtered, and used to activate swashplate actuators. System provided significant reduction in blade loads and desirable changes in hub forces and moments.
The results of an analytical study to evaluate the general response characteristics of a helicopter subjected to various types of discrete gust encounters are presented. The analysis employed was a nonlinear coupled, multi-blade rotorfuselage analysis including the effects of blade flexibility and unsteady aerodynamic stall. Only the controls-fixed response of the basic aircraft without any aircraft stability augmentation was considered. A discussion of the basic differences between gust sensitivity of fixed and rotary wing aircraft is presented. The effects of several rotor configuration and aircraft operating parameters on initial gust-induced load factor and blade vibratory stress and pushrod loads are discussed.
The procedures required to operate the thick wing and pylon-fuselage-fanpod-nacelle computer program are presented. The program computes surface velocities and pressure, section loads, and total configuration loads and pitching moment. Potential flow theory is used to compute the surface pressures and the associated lift, moment, and vortex drag. The skin friction drag is also computed.
For several decades, calibrated strain gages have been used to measure loads on airplanes. The accuracy of the equations used to relate the strain gage measurements to the applied loads has been based primarily on the results of the load calibration. An approach is presented for studying the effect of widely varying load distributions on strain gage load equations. The computational procedure provides a link between the load calibration and the load to be measured in flight. A matrix approach to equation selection is presented, which is based on equation standard error, load distribution, and influence coefficient plots of the strain gage equations, and is applied to a complex, delta-wing structure.
As an alternative to alleviating wing/store flutter by conventional passive methods or by more advanced active control methods, a quasi-passive concept, referred to as the decoupler pylon, is investigated which combines desirable features of both methods. Passive soft-spring/damper elements are used to decouple wing modes from store pitch modes, and a low-power control system maintains store alignment under changing mean loads. It is shown by analysis and wind tunnel tests that the decoupler pylon provides substantial increase in flutter speed and makes flutter virtually insensitive to inertia and center-of-gravity location of the store.
The buckling characteristics of Rene 41 tubular panels for a hypersonic aircraft wing were investigated. The panels were repeatedly tested for buckling characteristics using a hypersonic wing test structure and a universal tension/compression testing machine. The nondestructive buckling tests were carried out under different combined load conditions and in different temperature environments. The force/stiffness technique was used to determine the buckling loads of the panel. In spite of some data scattering, resulting from large extrapolations of the data fitting curve (because of the termination of applied loads at relatively low percentages of the buckling loads), the overall test data correlate fairly well with theoretically predicted buckling interaction curves. Also, the structural efficiency of the tubular panels was found to be slightly higher than that of beaded panels.
The buckling characteristics of Rene 41 tubular panels for a hypersonic aircraft wing were investigated. The panels were repeatedly tested for buckling characteristics using a hypersonic wing test structure and a universal tension/compression testing machine. The nondestructive buckling tests were carried out under different combined load conditions and in different temperature environments. The force/stiffness technique was used to determine the buckling loads of the panels. In spite of some data scattering resulting from large extrapolations of the data-fitting curve (because of the termination of applied loads at relatively low percentages of the buckling loads), the overall test data correlate fairly well with theoretically predicted buckling interaction curves. Also, the structural efficiency of the tubular panels was found to be slightly higher than that of beaded panels.
A laboratory-based test apparatus employing components typical of aircraft construction was developed that would allow the study of structure-borne noise transmission due to propeller induced wake/vortex excitation of in-wake structural appendages. The test apparatus was employed to evaluate several aircraft installation effects (power plant placement, engine/nacelle mass loading, and wing/fuselage attachment methods) and several structural response modifications for structure-borne noise control (the use of wing blocking mass/fuel, wing damping treaments, and tuned mechanical dampers). Most important was the development of in-flight structure-borne noise transmission detection techniques using a combination of ground-based frequency response function testing and in-flight structural response measurement. Propeller wake/vortex excitation simulation techniques for improved ground-based testing were also developed to support the in-flight structure-borne noise transmission detection development.
The design and testing of Natural Laminar Flow (NLF) airfoils is examined. The NLF airfoil was designed for low speed, having a low profile drag at high chord Reynolds numbers. The success of the low speed NLF airfoil sparked interest in a high speed NLF airfoil applied to a single engine business jet with an unswept wing. Work was also conducted on the two dimensional flap design. The airfoil was decambered by removing the aft loading, however, high design Mach numbers are possible by increasing the aft loading and reducing the camber overall on the airfoil. This approach would also allow for flatter acceleration regions which are more stabilizing for cross flow disturbances. Sweep could then be used to increase the design Mach number to a higher value also. There would be some degradation of high lift by decambering the airfoil overall, and this aspect would have to be considered in a final design.
In-flight, aircraft are exposed to a wide range of environments. One commonly exposed environment are clouds containing super-cooled water droplets. These water drop- lets exist in a metastable state below the freezing point of water, in the range of 0 to -20°C. As the vehicle impacts the droplets, latent heat is released and within milliseconds the droplets convert to ice. This process is referred to as impact icing or in-flight icing.1 Impact icing is a major concern for aircraft since it can lead to degraded aerodynamic performance and, if left un- treated, can lead to loss of the vehicle. Active approaches (i.e., pneumatic boots, heated air ducts) typically utilized in mitigating in-flight ice accretion significantly increases vehicle weight and cannot be applied to all aircraft.1-3 A passive approach based on coatings is desired, but durability issues are a concern, especially on the wing leading edge.3 Nanomaterials have been shown to afford significant improvement in coating and composite physical properties at low loading levels.4 In this study, Polyhedral Oligomeric Silsesquioxane (POSS) nanomaterials have been shown to increase coating durability. Also, with wide variety of functionalities present on the arm structure, POSS nanomaterials have been shown to readily alter coating surface chemistry to mitigate impact ice adhesion from -16 to -8°C in a simulated in-flight icing environment.
Composite materials have become ubiquitous in the aerospace industry due to their exceptionally light weight and high strength characteristics, as well as their unique ability to be engineered and tailored to meet specific loading conditions and performance requirements. These advanced materials offer superior strength-to-weight ratios compared to traditional metallic materials, making them particularly valuable in weight-critical aerospace applications where every pound saved translates to improved efficiency and performance. In currently operating fleets of commercial and military aircraft, composite materials have been successfully applied to critical structural components, including primary load-bearing elements such as the fuselage sections and flooring structures, which must withstand significant in-flight loads and provide passenger safety. Additionally, these materials have been specifically tailored and optimized for aerodynamic components such as wings and tail assemblies, where their ability to be molded into complex shapes while maintaining structural integrity is particularly advantageous. The application of composite materials extends beyond primary structural elements into the realm of internal cabin components, most notably in innovative seat designs where weight reduction and structural integration are paramount concerns. Modern composite seat structures can be designed to integrate multiple functions, including structural support, comfort features, and safety systems, all while maintaining the lightweight characteristics essential for aircraft performance.
This report documents a series of investigations to develop an approach for structural sizing of various morphing wing concepts. For the purposes of this report, a morphing wing is one whose planform can make significant shape changes in flight - increasing wing area by 50% or more from the lowest possible area, changing sweep 30 or more, and / or increasing aspect ratio by as much as 200% from the lowest possible value. These significant changes in geometry mean that the underlying load-bearing structure changes geometry. While most finite element analysis packages provide some sort of structural optimization capability, these codes are not amenable to making significant changes in the stiffness matrix to reflect the large morphing wing planform changes. The investigations presented here use a finite element code capable of aeroelastic analysis in three different optimization approaches -a "simultaneous analysis" approach, a "sequential" approach, and an "aggregate" approach.