Aerospace Engineering Theses and Dissertations

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Numerical Solutions for Two- and Three-Dimensional Non-Reacting Flowfields in an Internal Combustion Engine
(1977) Griffin, Michael Douglas; Anderson, John D. Jr; Jones, Everett; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md)
The numerical solution for the flowfield established in a spark- ignition internal combustion engine during the four-stroke (intake, compression, power, exhaust) cycle is considered. Only fluid-dynamic effects are treated with combustion simulated by constant- volume heat addition near top-dead-center on the compression stroke. The working fluid is assumed to be air of constant specific heat, with both viscous and inviscid models considered. Two- and three-dimensional engine models are examined, with the three-dimensional models including both rectangular and cylindrical geometries. The difficulties associated with obtaining numerical solutions in cylindrical coordinates for three-dimensional non-axisymmetric problems when the centerline is included in the region of interest are discussed. A new method which avoids the coordinate- singularity problems associated with such cases is presented and used to obtain the first known four-stroke inviscid-flow solution for a three- dimensional cylindrical engine model. Similar results are presented for a three-dimensional rectangular model, and for the first known two-dimensional four-stroke calculation for a viscous fluid. The inviscid three-dimensional results are compared with each other and with previously obtained two-dimensional inviscid-flow calculations. The use of two-dimensional models is found to be justified for the non- reacting flowfields considered, since the results obtained from a two-dimensional calculation in the valve plane are apparently not strongly dependent on the flowfield perpendicular to the valve plane. It is found that significant flowfields do exist in all I.C. engine models considered. It is shown that the unit-cell-Reynolds-number criterion limits viscous flow calculations to Reynolds numbers of approximately one ten-thousandth the realistic value, and that this produces flowfields which are strongly piston-dominated. In contrast, inviscid results show marked circulatory patterns, which are more realistic. The velocity patterns which develop in the three-dimensional cylindrical engine model are shown to exhibit a marked swirl in planes parallel and perpendicular to the cylinder axis.
An Experimental Investigation of the Effects of Leading Edge Modification on the Post-Stall Characteristics of an NACA 0015 Wing
(1979) Saini, Jugal Kishore; Jones, Everett; Winkelmann, Allen E.; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, MD)
The effects of leading edge modifications on the stalling characteristics of an NACA 0015 panel wing model were investigated in a series of low speed wind tunnel tests. The modification typically consisted of adding a 14% Clark Y glove onto a portion of the leading edge. Six-component balance data, pressure distribution measurements and oil flow visualization tests were completed at a Reynolds number based on chord of 2.0 x 10^6 for increasing and decreasing angles of attack from 0° to 50°. The leading edge modifications produce stabilizing vortices at stall and beyond. These vortices have the effect of fixing the stall pattern of the wing such that various portions of the wing upper surface stall nearly symmetrically. This results in a higher lift on the modified wing as compared to the lift on the unmodified wing in the post-stall region. The lift curve slope of the modified and unmodified wings remained essentially constant at 0.071 per degree. Two lift-coefficient peaks were obtained for the baseline NACA 0015 wing at angles of attack of 17° and 30°. The twin-peak behavior of the lift curve was also observed on the modified wings. The drag coefficient obtained with several modified configurations was smaller than the drag coefficient of the baseline NACA 0015 wing in the pre-stall region. Also a smaller center of pressure shift with angle of attack was observed with several modified configurations. Considering a smoother variation of lift, pitching moment, rolling moment at stall and a smaller drag and center of pressure movement to be desired criteria, the best configuration tested consisted of placing the glove on the entire leading edge except for a gap at 25% to 50% of the semispan.
Navier Stokes Solutions for Chemical Laser Flows: Steady and Unsteady Flows
(1979) Kothari, Ajay Prasannajit; Anderson, John D. Jr; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md)
This work consists of an overall effort to apply a detailed and accurate computational fluid dynamic technique to the solution of practical high energy laser flows. In particular, a third generation of super sonic diffusion chemical laser analysis is introduced, namely, the complete solution of the Navier-Stokes equations for the laminar, super sonic mixing flow fields fully coupled with chemical kinetics for both the hot and cold reactions for HF. Multicomponent diffusion is treated in a detailed fashion. Solutions are obtained, firstly, for "cold flows", where the effects of chemical reactions and vibrational relaxation are not included. Although such a situation is purely artificial, the results do isolate some of the fluid dynamic aspects of chemical laser flows, and provide a set of data to be compared later with hot flow calculations. A set of numerical experiments using four different time dependent finite difference schemes show that relatively minor changes in the differencing procedure can lead to major variations in the results. A modification of the well-known Maccormack approach appears to be the best suited for mixing flows associated with chemical lasers. A comparison is next made between cold flows (with fully coupled chemical kinetics). the results show that temperature distributions are affected the most and velocity distributions the least by chemical energy heat release. The results have an impact on the interpretation of cold flow aerodynamic experiments in the laboratory, and their proper extrapolation to the real chemical laser flows. also, comparisons between the present Navier Stokes results and other, more approximate, existing calculations are made. Gradients are calculated as a natural part of the Navier Stokes solutions. Results are given for steady flows with large pressure gradients where advantages of the Navier Stokes solutions are delineated. In addition, the effect of unsteady fluctuations intentionally introduced at the cavity inlet are studied. Specifically, sinusoidal fluctuations in one stream and then both streams (primary and secondary) in various quantities e.g. pressure, density, u velocity and v velocity were simulated. Of these, the oscillations in v velocity with approximate frequency and amplitude produced a remarkable improvement in mixing. Such unsteady fluctuations also yielded peak laser gain which were larger by almost a factor of two compared to the steady case. the flow at which the upstream boundary has so far, in the above mentioned cases been assumed to be uniform with real effects like Boundary Layer and Base Flow having been neglected. For comparison purposes these effects are next included. the boundary layer profile and velocity at the inlet is shown to feed production of gain substantially. Base flow calculations were attempted but were not successful
Dynamics of a Helicopter-Slung Load System
(1980) Sampath, Prasad; Barlow, Jewel B.; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md)
Stability of a tandem rotor helicopter (347/HLH) carrying a slung cargo container has been investigated. Lagrange equations were used to write the equations of motion. The cables of the sling were modeled as massless rigid extensible rods, which collapse under compressive loads. Extensibility was provided by considering the rods as linear spring with viscous damping. Aerodynamics of the cable were neglected. Tabulated static aerodynamic data were considered for the helicopter as well as the load. The equations were divided into two sets, one representing the towing vehicle (referred to as Subsystem 1) and the other representing the slung load (referred to as Subsystem 2). Subsystem 2 corresponds to a wind tunnel model of a slung load.
Experimental Evaluation of Circulation Control Aerodynamics on a Cylindrical Body
(1987) Ngo, Hieu Thien; Chopra, Inderjit; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, MD)
In this study, an experimental investigation is conducted on a two-dimensional circulation control cylinder with blowing taking place from a single spanwise slot to determine its aerodynamic characteristics. The results include detailed pressure distributions (both chordwise and spanwise) for a range of momentum coefficients and slot locations. The measured results showed that the lift coefficients up to 4.8 were produced at momentum coefficients of 0.14 in a turbulent flow condition. The experimental results of lift coeffficients Were correlated satisfactorily with analytical results. The surface flow patterns were observed using the oil and smoke techniques. Also flow field surveys of the model Were obtained using total pressure, yaw and pitch probes. A color video display technique was used to present the results of the flow field surveys. Based on this evidence, a flow field model of the circulation control cylinder is presented.
An Expert System for Helicopter Conceptual Design
(1987) Babuska, Vit; Fabunmi, James A.; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md)
The objective of this thesis is to demonstrate the applicability of expert systems in helicopter conceptual design by developing an expert assistant which aids the engineer in defining a feasible design configuration. The expert assistant combines some experiential knowledge of the design engineer with a typical conceptual design algorithm to guide the engineer to a reasonable baseline design. The expert assistant was developed on a personal computer using the expert system shell INSIGHT2+®. The design algorithm employed is SSPl, a helicopter weight and sizing program developed at the US Army Applied Technologies Laboratory. A set of heuristic rules was developed which attempts to simulate the thinking of an expert design engineer using SSP1 for helicopter conceptual design. The result, a Prototype expert assistant which aids an engineer in the conceptual design phase, demonstrates the feasibility of expert systems in helicopter design.
The Influence of Variable Flow Velocity on Unsteady Airfoil Behavior
(1991) van der Wall, Berend G.; Leishman, J. Gordon; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md)
The importance of unsteady aerodynamics for prediction of rotor dynamics is unquestioned today. The purpose of unsteady aerodynamic models is to represent the effect of unsteady airfoil motion on the lift, moment and drag characteristics of a blade section. This includes unsteady motion (arbitrary motion) of the airfoil in angle of attack (pitch) and vertical movement (plunge), as well as the effects of an airfoil traveling through a vertical gust field. However, the additional degrees of freedom, namely the fore-aft motion and the unsteady freestream variations commonly are acknowledged, but neglected in virtually all analyses. Since the effect of unsteady freestream results in a stretching and compressing of the shed wake vorticity distribution behind an airfoil, it will have an effect on the airfoil characteristics. The subject of this thesis is to provide a review of the analytic and experimental work done in the area of unsteady freestream and unsteady fore-aft motion, to clarify the limits of the various theories, and to show the differences between them. This will be limited to the attached flow regime since all theories are based on the small disturbance assumption in incompressible flow. As far as possible the theories are compared with experimental data, however most of the available experimental data are confined to stalled flow conditions and are not useful here. In addition to the theories, a semiempirical mathematical model will be used based on the aerodynamics of indicial functions. The purpose is to show the differences of using the theories of unsteady airfoil motion in a constant flow, and those accounting for unsteady freestream flow. This will help to justify whether it is necessary to include the unsteady freestream effect in comprehensive rotor codes. Finally, a generalisation of Isaacs unsteady aerodynamic theory for an airfoil undergoing a frequency spectra in pitch and plunge in a freestream oscillating with the fundamental frequency is presented here for the first time. Therein the axis of rotation of the airfoil is a free parameter.
MATHEMATICAL MODEL OF ADAPTIVE MOTOR CONTROL
(1999) Kosha, Makiko; Sanner, Robert M.; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md)
An adaptive control law incorporating a biologically inspired neural networks for robot control is used as a mathematical model of human motor control and the motor control adaptation. Modeling human motor control strategy is made difficult due to the redundancies in the human motor control system. This control model is able to overcome the difficulties of the human motor control modelling, and include the learning capability of the motor control strategy which was omitted in human motor control studies until now. By adaptively piecing together a collection of elementary computational elements, the proposed model develops complex internal models which are used to compensate for the effects of externally imposed forces or changes in the physical properties of the system. In order to examine the form of human motor control adaptation in detail, a computer simulation was developed with a two dimensional model of the human arm which utilized the proposed adaptive motor control model. The simulation result show that the model is able to capture the characteristics of the motor control adaptation seen in human experiments reported by [14], [46]. For cont inuation of this research, an experimental apparatus was designed and built for the human motor control study. This apparatus is a cable driven, two-dimensional manipulator which is used to apply specified disturbance forces to the human arm. The preliminary experiment conducted with this test apparatus show a strong correlation to the simulation data and other experimental data reported on human reaching motions.
Transient Dynamics of Helicopter Rotor Wakes Using a Time-Accurate Free-Vortex Method
(2001) Bhagwat, Mahendra J.; Leishman, J. Gordon; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md)
A second-order accurate predictor-corrector type algorithm has been developed to obtain a time-accurate solution of the vortical wake generated by a helicopter rotor. The rotor blade flapping solution was fully integrated with the wake geometry solution using the same time-marching algorithm. The analysis was used to predict the locations of wake vortex filaments under transient flight conditions, where the rotor wake may not be periodic at the rotational frequency. Applications of this analysis include prediction of the rotor induced velocity field and blade airloads during transient flight and maneuvers. The stability of the rotor wake structure is important from the perspective of free-vortex wake models. The wake stability was examined using a linearized stability analysis, and the rotor wake was shown to be physically unstable. Therefore, the stability of the numerical algorithm is an important consideration in developing robust wake methodologies. Both the stability and accuracy of the numerical wake solutions algorithms was rigorously examined. The straight-line vortex segmentation used in the present analysis was shown to be second-order accurate. The overall numerical solution was also demonstrated to converge with a second-order accuracy. A technique for increasing the order of accuracy for high resolution solutions is also described. Along with a formal (mathematical) verification of solution accuracy, the numerical solution for the rotor wake problem was compared with experimental results for both steady-state and transient operating conditions. The steady-state wake model was shown to give good predictions of rotor wake geometry, induced inflow distribution as well as performance trends. Under transient conditions, such as those following a pitch input during a maneuver, the time-accurate wake model was shown to correctly model the dynamic response of rotor wake. In axial descent passing through the vortex ring state, the present analysis was shown to properly model the associated power losses as shown by experimental results. The present analysis was also shown to give improved predictions of wake distortions during simulated maneuvering flight with various imposed angular rates of the rotor.
RBCC Engine-Airframe Integration on an Osculating Cone Waverider Vehicle
(2001) O'Brien, Timothy F.; Lewis, Mark J.; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md)
An analytical vehicle study is performed that integrates a rocket-based combined cycle engine with an osculating cones waverider-based fuselage. The integration of the two concepts brings about an interesting design challenge: predicting the aerodynamic performance of a high-speed fuselage design across the full range of Mach numbers from take-off to orbit that a rocket-based combined-cycle engine will operate. The aerodynamic performance of this class of vehicles is analyzed for on- and off-design Mach numbers and angles of attack. Analytical aerodynamic models are developed for the off-design behavior of both the fuselage of the vehicle and the engine. These models arc combined to predict the powered performance of this class of vehicle along a trajectory. The models developed arc rapid enough that they may be applied to initial design studies, optimization algorithms, or trajectory analyses. The aerodynamic model for the fuselage is based on the tangent-wedge, tangent-cone, and shock-expansion theories for hypersonic flow, and the linearized, small perturbation, velocity potential equations for supersonic and transonic flow. Each model is validated with numerical solutions for an example Mach 12 vehicle design. The results show an accurate prediction of the trends in lift and drag of the vehicle fuselage across a range of Mach numbers between 0.4 and 15. The aerodynamic engine model is based on Prandtl-Meyer flow and the oblique shock relations for the internal compression system, and quasi-one dimensional flow (including finite-rate chemistry) for the combustor flowfield. The strut-based compression model is validated with numerical solutions for a range of Mach numbers between 2.5 and 6. The combustor flowfield model is validated by comparison to two experimental hydrogen-fueled scramjet engines. The results showed that this class of geometry generates very little lift at low speeds (below Mach 3) and will require lift augmentation. The transonic drag rise is modeled analytically and numerically, with maximum inviscid drag coefficient occurring at Mach 1.2. Engine integration has a large effect on off-design behavior, including maximum lift-to-drag ratio and zero lift angle of attack.
FLOW INDUCED CAVITY RESONANCE FOR TURBULENT COMPRESSIBLE MIXING ENHANCEMENT IN SCRAMJETS
(2002) Nenmeni, Vijay Anand Raghavendran; Yu, Kenneth H.; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
In a Scramjet combustor, flow residence time is very short and fuel-air mixing can be adversely affected by compressibility effect. Thus, it is important to study mixing enhancement techniques for reducing the characteristic mixing time. It is also important to examine the feasibility of using them in practical settings. One of the promising mixing enhancement techniques is based on flow-induced cavity resonance, which generates large-scale coherent structures in the shear layer for faster mixing. Of particular interest is whether this technique, which is passive in nature, can be used over a wide range of flow conditions, expected in Scramjet operation. In this thesis, physical mechanisms governing the use of flow-induced cavity resonance were examined experimentally using Schlieren visualization of the flowfield and spectral analysis of resulting pressure oscillations. Various cavities with the length between 0.125 and 1.25 inch and the depth between 0.125 and 0.25 inch were placed inside a Mach 2 flow tunnel, which simulated the Scramjet internal flowfield. The properties of supersonic flow were further modified in the inlet, upstream of the cavity section, by changing the upstream stagnation pressure between 35 psig and 120 psig, which resulted in inlet shock trains of different strength. The objective was to characterize and compare the enhancement mechanism under various off-design conditions. In all, nine different cavity cases were tested under six different stagnation pressure settings. For each case, spark Schlieren images were taken and pressure oscillations inside the cavity were measured. The Schlieren images provided qualitative understanding of the physics while the pressure measurements were used to quantify the amplitude and frequency of dominant oscillations. Also from the images, inlet Mach number was deduced by measuring the Mach wave angles. The data were summarized to shed more light on reliability of the mixing enhancement mechanism under off-design inlet conditions. The results indicated that flow-induced cavity resonance mechanism was robust over a wide range of flow conditions. Also, mode-switching behavior of the cavities was observed, which could modify the mixing enhancement rate. Further, helium injection studies were conducted to gain qualitative assessment of the effect of cavity resonance on mixing.
Damping Augmentation of Helicopter Rotors Using Magnetorheological Dampers
(2003) Zhao, Yongsheng; Wereley, Norman; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, MD)
This thesis describes an investigation exploring the use of magnetorheological (MR) dampers to augment the stability of helicopter rotors. Helicopters with advanced soft in-plane rotors are susceptible to ground resonance instabilities due to the coupling of the lightly damped rotor lag modes and fuselage modes. Traditional passive lag dampers, such as hydraulic or elastomeric dampers, can be used to alleviate these instabilities. However, these passive dampers suffer from the disadvantages that they produce large damper loads in forward flight conditions. These damper forces increase fatigue loads and reduce component life. Thus, it is desirable to have lag dampers be controllable or adaptable, so that the damper can apply loads only when needed. MR fluid based dampers have recently been considered for helicopter lag damping augmentation because the forces generated by these dampers can be controlled by an applied magnetic field. In this dissertation, control schemes to integrate MR dampers with helicopters are developed and the influences of the MR dampers on rotorcraft ground resonance are studied.
CFD Based Unsteady Aerodynamic Modeling For Rotor Aeroelastic Analysis
(2003-12-02) Sitaraman, Jayanarayanan; Baeder, James D; Aerospace Engineering
A Computational Fluid Dynamics (CFD) analysis is developed for 3-D rotor unsteady aerodynamic load prediction. It is then coupled to a rotor structural analysis for predicting aeroelastic blade response, airloads and vibration. The CFD analysis accounts for the elastic deformations using a dynamically deforming mesh system. All the rotor blades are assumed to be identical, therefore to reduce the computational complexity the CFD calculations are performed for a single blade. This accounts for the near wake flow field. But the far wake effects because of the trailed tip vortices from all the blades have to be included separately. This is achieved by the use of the field velocity approach, which is a method for modeling unsteady flows via apparent grid movement. In this method, the induced velocity field caused by the trailed vortex wake is included by modifying the grid time metrics. The CFD method developed is systematically validated for a range of problems starting from simple 2-D model problems to full scale forward flight cases. The CFD analysis shows significant improvements in airloads prediction compared to a table lookup based lifting-line analysis. The CFD analysis is then used to investigate the fundamental mechanisms of rotor vibration. It is found that both the normal forces and pitching moments are dominated by three dimensional aerodynamic effects. The curvature introduced by the blade elasticity appears to play a key role in the generation of the vibratory harmonics in airloads. The pitching moments near the blade tip (85\% outboard) are significantly affected by transonic tip relief effects. The fundamental understanding of rotor vibrations gained from this study is then used to develop generic corrections for improving the accuracy of a lifting line analysis. Finally the CFD analysis developed is coupled with an advanced comprehensive rotor aeroelastic analysis. The coupling procedure is formulated in a way such that there is an exchange of information between the structural model and CFD model every rotor revolution. The coupled CFD/structure scheme is found to considerably improve the prediction of rotor vibratory airloads compared to the baseline rotor aeroelastic analysis which uses a lifting line based aerodynamic model.
A High-Order, Linear Time-Invariant Model for Application to Higher Harmonic Control and Flight Control System Interaction
(2003-12-04) Cheng, Rendy Po-Ren; Celi, Roberto; Aerospace Engineering
Helicopters can experience high vibration levels, which reduce passenger comfort and cause progressive damage to the aircraft structure and on-board equipment. Because the primary source of excitation is typically the main rotor, special rotor control systems have been proposed to reduce these vibrations at the source. This dissertation addresses one such system, generally known as ``Higher Harmonic Control" (HHC) because it consists of superimposing high frequency rotor inputs to the conventional low frequency ones used to control and maneuver the helicopter. Because both the primary flight control system and the HHC system act on the main rotor, the risk of adverse interactions between the two systems exists. This dissertation focuses on these interactions, which have never been studied before for the lack of suitable mathematical models. The key ingredient is an accurate linearized model of the helicopter, which includes the higher harmonic rotor response, and both the Automatic Flight Control System (AFCS) and the HHC system. Traditional linearization techniques lead to a system with periodic coefficients. Although Floquet theory can be used to study such periodic systems, there are far more control system design theories and software tools that are available for linear time-invariant systems than for periodic systems. Additionally, the theoretical evaluation of the handling qualities of the helicopter requires linear time-invariant systems. This research describes a new methodology for the extraction of a high-order, linear time invariant model, which allows the periodicity of the helicopter response to be accurately captured. This model provides the needed level of dynamic fidelity to permit an analysis and optimization of the AFCS and HHC algorithms. The key results of this study indicate that the closed-loop HHC system has little influence on the AFCS or on the vehicle handling qualities, which indicates that the AFCS does not need modification to work with the HHC system. On the other hand, the results show that the vibration response to maneuvers must be considered during the HHC design process, and this leads to much higher required HHC loop crossover frequencies. This research also demonstrates that the transient vibration responses during maneuvers can be reduced by optimizing the closed-loop higher harmonic control algorithm using conventional control system analyses.
comprehensive aeroelastic analysis of helicopter rotor with trailing-edge flap for primary control and vibration control
(2003-12-12) Shen, Jinwei; Chopra, Inderjit; Aerospace Engineering
A comprehensive aeroelastic analytical model of helicopter rotors with trailingedge flaps for primary and vibration controls has been developed. The derivation of system equations is based on Hamilton principles, and implemented with finite element method in space and time. The blade element consists of fifteen degrees of freedom representing blade flap, lag, torsional, and axial deformations. Three aerodynamic models of flapped airfoils were implemented in the present analysis, the unsteady Hariharan- Leishman model for trailing-edge flaps without aerodynamic balance, a quasi-steady Theodorsen theory for an aerodynamic balanced trailing-edge flap, and table lookup based on wind tunnel test data. The trailing-edge flap deflections may be modeled as a degree of freedom so that the actuator dynamics can be captured properly. The coupled trim procedures for swashplateless rotor are solved in either wind tunnel trim or free flight condition. A multicyclic controller is also implemented to calculate the flap control inputs for minimization of vibratory rotor hub loads. The coupled blade equations of motion are linearized by using small perturbations about a steady trimmed solution. The aeroelastic stability characteristics of trailing-edge flap rotors is then determined from an eigenanalysis of the homogeneous equations using Floquet method. The correlation studies of a typical bearingless rotor and an ultralight teetering rotor are respectively based on wind tunnel test data and simulations of another comprehensive analysis (CAMRAD II). Overall, good correlations are obtained. Parametric study identifies that the effect of actuator dynamics cannot be neglected, especially for a torsionally soft smart actuator system. Aeroelastic stability characteristics of a trailing-edge flap rotor system are shown to be sensitive to flap aerodynamic and mass balances. Key parameters of trailing-edge flap system for primary rotor control are identified as blade pitch index angle, torsional frequency, flap location, flap length, and overhang length. The swashplateless rotor is shown to achieve better rotor performance and overall more stable than the conventional configuration. Simulations of flaps performing both primary control and active vibration control are carried out, with the conclusion that trailing-edge flaps are capable of trimming the rotor and simultaneously minimizing vibratory rotor hub loads.
Design and Analysis of a Multi-Section Variable Camber Wing
(2004-03-31) Poonsong, Prasobchok; Pines, Darryll J; Aerospace Engineering
Minimizing fuel consumption is one of the major concerns in the aviation industry. In the past decade, there have been many attempts to improve the fuel efficiency of aircraft. One of the methods proposed is to vary the lift-to-drag ratio of the aircraft in different flight conditions. To achieve this, the wing of the airplane must be able to change its configuration during flight, corresponding to different flight regimes. In the research presented in this thesis, the aerodynamic characteristics of a multi-section, variable camber wing were investigated. The model used in this research had a 1-ft chord and a 1-ft wingspan, with the ribs divided into 6 sections. Two pneumatic actuators located at the main spar were used to morph the wing through mechanical linkages. The wing was tested in the free-jet wind tunnel at three different Reynolds numbers: 322000, 48000, and 636000. Static tests were performed to obtain lift and drag data for different configurations. Two rigid wings in baseline and camber configuration were built and tested to compare the test data with variable camber wing. The wind tunnel test results indicated that the multi-section variable camber wing provided a higher lift than the rigid wing in both configurations whereas high drag was also generated on the variable camber wing due to friction drag on the wing skin.
NONLINEAR OBSERVER/CONTROLLER DESIGNS FOR SPACECRAFT ATTITUDE CONTROL SYSTEMS WITH UNCALIBRATED GYROS
(2004-04-27) Thienel, Julie K.; Sanner, Robert M.; Aerospace Engineering
Gyroscopes, or gyros, are vital sensors in spacecraft onboard attitude control systems. Gyro measurements are corrupted, though, due to errors in alignment and scale factor, biases, and noise. This work proposes a class of adaptive nonlinear observers for calibration of spacecraft gyros. Observers for each of the calibration parameters are separately developed, then combined. Lyapunov stability analysis is used to demonstrate the stability and convergence properties of each design. First, an observer to estimate gyro bias is developed, both with and without added noise effects. The observer is shown to be exponentially stable without any additional conditions. Next a scale factor observer is developed, followed by an alignment observer. The scale factor and alignment observers are both shown to be Lyapunov stable. Additionally, if the angular velocity meets a persistency of excitation (PE) condition, the scale factor and alignment observers are exponentially stable. Finally, the three observers are combined, and the combination is shown to be stable, with exponential stability if the angular velocity is persistently exciting. The specific PE condition for each observer is given in detail. Next, the adaptive observers are combined with a class of nonlinear control algorithms designed to asymptotically track a general time-varying reference attitude. This algorithm requires feedback from rate sensors, such as gyros. The miscalibration discussed above will seriously degrade the performance of these controllers. While the adaptive observers can eliminate this miscalibration, it is not immediately clear that the observers can be safely combined with the controller in this case. There is, in general, no "separation principle" for nonlinear systems, as there is for linear systems. However, Lyapunov analysis of the coupled controller-observer dynamics shows that the closed-loop system will be stable for the class of observers proposed. With only gyro bias miscalibration, the closed-loop system is in fact asymptotically stable. For more general combinations of miscalibration, closed-loop stability is ensured with modest constraints on the observer/controller design parameters. These constraints are identified in detail. It is also shown that the constraints are not required if the angular velocity can be a priori guaranteed to be persistently exciting.
Quasi-Static Acoustic Mapping of Helicopter Blade-Vortex Interaction Noise
(2004-07-27) Gopalan, Gaurav; Schmitz, Fredric H; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
This research extends the applicability of storage-based noise prediction techniques to slowly maneuvering flight. The quasi-static equivalence between longitudinal decelerating flight and steady-state longitudinal descent flight, and its application to the estimation of BVI noise radiation under slow longitudinal maneuvering flight conditions, is investigated through various orders of flight dynamics modeling. The entire operating state of the helicopter is shown to be similar during equivalent flight conditions at the same flight velocity. This equivalence is also applied to the prediction of control requirements during longitudinal maneuvers. Inverse simulation based flight dynamics models of lower order are seen to capture many important trends associated with slow maneuvers, when compared with higher order modeling. The lower order flight dynamics model is used to design controlled maneuvers that may be practically flown during descent operations or as part of research flight testing. A version of a storage-based acoustic mapping technique, extended to slowly maneuvering longitudinal flight, is implemented for helicopter main rotor Blade-Vortex Interaction (BVI) noise. Various approach trajectories are formulated and analytical estimates of the BVI noise radiation characteristics associated with a full-scale two-bladed rotor are mapped to the ground using this quasi-static mapping approach. Multi-segment decelerating descent approaches are shown to be effective in ground noise abatement. The effects of steady longitudinal winds are investigated on radiated and ground noise. Piloting trim choices are seen to dominate the noise radiation under these flight conditions.
DESIGN, DEVELOPMENT AND TESTING OF A VARIABLE ASPECT RATIO WING USING PNEUMATIC TELESCOPIC SPARS
(2004-08-02) Blondeau, Julie Elizabeth; Pines, Darryll J; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
The purpose of this thesis is to discuss the design, development and testing of a pneumatic telescopic wing using pneumatic actuators that permit a change in the aspect ratio while simultaneously supporting structural wing loads. The key element of the wing is a pressurized telescopic spar that can undergo large-scale spanwise changes while under loadings in excess of 15 lbs/ft2. The wing cross-section is maintained by NACA0013 rib sections; telescopic skin sections preserve the span wise airfoil geometry and ensure compact storage and deployment of the telescopic wing. Several iterations led to a full-scale telescopic wing assembly that was tested in the Glenn L. Martin Wind Tunnel at the University of Maryland. These tests included measurements of Lift, Drag, and Lift to Drag ratio at a variety of Reynolds numbers. The telescopic wing was tested in several configurations and experimental results were compared to finite wing theory results. Preliminary aerodynamic results are promising for the variable aspect ratio telescopic wing.
Shock-Based Waverider Design with Pressure Gradient Corrections and Computational Simulations
(2004-08-10) Chauffour, Marie-Laure; Lewis, Mark J; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
Waveriders demonstrate good aerodynamic performance and thus are of special interest for hypersonic applications, especially for engine-airframe integration. The osculating cones waverider method is a generic shock-based derived waverider design method that allows prescribing a wide variety of flowfields at the inlet of the engine of the hypersonic vehicle. Previous osculating cones waveriders methods assumed that along the streamlines within the waverider shock layer, the pressure gradients in the azimuthal direction were negligible, and thus neglected it into the design process. The focus of this work is to investigate the magnitude of those pressure gradients, and integrate those into a new osculating cones waverider design method by modifying the derivation of the lower surface (streamsurface). The geometries resulting from the design code are to be compared with the previous solutions. The flowfield and aerodynamic performance predicted by the design code are compared with the results from Computational Fluid Dynamics simulations.