Aerospace Engineering Theses and Dissertations

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Abort Trajectories for Manned Lunar Missions
(2007-01-22) Beksinski Jr, Edward David; Lewis, Mark; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
With NASA's renewed focus towards a permanent human presence on the moon, comes the development of the Crew Exploration Vehicle. Unforeseen circumstances can induce emergency situations necessitating contingency plans to ensure crew safety. It is therefore desirable to define the feasibility of a direct abort from an outbound translunar trajectory. Thus an astrodynamic model for lunar transfer has been developed to allow for characterization of the abort feasibility envelope for conceivable transfer orbits. In addition the model allows for several trade studies involving differently executed abort options, factoring in fuel margins. Two optimization schemes were utilized; one to expedite return via any fuel in excess of that required for the abort, and one to explore the boundary region of direct abort infeasibility envelope searching for plausible abort trajectories. The characterization and optimization of translunar abort trajectories for the Crew Exploration Vehicle can ensure increased crew survivability in emergency situations.
Accurate SLAM With Application For Aerial Path Planning
(2013) Friedman, Chen; Chopra, Inderjit; Rand, Omri; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
This thesis focuses on operation of Micro Aerial Vehicles (MAVs), in previously unexplored, GPS-denied environments. For this purpose, a refined Simultaneous Localization And Mapping (SLAM) algorithm using a laser range scanner is developed, capable of producing a map of the traversed environment, and estimating the position of the MAV within the evolving map. The algorithm's accuracy is quantitatively assessed using several dedicated metrics, showing significant advantages over current methods. Repeatability and robustness are shown using a set of 12 repeated experiments in a benchmark scenario. The SLAM algorithm is primarily based on an innovative scan matching approach, dubbed Perimeter Based Polar Scan Matching (PB-PSM), which introduces a maximum overlap term to the cost function. This term, along with a tailored cost minimization technique, are found to yield highly accurate solutions for scan matching pairs of range scans. The algorithm is extensively tested on both ground and aerial platforms, in indoor as well as outdoor scenarios, using both in-house and previously published datasets, utilizing several different laser scanners. The SLAM algorithm is then coupled with a global A* path planner, and applied on a single rotor helicopter, performing targeted flight missions using a pilot-in-the- loop implementation. Targeted flight is defined as navigating to a goal position, defined by relative distance from a known initial position. It differs from the more common task of mapping, as it may not rely on loop closure opportunities to smooth out errors and optimize the generated map. Therefore, the importance of position estimates accuracy increases dramatically. The complete algorithm is then used for targeted flight experiments with a pilot in the loop. The algorithm presents the pilot with nothing but heading information. In order to further prevent the pilot from interfering with the obstacle avoidance task, the evolving map and position are not shown to the human pilot. Furthermore, the scenario is introduced with artificial (invisible) obstacles, apparent only to the path planner. The pilot therefore has to adhere to the path planner directions in order to reach the goal while avoiding all obstacles. The resulting paths show the helicopter successfully avoid both real and artificial obstacles, while following the planned path to the goal.
An Acoustic Analysis of Valveless Pulsejet Engines
(2015) Maqbool, Daanish; Cadou, Christopher P; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
Valveless pulsejets are extremely simple aircraft engines; essentially cleverly designed tubes with no moving parts. These engines utilize pressure waves, instead of machinery, for thrust generation, and have demonstrated thrust-to-weight ratios over 8 and thrust specific fuel consumption levels below 1 lbm/lbf-hr – performance levels that can rival many gas turbines. Despite their simplicity and competitive performance, they have not seen widespread application due to extremely high noise and vibration levels, which have persisted as an unresolved challenge primarily due to a lack of fundamental insight into the operation of these engines. This thesis develops two theories for pulsejet operation (both based on electro-acoustic analogies) that predict measurements better than any previous theory reported in the literature, and then uses them to devise and experimentally validate effective noise reduction strategies. The first theory analyzes valveless pulsejets as acoustic ducts with axially varying area and temperature. An electro-acoustic analogy is used to calculate longitudinal mode frequencies and shapes for prescribed area and temperature distributions inside an engine. Predicted operating frequencies match experimental values to within 6% with the use of appropriate end corrections. Mode shapes are predicted and used to develop strategies for suppressing higher modes that are responsible for much of the perceived noise. These strategies are verified experimentally and via comparison to existing models/data for valveless pulsejets in the literature. The second theory analyzes valveless pulsejets as acoustic systems/circuits in which each engine component is represented by an acoustic impedance. These are assembled to form an equivalent circuit for the engine that is solved to find the frequency response. The theory is used to predict the behavior of two interacting pulsejet engines. It is validated via comparison to experiment and data in the literature. The technique is then used to develop and experimentally verify a method for operating two engines in anti-phase without interfering with thrust production. Finally, Helmholtz resonators are used to suppress higher order modes that inhibit noise suppression via anti-phasing. Experiments show that the acoustic output of two resonator-equipped pulsejets operating in anti-phase is 9 dBA less than the acoustic output of a single pulsejet.
Active Control of Performance and Vibratory Loads using Trailing-Edge Flaps and Leading-Edge Slats
(2019) Ravichandran, Kumar; Chopra, Inderjit; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
The objective of this research work is to develop a comprehensive analysis version of UMARC (University of Maryland Advanced Rotor Code) to study the capabilities of trailing-edge flaps (TEFs)and leading-edge slats (LESs) for helicopter vibration reduction and performance improvement and rec- ommend flap and slat configurations for a typical helicopter rotor such as UH-60A rotor, which maximize these benefits. This study uses propulsive free-flight trim except in hover. Using TEFs , the rotor performance in hover was improved with a combination of torsionally softer blades and positive TEF deflections. For the vibration reduction studies, a multicyclic control algorithm was used to determine the actuation schedule . Suitable combi- nations of lower harmonic TEF inputs were shown capable of reducing the rotor power requirement by about 4-5 % at an advance ratio of μ = 0.4. The TEF was shown to be capable of suppressing the vibratory loads at a range of forward speeds, using half peak-to-peak deflections of about 5 °-10 ° . Softening the blades in torsion resulted in larger flap actuation requirements for vibration reduction. Parametric sweeps of TEF actuations were carried out to determine suitable combinations of steady and various frequencies of actuation of flaps , which yield overall power reductions and it is observed that a combination of 1, 2, 3, 4 and 5/rev TEF inputs resulted in a power reduction of 1.5% , while also reducing certain vibratory loads by more than 50% in high speed-forward flight. To explore the advantages of leading-edge slats, the slatted airfoils with configurations S0, S1 and S6 (used by Sikorsky) were used. The slatted blade sections had the SC2110 baseline/slatted airfoils in place of the baseline UH- 60A airfoils. Dynamic actuations are chosen to retain the high-lift benefits of the slats while seeking to minimize profile drag penalties over regions of the rotor disk operating at lower angles of attack, i.e., the advancing side. The effects of leading-edge slats extending over 20%, 30% and 40% of the blade span on rotor performance and vibratory hub loads were examined. The study uses propulsive free flight trim. In moderate to high-speed forward flight,leading-edge slats were shown to enhance the maximum rotor thrust by 15-30% at advance ratios larger than 0.2 and reduce power requirements by 10-20% at high thrust levels. 20% span slats offered a good compromise between power reductions and adverse effect on vibratory hub loads. The rotor with leading-edge slats could be trimmed at a maximum forward speed that was about 20 knots greater than for the baseline rotor with no slats. This study also shows that additional power reduction is achievable by suitable TEF deployments superimposed on certain slat actuations in high-speed forward flight.
ACTIVE JET ACOUSTIC CONTROL OF LOW FREQUENCY, IN-PLANE HELICOPTER HARMONIC NOISE
(2012) Sargent, Daniel Caleb; Schmitz, Fredric H.; Baeder, James D.; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
A new approach to reducing low frequency, in-plane harmonic noise of helicopter rotors is explored theoretically and experimentally in this dissertation. The active jet acoustic control methodology employs on-blade, tip located unsteady air blowing to produce an acoustic anti-noise waveform that reduces or cancels the observed noise at targeted positions in the acoustic far-field of the rotor system. This effectively reduces the distance at which the helicopter rotor can be aurally detected. An extended theoretical model of the subsonic air jet, which is modeled as both a source of mass and momentum, is presented. The model is applied to a baseline, full-scale, medium weight helicopter rotor for both steady and unsteady blowing. Significant reductions in low frequency, in-plane harmonic noise are shown to be possible for the theoretical rotor system by using physically reasonable unsteady jet velocities. A new model-scale active jet acoustic control experimental test rotor system is described in detail. Experimental measurements conducted in the University of Maryland Acoustic Chamber for the ~1/7th rotor, operated at a full-scale hover tip Mach number of 0.661, indicate that active jet acoustic control is a viable option for reducing low frequency, in-plane harmonic noise. Good correlation between theoretical predictions and measured data for four valve control cases are observed in both the time and frequency domains. Model-scale limitations of the tip-jet blowing experiment limited the peak noise level reductions to 30%. However, theory suggests that if the limitations of the model-scale controller are mitigated, much larger noise reductions are possible.
Active Spanwise Lift Control: A Distributed Parameter Approach
(2019) Dias, Joaquim; Hubbard, James E; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
Structural load alleviation has been a very active research topic since the 1950s for many reasons. By mitigating the effect of gusts on the wing, the maximum loads can be effectively reduced. This capability would lead to substantial benefits, such as reduced structural weight, better fuel burn performance, and improved passenger ride comfort. Instead of controlling the structural response, however, it can be argued that the aerodynamic behavior of the wing should be primarily controlled. Since the gust loads are caused by disturbances in the lift distribution, it is possible to mitigate the gust loads by controlling the shape of the lift distribution profile along the span. In contrast to previous approaches, this research builds on concepts from Distributed Parameter Systems (DPS), which is indeed the case of aerodynamic surfaces. The unsteady aerodynamic behavior of the 3D flow around a wing is modeled using two approaches: the Unsteady Lifting-Line Theory (ULLT) and the Unsteady Vortex-Lattice Method (UVLM). Then, modal identification techniques are used to identify spanwise aerodynamic mode shapes in terms of local lift coefficient along the span. These shapes provide an optimal basis for model order reduction and also for spatial control. The lift distribution is decomposed as a linear superposition of these shapes, with each weighted by a shape coefficient. By controlling a set of shape coefficients, the overall lift profile can be effectively controlled. In this work, the shape control problem is addressed using a Linear Quadratic Tracker to dynamically follow any desired reference lift profile. The gust alleviation problem is investigated using a similar controller with a special observer, able to decouple the state estimation from the gust input.
Active Suppression of Vortex-Driven Combustion Instability Using Controlled Liquid-Fuel Injection
(2005-09-09) Pang, Bin; Yu, Kenneth H; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
Combustion instabilities remain one of the most challenging problems encountered in developing propulsion and power systems. Large amplitude pressure oscillations, driven by unsteady heat release, can produce numerous detrimental effects. Most previous active control studies utilized gaseous fuels to suppress combustion instabilities. However, using liquid fuel to suppress combustion instabilities is more realistic for propulsion applications. Active instability suppression in vortex-driven combustors using a direct liquid fuel injection strategy was theoretically established and experimentally demonstrated in this dissertation work. Droplet size measurements revealed that with pulsed fuel injection management, fuel droplet size could be modulated periodically. Consequently, desired heat release fluctuation could be created. If this oscillatory heat release is coupled with the natural pressure oscillation in an out of phase manner, combustion instabilities can be suppressed. To identify proper locations of supplying additional liquid fuel for the purpose of achieving control, the natural heat release pattern in a vortex-driven combustor was characterized in this study. It was found that at high Damköhler number oscillatory heat release pattern closely followed the evolving vortex front. However, when Damköhler number became close to unity, heat release fluctuation wave no longer coincided with the coherent structures. A heat release deficit area was found near the dump plane when combustor was operated in lean premixed conditions. Active combustion instability suppression experiments were performed in a dump combustor using a controlled liquid fuel injection strategy. High-speed Schlieren results illustrated that vortex shedding plays an important role in maintaining self-sustained combustion instabilities. Complete combustion instability control requires total suppression of these large-scale coherent structures. The sound pressure level at the excited dominant frequency was reduced by more than 20 dB with controlled liquid fuel injection method. Scaling issues were also investigated in this dump combustor to test the effectiveness of using pulsed liquid fuel injection strategies to suppress instabilities at higher power output conditions. With the liquid fuel injection control method, it was possible to suppress strong instabilities with initial amplitude of 5 psi down to the background noise level. The stable combustor operating range was also expanded from equivalence ratio of 0.75 to beyond 0.9.
ACTIVE VIBRATION ATTENUATING SEAT SUSPENSION FOR AN ARMORED HELICOPTER CREW SEAT
(2013) Sztein, Pablo Javier; Wereley, Norman M; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
An Active Vibration Attenuating Seat Suspension (AVASS) for an MH-60S helicopter crew seat is designed to protect the occupants from harmful whole-body vibration (WBV). Magnetorheological (MR) suspension units are designed, fabricated and installed in a helicopter crew seat. These MR isolators are built to work in series with existing Variable Load Energy Absorbers (VLEAs), have minimal increase in weight, and maintain crashworthiness for the seat system. Refinements are discussed, based on testing, to minimize friction observed in the system. These refinements include the addition of roller bearings to replace friction bearings in the existing seat. Additionally, semi-active control of the MR dampers is achieved using special purpose built custom electronics integrated into the seat system. Experimental testing shows that an MH-60S retrofitted with AVASS provides up to 70.65% more vibration attenuation than the existing seat configuration as well as up to 81.1% reduction in vibration from the floor.
Adaptive Magnetorheological Seat Suspension for Shock Mitigation
(2014) Singh, Harinder Jit; Wereley, Norman M; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
This research focuses on theoretical and experimental analysis of an adaptive seat suspension employing magnetorheological energy absorber with the objective of minimizing injury potential to seated occupant of different weights subjected to broader crash intensities. The research was segmented into three tasks: (1) development of magnetorheological energy absorber, (2) biodynamic modeling of a seated occupant, and (3) control schemes for shock mitigation. A linear stroking semi-active magnetorheological energy absorber (MREA) was designed, fabricated and tested for intense impact conditions with piston velocities up to 8 m/s. MREA design was optimized on the basis of Bingham-plastic model (BPM model) in order to maximize the energy absorption capabilities at high impact velocities. Computational fluid dynamics and magnetic FE analysis were conducted to validate MREA performance. Subsequently, low-speed cyclic testing (0-2 Hz subjected to 0-5.5 A) and high-speed drop testing (0-4.5 m/s at 0 A) were conducted for quantitative comparison with the numerical simulations. Later, a nonlinear four degrees-of-freedom biodynamic model representing a seated 50th percentile male occupant was developed on the basis of experiments conducted on Hybrid II 50th percentile male anthropomorphic test device. The response of proposed biodynamic model was compared quantitatively against two different biodynamic models from the literature that are heavily implemented for obtaining biodynamic response under impact conditions. The proposed biodynamic model accurately predicts peak magnitude, overall shape and the duration of the biodynamic transient response, with minimal phase shift. The biodynamic model was further validated against 16 impact tests conducted on horizontal accelerator facility at NAVAIR for two different shock intensities. Compliance effects of human body were also investigated on the performance of adaptive seat suspension by comparing the proposed biodynamic model response with that of a rigid body response. Finally, three different control schemes were analyzed for maximizing shock attenuation using semi-active magnetorheological energy absorber. High-speed drop experiments were conducted by dropping two rigid payloads of 240 and 340 lb mass from heights of 35 and 60 inch to simulate different impact intensities. First control scheme called constant stroking load control offered inflexible stroking load irrespective of varying impact severity or occupant weight. The other two control schemes: terminal trajectory control and optimal control adapted stroking load as per the shock intensity. The control schemes were compared on the basis of their adaptability and ease of implementation. These tools can serve as the basis for future research and development of state-of-the-art crashworthy seat suspension designs that further enhance occupant protection compared to limited performance of existing passive crashworthy concepts.
Adaptive Magnetorheological Sliding Seat System for Ground Vehicles
(2011) Mao, Min; Wereley, Norman M.; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
Magnetorheological (MR) fluids (MRFs) are smart fluids that have reversible field dependent rheological properties that can change rapidly (typically 5 - 10 ms time constant). Such an MRF can be changed from a free flowing fluid into a semi-solid when exposed to a magnetic field. The rapid, reversible, and continuous field dependent variation in rheological properties can be exploited in an MRF-based damper or energy absorber to provide adaptive vibration and shock mitigation capabilities to varying payloads, vibration spectra, and shock pulses, as well as other environmental factors. Electronically controlled electromagnetic coils are typically used to activate the MR effect and tune the damping force so that feedback control implementation is practical and realizable. MR devices have been demonstrated as successful solutions in semi-active systems combining advantages of both passive and active systems for applications where piston velocities are relatively low (typically < 1 m/s), such as seismic mitigation, or vibration isolation. Recently strong interests have focused on employing magnetorheological energy absorbers (MREAs) for high speed impact loads, such as in helicopter cockpit seats for occupant protection in a vertical crash landing. This work presents another novel application of MREAs in this new trend - an adaptive magnetorheological sliding seat (AMSS) system utilizing controllable MREAs to mitigate impact load imparted to the occupant for a ground vehicle in the event of a low speed frontal impact (up to 15 mph). To accomplish this, a non-linear analytical MREA model based on the Bingham-plastic model and including minor loss effects (denoted as the BPM model) is developed. A design strategy is proposed for MREAs under impact conditions. Using the BPM model, an MREA is designed, fabricated and drop tested up to piston velocities of 5 m/s. The measured data is used to validate the BPM model and the design strategy. The MREA design is then modified for use in the AMSS system and a prototype is built. The prototype MREA is drop tested and its performance, as well as the dynamic behavior in the time domain, is described by the BPM model. Next, theoretical analysis of the AMSS system with two proposed control algorithms is carried out using two modeling approaches: (1) a single-degree-of-freedom (SDOF) rigid occupant (RO) model treating the seat and the occupant as a single rigid mass, and (2) a multi-degree-of-freedom (MDOF) compliant occupant (CO) model interpreting the occupant as three lumped parts - head, torso and pelvis. A general MREA is assumed and characterized by the Bingham-plastic model in the system model. The two control algorithms, named the constant Bingham number or Bic control and the constant stroking force or Fc control, are constructed in such a way that the control objective - to bring the payload to rest while fully utilizing the available stroke - is achieved. Numerical simulations for both rigid and compliant occupant models with assumed system parameter values and a 20 g rectangular crash pulse for initial impact speeds of up to 7 m/s (15.7 mph) show that overall decelerations of the payload are significantly reduced using the AMSS compared to the case of a traditional fixed seat. To experimentally verify the theoretical analysis, a prototype AMSS system is built. The prototype seat system is sled tested in the passive mode (i.e. without control) for initial impact speeds of up to 5.6 m/s and for the 5th percentile female and the 95th percentile male. Using the test data, the CO model is shown to be able to adequately describe the dynamic behavior of the prototype seat system. Utilizing the CO model, the control algorithms for the prototype seat system are developed and a prototype controller is formulated using the DSPACE and SIMULINK real time control environments. The prototype seat system with controller integrated is sled tested for initial impact speeds of up to 5.6 m/s for the 5th female and 95th male (only the 95th male is tested for the Bic control). The results show that the controllers of both control algorithms successfully bring the seat to rest while fully utilizing the available stroke and the decelerations measured at the seat are substantially mitigated. The CO model is shown to be effective and a useful tool to predict the control inputs of the control algorithms. Thus, the feasibility and effectiveness of the proposed adaptive sliding seat system is theoretically and experimentally verified.
Advancing Nitrous Oxide As A Monopropellant Using Inductively Heated Heat-Exchangers: Theory and Experiment
(2019) Saripalli, Pratik Sharma; Sedwick, Raymond J; Yu, Kenneth; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
Most monopropellant thrusters used for attitude control and station keeping employ hydrazine as their propellant. In recent years, significant effort has been focused on finding an alternative due to its high toxicity. This work focuses on advancing nitrous oxide, a green monopropellant with a strong performance capability, as a replacement for current monopropellant thrusters. A large emphasis is placed on trying to address catalyst degradation experienced in most thrusters due to the high temperatures from decomposition. The approach described here eliminates the dependence for a high catalytic surface area, typically decreased from degradation, and catalysts altogether by using high temperature porous heat exchangers. A 1-D numerical compressible fluid model was created to model a typical decomposition chamber and simulate self-sustained decomposition of nitrous oxide. It implements a preheated, thermally-conductive, metal foam as the heat exchanger. An extensive parameter study was conducted to help understand thermal and fluid effects on steady-state decompositions. Using a copper metal foam, steady-state solutions simulated successful nitrous oxide decomposition, with an exit gas temperature around 1345 K. Simulations were extended to other high temperature metal foams with different thermal conductivities and melting points. Modeling flow rate conditions more representative of current monopropellant thrusters required scaling of the decomposition chamber in order to be self-sustaining. Experiments were conducted using results from the numerical simulations as guidelines. Three different heat exchangers (copper metal foam, copper discs, and stainless-steel discs), all of which have significantly less effective surface area than nominal catalysts used in thrusters, were tested for nitrous oxide decomposition. These heat exchangers were preheated to thermal decomposition temperatures using an inductive heating system and placed in a vacuum bell jar to mitigate heat loss to the environment. Testing with copper metal foam resulted in complete degradation of the heat exchanger due to oxidation from nitrous oxide decomposition. A set of copper discs, uniquely designed to maximize tortuosity of the flow, was implemented in an attempt to address the oxidation issues. While the preliminary test did confirm steady-state decomposition of nitrous oxide within the heat exchanger, further tests resulted in temperatures exceeding the melting point of copper within the discs. The last heat exchanger was a set of stainless-steel discs of the same design. Repeated tests all successfully achieved steady-state decomposition of nitrous oxide within a two-minute interval.
Advancing the Multi-Solver Paradigm for Overset CFD Toward Heterogeneous Architectures
(2019) Jude, Dylan P; Baeder, James; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
A multi-solver, overset, computational fluid dynamics framework is developed for efficient, large-scale simulation of rotorcraft problems. Two primary features distinguish the developed framework from the current state of the art. First, the framework is designed for heterogeneous compute architectures, making use of both traditional codes run on the Central Processing Unit (CPU) as well as codes run on the Graphics Processing Unit (GPU). Second, a framework-level implementation of the Generalized Minimal Residual linear solver is used to consider all meshes from all solvers in a single linear system. The developed GPU flow solver and framework are validated against conventional implementations, achieving a 5.35× speedup for a single GPU compared to 24 CPU cores. Similarly, the overset linear solver is compared to traditional techniques, demonstrating the same convergence order can be achieved using as few as half the number of iterations. Applications of the developed methods are organized into two chapters. First, the heterogeneous, overset framework is applied to a notional helicopter configuration based on the ROBIN wind tunnel experiments. A tail rotor and hub are added to create a challenging case representative of a realistic, full-rotorcraft simulation. Interactional aerodynamics between the different components are reviewed in detail. The second application chapter focuses on performance of the overset linear solver for unsteady applications. The GPU solver is used along with an unstructured code to simulate laminar flow over a sphere as well as laminar coaxial rotors designed for a Mars helicopter. In all results, the overset linear solver out-performs the traditional, de-coupled approach. Conclusions drawn from both the full-rotorcraft and overset linear solver simulations can have a significant impact on improving modeling of complex rotorcraft aerodynamics.
Aero Database Development and Two-Dimensional Hypersonic Trajectory Optization for the High-speed Army Reference Vehicle
(2023) James, Brendan; Brehm, Christoph; Larsson, Johan; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
Steady-flow inviscid and simulations of the High-Speed Army Reference Vehicle geometry were performed within the CHAMPS solver framework at Mach numbers of 4, 6, and 8, and an integrated streamline method was used to apply viscous corrections for Reynolds numbers up to 2x10^8. For each flow Mach, angle of attack sweeps from -10° to +10° were used to determine baseline drag, lift, and moment coefficient alpha dependencies. Coefficient values were then interpolated across Mach, alpha, and Reynolds number parameter spaces to construct an aerodynamic force coefficient database for use in two-dimensional flight simulation and trajectory optimization. By simulating flight with a maximum lift-to-drag control input, sample trajectories for determining maximum vehicle range were produced. A proportional-navigation (PN) controller was implemented which allowed for the targeting of specific altitudes throughout the progression of a trajectory. The PN controller and simulation schemes were then utilized in genetic-algorithm optimization to produce trajectory profiles for achieving minimum time-to-target for gliding flight in standard atmospheric conditions. Over the examined range of initial altitudes, Mach numbers, and release angles, the fastest trajectories were consistently shown to be those which achieved or maintained stratospheric altitudes and consequently benefited from significantly reduced drag before performing a nose-over maneuver for an accurate ground strike.
Aero-Assisted Spacecraft Missions Using Hypersonic Waverider Aeroshells
(2015) Knittel, Jeremy; Yu, Kenneth; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
This work examines the use of high-lift, low drag vehicles which perform orbital transfers within a planet’s atmosphere to reduce propulsive requirements. For the foreseeable future, spacecraft mission design will include the objective of limiting the mass of fuel required. One means of accomplishing this is using aerodynamics as a supplemental force, with what is termed an aero-assist maneuver. Further, the use of a lifting body enables a mission designer to explore candidate trajectory types wholly unavailable to non-lifting analogs. Examples include missions to outer planets by way of an aero-gravity assist, aero-assisted plane change, aero-capture, and steady atmospheric periapsis probing missions. Engineering level models are created in order to simulate both atmospheric and extra-atmospheric space flight. Each mission is parameterized using discrete variables which control multiple areas of design. This work combines the areas of hypersonic aerodynamics, re-entry aerothermodynamics, spacecraft orbital mechanics, and vehicle shape optimization. In particular, emphasis is given to the parametric design of vehicles known as “waveriders” which are inversely designed from known shock flowfields. An entirely novel means of generating a class of waveriders known as “starbodies” is presented. A complete analysis is performed of asymmetric starbody forms and compared to a better understood parameterization, “osculating cone” waveriders. This analysis includes characterization of stability behavior, a critical discipline within hypersonic flight. It is shown that asymmetric starbodies have significant stability improvement with only a 10% reduction in the lift-to-drag ratio. By combining the optimization of both the shape of the vehicle and the trajectory it flies, much is learned about the benefit that can be expected from lifting aero-assist missions. While previous studies have conceptually proven the viability, this work provides thorough quantification of the optimized outcome. In examining an aero-capture of Mars, it was found that with a lifting body, the increased maneuverability can allow completion of multiple mission objectives along with the aero-capture, such as atmospheric profiling or up to 80 degrees of orbital plane change. Completing a combined orbital plane change and aero-capture might save as much as 4.5 km/s of velocity increment while increasing the feasible entry corridor by an order of magnitude. Analyzing a higher energy mission type, a database of maximum aero-gravity assist performance is developed at Mars, Earth and Venus. Finally, a methodology is presented for designing end-toend interplanetary missions using aero-gravity assists. As a means of demonstrating the method, promising trajectories are propagated which reduce the time of flight of an interstellar probe mission by up to 50%.
Aeroacoustic Analysis of Asymmetric Lift-Offset Helicopter in Forward Flight
(2021) Arias, Paulo; Baeder, James; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
In recent years, the University of Maryland has worked on an asymmetric lift-offset compound helicopter. The configuration consists of a single main rotor helicopter with the addition of two key ways to increase the forward speed: a stubbed wing on the retreating fuselage side, and a slowed down rotor. Experiments and simulations have shown that the novel concept provides improved thrust potential and lift-to-drag ratios in high-speed forward flight. This study aims to determine whether the asymmetric lift-offset configuration also provides aeroacoustic benefits in forward flight in addition to its aerodynamic advantages. The aerodynamic results from previous computational and experimental studies are recreated using the Mercury framework, in-house Computational Fluid Dynamics solver based on Reynolds-Averaged Navier-Stokes (RANS) coupled to a comprehensive rotor analysis for structural deformations and trim. The acoustic analysis is performed using an acoustic code based on the Ffowcs William-Hawkings equation to solve for the tonal noise propagating from the surfaces of the aircraft. The BPM model is used for broadband noise prediction. It was found that for an advance ratio of 0.5 the wing-lift offset configuration can produce 56.8% more thrust at the same collective angle without any penalties in total noise. When the configurations produce equal thrust it was found that the wing-lift offset case has a 4 dB reduction in maximum overall sound pressure level. At an advance ratio of 0.3 with trim for equivalentthrust between configurations, a 3 dB maximum OASPL reduction was obtained with the inclusion of the wing. The rotor of the wing-lift offset case was also slowed down while maintaining equal thrust to find a 6 dB reduction at an advance ratio of 0.55. Blade flap and lag bending moments near the root were also significantly reduced for the wing-lift offset configuration with equal thrust.
Aeroacoustic Implications of Installed Propeller Interactional Aerodynamics and Transient Propeller Motions
(2023) Jayasundara, Dilhara; Baeder, James; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
The emergence of advanced air mobility and sustainable aviation concepts have revived the interest in propeller-driven aircraft. A number of electric vertical take-off and landing (eVTOL) aircraft have been developed to cater to the demands of urban air mobility (UAM) and significant advancements have been made in unmanned aerial vehicles (UAV) equipped with vertical take-off and landing capabilities. However, the community acceptance of these new aircraft configurations highly depends on having a low noise footprint as they will operate in dense urban environments. Propeller noise is considered the major source of noise in these aircraft with the introduction of electric propulsion and it can significantly increase with the effects of installation and transient propeller motions. This study aims to comprehend the complex aerodynamic interactions within such aircraft that result from propeller installation and contribute to the generation of high noise levels. To understand the physics of propeller installation, a wingtip-mounted propeller was analyzed at several angles of attack using computational fluid dynamics (CFD) based on Reynolds-averaged Navier-Stokes (RANS) equations and computational aeroacoustics based on the Ffowcs Williams - Hawkings equation. The aeroacoustic implications of the propeller axis inclination and the propeller-wing aerodynamic interaction were studied in-depth. The propeller-wing interaction leads to a significant increase in propeller noise (~20 to 30 dB increase along the rotational axis) and causes the wing to generate a loading noise in the same order of magnitude as the propeller noise. To extrapolate the understanding of installation effects to a full aircraft, the aeroacoustic characteristics of a quadrotor biplane tailsitter were analyzed in both hover and forward flight focusing on the rotor-rotor and rotor-airframe aerodynamic interaction. The rotor-rotor interaction was found to be a significant source of loading noise in hover but having the fuselage as a physical barrier between the rotors largely reduces its effect. The airframe loading noise and rotor broadband noise are equally dominant as the rotor tonal noise when the aircraft is in forward flight. Moreover, the study evaluated the effectiveness of rotor synchrophasing in reducing the aircraft noise footprint and it showed promising results in hover, causing a reduction of aircraft noise by more than 10 dB. Furthermore, an efficient computational aeroacoustics framework was developed to facilitate the computations, ensuring optimal utilization of the computational resources. The CPU and GPU parallelization and other optimization techniques were able to achieve a 98% reduction in computation time for an isolated propeller case. This enabled the rapid aeroacoustic computations of periodic and non-periodic problems. This was used to analyze the aeroacoustics of an isolated propeller undergoing a transition from hover to forward flight. The aerodynamic and acoustic results of the unsteady case were compared with quasi-steady cases performed at intermediate tilt angles. The quasi-steady CFD simulations predicted the unsteady transition aeroacoustics with reasonable accuracy. A tilting quasi-steady approach was proposed to better capture the aerodynamics and acoustics of the unsteady transition.
Aerodynamic Analysis and Simulation of a Twin-Tail Tilt-Duct Unmanned Aerial Vehicle
(2010) Abdollahi, Cyrus; Humbert, James S; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
The tilt-duct vertical takeoff and landing (VTOL) concept has been around since the early 1960s; however, to date the design has never passed the research and development phase. Nearly 50 years later, American Dynamics Flight Systems (ADFS) is developing the AD-150, a 2,250lb weight class unmanned aerial vehicle (UAV) configured with rotating ducts on each wingtip. Unlike its predecessor, the Doak VZ-4, the AD-150 features a V tail and wing sweep- both of which affect the aerodynamic behavior of the aircraft. Because no aircraft of this type has been built and tested, vital aerodynamic research was conducted on the bare airframe behavior (without wingtip ducts). Two weeks of static and dynamic test were performed on a 3/10th scale model at the University of Maryland's 7'x10' low speed wind tunnel to facilitate the construction of a nonlinear flight simulator. A total of 70 dynamic tests were performed to obtain damping parameter estimates using the ordinary least squares methodology. Validation, based on agreement between static and dynamic estimates of the pitch and yaw stiffness terms, showed an average percent error of 14.0% and 39.6%, respectively. These inconsistencies were attributed to: large dynamic displacements not encountered during static testing, regressor collinearity, and, while not conclusively proven, dierences in static and dynamic boundary layer development. Overall, the damping estimates were consistent and repeatable, with low scatter over a 95% confidence interval. Finally, a basic open loop simulation was executed to demonstrate the instability of the aircraft. As a result, it is recommended that future work be performed to determine trim points and linear models for controls development.
Aerodynamic Analysis of an MAV-Scale Cycloidal Rotor System Using a Stuctured Overset RANS Solver
(2010) Yang, Kan; Baeder, James D; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
A compressible Reynolds-Averaged Navier-Stokes solver was used to investigate the performance and flow physics of the cycloidal rotor (cyclocopter). This work employed a computational methodology to understand the complex aerodynamics of the cyclocopter and its relatively unexplored application for MAVs. The numerical code was compared against performance measurements obtained from experiment and was seen to exhibit reasonable accuracy. With validation of the flow solver, CFD predictions were used to gain qualitative insight into the flowfield. Time histories revealed large periodic variations in thrust and power. In particular, the virtual camber effect was found to significantly influence the vertical force time history. Spanwise thrust and flow visualizations showed a highly three-dimensional flowfield with large amounts of blade shedding and blade-vortex interaction. Overall, the current work seeks to provide unprecedented insight into the cyclocopter flowfield with the goal of developing an accurate predictive tool to refine the design of future cyclocopter configurations.
Aerodynamic Design Optimization of Proprotors for Convertible-Rotor Concepts
(2012) Stahlhut, Conor; Leishman, J. Gordon; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
Trades in the aerodynamic design of proprotors that could be used to power convertible-rotor aircraft have been examined. The key design challenge is to maximize overall aerodynamic efficiency of the proprotor in both hover and forward flight, while preserving adequate stall margins for maneuvering flight and compressibility margins for high speed flight. To better assess proprotor performance, a new formulation of the blade element momentum theory for high-speed propellers and proprotors was developed. This approach uses an efficient and robust numerical method to solve simultaneously for the axial and swirl induced velocity components in the wake of the proprotor. The efficacy of the approach was validated against measurements of the performance of two NACA high-speed propellers at advance ratios up to 2.5 and tip Mach numbers up to supersonic conditions. The importance of calculating accurately the swirl component of the induced velocity is emphasized. Parametric studies and design optimization studies were performed for different convertible-rotor aircraft platforms with the goal of developing a better understanding of the tradeoffs that would be needed for the development of advanced proprotors to power such convertible-rotor aircraft. The effects that solidity, diameter, rotational speed, blade twist and taper, number of blades, tip sweep, and airfoil characteristics have on proprotor performance were all explored. Particular importance was given to proprotors that may have variable tip speed, and assessing the relative advantages of variable diameter versus variable rotational shaft speed concepts. Proprotors with variable blade twist were also considered. It was found that significant improvements in proprotor performance may only be practically realized by varying one or more of diameter, shaft speed, or blade twist during flight.
Aerodynamic Modeling of a Flapping Membrane Wing Using Motion Tracking Experiments
(2008) Harmon, Robyn Lynn; Hubbard Jr., James E; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
An analytical model of flapping membrane wing aerodynamics using experimental kinematic data is presented. An alternative to computational fluid dynamics, this experimental method tracks small reflective markers placed on two ornithopter membrane wings. Time varying three dimensional data of the wing kinematics and the corresponding aerodynamic loads were recorded for various flapping frequencies. The wing shape data was used to form an analytical aerodynamic model that uses blade element theory and quasi-steady aerodynamics to account for the local twist, stroke angle, membrane shape, wing velocity and acceleration. Results from the aerodynamic model show adequate correlation between the magnitude of lift and thrust produced but some phase errors exist between the measured and calculated force curves. This analytical model can be improved by comparison with a RANS CFD solver which provides insight into the fluid behavior. Implications on the membrane wing design are also presented.