This article presents an extended state observer for a vehicle modeled as a rigid body in three-dimensional translational and rotational motions. The extended state observer is applicable to a multi-rotor aerial vehicle with a fixed plane of rotors, modeled as an under-actuated system on the state-space TSE(3), the tangent bundle of the six-dimensional Lie group SE(3). This state-space representation globally represents rigid body motions without singularities. The extended state observer is designed to estimate the resultant external disturbance force and disturbance torque acting on the vehicle. It guarantees stable convergence of disturbance estimation errors in finite time when the disturbances are constant, and finite time convergence to a bounded neighborhood of zero errors for time-varying disturbances. This extended state observer design is based on a H & ouml;lder-continuous fast finite time stable differentiator that is similar to the super-twisting algorithm, to obtain fast convergence. Numerical simulations are conducted to validate the proposed extended state observer. The proposed extended state observer is compared with other existing research to show its advantages. A set of experimental results implementing disturbance rejection control using feedback of disturbance estimates from this extended state observer is also presented.

We analyze a supersonic multi-stream rectangular jet flow using the linear-operator-based input-output method. The primary objective is to unravel the fundamental dynamics of perturbations and generate valuable insights for designing flow control strategies to minimize far-field noise and suppress near-field unsteadiness. We conduct a comprehensive resolvent analysis by leveraging a Large Eddy Simulation-derived time-averaged flowfield on the center plane as the base flow. A discounted resolvent analysis is performed over a finite-time window, revealing a reduction in flow response amplification with a higher discounted parameter. Exploring a wide range of frequencies and spanwise wavenumbers, we observe the maximum amplification near the dominant frequency of the flowfield. The leading forcing-response mode is located in different shear layer regions, subject to the combination of frequency and wavenumber. Furthermore, an input-output analysis provides insights into the componentwise amplification of the flow response, considering state variables and spatial restrictions. The streamwise input demonstrates maximum amplification near the dominant frequency and zero spanwise wavenumber across all cases. Intriguingly, a shift in the gain distribution curve is observed, indicating an optimal receptive region switch among shear layers based on input frequency at a specific wavenumber. Further, introducing spatial restrictions at the splitter plate trailing edge (SPTE) in an input yields distinct gain behaviors compared to cases without spatial constraints. The leading 2-D response mode at the dominant frequency resembles Kelvin-Helmholtz instabilities when input is restricted at SPTE. The findings from the resolvent analysis offer valuable guidance for parameter selection in active flow control design.

Synchronized time-resolved schlieren and far-field pressure measurements were used to examine control of a supersonic multi-aperture rectangular single- expansion ramp nozzle with an aft deck. The design contains core (Mach 1.6) and bypass (sonic) streams mixing behind a splitter plate, generating a high-frequency tone. Control was applied through splitter-plate trailing-edge spanwise waviness (passive), spanwise-arranged microjet blowing actuators (active), and their combination (hybrid). Passive and hybrid control successfully alleviate the tone, whereas active control amplifies and alters the tone frequency. Time-averaged schlieren data show changes to mean shock structures with control. Spectral proper orthogonal decomposition (SPOD) spectra from the raw schlieren data are consistent with the far-field measurements and reveal coherent structural details. Radiation associated with high-frequency structures is inhibited by passive and hybrid control that enhance mixing, but not by active control. Momentum potential theory filtering prior to SPOD application augments the analysis by isolating irrotational content, providing additional clues on acoustic vs hydrodynamic mechanisms. The results indicate shock–shear-layer interaction and shedding as sources of low- and high-frequency content, respectively. As such, through suitable processing, relatively easy-to-obtain high-speed schlieren provides a powerful tool to assess complex flow-control physics.

The present work performs a numerical investigation of a two-dimensional multi-stream supersonic jet flow with steady-blowing control. The nozzle flow of interest contains a main supersonic core flow at M=1.6 and a bypass flow at M=1.0, separated by a thick splitter plate, and an aft-deck is implemented to represent the structure of an aircraft. The mixing of the two flows forms a shear layer which breaks down to form vortical structures convecting downstream along the aft-deck. The shedding vortices produce tones that persist through out the flow field and generate intense surface pressure fluctuations. A relatively strong shock emanating from the proximity of the splitter plate trailing edge induces a region of flow separation along the nozzle wall. As a form of active flow control, a steady-blowing actuator, with injection at different angles, is introduced along the surfaces of the splitter plate. This control strategy aims to diminish the dominant shedding frequency, eliminate flow separation by weakening the shock strength, and reduce the high surface loading along the aft-deck. It is found that there is an an optimal actuation angle at each location that reduces the strength of the primary shock and mitigates flow separation. When actuation is introduced at the splitter plate trailing edge surface, the surface loading along the aft-deck is reduced for all actuation angles considered. Spectral analysis additionally confirms the dominant tone is reduced. Spectral proper orthogonal decomposition is applied to understand changes in the energy distribution when control is introduced. These results demonstrate the efficacy of steady-blowing control, prompting future investigation into unsteady control strategies to achieve the same control outcomes.

Experiments are conducted on a supersonic Multi-Aperture Rectangular Single Expansion Ramp Nozzle with both passive and active flow control techniques. The baseline flow is comprised of two physical streams, a core (M = 1.6) and bypass (M = 1.0) stream that coalesce behind a splitter plate, merge, and exit through a rectangular orifice onto an aft-deck plate meant to mimic the nozzle being integrated into an air-frame. This interaction between the streams produces a high-frequency tonal phenomena that propagates throughout the flow-field. Passive control via a geometric modification to the splitter plates trailing edge (SPTE) is implemented to alleviate this tone. Active control is implemented via steady blowing through a span-wise array of micro-jets in the near-splitter plate region. A hybrid control case combining the active and passive methods is also investigated and compared to the baseline flow. All tests were conducted at the jets nominal operating conditions and measured via time-resolved schlieren images with synchronized far-field pressure measurements. Spectral Proper Orthogonal Decomposition (SPOD) of the schlieren data sets breaks down the spatio-temporal scales and reveals an eradication of the high-frequency tone in the cases with the modified (passive) SPTE. Active control with the nominal SPTE amplifies the tone and reduces it’s frequency, while simultaneously affecting the mean shock structures. The far-field acoustic measurements confirm these observations. Momentum potential theory based filtering is applied to the schlieren data-sets to isolate the acoustic phenomena ejecting from the nozzle exit. The filtered data-sets are then analyzed via SPOD and reveal a shift in the spectra between the different control cases with the first shock-shear layer interaction location acting as a source for noise generation in all cases.

A framework for active control of a modern rectangular adaptive nozzle design is considered for analysis and comparison to a baseline flow. The nozzle of interest is a multi stream rectangular nozzle characterized by a core (M = 1.23) and bypass (M = 1.0) streams that coalesce behind a splitter and expand to achieve M = 1.6 flow before exiting onto an aft deck plate meant to simulate airframe integration. The control mechanism involves the introduction of a microjet array to blow sonic air into the region where the two streams meet. Steady blowing of the jet array seeks to alleviate the unsteadiness found to be associated with the complex instability that results from the two streams interacting, perturbing the shock train and propagating throughout the flowfield. Far-field pressure as well as Schlieren imaging, are used to fully quantify the baseline as well as the controlled flow for several actuation configurations. Steady blowing was shown to effect the near field shock structures as well as reduce the frequency of the dominant peak. Actuating perpendicular to the flow was shown to amplify the tone by up to 8.1 dB, whereas actuation at an angle to the incoming flow was shown to diminish the tone by 7.4 dB.

This article presents a tracking control framework enhanced by an extended state observer for a rotorcraft aerial vehicle modeled as a rigid body in three-dimensional translational and rotational motions. The system is considered as an underactuated system on the tangent bundle of the six-dimensional Lie group of rigid body motions, S E ( 3 ) . The extended state observer is designed to estimate the resultant external disturbance force and disturbance torque acting on the vehicle. It guarantees stable convergence of disturbance estimation errors in finite time when the disturbances are constant and finite time convergence to a bounded neighborhood of zero errors for time-varying disturbances. This extended state observer design is based on a Hölder-continuous fast finite time stable differentiator that is similar to the super-twisting algorithm, to obtain fast convergence. A tracking control scheme that uses the estimated disturbances from extended state observer for disturbance rejection, is designed to achieve fast finite-time stable tracking control. Numerical simulations are conducted to validate the proposed extended state observer and tracking control scheme with disturbance rejection. The proposed extended state observer is compared with other existing research to show its supremacy.

Weather, winds, thermals, and turbulence pose an ever-present challenge to small UAS. These challenges become magnified in rough terrain and especially within urban canyons. As the industry moves towards Beyond Visual Line of Sight (BVLOS) and fully autonomous operations, resilience to weather perturbations will be key. As the human decision-maker is removed from the in-situ environment, producing control systems that are robust will be paramount to the preservation of any Airspace System. Safety requirements and regulations require quantifiable performance metrics to guarantee a safe aerial environment with ever- increasing traffic. In this regards, the effect of wind and weather disturbances on a UAS and its ability to reject these disturbances present some unique concerns. Currently, drone manufacturers and operators rely on outdoor testing during windy days (or in windy locations) and onboard logging to evaluate and improve the flight worthiness, reliability and perturbation rejection capability of their vehicles. Waiting for the desired weather or travelling to a windier location is cost- and time-inefficient. Moreover, the conditions found on outdoor test sites are difficult to quantify and repeatability is non-existent. To address this situation, a novel testing methodology is proposed, combining artificial wind generation thanks to a multi-fan array wind generator (windshaper), coherent GNSS signal generation and accurate tracking of the test subject thanks to motion capture cameras. In this environment, the drone being tested can fly freely, follow missions and experience wind perturbations whilst staying in a modest indoor volume. By coordinating the windshaper, the motion tracking feedback and the position emulated by the GNSS signal generator with the drone’s mission profile, it was demonstrated that outdoor flight conditions can be reliably recreated in a controlled and repeatable environment. Specifically, thanks to real-time update of the position simulated by the GNSS signal generator, it was possible to demonstrate that the drone’s perception of the situation is similar to a corresponding mission being executed outdoor. In this work, the drone was subjected to three distinct flight cases: (1) hover in 2 m s−1 wind, (2) forward flight at 2 m s−1 without wind and (3) forward flight at 2 m s−1 with 2 m s−1 headwind. In each case, it could be demonstrated that by using indoor GNSS signal simulation and wind generation, the drone displays the characteristics of a 20 m move forward, while actually staying stationary in the test volume, within ±1 m. Further development of this methodology opens the door for fully integrated hardware-in- the-loop simulation of drone flight operations.

Passive control is applied to a modern adaptive engine design and its effectiveness compared to an unaltered configuration. The adaptive engine is represented by a supersonic multi-stream rectangular jet nozzle comprised of a core stream (M = 1.6) and bypass (M = 1.0) who merge behind a splitter plate and exit through a SERN onto an aft-deck plate. LES performed at The Ohio State University have identified the splitter plate trailing edge (SPTE) as a highly sensitive region for altering the overall nature of the flow. It has been shown that the shedding frequency initiated at the SPTE acts as a dominating tone throughout the flowfield. As a means for passive control, and an attempt to mitigate this tone, a spanwise perturbation has been introduced to the SPTE and implemented experimentally at Syracuse University. Far-field acoustics and near-field pressure measurements have confirmed the diminishment of the dominant tone for the wavy SPTE. To simulate the nozzle being integrated into an existing airframe, tests are conducted with three different aft deck geometries: a nominal length deck, half nominal, and no deck. PIV of the nominal configuration showed a slight change in the deflection angle of the jet plume. The differences for the half and no deck cases were less significant. These results along with future analyses will be used as a point of comparison between the two splitter plate configurations and act as means to validify the accuracy of the LES.

Supersonic multistream engines with single-expansion ramp nozzles use splitter plates to separate the core from bypass streams. The resulting shear-layer instability imposes unsteady loading on proximal surfaces and generates acoustic tones. The present work develops a passive control strategy to alleviate these effects. The thick splitter plate subcomponent of the full configuration is isolated, and Mach 1.6 and sonic boundary layers are allowed to develop on either side to form the shear layer. Instability mechanisms and optimal forcing-response characteristics are identified from linear analysis of time-averaged large-eddy simulations (LESs). These results are then used to guide splitter plate trailing-edge modifications with different wave numbers and amplitudes to interfere with the internal forcing mechanism. LESs of the altered configurations at small wave numbers reveal a diamond-grid-like pattern of cells and dislocations, similar to those observed in low-speed active control studies. Modal decomposition displays an increase in rank behavior, with the dominant mode in good agreement with linear predictions. At large wave numbers and amplitudes, where nonlinear effects are stronger, LESs show that control introduces streamwise vorticity into the wake, reduces coherent structure size (streamwise correlation length) and inhibits the tonal content. These effects are confirmed in companion experiments.

Experimental measurements are performed to analyze the effects of passive control on a supersonic multi-stream rectangular nozzle. The configuration explored consists of a supersonic core stream (M = 1.6) and a sonic (M = 1) bypass stream which merge behind a splitter plate exiting into a Single Expansion Ramp Nozzle (SERN) and onto an aft-deck. Previous studies have deduced that the aft-deck geometry can alter the plume deflection and farfield acoustics, while the splitter plate has an influence on the shock train development and unsteady loading on the aft-deck due to the shedding instability behind the splitter plate. This campaign seeks to exploit the inherent receptivity of these regions by performing geometric modifications as a form of passive control. The study is broken down into two separate parts, the first being the aft-deck changes. Aft-decks explored vary parameters such as length, width, and chamfer from the nominal design to observe the influence of each on the flow. Comparison to the nominal, half nominal, and no deck cases are in agreement with previous studies and show the plume deflection being a result of the shock train development. All deck modifications showed a slight upward deflection of the jet plume. The second effort of this study is the exploration of a sinusoidal spanwise wavenumber to the splitter plate trailing edge. Experimental design of the splitter plate is guided by Large Eddy Simulations (LES) performed by The Ohio State University. Particle Image Velocimetry and farfield acoustics are recorded for a wavenumber of 0.8 to match that of the LES. Both simulations and experiments show a reduction in the dominating tone. Results from both forms of passive control are compared with and used for validation of simulations.

A method for the prediction of far-field acoustics emanating from a complex supersonic rectangular nozzle using machine learning techniques is presented. Complexity and lengthy times associated with traditional experimental and numerical procedures motivate the use of a rapid noise prediction system based on modern machine learning computational methods. An artificial neural network (ANN) is employed to predict the far-field overall sound pressure levels (OASPL) for a Multi Aperture Rectangular Single Expansion Ramp Nozzle (MARS). A combination of experimental techniques and machine learning algorithms are utilized. The results yield a simple, accurate, and computationally frugal method for the prediction of far-field acoustic magnitudes and directions. Realistic operating conditions and geometric properties serve as inputs to this model which calculates OASPL at multiple far-field locations. This provides a basis for further exploration of geometric modifications and passive control schemes while continuing to explore the effects of a newly installed splitter plate trailing edge (SPTE) geometry. Preliminary Large Eddy Simulation (LES) data performed by The Ohio State University have shown that the introduction of a spanwise wavenumber to the SPTE results in a reduction of the 34 kHz tone associated with the shedding frequency of the splitter plate. With this dominating tone greatly reduced, a more in depth study into the acoustic effects of different aft deck lengths can be performed and ultimately optimized for a desired acoustic output.

The Large Eddy Simulation (LES) mean data of a flow past a splitter plate is analyzed using resolvent analysis with the objective of motivating a passive control strategy. The configuration of interest consists of two streams; a supersonic (Mach 1.23) upper stream and a sonic (Mach 1) lower stream, separated by a splitter plate of non-negligible thickness. Canonically, these two streams and the plate represent the supersonic multi-stream nozzle flow in an airframe-integrated variable-cycle engine architecture, with the upper stream being analogous to the power-producing core stream, and the lower stream being analogous to the bypass stream which shields the airframe from the thermal and acoustic loading of the core while providing some other design benefits. Numerical and experimental campaigns in previous studies have shown that this flow characteristically exhibits a two-dimensional vortex-shedding-like instability, which has a potentially detrimental first-order impact on performance. To gain insight into the instability mechanism, resolvent analysis is used to obtain the spatial structure of the optimal forcing of the mean flow. Modifications to the trailing edge of the plate, in the form of sinusoidal spanwise crests and troughs, are then introduced to interfere with this internal forcing mechanism and the resulting flows are simulated with LES. Proper Orthogonal Decomposition (POD) using snapshots of the LES data along the streamwise direction reveal a substantial increase in the rank of the modified flow, with a diminishing of the shedding instability as the spanwise wavenumber of the trailing-edge features is increased. Additionally, Dynamic Mode Decomposition (DMD) results reveal that the trailing-edge features induce a superposition of several resolvent response modes which rapidly decay downstream due to the non-linear action of turbulence.

Large Eddy Simulations (LES) and experimental measurements are performed to examine different aspects of a supersonic multi-stream jet representing a canonical airframe-integrated variable cycle engine architecture. The flowfield consists of two streams separated by a splitter plate; an upper supersonic core stream (Mach = 1.6) and a lower bypass stream (Mach = 1), which exit onto an aft-deck plate. Previous experimental and numerical efforts have shown that an instability associated with the splitter plate trailing edge exhibits a global influence on the flow because of its interaction with the shock train. The instability is also responsible for impairing the effectiveness of the bypass stream as a thermal and acoustic barrier between the core jet and the airframe. In the first part of this study, the potential for passive flow control by introducing sinusoidal spanwise modifications along the splitter plate edge is explored. The underlying mechanisms are explored computationally using LES on a simplified configuration that isolates the splitter plate. Results on two different spanwise wave numbers indicate that the sinusoidal trailing edge induces streamwise vorticity, which enhances mixing between the two streams and breaks up the shed structures seen previously. The tone is suppressed, together with small changes in the mean shock locations and shear layer trajectories. These spanwise modifications are in the process of being tested in the experimental facility using the full multistreamnozzle. In a complementary effort, geometricmodifications to the aft-deck and its effects on the flowfield are examined experimentally using the actual test rig. Results on the nominal deck length are compared with those from a half and double length deck, and without the deck. In each case, the shock train initiated at the SERN persists, and influences the downstream plume development. Although the impact on the tone is relatively small, the plume itself shows perceptible change. The extended deck results in a small downward movement of the plume relative to the nominal length, while the half nominal length displays larger spread of the lower shear layer initiated at the deck edge. Current efforts are focused on using the results to target complementary simulations and experiments for each tested modification.

The desire to develop faster, stealthy next-generation aircraft has led to the use of complex nozzles for aircraft integration. These nozzles contain multiple high-velocity streams that exit from non-axisymmetric areas. Due to the complexity of these nozzles, acoustic experiments and simulations are computationally and time expensive, making them not ideal for the design process. This study employs an artificial neural network (ANN) as a tool for rapid noise pre- diction of far-field acoustics based on geometric and flow parameters. The ANN is configured to predict the input features that minimize resulting far-field noise. To achieve this, the optimization strategies Genetic Algorithm (GA) and Particle Swarm Optimization (PSO) were used. The model was restricted to a third stream nozzle pressure ratio (NPR3) of 1.89 due to a low noise bucket that has been shown to occur, which correlates to perfectly expanded flow leaving the third stream. For this constraint, the model predicted that the lowest noise configuration occurs with a rectangular deck plate which extends as far as allowed by the network. With ESPL predictions within .15 dB of measured values, this study shows that an ANN can be utilized as a rapid noise prediction tool in the design process to determine low noise configurations when provided flow and geometry parameters of a supersonic jet nozzle.

The upstream turbulent velocity field of a quadrotor unmanned aerial vehicle (UAV) and itsrotor performance are investigated in experiments conducted in an outdoor wind tunnel. Thisresearch targets details on the dynamic interaction between upstream turbulent flow and UAV’s attitude control performance to support the development of more robust quadrotor controllers when exposed to turbulent and gusty environments. The quadrotor is exposed to a primary stream of free-shear flow with different turbulence intensities along with varying length and time scales in order to gain interaction information. Characterization of the upstream flow field is empirically calculated based on previous research. The UAV’s translational degree-of-freedom (DOF) is fixed downstream of the turbulent flow to test its rotational DOF (the attitude) controller performance within a turbulent environment. The Euler angles and their respective angular velocities are measured with the Inertial Measurement Unit (IMU) built inside the UAV flight controller.