This paper is the result of a master thesis project on helicopter rotor downwash effects using computational fluid dynamics (CFD). The work was performed at the department of Aerodynamics and Flight Mechanics at Saab AB, Linköping in 2008. It completes the author’s studies for a M.Sc degree in Applied Physics and Electrical Engineering at the Department of Electrical Engineering at the Linköping institute of technology (LiTH), Linköping, Sweden.   The aim of the project was to study the rotor downwash effects and its influence on the helicopter fuselage. To fulfil this purpose, several CFD calculations were carried out and the aerodynamic forces and moments resulting from the calculations were implemented in an existing simulation model, developed in-house at Saab. The original (existing) model was compared to the updated model by studying step responses in MATLAB, Simulink. For some step commands, the comparisions indicated that the updated model was more damped in yaw compared to the original model for the hovering helicopter. When the helicopter was trimmed for a steady turn, the states in the updated model diverged much faster than the states in the original model for any given step command.     In order to investigate the differences between the original helicopter model and the updated model from a controlling perspective, a linear quadratic (LQ) state feedback controller was synthesized to stabilize the vehicle in a steady turn. The LQ method was chosen as it is a modern design technique with good robustness and sensitivity properties and since it is easily implemented in MATLAB.  Before synthesising, a simplification of the helicopter model was made by reducing states and splitting them into lateral and longitudinal ones. Step responses from simulations with the original and the updated model were studied, showing an almost identical behavior.   It can be concluded that the aerodynamic coefficients obtained from the CFD calculations can be used for determining the aerodynamic characteristics of the helicopter. Some further validation is needed though, for example by comparing the results with flight test data. In order to build an aerodynamic data base that covers the whole flight envelop, additional CFD calculations are required.

A helicopter based system equipped with a millimetre-wave fire control radar is studied for improvement of the detection probability for moving ground targets. The system is a pulse doppler radar that sends out the electromagnetic wave in repetitive pulses. The frequency in which radar sends this pulse is called Pulse Repetition Frequency (PRF). The velocity of the moving targets is measured by spectral analysis of the pulses received. With this sort of system some targets will not be possible to detect. These targets are said to be blind for the radar. Also the measured targets can be ambiguous. A solution to this is to shift PRF during the time the radar emits itspulses on the target. This method is called PRF block staggering. The antenna is mounted on a mechanical sweep on top of the rotor of the helicopter. Because of this the time on target is limited. This means that the measured velocity resolution of the radar is also limited. With the studied method it is possible to resolve range and velocity ambiguities. This is under the condition that it is acceptable with a reduction in the velocity resolution by a factor equal to the number of PRF blocks used. The work leading to this thesis included developing a program in Mathworks Matlab for studying different configurations of PRFs and the effect on the detection probability.

The helicopter is a Multiple-Input Multiple-Output (MIMO) system with highly coupled characteristics, which increases the complexity of the system dynamics. In addition, the system dynamics of the helicopter are unstable, referring to its tendency to deviate from an equilibrium when disturbed. Despite the complexity in its modelling and control, the benefit of using a helicopter for unmanned, autonomous applications can be tremendous. One particular application that motivates this research is the use of an unmanned small-scale helicopter in an autonomous survey mission over an area struck by disaster, such as an earthquake. The work presented in this thesis provides a framework for utilizing a platform system for research and development of small-scale helicopter systems. A platform system enables testing and analysis to be performed indoor in a controlled environment. This can provide a more convenient mean for helicopter research since the system is not affected by environmental elements, such as wind, rain or snow condition. However, the presence of the platform linkages poses challenges for analysis and controller design as it alters the helicopter system flight dynamics. Through a six degree-of-freedom (6 DOF) platform model derived in this research, the criteria for matching the trim conditions between the platform system and a stand alone helicopter have been identified. With the matched trim conditions, linearization is applied to perform analysis on the effects that the platform has on the system dynamics. The results of the analysis provide insights into both the limitations and benefits of utilizing the platform system for helicopter research. Finally, a Virtual Joint Control scheme is proposed as an unified control strategy for both the platform and the stand alone helicopter systems. Having a consistent control scheme between the two systems allows for comparisons between simulation and experimental results for the two systems to be made more readily. Furthermore, the Virtual Joint Control scheme represents a novel flight control strategy for stand alone helicopter systems.

<p>Spatial variability of the Earth's surface has a considerable impact on the atmosphere at all scales and understanding the mechanisms involved in land-atmosphere interactions is hindered by the scarcity of appropriate observations. A measurement gap exists between traditional point sensors and large aircraft and satellite-based sensors in collecting measurements of atmospheric quantities. Point sensors are capable of making long time series of measurements, but cannot make measurements of spatial variability. Large aircraft and satellites make measurements over large spatial areas, but with poor spatial and temporal resolution. A helicopter-based platform can make measurements on scales relevant for towers, especially close to the Earth's surface, and can extend these measurements to account for spatial variability. Thus, the Duke University Helicopter Observation Platform (HOP) is designed to fill the existing measurement gap. </p><p>Because measurements must be made in such a way that they are as uncontaminated by the platform itself as much as is possible, it is necessary to quantify the aerodynamic envelope of the HOP. The results of an analytical analysis of the location of the main rotor wake at various airspeeds are shown. Similarly, the results of a numerical analysis using the commercial Computational Fluid Dynamics software Fluent are shown. The optimal flight speed for the sampling of turbulent fluxes is found to be around 30 m/s. At this airspeed, the sensors located in front of the nose of the HOP are in advance of the wake generated by the main rotor. This airspeed is also low enough that the region of high pressure due to the stagnation point on the nose of the HOP does not protrude far enough forward to affect the sensors. Measurements of differential pressures, variables and turbulent fluxes made while flying the HOP at different airspeeds support these results. No systematic effects of the platform are seen at airspeeds above about 10 m/s.</p><p>Processing of HOP data collected using the current set of sensors is discussed, including the novel use of the Empirical Mode Decomposition (EMD) to detrend and filter the data. The EMD separates the data into a finite number of Impirical Mode Functions (IMFs), each of which is unique and orthogonal. The basis is determined by the data itself, so that it need not be known a priori, and it is adaptive. The EMD is shown to be an ideal tool for the filtering and detrending of HOP data using data gathered during the Cloud and Land Surface Interaction Campaign (CLASIC). </p><p>The ability of the HOP to accurately measure atmospheric profiles of potential temperature is demonstrated. During experiments conducted in the marine boundary layer (MBL) and the convective boundary layer (CBL), HOP profiles are evaluated using profiles from an elastic backscatter lidar. The HOP and the lidar agree on the height of the boundary layer in both cases, and the HOP effectively locates other atmospheric structures.</p><p>Atmospheric sensible and latent heat fluxes, turbulence kinetic energy (TKE) and horizontal momentum fluxes are also measured, and the resulting information is used to provide context to tower-based data collected concurrently. A brief comparison made over homogeneous ocean conditions yields good results. A more exhaustive evaluation is made using short HOP flights made over an orchard during the Canopy Horizontal Turbulence Study (CHATS).</p> / Dissertation

The transformation of Joint forces to be lighter, more lethal, and capable of deploying from multiple dispersed locations free of prepared landing zones requires a dedicated heavy lift VTOL aircraft capable of rapidly delivering large payloads, such as the 20 to 26 ton Future Combat System, at extended ranges in demanding terrain and environmental conditions. Current estimates for a single main rotor configuration place the design weight over 130,000 pounds with an installed power of approximately 30,000 horsepower. Helicopter drive systems capable of delivering torque of this magnitude succeeded in the Russian Mi-26 helicopters split-torque design and the Boeing VERTOL Heavy Lift Helicopter (HLH) prototypes traditional multi-stage planetary design. The square-cube law and historical trends show that the transmission stage weight varies approximately as the two-thirds power of torque; hence, as the size and weight of the vehicle grows, the transmissions weight becomes an ever-increasing portion of total gross weight. At this scale, optimal gearbox configuration and component design holds great potential to save significant weight and reduce the required installed power. The drive system design methodology creates a set of integrated tools to estimate system weight and rapidly model the preliminary design of drives system components. Tools are provided for gearbox weight estimation and efficiency, gearing, shafting, and cooling. Within the same architecture, the designer may add similar tools to model subcomponents such as support bearings, gearbox housing, freewheeling units, and rotor brakes. Measuring the relationships between key design variables and system performance metrics reveals insight into the performance and behavior of a heavy lift drive system. A parametric study of select design variables is accomplished through an intelligent Design of Experiments that utilizes Response Surface Methodology to build a multivariate regression weight model. The model permits visualization of the design space and assists in optimization of the drive system preliminary design. This methodology is applied to both the Boeing HLH and the Russian Mi-26 main gearboxes. This study applies the drive system design methodology to compare the Mi-26 split-torque gearbox over the Boeing HLH multi-stage planetary gearbox in a single main rotor heavy lift helicopter.

This thesis extends the use of neural-network-based model reference adaptive control to systems that occur as cascades. In general, these systems are not feedback linearizable. The approach taken is that of approximate feedback linearization of upper subsystems whilst treating the lower-subsystem states as virtual actuators. Similarly, lower-subsystems are also feedback linearized. Typically, approximate inverses are used for linearization purposes. Model error arising from the use of an approximate inverse is minimized using a neural-network as an adaptive element. Incorrect adaptation due to (virtual) actuator saturation and dynamics is avoided using the Pseudocontrol Hedging method. Using linear approximate inverses and linear reference models generally result in large desired pseudocontrol for large external commands. Even if the provided external command is feasible (null-controllable), there is no guarantee that the reference model trajectory is feasible. In order to mitigate this, nonlinear reference models based on nested-saturation methods are used to constrain the evolution of the reference model and thus the plant states. The method presented in this thesis lends itself to the inner-outer loop control of air vehicles, where the inner-loop controls attitude dynamics and the outer-loop controls the translational dynamics of the vehicle. The outer-loop treats the closed loop attitude dynamics as an actuator. Adaptation to uncertainty in the attitude, as well as the translational dynamics, is introduced, thus minimizing the effects of model error in all six degrees of freedom and leading to more accurate position tracking. A pole-placement approach is used to choose compensator gains for the tracking error dynamics. This alleviates timescale separation requirements, allowing the outer loop bandwidth to be closer to that of the inner loop, thus increasing position tracking performance. A poor model of the attitude dynamics and a basic kinematics model is shown to be sufficient for accurate position tracking. In particular, the inner-outer loop method was used to control an unmanned helicopter and has subsequently been applied to a ducted-fan, a fixed-wing aircraft that transitions in and out of hover, and a full-scale rotorcraft. Experimental flight test results are also provided for a subset of these vehicles.

Recent increases in computing power and memory have created renewed interest in alternative grid schemes such as unstructured grids, which facilitate rapid grid generation by relaxing restrictions on grid structure. Three rotor models have been incorporated into a popular fixed-wing unstructured computational fluid dynamics (CFD) solver to increase its capability and facilitate availability to the rotorcraft community. The benefit of unstructured grid methods is demonstrated through rapid generation of high fidelity configuration models. The simplest rotor model is the steady state actuator disk approximation. By transforming the unsteady rotor problem into a steady state one, the actuator disk can provide rapid predictions of performance parameters such as lift and drag. The actuator blade and overset blade models provide a depiction of the unsteady rotor wake, but incur a larger computational cost than the actuator disk. The actuator blade model is convenient when the unsteady aerodynamic behavior needs to be investigated, but the computational cost of the overset approach is too large. The overset or chimera method allows the blades loads to be computed from first-principles and therefore provides the most accurate prediction of the rotor wake for the models investigated. The physics of the flow fields of these models for rotor/fuselage interaction are explored, along with efficiencies and limitations of each methodology.

A helicopter rotor in descent flight encounters its own wake, resulting in a doughnut-shaped ring around the rotor disk, known as the Vortex Ring State (VRS). Flight in VRS condition can be dangerous as it may cause uncommanded drop in descent rate, loss of control effectiveness, power settling, excessive thrust and torque fluctuations, and vibration. As simple momentum theory is no longer valid for a rotor in VRS, modeling of rotor inflow in VRS continues to challenge researchers, especially for flight simulation applications. In this dissertation, a simplified inflow model, called the ring vortex model, is developed for a helicopter rotor operating in descent condition. By creating a series of vortex rings near the rotor disk, the ring vortex model addresses the strong flow interaction between the rotor wake and the surrounding airflow in descent flight. In addition, the total mass flow parameter in the existing inflow models is augmented to create a steady state transition between the helicopter and the windmill branches. With the ring vortex model, rotor inflow can now be adequately predicted over a wide range of descent rates. Validations of the ring vortex model for helicopter rotors are conducted extensively in axial and inclined descent. Effects from blade taper, blade twist, and rotor thrust are also investigated with further application of the finite-state inflow model. The ring vortex model is applied to a single main-rotor helicopter. The main effort is to establish VRS boundary based on heave stability criterion. In addition, two important phenomena observed in the descent flight tests are addressed in the dynamic simulation, including uncommanded drop in descent rate and loss of collective control effectiveness. The ring vortex model is further applied to a side-by-side rotor configuration. Lateral thrust asymmetry on the side-by-side rotor configuration can be reproduced through uneven distribution of vortex rings at the two rotors. Two important issues are investigated, including the impact of vortex rings on lateral thrust deficit and on lateral AFCS limit.

Assembling of the compliant parts used in aviation industry is a challenging process. Assembly fixtures are quite important tools in this effort and widely used in industry. In fixturing of easily deformable sheet metal parts, besides restraining the rigid body motion of the parts, the possible deformations that may occur during the assembly process and the spring-back effect on the final product need to be taken in to consideration. In order to guarantee a successful assembling, in other words, to obtain the final product within specified tolerances, a systematic approach to the fixture design problem is required. The designer should predict the correlation between the input variations and the final assembly variation, especially, for the complex assemblies. This study proposes a design and analysis approach in fixturing of sheet metal assemblies for helicopter components. The design of an assembly fixture for a particular tail cone has been completed convenient to the existing locating principles. Finite Element Analysis (FEA) has been realized in simulating the assembling process in order to predict the possible variation of the interested feature on a complex assembly due to deformations.

In this thesis the development of a nonlinear simulation model of a utility helicopter and the design of its automatic flight control system is addressed. In the first part of this thesis, the nonlinear dynamic model for a full size helicopter is developed using the MATLAB/SIMULINK environment. The main rotor (composed of inflow and flapping dynamics parts), tail rotor, fuselage, vertical stabilizer, horizontal stabilizer of the helicopter are modeled in order to obtain the total forces and moments needed for the flight simulation of the helicopter. Total forces and moments are used in 6 degrees of freedom equations of motion model and helicopter states are calculated for the specified flight conditions such as hover and forward flight. Trim and linearization programs are developed. The linearized models of hover and forward flight conditions are used for the automatic flight control system design. Automatic flight control system model consists of necessary systems in order to ease the pilot control of the helicopter. A classical inner stability loop and outer flight directory mode approach is taken to design the automatic flight control system. For the inner stability loop both classical rate feedback and truncated system state feedback control approaches are used. The outer loop modes implemented are heading hold, attitude hold (pitch, roll), altitude acquire and hold mode for hover condition and heading hold, attitude hold (pitch, roll), altitude acquire and hold mode and airspeed hold for forward flight condition. Finally, the success of the controllers are demonstrated through nonlinear simulations for different flight directory modes in hover and forward flight conditions.