Thesis (M. Phil.)--University of Hong Kong, 2010. / Includes bibliographical references (p. 83-85). Also available in print.

published_or_final_version / Mechanical Engineering / Master / Master of Philosophy

Ducted fan VTOL UAVs are currently being designed for use by the Army for surveillance and reconnaissance in the battlefield. The vehicle tested in this research is part of the Honeywell MAV program. Able to be carried in a backpack by a single soldier, it has a duct diameter of 11.5-in. and weighs approximately 20-lbs at takeoff. It is designed for flight up to 50-knots with fan speeds varying between 5000 and 8500 RPM. Reynolds numbers, based on the duct diameter, were on the order of Re = 0.96 x 106 to 4.6 x 106. Various control effectors were mounted to the vehicle and tested for maximum control authority to reduce the nose-up pitch moment created during forward flight or with crosswinds present. Static and wind tunnel tests were conducted to measure each control effector's performance. Box vanes, mounted downstream of the duct, utilized four assemblies, each with three vanes and one flap. This configuration showed adequate results in generating a nose-down pitch moment. Duct deflectors, mounted on the windward strut upstream of the duct, reduced the windward lift produced by both the fan and duct, resulting in a nose-down pitch moment. Opposed vanes, utilizing two vanes and flaps for each of the four assemblies, were designed with the same surface area as the box vanes. With each pair capable of independent vane movement, the opposed vanes were tested with the vanes rotating in tandem for a basic elevator deflection and with the leading edges touching, disrupting the flow at the duct exit. Opposed vanes combined the capabilities of the box vanes and duct deflector. Results show that the opposed vanes were the most practical control effector, offering the most control authority for maximum nose-down pitch moment (up to 80% better) with minimal loss in thrust. / Master of Science

In order to improve the flight performance of the Martin Jetpack research was undertaken to investigate the aerodynamic issues that were limiting the P-11A Jetpack's flight envelope. Through research of existing ducted-fan aircraft, a flight model describing the unique aerodynamics of the Martin Jetpack was developed using Matlab®/Simulink® software. The dynamic flight model, which can be ran in real time, includes the reactions from: ducted-fans, aircraft body aerodynamics, control surfaces, gyration and landing gear interactions. Abstract Numerous experiments were designed to quantify and validate assumptions used in the development of the model equations. The experiments took advantage of the small size of the Jetpack by designing and building test apparatuses that measured reactions directly on the actual aircraft. This avoided scaling issues that are traditionally encountered when employing wind tunnels for aerodynamic measurements. Abstract Implementing the experimental results into the model led to the modifications of the existing Jetpack airframe to produce the P-11C Jetpack prototype, which significantly improved the performance of the aircraft. The collected flight data was used to validate the model and good agreement was achieved. Abstract Based on this research a new Jetpack prototype (P-12) was developed that combined the flight performance of the P-11C Jetpack with the ability to carry a man or manned sized payload. The model was used to design the layout and to size the control vanes for the P-12 Jetpack. Further research was performed to design larger rotor and stator blades required for the P-12 Jetpack prototype. Abstract The developed model allows the user to efficiently evaluate various control methodologies and changes to key aerodynamic features of the aircraft to aid in the design and flying of the Martin Jetpack. Abstract The outcome of this research is a better understanding of the ducted-fan technology, and via the development of the Jetpack flight model, correctly applying this understanding to improve the Jetpack's flight performance.

Unmanned Aerial Vehicle (UAV) is any flying vehicle which is not controlled by actual human pilots sitting in the cockpit but is installed with proper avionics that can either fly autonomously or by using the commands from its base. Some rotorcraft UAVs use a ducted propeller for two main reasons- safety and to increase the thrust produced by the propellers. While ducted rotors can increase the thrust produced, it also adds weight to the UAV. It was therefore hypothesized that by removing part of the duct materials (i.e. adding perforations in the duct) would benefit from both decreased duct weight and increased thrust. However, it is not clear how much trade-off would be between these two factors. Hence, the objective of this study is to explore the relationship between the change of thrust and addition of different numbers or sizes of perforations. Cases with and without duct, and duct with perforations were simulated using a commercial computational fluid dynamic (CFD) software Ansys/Fluent. The physics of the rotating propeller was modeled by a simplified disc with a pressure jump across an infinitesimal volume. Three different RPM speeds of the propellers were simulated by varying the strength of the pressure jump. The results show that the thrust decreases as the duct is added. As perforations are added, the result shows that with more perforations (i.e. more open area on the duct wall), the thrust increases accordingly until the thrust reaches a maximum value without the duct. The result is in contrast to a published experimental data stating that installation of duct can increase thrust. It is speculated that the current duct with a flat wall has caused such difference from the experimental data. Further study is recommended to continue more detailed computational simulation using a duct with cambered airfoil configuration to reduce the aerodynamic losses.

Experiments to generate multiple shock waves in an axisymmetric model at hypersonic speeds were conducted in a small reflected shock tunnel. Conical surfaces were used to generate shock waves inside a circular duct chosen to be representative of a scramjet combustor. These shock waves impinged on turbulent boundary layers to produce shock wave/boundary layer interactions (SWBLIs). In the process of observing this phenomenon, the commonly used empirical correlations of Korkegi were tested for accuracy, i.e. the combined pressure ratio across these shocks can be measured and compared to that predicted by these correlations. Korkegi correlates only with Mach number, and is independent of Reynolds number and on how the pressure is applied. A major contribution of this work is to examine how the details of the compression process effect separation. In this study, the history of applying the compression was varied. An analytical method was developed for theoretically estimating the onset of incipient separation using an integrated computation of the momentum flux contained in the boundary layer. By including the summed (negative) contribution of wall shear stress on the integrated momentum flux, the upstream history of the boundary layer was considered. The overall result has a form similar to the Korkegi correlations, plus an additional correction term relating to momentum loss through wall shear stress. The correction term was determined to be a second order effect, which explains why the Reynolds number independent Korkegi correlations work so well over such a large range of conditions. A hypersonic flow test condition conducive to the generation of high Reynolds number flows and turbulent boundary layer production was developed in a small reflected shock tunnel. The experimentally measured flow parameters were matched by numerical simulation using a number of in-house codes at The University of Queensland. This has allowed the unmeasured parameters which are numerically derived to be stated with greater confidence. An internal centre-body with a conical forebody was used to generate conditions of incipient separation. This provided benchmark data for comparison with subsequent experiments with multiple compressions. A semi-vertex angle of 15o was selected based on Large Eddy Simulation (LES) numerical results once the experimental and numerical static wall pressure and heat flux were matched. A two-cone experimental model, which provided for adjustment of the axial separation between the two shock systems, was tested at the same flow conditions as used in the single-cone experiments. A technique of incrementally moving the instrumentation (relative to the centre-body) and repeating the same condition to achieve high resolution in pressure and heat flux distributions with a limited number of transducers was successful. The results verified that it was possible to subject a hypersonic turbulent boundary layer to two quantified compression-expansion systems with an adjustable axial separation between them and capture the first reflected shock in a “shock trap” to remove it's influence from the second SWBLI. The data from this initial two-cone model provided non-separated pressure and heat flux data which was used as a reference to help interpret data from separated flows. The commercially available Reynolds Averaged Navier-Stokes (RANS) numerical code, CFD-Fastran, was used to help design an experimental model which produces boundary layer separation. Algebraic and two-equation turbulence models were applied to a modified two-cone model to show greater pressure rises which would produce boundary layer separation. A modified two-cone model was tested and demonstrated boundary layer separation. Three configurations with varying axial separation between SWBLIs were tested which all produced separation. The configuration that produced the largest pressure ratio and largest separation region at the second SWBLI may represent a geometry whereby the distance from the hollow cylinder inlet and the second cone may represent a critical value. The amount of viscous interaction, generated from the leading edge of the shock trap, and the proximity of the two interactions may be coupled to produce higher than expected values. It is postulated that the boundary layer momentum recovery for the configuration where the second SWBLI was furthest downstream (30 mm configuration), prevented severe separation from occurring. An in-house RANS code, elmer3, was used to simulate the flow of the modified two-cone model. An algebraic turbulence model was applied to this model and comparisons of experimentally measured static wall pressure and heat flux have given good agreement. The wall shear stress was investigated to provide further information concerning the position and size of flow reversal regions. The use of the numerical codes utilised in this study has reinforced their effectiveness for model design and comparison of experimental results.

<p>The ducted fan UAV is an unmanned aerial vehicle consisting mainly of a propeller enclosed in a open ended tube. The UAV has the same basic functions as an ordinary helicopter UAV but has several advantages to the same.</p><p>This thesis aims to estimate the performance of the concept of the ducted fan UAV. The company where this thesis has been written, DST Control AB, is currently investigating the economical possibilities to continue the development of this kind of UAV. This thesis shall provide DST Control AB with a theoretical as well as experimental ground for the investigation by estimation the lift capacity, position accuracy and wind tolerance.</p><p>A ducted fan UAV prototype and a mathematical model for that UAV have been developed by DST Control AB and a student project at Linköping University. The model is constructed through pure physical modeling. Several noise sources have been added to better fit the reality. Several experiments have been conducted to validate the model with satisfying results. Experiments to determine the lift capacity of the craft have also been conducted. These experiments showed a slightly smaller lift capacity than the theoretically calculated lift capacity. The wind tolerance has not been tested in experiments because of the lack of available wind tunnels but simulations have given an estimation of this tolerance.</p><p>To estimate the position accuracy, two different control systems have been implemented. The simplest control system is a system consisting of several PID controllers. The system is divided into two separate subsystems connected in cascade. The inner subsystem takes the pitch, roll and yaw angle as inputs and gives the rudder angles as outputs. The outer subsystem takes the inertial position as input and gives roll, pitch and yaw as outputs. Together, the two subsystems can be used to control the entire craft. The inner subsystem has also been replaced with a small LQ Compensator. An LQ Compensator for the entire system is also implemented giving about as good performance as the PID controller and better performance than the PID/LQ combination.</p>

The ducted fan UAV is an unmanned aerial vehicle consisting mainly of a propeller enclosed in a open ended tube. The UAV has the same basic functions as an ordinary helicopter UAV but has several advantages to the same. This thesis aims to estimate the performance of the concept of the ducted fan UAV. The company where this thesis has been written, DST Control AB, is currently investigating the economical possibilities to continue the development of this kind of UAV. This thesis shall provide DST Control AB with a theoretical as well as experimental ground for the investigation by estimation the lift capacity, position accuracy and wind tolerance. A ducted fan UAV prototype and a mathematical model for that UAV have been developed by DST Control AB and a student project at Linköping University. The model is constructed through pure physical modeling. Several noise sources have been added to better fit the reality. Several experiments have been conducted to validate the model with satisfying results. Experiments to determine the lift capacity of the craft have also been conducted. These experiments showed a slightly smaller lift capacity than the theoretically calculated lift capacity. The wind tolerance has not been tested in experiments because of the lack of available wind tunnels but simulations have given an estimation of this tolerance. To estimate the position accuracy, two different control systems have been implemented. The simplest control system is a system consisting of several PID controllers. The system is divided into two separate subsystems connected in cascade. The inner subsystem takes the pitch, roll and yaw angle as inputs and gives the rudder angles as outputs. The outer subsystem takes the inertial position as input and gives roll, pitch and yaw as outputs. Together, the two subsystems can be used to control the entire craft. The inner subsystem has also been replaced with a small LQ Compensator. An LQ Compensator for the entire system is also implemented giving about as good performance as the PID controller and better performance than the PID/LQ combination.

A two-dimensional viscous-inviscid interactive boundary layer method is applied to three dimensional problems of flow around ducts and ducted propellers. The idea is to predict the effects of fluid viscosity on three dimensional geometries, like ducts, using a two-dimensional boundary layer solver to avoid solving the fully three dimensional boundary layer equations, assuming that the flow is two-dimensional on individual sections of the geometry. The viscous-inviscid interactive method couples a perturbation potential based inviscid panel method with a two-dimensional viscous boundary layer solver using the wall transpiration model. The boundary layer solver used in the study solves for the integral boundary layer characteristics given the edge velocity distribution on the geometry. The viscous-inviscid coupling is applied in a stripwise manner but by including the interaction e ffects from other strips. An important development in this thesis is the consideration of eff ects of other strips in a more rational and accurate manner, leading to improved results in the cases examined when compared to the results of a previous method. In particular, the effects of potentials due to other strips arising out of the three dimensional formulation are considered in this thesis. The validity of assuming two-dimensional flow along individual sections for application of viscous-inviscid coupling is investigated for the case of an open propeller by calculating the boundary layer characteristics in the direction normal to the assumed direction of two-dimensional flow from data obtained by RANS simulations. Also, a previous method which models the flow around the trailing edge of blunt hydrofoils has been improved and extended to three dimensional axisymmetric ducts. This method is applied to ducts with blunt and sharp trailing edges and to a ducted propeller. Correlations of results with experiments and simulations from RANS are shown. / text

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