The effective and efficient evaluation of conceptual product designs prior to manufacturing of prototypes, is of utmost importance for any industrial ac¬tivity. The rotorcraft industry is no different. Accurate, thorough and cost-effective evaluation of conceptual rotorcraft designs requires an equivalently rigorous and simultaneously affordable methodology covering all aspects rel¬ative. The associated issues to be tackled are indeed multi-dimensional and include trim performance, rotor blade structural loads prediction, engine per¬formance, mission analysis and associated environmental impact. The aforementioned topics are now raising even more interest as rotorcraft traffic is expected to grow sharply within the next 20 years. Current rotorcraft operations imply the consumption of the equivalent of 400,000 tons of aviation fuel per year, only with regards to the European region. Maintaining current rotorcraft technologies is expected to lead to a quadruplicating of this figure. This is a direct result of the expected traffic augmentation. In recognition of this trend, a wide range of research & development activities is currently being undertaken at national and European levels. The objective is to effectively return within 20 years to the present global level of environmental impact, while sustaining the expected growth of rotorcraft services. The objective set above is indeed an ambitious one. Due to the rather short time-scales involved, its realization requires for focus to be placed predomi¬nantly on the design of operational procedures. It is however realized that, in order to manage the environmental impact of civil rotorcraft aviation within larger time-scales, options concerning the design of conceptual configurations as well as incorporated operational procedures, need to be explored. Two fundamental requirements are therefore identified and addressed within this work. The 1st lies with regards to a generic methodology capable of design¬ing optimum operational procedures in terms of fuel burn, gaseous emissions and ground noise impact. The 2nd can be designated as a design assessment approach, capable of estimating the overall fuel consumption with regards to any designated operation. The employed method has to be capable of being utilized within the task of design while simultaneously maintaining a reasonable computational overhead so as to be applicable in the context of multidisciplinary optimization. This work elaborates on the development and application of an integrated approach, targeting the comprehensive assessment of combined helicopter– engine designs, within complete, three-dimensional operations. A series of individual modeling methods has been developed, each applicable to a dif-ferent aspect of helicopter flight dynamics and performance. These comprise rotor blade modal analysis, aeroelasticity, flight dynamics trim solution, en-gine performance and three-dimensional flight path definition. The individual mathematical models are elaborately integrated within a numerical procedure solving for the total mission fuel consumption. Extensive validation with ex-isting experimental and numerical data has been waged during each step of the development process. The aspect of multidisciplinary design of optimum rotorcraft operations in terms of fuel burn and environmental impact is also tackled within the con-text of this work. An existing integrated tool capable of estimating the per-formance and emitted noise of any defined rotorcraft configuration within any designated mission has been incorporated. A comprehensive and cost-effective optimization strategy has been structured. The methodology has been applied to two generic – baseline missions representative of current rotorcraft opera¬tions. Missions optimally designed in a multidisciplinary manner for fuel burn, gaseous emissions and ground noise impact have been obtained. The contribution to knowledge arising from the successful completion of this work, broadly comprises the development of methodologies applicable to the following aspects of civil rotorcraft aviation: (i) Comprehensive analysis of the overall flight dynamics and performance of any designated helicopter–engine integrated system, within realistically defined three-dimensional missions; (ii) Multidisciplinary design of optimum rotorcraft operations in terms of fuel consumption, gaseous and ground noise impact; Further to the above, con-tribution has been made through the analytical development of modeling ap-proaches with application to: (i) Rotor blade modal analysis; (ii) Treatment of rotor blade flexibility; (iii) Rotor–fuselage aerodynamic interaction. The developed analytical methods have been utilized within this work to facilitate the achievement of the set objectives. The potential to comprehensively eval-uate integrated helicopter–engine systems within complete three-dimensional operations, based on the solution of the aeroelastic behavior of the main rotor is demonstrated. The ability to design optimum operations in a multidisci¬plinary fashion using only a single design criterion has been exhibited.

Helicopter fuselage vibrations are of significant levels and produce a unique vibration problem. Reducing the vibration levels increases passenger comfort, reduces crew fatigue, allows higher cruise speeds to be achieved and improves equipment reliability. Vibration Reduction Techniques can be divided into two distinct categories, either passive or active techniques. Active control techniques offer the potential for good vibration reduction performance over significant areas of the fuselage. This study describes strategies for the Active Control of Structural Response. The technique aims to minimise the structural vibration of a helicopter. Accelerometers measure the vibration at several key points on the fuselage. A multivariable control algorithm processes this information and calculates a set of control forces for a set of hydraulic actuators, located at strategic points in the structure. Vibration reductions result from the superposition of the actuator induced vibrations forces with those induced in the fuselage by the rotor. This research presents a number of different active control strategies for the reduction of helicopter fuselage vibration. Two distinct active vibration control approaches are a frequency domain controller and a time domain controller, and the Thesis establishes the advantages and disadvantages of each of the control strategies. The time domain option is based upon direct feedback of vibration through constant gain matrices. The subsequent vibration waveform contains information over a wide spectrum of frequencies and consequently control is possible over a range of frequencies. Alternatively the frequency algorithms are specifically concerned with the control of discrete frequencies, the blade-passing frequency being dominant in the case of a helicopter. This thesis describes a third novel approach to the design of an adaptive controller for the reduction < of Helicopter vibration. This new technique is a hybrid time/frequency domain solution combining the advantages from both the time domain linear quadratic feedback controller and the frequency domain quasi-static controller. Both fixed gain and adaptive control designs have been implemented, and comparisons of the performance of the various control approaches to the problem of minimising vibration in helicopter structures are made. An estimator provides control system adaptability that permits the periodic update of the fuselage model and produces robustness to changes in the structural dynamics. A simulation study presents results for the performance and robustness of the control strategies. Experimental investigations considered the effects of linear and nonlinear actuator dynamics, the performance of the strategies during aircraft manoeuvres and the robustness the strategies to changes in the structural response. Experimental validation of the strategies was achieved by testing on a helicopter airframe test rig at the premises of Westland Helicopter Ltd, who collaborated in the SERC funded project within which this work was carried out.

A Cool Air unit (C. A. U. ) is powered by air bled from the engine and supplies cool air to the avionics and cockpit of the aircraft. Essentially a small turbo machine, the rotor is mounted upon two plain (gas) air bearings, which are flexibly supported by elastomeric '0'-Rings. The project objectives were to investigate, experimentally and theoretically, the mechanism of dynamic instability, half frequency whirl (HFW) observed in the C. A. U design, and hence evaluate methods by which it can be best controlled. Two main areas of study were embarked upon. Firstly a steady state aerodynamic study, based on an existing single bearing rig, to evaluate the extent of aerodynamic operation and single bearing instability (HFW). This Test Rig did not have '0'-Ring flexibility. Bearing L/D tested were 2,1.5 and 1.0, with variable loading of 18-1 -º 51.2N and speeds of 6000 -º 40000 Rpm in approximately 5000 Rpm intervals. the theoretical study investigated non-linear effects of air film pressure distribution. Secondly a Dynamic instability study of HFW was carried out on a newly designed Test Rig, simulating small turbo machines and the C. A. U. Three bearing types were evaluated, Aerodynamic, Hybrid and Hybrid porous. With L/D of 1.0. '0'-Ring Centres varied from 0.17 4 0-83 of bearing length. Viton and Silicon materials were tested with 70 shore hardness. Rotor mass, inertia, asymmetry and unbalance were investigated along with '0'- Ring stiffness and damping variation from air pressure (0 4 120psi). Theory was developed to determine whirl onsets and effects of unbalance, and damping of the support. Non-linear (stagnant areas) of air film detected experimentally, are not explained by Raimondi's theory. Experiments show that Raimondi's theory seriously over estimates the applicable area of fully developed aerodynamic operation. Pressure Profile at onset of Instability tends to a Sommerfeld condition. Temperature was a good indication of lubrication regime. Linear temperature rise curve denotes Aerodynamic operation, and transition to non-linear curve represents onset of Instability (HFW). Authors theory considers non-linearity of air film and suggests a new method of evaluation to improve convergence. Dynamic study of dual bearing rig concludes, first instability onset speed (RSW) can be passed through due to residual unbalance and damping in the '0'-Ring support. Theory developed shows relationship of RSW and HFW effects due to unbalance and damping and results agree well. Viton offered better damping than Silicon. RSW not present in Viton at'0'-Ring centres above 10 mm. HFW not present with Viton below 55,000 rpm. RSW observed with ui /cu = 1.0 and HFW occurred with m /m = 0-493 4 0.58. hybrid porous bearings had lower performance than hybrid, but better than aerodynamic. Hybrid bearings mounted in Viton '0'-Rings offer best performance long term. Aerodynamic bearings can be concluded to be inherently unstable and have limited aerodynamic operation, so should see little use in high speed turbo machinery, including C. A. U. or aircraft applications where zero g loading likely.

This study has shown that the main characteristics of axial velocity decay and recirculation in multiple enclosed jets can be related to those of single enclosed jets. This is achieved by considering each jet to be bounded by an imaginary duct whose dimensions can be related to the nozzle spacing parameters. If the nozzles are close together compared with the surrounding chamber, the above treatment is only valid near- to the nozzles. Further downstream the jets coalesce and their behaviour can be related to that of a single jet by an equivalent nozzle radius.

An effective collision avoidance logic should prevent collision without excessive alerting. This requirement would be even more stringent for an automatic collision avoidance logic, which is probably required by Unmanned Aerial Vehicles to mitigate the impact of delayed or lost link issues. In order to improve the safety performance and reduce the frequency of false alarms, this thesis proposes a novel collision avoidance logic based on the three-layer architecture and a real-time trajectory planning method. The aim of this thesis is to develop a real-time trajectory planning algorithm for the proposed collision avoidance logic and to determine the integrated logic’s feasibility, merits and limitations for practical applications. To develop the trajectory planning algorithm, an optimal control problem is formulated and an inverse-dynamic direct method along with a two stage, derivative-free pattern search method is used as the solution approach. The developed algorithm is able to take into account the flyability of three dimensional manoeuvres, the robustness to the intruder state uncertainty and the field-of-regard restriction of surveillance sensors. The testing results show that the standalone executable of the algorithm is able to provide a flyable avoidance trajectory with a maximum computation time less than 0.5 seconds. To evaluate the performance of the proposed logic, an evaluation framework for Monte Carlo simulations and a baseline approach for comparison are constructed. Based on five Monte Carlo simulation experiments, it is found that the proposed logic should be feasible as 1) it is able to achieve an update rate of 2Hz, 2) its safety performance is comparable with a reference requirement from another initial feasibility study, and 3) despite a 0.5 seconds computation latency, it outperforms the baseline approach in terms of safety performance and robustness to sensor and feedback error.

Amongst the objectives of the EU's Green Rotorcraft programme is the development of a functional and airworthy Active Gurney Flap (AGF) for a full-scale helicopter rotor blade. Interest in the development of this 'smart adaptive rotor blade' technology lies in its potential to provide a number of aerodynamic benefits, which would in turn translate into a reduction in fuel consumption and noise levels. With a technical specification informed by helicopter manufacturer Agusta Westland as a starting point, the AGF concept developed emerged from the application of a design methodology consisting of an initial concept generation phase and a subsequent selection process based on a series of problem-specific qualitative and quantitative performance criteria. This methodology resulted in a novel AGF concept design where the use of flexural pivots was favoured over bearing-type joints. As a first step towards full validation of both the concept design and the theoretical aerodynamic benefits of the AGF, a baseline design of the mechanism was successfully tested both in a fatigue rig and in a 2D wind tunnel environment at flight-representative deployment schedules. This baseline design was then reengineered with a view to making it fit for flight test. However, analysis of the flight test baseline design under full in-flight loading and blade deformations revealed that the stresses arising in the flexures exceeded the allowable limits. In order to overcome this problem, two complementary alternatives were investigated. Initially, a generic 2D and 3D shape optimisation of leaf-type crossed flexure pivot springs was carried out considering the pivots as individual isolated elements. This route produced important novel results in the field of crossed flexure pivots and proved that there is great scope for stress reduction through shape optimisation. Furthermore, a parametric optimisation of the AGF mechanism as a whole was performed, where the effect of a range of topological parameters was investigated. This second approach resulted in very significant stress reductions" sufficient so as to conclude that the proposed AGF concept has potential to become an airworthy design, although further work is needed to achieve a sufficiently mature technology readiness level.

Active flow control systems have the potential to allow future designs of ground vehicles and aircraft to realise increased operational efficiency through improved optimisation of the flow, leading to decreased fuel use and reduced environmental impact. To achieve this, dynamic actuators are required that can adapt to the changing conditions experienced over low-speed, high-lift aerofoils where separation control can be particularly advantageous. Synthetic Jet Actuators (SJAs) are a form of the technology that shows promise; they are small, low-mass and low-power devices, which means that they can potentially realise the system efficiency a vehicular application of active flow control requires. The aims of this research were directed to achieving better understanding and robustness of the control authority from SJA systems at Reynolds Numbers close to real-world operations. The characteristics of the control authority from a round-orifice SJA array positioned near the leading-edge position of an NACA0015 aerofoil have been investigated at Re = O(10^6). Measurements demonstrated how the jet flow imparts a controlling mechanism over the separating boundary layer flow, and hence can be used to improve the overall efficiency of the wing. A series of parametric alterations to the test conditions was made in order to understand the robustness of the control effect. The forcing frequency was decoupled from the dynamic response of the actuators themselves by means of amplitude modulation. The results demonstrated that successful control could be achieved with significantly reduced input power requirements, improving net efficiency. The effectiveness was shown to be largely independent of the frequency when used in this way. Using a counterstreamwise jet orientation to control the same basic separated flow condition was not found to generate significant improvements in operational efficiency. The results suggest an in-depth understanding of the jet flow, and the excitation location in relation to the point where the flow separates is important when designing the actuators. Tests also considered a different flow condition with a stronger adverse pressure gradient, by generating a ground-effect flow over the suction surface. The control authority afforded was diminished in the more adverse flow states. The performance of the actuators was considered, and the system achieved a larger than unity Figure of Merit, indicating the overall benefit of control shows direct relevance for realising practical flow control systems. The results indicate arrays of SJA’s are capable of delivering energy savings when managing this type of flow.

Landing on lighting columns like nature’s birds is a desirable capability which can only extend the uses of unmanned aerial systems. This thesis investigates what the most effective form of perching on existing street furniture with a VTOL UAV and how the perch site can be recognised using low cost off the shelf sensors. Additionally to this, the UAV in question will have to execute the perch without relying on GPS data. The work conducted here covers an extensive design review which selects a bird claw like gripper to sustain the perch. In order for the UAV to know where it is in relation to the perch site without relying on GPS data, a Raspberry Pi and PiCamera were used to detect common features which are found on top of a lighting column. Using a search and perch algorithm which was developed specifically for the task of perching on lamp post projection brackets, the on-board microprocessor controlled the UAV over the perch site and gradually descended into the perch position. The perched position and approach was also tested to ensure the perching element could cope with various weather conditions. The testing was conducted in a wind tunnel with the UAV mounted in various perched positions and the moment the UAV would slip, the wind speed were measured and analysed which highlighted an interesting prediction method. During the perching development the addition of a gantry style test-rig was also developed to ease the algorithm development with minimal incidents. The final result is a search and perch algorithm which is initiated when the VTOL UAV is within the vicinity of a lamp post at which point the on-board vision processing and control system takes over, removing the burden from the UAV operator to ensure a collision free perch.

There is a growing need for new sources of energy due to the rise in global energy demand, the decline in fossil fuels, and the increasing, negative consequences of climate change. Renewable energy resources are sustainable but they are also intermittent, meaning that they cannot supply energy on demand unless it is stored. Hydrogen is one potential chemical method of storing this energy; however, it has a very low energy density per unit volume, meaning that storage in low mass and volume containers can be problematic. One solution is to adsorb hydrogen onto highly porous materials. This thesis presents an improved methodology for analysing hydrogen adsorbed inside porous materials, and how it can be utilised to determine the potential use of storing hydrogen via physisorption for aviation. Preliminary studies are conducted on pressure and temperature dependencies of both the pore volume and the adsorbate density, and a comparison is also made between the utilisation of different Type 1 isotherms for the fractional filling of hydrogen, with the use of the Tόth equation resulting in the best quality of fit to the isotherms overall. The model is verified using inelastic neutron scattering and computer simulations. The model is then utilised to calculate the amount of hydrogen within a tank containing varying quantities of adsorbent, and comparing this to the amount of hydrogen that can be stored via direct compression at the same conditions. This is then expanded to be compared to alternative energy systems, and a preliminary investigation on the use of adsorbed hydrogen within aviation is conducted. The results show hydrogen adsorption to always have a higher energy density than compressed hydrogen up to a certain pressure, and for both to have a comparable energy density to battery storage at certain conditions, but not to standard jet fuels.

The aircraft industry has long recognised the importance of climate protection and the benefits of reducing weight for the production of cost effective and fuel efficient aircraft structures. Fibre laser welding provides advantages over conventional riveting, mainly in terms of weight reduction and time saving. However, significant changes in microstructure, metallurgical state and associated mechanical properties occur in welded joints. Such changes can result in residual stresses, distortions and defects formation in the welded structure, thus significantly influencing the performance and service life. In order to maintain structural integrity of welded structures, the relationship between welding process and performance of the structure needs to be fully assessed. In this thesis, comprehensive relationships between materials, welding process, microstructure and mechanical properties of welded joints were established. Welding parameters including power density, laser power, welding speed, focal position, filler metal feed rate and shielding gas composition were optimised to produce high quality full penetration welds. Solidification cracking was found to be a critical issue in AA 2024-T3 when welding without filler metal. The addition of filler metal reduced its crack sensitivity but it was also necessary to provide the optimum feed rate to avoid welding defects and keyhole instability. Sufficiently high laser power and low welding speed were required for full penetration and also to minimise welding defects. Both argon and helium shielding gases were found to be effective since only weakly ionised laser induced vapour plume was formed rather than strongly ionised plasma. Softening in AA204-T3 deteriorated the plastic straining capacity of the weld due to confined plasticity development within the weld. A poor weld quality resulted in a mixed mode of brittle and ductile failure and contained micro porosities and hot cracks, whereas, a good weld quality led to a ductile mode with significantly less welding defects. In the case of Ti-6Al-4V, the strength was the greatest in the weld as a result of martensitic microstructure formed during fast cooling rates. Local plastic deformation was the lowest in the weld and therefore, failed in the parent material but at the cost of reduced ductility relative to the unwelded parent tensile specimens. The residual stresses and distortions due to time dependent and localised heating imposed during fibre laser welding were numerically simulated with thermal and mechanical boundary conditions integrated in the finite element models including post weld heat treatment, mechanical stress relieving treatment and various clamping arrangements. Mechanical boundary conditions had relatively small influence on residual stresses in thin sheets of butt welded specimens, whereas, greater restraints led to higher residual stresses and lower restraints led to lower residual stresses in T-joint specimens. Non-isothermal diffusional and diffusionless phase transformations in Ti-6Al-4V were modelled and their influence on residual stresses and distortions was examined. Phase transformations only had a small influence on the magnitude and distribution of residual stresses and distortions because the level of internal stresses due to phase transformation remained low unlike other materials which exhibit greater differences in the specific volumes between phases. Post weld heat treatment (PWHT) induced diffusional phase transformations via decomposition of martensite into α. It also decreased the magnitude of y stresses to the yield strength of Ti-6Al-4V at the treatment temperature by releasing the locked-in stresses. Mechanical stress relieving was also studied for reducing residual stresses and distortions, by means of plastic deformation applied during as well as after welding. When the load reached more than 50% of its yield strength, the stresses became compressive. Residual stresses were experimentally measured using X-ray and neutron diffraction techniques They were found to be dependent on the crystallographic hkl plane due to the presence of microscopic stresses. In the case of Ti-6Al-4V, the reflections were weak and only few times larger than the background due its highly incoherent cross-section. In addition, texture in Ti-6Al-4V weld also contributed to lower intensity counts observed during measurements. As a result, only certain peaks were detected in certain orientations. The Y residual stresses in the welding direction were very high but not as high as the yield strength of the material at room temperature for both AA 2024-T3 and Ti-6Al-4V. They were largely tensile in nature only within the weld and tended to be weakly compressive in the rest of the specimen. Comparative analyses between experimental and numerical results showed good agreements, proving the robustness of the finite element models.