The BOSS symbology for rotorcraft is specifically designed to provide the pilot with the necessary information and guidance to safely land in brownout environments. From the last BOSS study, issues were brought forth regarding the longitudinal velocity algorithm, which sets up a deceleration profile and commands the forward speed of the aircraft throughout the approach. Pilots commented that the algorithm lead the aircraft to be too slow for too long, effectively prolonging the brownout. Thus the purpose of this study was to investigate new algorithms to enable a faster approach with less time spent in brownout. The previous deceleration algorithm was also not robust in its ability to provide consistent guidance at variable starting distances and starting velocities. Therefore a new algorithm was developed capable of providing more consistent guidance from various starting positions and velocities. Additionally, through manipulation of its parameters, it was found possible to reduce the amount of time spent at low speeds in the approach. Four algorithms were subsequently developed with varying levels of aggressiveness. Eight highly skilled pilots participated in a simulation study using a generic fixed-base simulator with a high-fidelity rotorcraft H-60 model. Results found that as the aggressiveness of the algorithm increased, the time spent at low speeds and in brownout significantly decreased. Concurrently the pitch of the aircraft (and resulting deceleration) significantly increased, though the pitch values were within reasonable limits for IMC flight according to previous literature. One of the new algorithms was found to significantly reduce the amount of time spent at low speeds by 24% and also received the highest preference ranking and the highest comfort ratings.

Helicopters play a unique role in modern aviation providing a varied range of benefits to society and satisfying the need for fast mobility, particularly in metropolitan areas. However, environmental concerns associated with the operation of rotorcraft have increased due to envisaged growth of air traffic. Even though helicopter operations represent a small percentage of the total greenhouse gas emissions resulting from all human activities, helicopters are categorised as a main source of local air pollution around airports and urban areas. New rotorcraft designs, innovative aero engines and all-electrical systems are being developed in order to diminish the impact that aviation has on the global and local environment. However, advanced rotorcraft designs and breakthrough technologies might take decades to be in service. Additionally, there is a large number of polluting rotorcraft that are in use and must be progressively replaced. Therefore, in the near-term, improvements to minimise air quality degradation (around airports and metropolitan areas) may be possible from better use of existing rotorcraft by focusing on trajectory and mission profile management. In this research project, a parametric study was carried out in order to assess the environmental impact, in terms of fuel burn and emissions, that the operation of light single-engine helicopters causes under different flight conditions. The results of this assessment were used as a basis to carry out a single and multi-objective optimisation for minimum fuel consumption and air pollutant emissions. Oxides of nitrogen, carbon monoxide and unburnt hydrocarbons were considered as trade-off parameters. In order to achieve this, a multidisciplinary assessment framework, intended to generate outputs for estimating the fuel burn and emissions during the operation of conventional helicopters, was developed. Simulink® Design Optimization™ software was incorporated into the framework in order to enhance the benefits of this tool.A baseline mission profile was proposed in order to validate the potential of mission profile management. Different case studies were carried out changing flight parameters at every segment of the baseline mission. The single and multi-objective optimisation proved that favourable reductions in fuel burn may be attainable at the expense of a slight increase of NOX emissions during the entire mission. If reductions of more than 3% in block fuel burn are to be achievable in the short term for a single helicopter, savings for air transport companies are expected to be significant if mission profile management is considered for a whole fleet of helicopters.

Throughout the evolution of rotorcraft design, great advancements have been made in developing performance analysis and sizing tools to assist designers during the preliminary and detailed design phases. However, very few tools exist to assist designers during the conceptual design phase. Most performance analysis tools are very discipline or concept specific, and many are far too cumbersome to use for comparing vastly different concepts in a timely manner. Consequently, many conceptual decisions must be made qualitatively. A need exists to develop a single software tool which is capable of modeling any type of feasible rotorcraft concept using different levels of detail and accuracy in order to assist in the decision making throughout the conceptual and preliminary design phases. This software should have a very intuitive and configurable user interface which allows users of different backgrounds and experience levels to use it, while providing a broad capability of modeling traditional, innovative, and highly complex design concepts. As an illustration, a newly developed Concept Independent Rotorcraft Analysis and Design Software (CIRADS) will be presented to prove the applicability of such software tools. CIRADS is an object oriented application with a Graphical User Interface (GUI) for specifying mission requirements, aircraft configurations, weight component breakdowns, engine performance, and airfoil characteristics. Input files from the GUI are assembled to form analysis and design project files which are processed using algorithms developed in MATLAB but compiled as a stand alone executable and imbedded in the GUI. The performance calculations are based primarily upon a modified momentum theory with empirical correction factors and simplified blade stall models. The ratio of fuel (RF) sizing methodology is used to size the aircraft based on the mission requirements specified by the user. The results of the analysis/design simulations are then displayed in tables and Text Fields in the GUI. The intent for CIRADS is to become a primary conceptual sizing and performance estimation tool for the Georgia Institute of Technology rotorcraft design teams for use in the annual American Helicopter Society Rotorcraft Design Competition.

The present work focuses on the system identification method of aerodynamic parameter estimation which is used to calculate the stability and control derivatives required for aircraft flight mechanics. A new rotorcraft parameter estimation technique is proposed which uses a type of artificial neural network (ANN) called radial basis function network (RBFN). Rotorcraft parameter estimation using ANN is an unexplored research topic and the earlier works in this area have used the output error, equation error and filter error methods which are conventional parameter estimation methods. However, the conventional methods require an accurate non-linear rotorcraft simulation model which is not required by the ANN based method. The application of RBFN overcomes the drawbacks of multilayer perceptron (MLP) based delta method of parameter estimation and gives satisfactory results at either end of the ordered set of estimates. This makes the RBFN based delta method for parameter estimation suitable for rotorcraft studies, as both transition and high speed flight regime characteristics can be studied. The RBFN based delta method for parameter estimation is used for computation of aerodynamic parameters from both simulated and real time flight data. The simulated data is generated from an 8-DoF non-linear simulation model based on the Level-1 criteria of rotorcraft simulation modeling. The generated simulated data is used for computation of the quasi-steady and the time-variant stability and control parameters for different flight conditions using the RBFN based delta method. The performance of RBFN based delta method is also analyzed in the presence of state and measurement noise as well as outliers. The established methodology is then applied to compute parameters directly from real time flight test data for a BO 105 S123 helicopter obtained from DLR (German Aerospace Center). The parameters identified using the RBFN based delta method are compared with the identified values for the BO 105 helicopter from published literature which have used conventional parameter estimation techniques for parameter estimation using a 6-DoF and a 9-DoF rotorcraft simulation model. Finally, the estimated parameters are verified from the flight data generated by a frequency sweep pilot control input for assessing the predictive capability of the RBFN based delta method. Since the approach directly computes the parameters from flight data, it can be used for a reliable description of the higher frequency range, which is needed for high bandwidth flight control and in-flight simulation.

The aerospace industry has always been looking forward new developments with the aim to create more environmental friendly aircraft, as well as to improve their performance. Over the last few years, a prominent research topic to achieve these challenging goals has been focussed on the incorporation of more electric Secondary Power Systems (SPS), this concept is known as More Electric Aircraft (MEA) or All Electric Aircraft (AEA) when the internal combustion engine is also replaced. Among others, Airbus is using Electro-hydrostatic Actuators, (EHAs) to combine hydraulic and electric power in A320 and A340 for flight tests since 1993. The company TTTECH applied the same concept by working on the development of an electrical steering system for an aircraft nose landing gear, and power source rationalization and electrical power flexibility in aircraft. Some of the advantages stated when the MEA concept is applied are: reduction in aircraft weight and performance penalties related to conventional SPS. Although the More/All electric aircraft concept provided satisfactory results for fixed-wing aircraft, research for rotary-wing aircraft is less common. This encourages the assessment of fuel consumption and performance penalties due to conventional and more electric SPS at conceptual level, which could achieve similar outcomes, while finding the best configuration possible. This project takes into account the previous research focused on fixed-wing aircraft and studies on new technologies for SPS within Cranfield University, this includes electrical Ice Protection System (IPS), Environmental Control System (ECS) and Actuation System (AS). Additionally, Fuel System (FS) and Electrical System (ES) capabilities were added, developing a generic tool able to predict the total power requirements depending on the flight conditions. This generic tool was then integrated with a performance model, where overall fuel consumption is calculated for a flight mission, giving continuity and improvement to the work already done. Secondary systems configuration and operating characteristics for a representative light single-engine rotary-wing aircraft were tailored, and the systems behaviour is presented. Finally, fuel consumption was calculated for a baseline mission profile, and compared to the fuel consumption when the systems are not included. The baseline mission set the initial flight conditions from which a parametric study was carried out; by varying these conditions the parametric study determined total fuel requirements for the analysed flight segments. An increment of up to %1.9 in the fuel consumption was found by integrating the proposed systems to the performance model, showing the impact produced by the systems, and the importance of studying different technologies to minimise it.

abstract: This thesis discusses the equilibrium conditions and static stability of a rotorcraft kite with a single main tether flying in steady wind conditions. A dynamic model with five degrees of freedom is derived using Lagrangian formulation, which explicitly avoids any constraint force in the equations of motion. The longitudinal static stability of the steady flight under constant wind conditions is analyzed analytically from the equilibrium conditions. The rotorcraft kite orientation and tether angle are correlated through the equation Γ=δ-ϑ, a necessary condition for equilibrium. A rotorcraft kite design with 3kg mass and 1.25m rotor radius is found to be longitudinally statically stable at 25,000ft with Γ>〖65〗^0 for wind speeds above 19m/s. / Dissertation/Thesis / Masters Thesis Aerospace Engineering 2017

A Rotorcraft UAV provides an ideal experimental platform for vision-based navigation. This thesis describes the flight tests of the US Army PALACE pro ject, which implements Moravec's pseudo-normalized correlation tracking algorithm. The tracker uses the movement of the landing site in the camera, a laser range, and the aircraft attitude from an IMU to estimate the relative motion of the UAV. The position estimate functions as a GPS equivalent to enable the rotorcraft to maneuver without the aid of GPS. Flight tests were performed with obstacles and over concrete, asphalt, and grass in daylight conditions with a safe landing area determined by a separate method. The tracking algorithm and position estimation performance are compared to GPS. Accurate time synchronization of the inputs to the position estimation algorithm directly affect the closed-loop stability of the system, proportional with altitude. By identifying the frequency response of each input and adding filters to delay some of the inputs, the closed-loop system maintains stable flight above 18 m above ground, where the system was unstable without the additional filters.

In recent years, the use of rotorcraft uninhabited aerial vehicles (UAVs) for cargo delivery has become of particular interest to private companies and humanitarian organizations, namely due to their reduced operational costs, ability to reach remote locations and to take off and land vertically. The slung configuration, where the cargo is suspended below the vehicle by a cable, is slowly being favoured for its ability to transport different sized loads without the need for the vehicle to land. However, such configurations require complex control systems in order to stabilize the swing of the suspended load. The goal of this research is to design a control system which will be able to bring a slung payload transported by a rotorcraft UAV back to its stable equilibrium in the event of a disturbance. A simple model of the system is first derived from first principles for the purpose of simulating a control algorithm. A controller based in model-free, policy-gradient reinforcement learning is then derived and implemented on the simulator in order to tune the learning parameters and reach a first stable solution for load stabilization in a single plane. An experimental testbed is then constructed to test the performance of the controller in a practical setting. The testbed consists of a quadcopter carrying a weight suspended on a string and of a newly designed on-board load-angle sensing device, to allow the algorithm to operate using only on-board sensing and computation. While the load-angle sensing design was found to be sensitive to the aggressive manoeuvres of the vehicle and require reworking, the proposed control algorithm was found to successfully stabilize the slung payload and adapt in real-time to the dynamics of the physical testbed, accounting for model uncertainties. The algorithm also works within the framework of the widely-used, open-source autopilot program ArduCopter, making it straightforward to implement on existing rotorcraft platforms. In the future, improvements to the load angle sensor should be made to enable the algorithm to run fully on-board and allow the vehicle to operate outdoors. Further studies should also be conducted to limit the amount of vehicle drift observed during testing of the load stabilization.

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