Browsing by Author "Lawson, Craig"

Now showing 1 - 10 of 10
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    'A Holistic Cabin Conceptualisation Approach (HCCA) framework'
    (Cranfield University, 2023-11-01 09:50) Kirensky, Roman; Lawson, Craig; Orson, Ben
    The presented toolset defines the stakeholders involved in aircraft passenger cabin interior development projects, and their pursued design drivers. - Stakeholder Definition sheet presents all types of organisations involved in the design of aircraft cabin interiors, disregarding of the involvement extent. - Stakeholder Analysis sheet contains the pairwise comparison matrices for all stakeholders along the influence and interest axes, and the resulting plot. - Cabin Design Drivers sheet contains the definition of drivers pursued by the cabin interior projects. These are defined as a multi-level structure including: -- Top-level profitability-based design drivers, -- Their constituent factors representing the design considerations and themes, and -- Bottom-level design criteria representing the detailed product specifications, requirements, features, qualities, performance targets etc. - Design Factor Weights present a set of pairwise comparison matrices for deriving the relative importance weights at the factor and driver levels. - R Input sheet contains R code template to use with R Studio software to retrieve Eigenvector values. - Design Factors Map sheet contains factor applicability mapping to represent stakeholder concern and interaction points on a cabin interior project. - Cabin Product Breakdown sheet presents the composition of state-of-the-art cabin interiors by listing out the components it may have. The presented list is an all-encompassing version and does not represent the product line of any specific cabin manufacturer or equipment supplier. The presented sheets may be used as information source, or be amended to reflect the needs of their specific project. Amendments may be performed in the white cells of the pairwise comparison matrices on Stakeholder Analysis and Design Factor Weights; and applicability indicators in Design Factors Map and Cabin Product Breakdown sheets.
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    Aircraft assembly process design for complex systems installation and test integration.
    (Cranfield University, 2019-04) Li, Tao; Lockett, Helen L.; Lawson, Craig
    The assembly line planning process connects product design and manufacturing through translating design information to assembly integration sequence. The assembly integration sequence defines the aircraft system components installation and test precedence of an assembly process. From a systems engineering view point, this activity is part of the complex systems integration and verification process. At the early conceptual design phase of assembly line planning, the priority task of assembly process planning is to understand product complexities in terms of systems interactions, and generate the installation and test sequence to satisfy the designed system function and meet design requirements. This research proposes to define these interactions by using systems engineering concept based on traceable RFLP (Requirement, Functional, Logical and Physical) models and generate the assembly integration sequence through a structured approach. A new method based on systems engineering RFLP framework is proposed to generate aircraft installation and test sequence of complex systems. The proposed method integrates aircraft system functional and physical information in RFLP models and considers these associated models as new engineering data sources at the aircraft early development stage. RFLP modelling rules are created to allow requirements, functional, logical and physical modes be reused in assembly sequence planning. Two case studies are created to examine the method. Semi- structured interviews are used for research validation. The results show that the proposed method can produce a feasible assembly integration sequence with requirements traceability, which ensures consistency between design requirements and assembly sequences.
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    Aircraft icing and thermo-mechanical expulsion de-icing technology
    (Cranfield University, 2010-12) Ma, Qinglin; Lawson, Craig
    The topic of this thesis is Aircraft Icing and Aircraft Icing and Thermo-Mechanical Expulsion De-icing Technology. The main objectives are to investigate aircraft icing meteorology and effects on aircraft, ice protection systems and thermo-mechanical expulsion de-icing technology. Initially, the research project focuses on aircraft icing meteorology, ice accumulation and icing effects on flight safety. A basic understanding of aircraft icing is explained, including icing conditions and parameters, ice detection, ice accretion shapes, and icing effects in terms of different airframe surfaces. Ice protection systems are employed to prevent the icing hazards and ensure flight safety. The conventional and novel ice protection technologies are investigated and their characteristics are explained. Thermo-mechanical expulsion de-icing technology is selected as the research topic because of its typical characteristics of lower power required and high de-icing efficiency. Methods are explored to assess power consumption of thermo-mechanical expulsion de-icing system. In addition, a thermo-mechanical expulsion de-icing system architecture is explored for a civil aircraft and system power consumption is calculated. In order to explain the system design characteristics, conventional electro-thermal de-icing system architecture is designed to show the contrasts in terms of power consumption, de-icing efficiency, cost and mass. This novel ice protection system has better results in power consumption and de-icing performance, allowing long time flight in icing conditions. It can compensate the disadvantages of heavier mass and higher cost because there is no need to escape from icing encounters immediately and fuel penalties are reduced sequentially.
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    Development of tail rotor power analysis model with feasibility study of electrical tail rotor.
    (2016-09) Subramanian, Nithya; Lawson, Craig
    In recent years, there has been significant work undertaken by the aviation industry to increase the overall rotorcraft performance, and eventually, eliminate leak prone hydraulic fluids and to reduce CO₂ emissions. Even though a mechanical-gearbox-driven tail rotor has been extensively used in several applications, it comes at the expense of high life cost of the gearbox and shaft gear mechanism. This thesis concentrates on the developing a model to analyse the power requirement for the tail rotor drive and feasibility investigation of an electrical tail rotor to substitute the shaft geared system and the conventional tail rotor power transmission gearbox. A case study is conducted on the Sikosky UH-60A rotorcraft to assess the conventional tail rotor power requirement and Electrical systems. A mathematical model based on Rankine Froude’s momentum theory is created to analyse the power required to drive the anti-torque system, which could be adapted to any conventional drive train (with the main rotor and a tail rotor) rotorcraft. A mission profile and trajectory are created and implemented into Excel based mathematical model. The challenges in implementing electrical drivetrain (electrical generation, energy conversion and electric transmission) are briefly discussed in this thesis. Electrical load analysis database is generated to find the electrical load of the generator for the entire flight phases and utilised to up-scaled the generator to compensate the new load from Electrical tail rotor. The electrical powertrain system is designed with a Brushless DC motor attached to the tail rotor and the generator and the battery for redundancy purposes. The research thesis develops an understanding of current electric motor and battery technology to create a novel design of electric tail drive that increases the reliability of the helicopter system.
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    Health monitoring of cavitation erosion on airframe fuel pumps
    (Cranfield University, 2023-09) Verhulst, Tedja; Lawson, Craig; Judt, David
    Aircraft maintenance is a critical aspect of operations that can lead to substantial financial savings if improved. The fuel pump is a vital component of the entire aircraft, ensuring fuel delivery to the engine. Cavitation is common in these pumps, leading to erosion and reducing the pump's remaining useful life. Therefore, the main objective of this work is to develop a viable health monitoring method to diagnose cavitation erosion, where few solutions exist. Initially, a literature review is conducted to identify knowledge gaps and opportunities for technology transfer related to current Health Monitoring (HM) technologies for airborne pumps. Four sensing methods, pressure, flow, current, and temperature sensing, are shortlisted based on their past applications and suitability for an aircraft fuel system installation. A hybrid health monitoring scheme consists of a Computational Fluid Dynamics (CFD) simulation, a model running on Simscape, and an experimental test rig. Live experiments are conducted to validate the simulation methods, enabling the testing of scenarios on a wide range of boundary conditions. The simulations demonstrate strong alignment with the experimental data and successfully distinguish the different levels of erosion. Three out of the four tested sensing methods are sensitive enough to distinguish the different levels of erosion, but each method has its advantages and limitations. Temperature sensing is not useful for health monitoring as the ambient environment strongly influences its results. Despite the success of the developed health monitoring schemes, there is a need for further research and development into more sophisticated health monitoring algorithms before the technologies can be widely implemented on aircraft.
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    Novel swing arm mechanism design for trailing edge flaps on commercial airliner
    (Cranfield University, 2008) Yu, Jie; Stocking, Philip; Lawson, Craig
    This thesis will describe the works had been done by the author in the Flying Crane aircraft group design project and the new design of a novel swing arm mechanism which can be applied in the trailing edge high lift devices for this aircraft concept. Flying Crane aircraft is a new generation commercial airliner concept as the result of group design project conducted by China Aviation Industry Corporation I (AVIC I) and Cranfield University. At the end of the group design project, parameters such as take-off and landing distance, trailing edge flap type and deflection in take-off and landing configuration of the Flying Crane concept have been determined. These parameters are design input of the novel trailing edge high lift device mechanism for this aircraft concept. The idea of this innovative mechanism comes from the research achievement of a previous MSc student, Thomas Baxter, which applied swing arm mechanism into a passenger aircraft's leading edge slat. This thesis applied this idea to trailing edge flap and modeled the mechanism on CATIA software to yield a kinematic simulation for the purpose of check motion trail and force transfer in this mechanism. Relevant works such as actuation, mass and stress analysis are also involved. As the result of this research project, it was found that swing arm mechanism trends to require relatively small fairings for supports and attachments due to its high stowed space utilizing efficiency. Initial mass estimation carried out in this thesis also indicates that the new design takes advantage in terms of weight comparing with traditional trailing edge flap mechanisms. Thus. swing arm mechanism is supposed to show great competitive potential for commercial airliner's trailing edge flaps after further analysis has been done in the detail design phase.
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    ProBCA Taxonomy of Discrete MCDA Methods for Ranking
    (Cranfield University, 2023-11-01 09:46) Kirensky, Roman; Lawson, Craig; Salonitis, Konstantinos
    This book is accompanied by the publication titled 'A SYNOPTIC TAXONOMY OF DISCRETE MULTI-CRITERIA DECISION ANALYSIS METHODS FOR RANKING', which introduces the presented taxonomy for MADM (Multi-Attribute Decision-Making) methods for ranking tasks. The taxonomy consists of: - the collection of 300 MADM methods covering the various parts of a typical MCDA (Multi-Criteria Decision Analysis) process; - the characterisation system for the recorded methods called ProBCA (Problem-Based Characterisation Approach). The title (ProBCA) reflects an application-oriented mindset that the presented taxonomy is based on. It focuses on the DP (Decision Problem) parameters and how the DM (Decision Maker) deals with it to describe the presented methods, rather than the intrinsic characteristics of the methods itself. The taxonomy is operated by picking from the list of available values for each of the 17 descriptor parameters characterising the possible DP context specifics and DM constraints. If a method (or several) matching the provided DP characterisation is available in the presented collection, it will remain visible after filtering for appropriate values while the remaining methods will become hidden. It is possible to use partial DP characterisation to identify a range of potentially suitable methods if the DM is flexible about theof defining define the DP and how to approach its solution. The number of methods matching each of the available characterising values is always shown next to these values in the top section of the taonomy, and is progressively updated as the DM proceeds with value selection. The taxonomy is dedicated to allow a broad spectrum of DMs to efficiently select the most appropriate MADM method for their ranking DP at hand.
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    Rapid design of aircraft fuel quantity indication systems via multi-objective evolutionary algorithms
    (IOS Press, 2020-12-11) Judt, David M.; Lawson, Craig; van Heerden, Albert S. J.
    The design of electrical, mechanical and fluid systems on aircraft is becoming increasingly integrated with the aircraft structure definition process. An example is the aircraft fuel quantity indication (FQI) system, of which the design is strongly dependent on the tank geometry definition. Flexible FQI design methods are therefore desirable to swiftly assess system-level impact due to aircraft level changes. For this purpose, a genetic algorithm with a two-stage fitness assignment and FQI specific crossover procedure is proposed (FQI-GA). It can handle multiple measurement accuracy constraints, is coupled to a parametric definition of the wing tank geometry and is tested with two performance objectives. A range of crossover procedures of comparable node placement problems were tested for FQI-GA. Results show that the combinatorial nature of the probe architecture and accuracy constraints require a probe set selection mechanism before any crossover process. A case study, using approximated Airbus A320 requirements and tank geometry, is conducted and shows good agreement with the probe position results obtained with the FQI-GA. For the objectives of accessibility and probe mass, the Pareto front is linear, with little variation in mass. The case study confirms that the FQI-GA method can incorporate complex requirements and that designers can employ it to swiftly investigate FQI probe layouts and trade-offs.
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    Simulation for temperature control of a military aircraft cockpit to avoid pilot’s thermal stress
    (Springer, 2015-01-30) Shetty, Janardhana; Lawson, Craig; Zare Shahneh, Amir
    During flying, military pilots are normally subjected to a number of stresses like mild hypoxia, high accelerations, vibrations and thermal discomfort. Among all of these, thermal stress is the most predominant factor while operating in highly tropical regions. Despite the use of aircraft’s environmental control system, the temperature inside the cockpit may easily reach more than 10 °C above ambient temperature and sometimes it may even exceed 45 °C. When these extreme temperatures are coupled with high relative humidity, causes for the degradation of both mental and physical performance of the pilots are present. This situation becomes severe, especially during low altitude and high-speed operations due to aerodynamic heating of the external surfaces. Sometimes, at high altitude and low-speed operations, the cockpit temperature falls and cold stress can pose a serious problem on the health of pilots. It is necessary to protect the pilots from high thermal stress to keep them under safe thermoregulatory limits and also help them perform an intended mission. This paper develops and deploys a basic method that can be used at an early design stage of any military aircraft to analyse the environmental control system’s performance in avoiding pilot thermal stress. The method is also applicable to a design study for an enhanced environmental control system on an existing aircraft. Results present the effect of parameters including Mach number, altitude, ambient temperature, cockpit geometry, and solar radiation on cockpit thermal balance which have a direct impact on the thermal stress on pilots. A military aircraft with a cockpit volume of 1.5 m3 is considered for performing the thermal balance simulation studies. This paper also addresses the effects on engine bleed flow requirements, and corresponding air inlet temperatures to maintain the cockpit target wet bulb globe temperature of 28 °C as well as a pilot mean skin temperature target of 33 °C. These are some of the thermal stress indicators proposed by different aero-medical authorities. These requirements are to be maintained if the physiological stress and impairment of performance of the pilots are to be avoided while operating in hot and humid environments
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    Stability and control effectiveness of a seamless aeroelastic wing aircraft.
    (2011-04) Yang, Da Qing; Guo, Shijun J.; Lawson, Craig
    Today’s aerospace engineers can design a much lighter aircraft than before thanks to the improved accuracy in the estimation of aerodynamic load and structural strength. But reduction of airframe material has also made the aircraft more flexible, giving rise to aeroelastic phenomena which would have profound effect on flight stability and control. From stability point of view, redistributed aerodynamic force due to wing deformation would influence the static stability of the aircraft. The structural flexibility and dynamic behaviour reduces the flight dynamic damping and influences the flight dynamic modes, which in turn could cause dynamic stability problems. From a control effectiveness perspective, control surface deflection exerts local aerodynamic force and deformation, which generally reduces intended control authority. The magnitude of control effectiveness reduction mainly depends on the structure and control surfaces design. An aircraft of special planform such as swept back wing makes the problem more complicated. This research project was motivated by the Active Aeroelastic Wing Technology (AAWT) demonstrated by the F/A-18AAW (later known as NASA X-53). The proposed research aircraft was a small flying wing type aircraft with 40 degree swept back angle and centrally mounted high swept fin. Powered by a single micro jet engine, this technology demonstrator features a Seamless Aeroelastic Wing (SAW). This study follows the AAWT approach, but extends the AAWT discrete rigid body control surfaces to spanwise continuous and chordwise hingeless control surfaces, integrated with internal actuation system. This research is aimed to achieve the stability and controllability of an aircraft with a specifically designed flexible seamless aeroelastic wing, which had control effectiveness problem. Firstly an investigation was made to mitigate the stability problems of the SAW aircraft. Then secondly, based on the SAW feature, the research proposes a systematic solution to the control effectiveness problem of this SAW aircraft. The solution would combine multidisciplinary efforts including aeroelastic tailoring, novel control surface actuation system design, distributed flight control management and active flight control algorithm. In the study, the effect of structural flexibility and dynamic behaviour of the wing on the flight dynamics has been considered. This thesis covers design, modelling, simulation and control of the SAW aircraft. The design, modelling, simulation, prototype construction and testing of the SAW were undertaken with collaboration of another PhD student and part of the work is also included in this thesis. The study has shown that the stability of the SAW aircraft can satisfy flying quality requirements with the compensation of stability augmentation system. The control effectiveness can be increased by SAW technology comparing with a conventional hinged control surface design. Flight envelope can be extended beyond control reversal speed with the aid of successfully designed gain scheduling control algorithm. The SAW aircraft shares many common aeroelastic problems with other flexible aircraft, but also has aeroelastic problems specific to its own design. If the design is changed, detailed solution would be altered accordingly, but the problem solving approach would be applicable. This research views these problems from a system level and endeavour to lead the way forward in achieving and demonstrating a systematic approach to a series of aeroelastic problems.