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In 1991, shortly after acquiring Learjet, Bombardier consolidated all flight testing of new aircraft at the Wichita, Kansas facility. Since then, nine new aircraft were certified, and the Flight Test Center grew from 20 dedicated flight test personnel, to nearly 500 dedicated flight test personnel. The Canadian based company in conjunction with several international risk sharing partners, has created a highly dynamic flight test environment, tasking the Flight Test Center with the challenge of bringing a new product to market each year. This rapid growth was centered on supporting three aircraft product lines; Learjet, Canadair, and DeHavilland. New hangars, telemetry, and ground support facilities were built to accommodate the increased flight test demands. The Bombardier Flight Test Center, otherwise known as BFTC, conducts flight test operations on a seven day per week schedule, and in 1999, flew over 5000 flight test hours in development and certification testing.

The hydraulic actuators are used to power flight control surfaces of the aircraft and to ensure surface movement. A system of two or three actuators is usually designed depending on the surface and intuitively these actuators are considered as a redundant architecture from a reliability and functionality point of view. The proper reliability modeling of the system of actuators must consider the system's functionality and design constraints for the remaining available actuator hinge-moment in the event of a partial or total actuator failure. As a result, this will affect the reliability assessment of that design. Furthermore, this system of actuators is also designed to provide a second function involving an assurance of the surface stiffness and damping. Generally, this second function does not require necessarily the same number of available actuators in order to be fully provided.

Traceability has always been considered a useful but costly activity and different methods have been applied to reduce this cost. The current paper constitutes an attempt to improve these methods by introducing an optimal traceability process to be used in the context of RTCA DO-297 “Integrated Modular Avionics (IMA) Development Guidance and Certification Considerations”. The paper starts by comparing the definitions of traceability from DO-297 and the related development guidelines (i.e. ARP4754A, DO-254 and DO-178B). The paper continues by classifying the traceability methods recommended by the guidelines and introducing a performance criterion for optimal traceability based on category theory. This criterion addresses the possibility of information loss present in the current traceability methods. The paper proposes an optimal traceability process (i.e. that guarantees that information is not lost) and exemplifies it. The paper ends by recommending further enhancements.

A multi-axis serially redundant, single channel, multi-path FBW (FBW) control system comprising: serially redundant flight control computers in a single channel where only one “primary” flight control computer is active and controlling at any given time; a matrix of parallel flight control surface controllers including stabilizer motor control units (SMCU) and actuator electronics control modules (AECM) define multiple control paths within the single channel, each implemented with dissimilar hardware and which each control the movement of a distributed set of flight control surfaces on the aircraft in response to flight control surface commands from the primary flight control computer, and a set of (pilot and co-pilot) controls and aircraft surface/reference/navigation sensors and systems which provide input to a primary flight control computer and are used to generate the flight control surface commands in accordance with the control law algorithms implemented in the flight control computers.

Integrated modular avionics (IMA) architectures host multiple federated avionics applications on a single platform and provide benefits in terms of size, weight, and power, which, however, leads to increased complexity, especially during the development process. To cope efficiently with the high level of complexity, a novel, structured development methodology is required. This paper presents a model-based systems engineering (MBSE) development approach for the so-called “distributed integrated modular architecture” (DIMA). The proposed methodology adapts the open-source Capella tool, based on the Architecture Analysis & Design Integrated Approach (ARCADIA) methodology, to implement a complete design cycle, starting with requirements captured from the aircraft level to streamline the development, culminating in the integration of an avionics application into an ARINC 653 platform.

The aircraft components' lifetime is a key decision–making metric for the performance of safety–critical items. The piece–part degradation and age–related changes are critical from the perspective of design and continued airworthiness. The most obvious issue during design development is to establish the need for planned replacement for components that are known to have a limited life. During investigation of an airworthiness issue, it is necessary to determine if the anomaly is time–dependent. If it is, then the anticipated failure probability as a function of time must be estimated such that a decision regarding corrective action can be made. For both cases, an analysis must be performed to determine if and when planned replacement is necessary. Because unanticipated retrofits are costly and difficult, credible and accurate lifetime prediction is essential.

Advanced commercial aircraft increasingly use more composite or hybrid (metal and composite) materials in structural elements and, despite technological challenges to be overcome, composites remain the future of the aviation industry. Composite and hybrid aircraft today are equipped with digital systems such as fly by wire for reliable operations no matter what the flying environment is. These systems are however very sensitive to electromagnetic energy. During flight, aircraft can face High Intensity Radiated Fields (HIRF), static electricity, or lightning. The coupling of any of these threats with airframe structure induces electromagnetic energy that can impair the operation of avionics and navigation systems. This paper focuses on systems susceptibility in composite aircraft and concludes that the same electromagnetic rules dedicated to all metal aircraft for systems and wiring integration cannot be applied directly as such for composite aircraft.

This paper discusses the correlation of IVHM (Integrated Vehicle Health Management) as an emerging aerospace discipline and the Big Data paradigm widely discussed in the Information Technology industry. The 4-V model is discussed to qualify a Big Data problem in terms of the volume, variety, velocity and veracity of the data involved. Big Data management allows, for example, correlations to be found to “spot business trends, determine quality of research, prevent diseases, combat crime, and determine real-time roadway traffic conditions”. Examining these two fields side by side is necessary and desirable because innovation is very likely to occur when and where different but correlated domains interface. This paper compares the most significant technical components required for Big Data Analytics and IVHM to work.

The lightning represents a fundamental threat to the proper operation of aircraft systems. For aircraft protection, Electromagnetic Compatibility requires conductive structure that will provide among all, electromagnetic shielding and protection from HIRF and atmospheric electricity threat. The interaction of lightning with aircraft structure, and the coupling of induced energy with harnesses and systems inside the airframe, is a complex subject mainly for composite aircraft. The immunity of systems is governed by their susceptibility to radiated or conducted electromagnetic energy. The driving mechanism of such susceptibility to lightning energy is the exposure to the changing magnetic field inside the aircraft and IR voltage produced by the flow of current through the structural resistance of the aircraft. The amplitude of such magnetic field and IR voltage is related to the shielding effectiveness of the aircraft skin (wiremesh, composite conductivity).

During the development of an aircraft, a comprehensive understanding of the electrical load profile is essential to properly estimate the required electrical power to be generated and distributed by the electrical system, also known as EPGDS - Electrical Power Generation and Distribution System. By sizing the EPGDS early in the development process, system parameters like weight and volume can be estimated and applied to the multidisciplinary design optimization process, in search for optimized design solutions at the conceptual aircraft level when developing integrated aircraft systems. With this in mind, a methodology was developed to estimate the amount of electrical power required by the aircraft systems during a typical mission flight cycle.

This paper presents a generic Redundancy Management (RM) requirements definition process that is applicable to a complex system RM requirements development. In the aerospace industry, the ‘Aerospace Recommended Practices’ (ARP) 4754 and 4761 are typically used processes to ensure given safety and availability goals for complex systems. The process proposed in this paper is based on these standard guidelines and enhances them to provide a standardized process for the development of RM requirements with interactions between the system requirements development and the preliminary system safety assessment processes. The output of this process will help to achieve the following objectives: The system RM/failure monitoring requirements are defined commensurate with the system safety and availability requirements; the system is fault-tolerant to the degree necessary to meet the system safety and availability requirements; the system is robust and the system architecture is optimized.

This paper presents a methodology for conceptual aircraft design to evaluate the space available for systems (top-down approach) and to estimate the space required for critical components impacting the aircraft configuration (bottom-up approach). The presented top-down approach introduces the concept of “equivalent design volume”, including the space required for systems and the associated empty space to access, maintain and ventilate them. This approach enables an early feasibility check for aircraft configuration exploration regarding the integration and installation of systems, without having to detail the system architecture. In complement, the bottom-up approach introduces the estimation of the required dimensions for critical components. Here, the example of the flight control actuators integration in the wing tip is presented.