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The performance of a more-electric aircraft (MEA) power system electrical accumulator unit (EAU) architecture consisting of a 57000 rpm induction machine (IM) coupled to a controllable shaft load and controlled using direct torque control (DTC) is examined through transient modeling and simulation. The simplicity and extremely fast dynamic torque response of DTC make it an attractive choice for this application. Additionally, the key components required for this EAU system may already exist on certain MEA, therefore allowing the benefits of EAU technology in the power system without incurring a significant weight penalty. Simulation results indicate that this architecture is capable of quickly tracking system bus power steps from full regenerative events to peak load events while maintaining the IM's speed within 5% of its nominal value.

As the cost and complexity of modern aircraft systems increases, emphasis has been placed on model-based design as a means for reducing development cost and optimizing performance. To facilitate this, an appropriate modeling environment is required that allows developers to rapidly explore a wider design space than can cost effectively be considered through hardware construction and testing. This wide design space can then yield solutions that are far more energy efficient than previous generation designs. In addition, non-intuitive cross-coupled subsystem behavior can also be explored to ensure integrated system stability prior to hardware fabrication and testing. In recent years, optimization of control strategies between coupled subsystems has necessitated the understanding of the integrated system dynamics.

Minimizing energy use on more electric aircraft (MEA) requires examining in detail the important decision of whether and when to use engine bleed air, ram air, electric, hydraulic, or other sources of power. Further, due to the large variance in mission segments, it is unlikely that a single energy source is the most efficient over an entire mission. Thus, hybrid combinations of sources must be considered. An important system in an advanced MEA is the adaptive power and thermal management system (APTMS), which is designed to provide main engine start, auxiliary and emergency power, and vehicle thermal management including environmental cooling. Additionally, peak and regenerative power management capabilities can be achieved with appropriate control. The APTMS is intended to be adaptive, adjusting its operation in order to serve its function in the most efficient and least costly way to the aircraft as a whole.

In this paper, a MATLAB-Simulink based general co-simulation approach is presented which supports multi-resolution simulation of distributed models in an integrated architecture. This approach was applied to simulating aircraft thermal performance in our Vehicle Systems Model Integration (VSMI) framework. A representative advanced aircraft thermal management system consisting of an engine, engine fuel thermal management system, aircraft fuel thermal management system and a power and thermal management system was used to evaluate the advantages and tradeoffs in using a co-simulation approach to system integration modeling. For a system constituting of multiple interacting sub-systems, an integrated model architecture can rapidly, and cost effectively address technology insertions and system evaluations. Utilizing standalone sub-system models with table-based boundary conditions often fails to effectively capture dynamic subsystem interactions that occurs in an integrated system.

Design of modern aircraft relies heavily on modeling and simulation for reducing cost and improving performance. However, the complexity of aircraft architectures requires accurate modeling of dynamic components across many subsystems. Integrated power and thermal modeling necessitates dynamic simulations of liquid, air, and two-phase fluids within vapor cycle system components, air cycle machine and propulsion components, hydraulic components, and more while heat generation of many on-board electrical components must also be precisely calculated as well. Integration of these highly complex subsystems may result in simulations which are too computationally expensive for quickly modeling extensive variations of aircraft architecture, or will require simulations with reduced accuracy in order to provide computationally inexpensive models.

In this paper, first an analytical model of a synchronous generator with a pulsed constant power load (CPL) is developed and numerically compared with a detailed simulation model. The analytical model is shown to possess good predictive abilities, thus enabling its use for control purposes. Second, the generator has a proportionalintegral (PI) control inner-loop, whose task is to regulate the generator’s output voltage to a desired reference. A novel outer-loop predictive reference governor (PRG) is designed and tested via simulation. The PRG uses the analytical model to predict the output behavior of the generator over a short time window, and continuously modifies the reference given to the inner-loop in order to maintain stringent steady-state requirements, in spite of demanding power requirements at the CPL. Simulation results illustrate the significant performance advantages of using the PRG versus using the inner-loop PI controller alone.

Future aircraft systems are projected to have order of magnitude greater power and thermal demands, along with tighter constraints on the performance of the power and thermal management subsystems. This trend has led to the need for a fully integrated design process where power and thermal systems, and their interactions, are considered simultaneously. To support this new design paradigm, a general framework for codifying and checking specifications and requirements is presented. This framework is domain independent and can be used to translate requirement language into a structured definition that can be quickly queried and applied to simulation and measurement data. It is constructed by generalizing a previously developed power quality analysis framework. The application of this framework is demonstrated through the translation of thermal specifications for airborne electrical equipment, into the SPecification And Requirement Evaluation (SPARE) Tool.

Model based design is a standard practice within the aerospace industry. However, the accuracies of these models are only as good as the parameters used to define them and as a result a great deal of effort is spent on model tuning and parameter identification. This process can be very challenging and with the growing complexity and size of these models, manual tuning is often ineffective. Many methods for automated parameter tuning exist. However, for aircraft systems this often leads to large parameter search problems since frequency based identification and direct gradient search schemes are generally not suitable. Furthermore, the cost of experimentation often limits one to sparse data sets which adds an additional layer of difficulty. As a result, these search problems can be highly sensitive to the definition of the model fitness function, the choice of algorithm, and the criteria for convergence.

In this work we present our recent effort in developing a novel heat exchanger based on endothermic chemical reaction (HEX reactor). The proposed HEX reactor is designed to provide additional heat sink capability for aircraft thermal management systems. Ammonium carbamate (AC) which has a decomposition enthalpy of 1.8 MJ/kg is suspended in propylene glycol and used as the heat exchanger working fluid. The decomposition temperature of AC is pressure dependent (60°C at 1 atmosphere; lower temperatures at lower pressures) and as the heat load on the HEX increases and the glycol temperature reaches AC decomposition temperature, AC decomposes and isothermally absorbs energy from the glycol. The reaction, and therefore the heat transfer rate, is controlled by regulating the pressure within the reactor side of the heat exchanger. The experiment is designed to demonstrate continuous replenishment of AC.

As More Electric Aircraft design becomes the preferred system concept for several aerospace platforms, the electro-mechanical actuator (EMA) is emerging as a solution of choice for the primary flight control actuation system. This paper will give a brief history of electric actuation for flight systems, diagnosis and prognosis demonstrations and current state of health management research. AFRL and NASA working with industry and academic partners have been developing health management technologies that will help prevent the occurrence of some inherent EMA failure modes. Advanced fault diagnostics and failure prognostics were applied to the critical failure modes identified in the Failure Mode, Effects, and Criticality Analysis (FMECA). Modeling and simulation of EMA with degraded components were developed to support the design and evaluation of physics-based algorithms. Test data were generated using EMA hardware to validate high-fidelity EMA and physics-of-failure models.

The Air Force Research Laboratory has interest in automatically generating the extensive aerodynamic databases essential for six degree of freedom (6DOF) models and the use of 6DOF models for design. To be most useful, automation must include all aspects of producing the database including meshing, control surface deflections, running the CFD solution, and storage of the results. This effort applies newly-developed software to produce the desired results. Firstly, AFRL software called Computational Aircraft Prototype Syntheses (CAPS) allows automated meshing using the Advancing Front Local Reconnection (AFLR) software from Mississippi State University1 and automated control surface deflection using Engineering Sketch Pad (ESP) software from MIT/Syracuse. CAPS includes the ability to run the NASA CFD code FUN3D and interpret the FUN3D results via an Application Interface Module (AIM). This may sound like a complicated process.

The trend of moving towards model-based design and analysis of new and upgraded aircraft platforms requires integrated component and subsystem models. To support integrated system trades and design studies, these models must satisfy modeling and performance guidelines regarding interfaces, implementation, verification, and validation. As part of the Air Force Research Laboratory's (AFRL) Integrated Vehicle and Energy Technology (INVENT) Program, standardized modeling and performance guidelines have been established and documented in the Modeling Requirement and Implementation Plan (MRIP). Although these guidelines address interfaces and suggested implementation approaches, system integration challenges remain with respect to computational stability and predicted performance over the entire operating region for a given component. This paper discusses standardized model evaluation tools aimed to address these challenges at a component/subsystem level prior to system integration.

In this paper, a parametric average-value modeling approach is applied to a high-frequency six-phase aircraft generation subsystem. This approach utilizes a detailed switch-level model of the system to numerically establish the averaged dynamic relationships between the ac inputs of the rectifier and the dc-link outputs. A comparison between the average-value and detailed models is presented, wherein, the average-value model is shown to accurately portray both the large-signal time-domain transients and the small-signal frequency-domain characteristics. Since the discontinuous switching events are not present in the average-value model, significant gains can be realized in the computational performance. For the study system, the developed average-value simulation executed more than two orders of magnitude faster than the detailed simulation.

Numerous previous studies have highlighted the potential efficiency improvements which can be provided to aircraft thermal management systems by the incorporation of vapor cycle systems (VCS), either in place of, or in conjunction with, standard air cycle systems, for providing the needed thermal management for aircraft equipment and crews. This paper summarizes the results of a cycle-based VCS control architecture as tested using the Vapor Cycle System Research Facility (VCSRF) in the Aerospace Systems Directorate of the Air Force Research Laboratory at Wright-Patterson Air Force Base. VCSRF is a flexible, dynamic, multi-evaporator VCS which incorporates electronic expansion valves and a variable speed compressor allowing the flexibility to test both components and control schemes. The goal of this facility is to reduce the risk of incorporating VCS into the thermal management systems (TMS) of future advanced aircraft.

A reduced order dynamic aircraft model has been created for the purpose of enabling constructive simulation studies involving integrated thermal management subsystems. Such studies are motivated by the increasing impact of on-board power and thermal subsystems to the overall performance and mission effectiveness of modern aircraft. Previous higher-order models that have been used for this purpose have the drawbacks of much higher development time, along with much higher execution times in the simulation studies. The new formulation allows for climbs, accelerations and turns without incurring computationally expensive stability considerations; a dynamic inversion control law provides tracking of user-specified mission data. To assess the trade-off of improved run-time performance against model capability, the reduced order formulation is compared to a traditional six degree-of-freedom model of the same air vehicle.

Future Air Force intelligence, surveillance, and reconnaissance (ISR) platforms, such as high-altitude Uninhabited Aerial Vehicles (UAV), may drastically change the requirements of aircraft power systems. For example, there are potential interactions between large pulsed-power payloads and the turbine engine that could compromise the operation of the power system within certain flight envelopes. Until now, the development of large-scale, multi-disciplinary (propulsion, electrical, mechanical, hydraulic, thermal, etc.) simulations to investigate such interactions has been prohibitive due to the size of the system and the computational power required. Moreover, the subsystem simulations that are developed separately often are written in different commercial-off-the-shelf simulation programs.

As aircraft continue to increase their power and thermal demands, transient operation of the power and propulsion subsystems can no longer be neglected at the aircraft system level. The performance of the whole aircraft must be considered by examining the dynamic interactions between the power, propulsion, and airframe subsystems. Larger loading demands placed on the power and propulsion subsystems result in thrust, speed, and altitude transients that affect the aircraft performance and capability. This results in different operating and control parameters for the engine that can be properly captured only in an integrated system-level test. While it is possible to capture the dynamic interactions between these aircraft subsystems by using simulations alone, the complexity of the resulting system model has a high computational cost.

The movement to more-electric architectures during the past decade in military and commercial airborne systems continues to increase the complexity of designing and specifying the electric power system. In particular, the electrical power system (EPS) faces challenges in meeting the highly dynamic power demands of advanced power electronics based loads. This paper explores one approach to addressing these demands by proposing an electrical equivalent of the widely utilized hydraulic accumulator which has successfully been employed in hydraulic power system on aircraft for more than 50 years.

The diverse and complex requirements of next-generation energy optimized aircraft (EOA) demand detailed transient and dynamic model-based design (MBD) to ensure the proper operation of numerous interconnected and interacting subsystems across multiple disciplines. In support of the U.S. Air Force's Integrated Vehicle Energy Technology (INVENT) program, several MBD-derived software tools, including models of EOA technologies, have been developed. To validate these models and demonstrate the performance of EOA technologies, a series of Integrated Ground Demonstration (IGD) hardware tests are planned. Several of the numerous EOA software tools and MBD-based processes have been updated and adapted to support this activity.

Detailed machine models are, and will continue to be, a critical component of both the design and validation processes for engineering future aircraft, which will undoubtedly continue to push the boundaries for the demand of electric power. This paper presents a survey of experimental testing procedures for typical synchronous machines that are applied to brushless synchronous machines with rotating rectifiers to characterize their operational impedances. The relevance and limitations of these procedures are discussed, which include steady-state drive stand tests, sudden short-circuit transient (SSC) tests, and standstill frequency response (SSFR) tests. Then, results captured in laboratory of the aforementioned tests are presented.