Propulsion System Structural Response to Variation in Aircraft Architecture

This paper presents the design and performance details of the Chevrolet Voltec electric propulsion system. The propulsion system has two machines, one machine is operating mostly as a motor while the other machine is operating mostly as a generator. Both machines of the Voltec electric drive system are permanent magnet AC synchronous machine types with the magnets buried inside the rotor. The motor has distributed windings. However, as opposed to a conventional stranded winding the Chevrolet Volt motor has bar-wound construction to improve the motor performance, especially in the low to medium speed range. At higher speed the skin and proximity effects in the stator bars lead to increased stator winding losses but are addressed in the design. The bar-wound construction also has excellent thermal performance, in both the steady-state and the transient conditions, necessary for full EV driving. The generator uses concentrated windings. The concentrated winding construction has good slot fill and short end-turn length. These features resulted in good performance in the intended operational region and were an enabler for machine packaging inside the transmission. Both the machines exhibit excellent efficiency and exceptionally smooth and quiet operation. Machine design and construction details as well as the measured thermal, electromagnetic and acoustic noise performances are presented in the paper.

Current research trends have advanced the use of “green propellants” on a wide scale for spacecraft in various space missions; mainly for environmental sustainability and safety concerns. Small satellites, particularly micro and nanosatellites, evolved from passive planetary-orbiting to being able to perform active orbital operations that may require high-thrust impulsive capabilities. Thus, onboard primary and auxiliary propulsion systems capable of performing such orbital operations are required. Novelty in primary propulsion systems design calls for specific attention to miniaturization, which can be achieved, along the above-mentioned orbital transfer capabilities, by utilizing green monopropellants due to their relative high performance together with simplicity, and better storability when compared to gaseous and bi-propellants, especially for miniaturized systems. Owing to the ongoing rapid research activities in the green-propulsion field, it was necessary to extensively study and collect various data of green monopropellants properties and performance that would further assist analysts and designers in the research and development of liquid propulsion systems. This review traces the history and origins of green monopropellants and after intensive study of physicochemical properties of such propellants it was possible to classify green monopropellants to three main classes: Energetic Ionic Liquids (EILs), Liquid NOx Monopropellants, and Hydrogen Peroxide Aqueous Solutions (HPAS). Further, the tabulated data and performance comparisons will provide substantial assistance in using analysis tools—such as: Rocket Propulsion Analysis (RPA) and NASA CEA—for engineers and scientists dealing with chemical propulsion systems analysis and design. Some applications of green monopropellants were discussed through different propulsion systems configurations such as: multi-mode, dual mode, and combined chemical–electric propulsion. Although the in-space demonstrated EILs (i.e., AF-M315E and LMP-103S) are widely proposed and utilized in many space applications, the investigation transpired that NOx fuel blends possess the highest performance, while HPAS yield the lowest performance even compared to hydrazine.

A new scramjet engine model has been developed to support hypersonic vehicle design studies and flight dynamics and control system analysis. This paper explains the methodology and the governing equations for the new propulsion system model that is suitable for use with a control oriented dynamic model of a hypersonic vehicle. Previous propulsion models used for this purpose were based on simple Rayleigh flow for the combustion process, but despite this, captured the propulsion system interactions with the vehicle aerodynamics and structural dynamics. A new, higher fidelity propulsion system model is constructed that simulates numerous phenomena that were neglected in the Rayleigh flow approach. The new model is of higher fidelity, and therefore it is not designed to calculate the flow physics on a timescale that is suitable for dynamics and control simulations. Instead it will be used as a truth model and the starting point for the derivation of a reduced-order model. Specific phenomena that are included in the new model are: a pre-combustion shock train within the isolator and its interactions with the combustor, the loss of stagnation pressure due to gas dissociation and recombination, wall heat transfer and skin friction, a fuel-air mixing submodel, and a finite-rate chemistry and autoignition reaction mechanism. It is shown that the new propulsion system model expands the operability envelope as compared to the previous model by accommodating ramjet combustion, which occurs at high supersonic/low hypersonic flight Mach numbers.

Adelis-SAMSON is a nano-satellite mission aimed at performing geo-location of target signals on Earth using a tight three-satellite formation in space. To maintain formation, each nano-satellite is equipped with a cold gas propulsion system. The design, qualification, and integration of the Adelis-SAMSON nano-satellite propulsion system is presented in this paper. The cold gas propulsion system mass is approximately 2 kg, takes a volume of 2U, and generates a thrust of 80 mN from four thrusters using krypton as a propellant. We first present the propulsion system requirements and corresponding system configuration conceived to meet the mission requirements. Subsequently, we present the system architecture while listing all the components. We overview the particular role and qualification process of four of the propulsion system’s components: the propellant tank, thruster assembly, pressure regulators, and fill and vent valve. We detail the tests performed on each component, such as proof pressure tests, vibration tests, and external leak tests. Finally, we present the propulsion system level tests before delivery for satellite integration and discuss the propulsion system’s concept of operations.

Achieving noble and more practical propulsion systems is still a major concern for engineers and researchers. Accordingly, many researchers have strived to opt for better and suitable system for the targeted vessels. One of the marine propulsion systems is linear-jet propulsion which includes rotor and stator. In this paper, the effects of geometrical shape of the duct in the form of different length and angle of attack for two types of the accelerating and decelerating ducts are comprehensively explored. Furthermore, section of rotor from Kaplan series are selected and the geometrical shape effects of the rotor on hydrodynamic performance of linear-jet propulsion system are also investigated. Therefore, to obtain the effects of the expressed parameters on the hydrodynamic characteristics of this propulsion system, Ansys-CFX software has been used. The governing equations of the studied problem are RANS equations and SST k-omega turbulent model is utilized. The obtained results are presented in the form of dimensionless coefficients of the forces applied on the propulsion system. The existing data for a ducted propeller are used to validate the numerical solution. Based on the results of the numerical solution, it is concluded that use of a decelerating duct increases the efficiency of the linear-jet system. On the other hand, the use of accelerating ducts reduces the hydrodynamic efficiency of the linear jet, so the performance is not desirable at high advance ratio. Therefore, the use of decelerating ducts in the linear-jet system is found to be the appropriate choice. Results further illustrate that, if the duct length increases, the generated thrust by the rotor increases. It is also concluded that, by increasing the angle of attack of the decelerating duct, the overall efficiency of the linear-jet propulsion system decreases.

The importance of underwater robots is evident in ocean exploration, resource development, and environmental monitoring. However, the harsh underwater environment requires higher efficiency, stability, and intelligence from their propulsion systems. The challenges faced by operational underwater robots today include low propulsion efficiency, poor adaptability to extreme environments, and a lack of sufficient autonomous control capabilities. To address these issues, this paper reviews the definition, requirements, core technologies, and key performance indicators of underwater robot propulsion systems by analyzing relevant literature from 2016 to 2024. It emphasizes optimization strategies aimed at enhancing propulsion efficiency, fault diagnosis and identification, reliability, durability, and adaptive control. Besides, it summarizes the current technical challenges and provides a reference for subsequent research. The results show that optimizing the propulsion system of operational underwater robots relies primarily on bionic design, new materials, adaptive control, deep learning, and fault diagnosis technologies to enhance propulsion efficiency, stability, durability, and environmental adaptability. However, optimizing the propulsion system involves challenges such as energy control, cost, and multi-objective optimization. Future research should prioritize efficient, low-energy propulsion, multi-modal perception, and intelligent adaptive control to advance underwater robot technology.

Operational requirements of ships that navigate in ice‐covered waters are important drivers for the design of the ships' propulsion systems and powering. Using typical operational requirements to illustrate, connections are made to the design of the propeller, propulsion system machinery, and powering. Propeller–ice interaction is a unique feature of ships that operate in ice. It affects elements of propulsion efficiency as well as the strength of the propulsion system components, particularly the propellers. The physical processes involved in propeller–ice interaction are described, including how ice contact forces are imposed on propellers and how the propellers are consequently loaded. Design implications of propeller–ice interaction and ice loads are discussed. A method for estimating propulsion power at the early stages of design is presented and some machinery types are identified. The article begins with a review of the elements of propulsion efficiency and power to put the discussion of propulsion in ice into context.

An economic model is presented which relates engineering performance characteristics to product profitability for conceptual and preliminary design of propulsion systems and aircraft. The model represents the competitive marketplace rigorously, which sets it apart from the prevailing state of the art. The model is shown to produce significantly different design results than current approaches. INTRODUCTION Today the design, certification and tooling costs to introduce a new aircraft engine can exceed one billion dollars. To make such an investment responsibly, design engineers must be focused on creating a product that will succeed in the marketplace, providing a good value for prospective customers [Collopy, 1997]. However, today’s design teams lack the economic tools to translate among engineering parameters, market needs, and costs. Therefore, new propulsion systems are based on goals such as “reduce aircraft total operating cost by 10%” [Aviation Week & Space Technology, 1996] that neglect airline revenue and engine manufacturing costs. Attempts to incorporate economic values, such as Altman [1994], although laudable, do not incorporate rigorous, complete representation of market operation and therefore do not generally lead to optimal designs. This paper presents an economic model that relates aircraft performance and manufacturing cost to aircraft, airline, and engine profitability. The model incorporates a rigorous representation of the aircraft market. It is shown that the total profit of the airframer, engine manufacturer, and purchasing airline, referred to as the surplus value, is set by the aircraft and engine design; the market functions only to divide the profit among the parties. Furthermore, it will be shown that the optimal design for airframe and engine is that design which maximizes surplus value. This same design maximizes profit for the airline, the airframer and the engine manufacturer—a true win-win solution. An example is provided in which the economic model is employed in a simplified design study to optimize bypass ratio. The results using the model are significantly different than with prior techniques. WHY AIRLINES PURCHASE AIRCRAFT Aircraft purchases are complex decisions. Salesmanship, maintenance agreements, manufacturing offsets and international politics all come into play. But virtually all purchases rest on a fundamentally rational financial base: the intent of the airline to make money by operating the aircraft. A simple economic model of the aircraft purchase decision can be made by assuming that operating profit is the airline’s only objective in purchasing an aircraft. The influence of this model on propulsion system design is instructive. Revenue and Operating Cost Airline operating profit is the difference between revenue and operating cost. The revenue that an aircraft can earn for an airline depends on the payload and range capability of the aircraft, in conjunction with airline’s route structure and the passenger and freight demand on those routes. Note in Figure 1 that payload and range are not separate attributes—the usual measure is in fact a function of maximum payload which depends on range. Economic probabilistic models can be used to estimate the

In modern times, the design and optimization of different actuator systems for controlling a high-precision position control system represent a popular interdisciplinary research area. Initially, only single-stage actuator systems were used to control most of the motion control applications. Currently, dual-stage actuation systems are widely applied to high-precision position control systems such as hard disk drive (HDD) servo systems. In the dual-stage system, a voice coil motor (VCM) actuator is used as the primary stage and a piezoelectric micro-actuator is applied as the secondary stage. However, a dual-stage control architecture does not show significant performance improvements to achieve the next-generation high-capacity HDD servo system. Research continues on how to fabricate a tertiary actuator for a triple-stage HDD servo system. A thermal positioning controller (TPC) actuator is considered promising as the tertiary stage. The triple-stage system aims to achieve greater bandwidth, track density, and disk speed, with minimum sensitivity and greater error minimization. In this work, these three actuation systems with different combinations of proportional plus integral (PI), proportional plus derivative (PD), and proportional plus integral plus derivative (PID) controller, lag-lead controller, lag filter, and inverse lead plus a PI controller were designed and analyzed through simulation to achieve high-precision positioning. The comparative analyses were done on the MATLAB/Simulink simulation platform.

With the advancements in aircraft gas turbine engine technology over the past decade (in materials, structures, and analytical design methods) along with those anticipated for the next decade, there is renewed interest in the Mach 3.0 to 6.0 operating regime for both commercial transports and military weapon systems. The propulsion systems for these airplanes will in all probability be a function of the flight regime selected for a particular application; i.e., a turbofan engine for Mach<4.0, and a dual mode turboramjet for the Mach 4.0 to 6.0 aircraft. There are a number of component technologies that will have to be developed for the propulsion system before these applications can become a reality. This paper addresses some of the tasks that will have to be accomplished relative to one of those components: the exhaust nozzle system. Introduction The exhaust nozzle system is a unique component which requires integration with the airplane system components (engine, winglbody, inlet) as well as special design considerations for acoustics and off-design point operation in order to optimize the overall system performance. The design requirements for the exhaust system stem from two sources: the airplane and the engine (see Figure 1). Requirement 0 Airplane Sustained supersonic/subsonic operation Propulsion system integration Community noiw regulations Landing WeatherItiW brakes 0 Advanced engine technology Dual mode cycle High exhaust system temperature Figure 1. Factors Affecting Nozzle Design. EWct on nozzle Multiple design points. high WA,, and eiector nozzles Drag and propulsion system performance Acoustic design considerations Thrust reverser Bypau ducWvalw, variable area mixers, and secondary air systems Cooling/materia$ High exhaust system pressure ratios With reference to Figure 1, airplane performance and operational requirements have a direct impact on the exhaust system design. For example: sustained high Mach operation requires high nozzle area ratio capability for best efficiency with the nozzle possibly being an integral part of the airplane; for commercial transports, FAA noise regulations impact engine size and require noise ASIA8 excunion/seals/ st~ctures/~~1uaton suppression and treatment techniques; inletlengine airflow mismatching at transonic speeds may necessitate an ejector nozzle to reduce large spillage drags; and, landing operations may call for a thrust reverser for safety as well as prolonging the life of tires and brakes. Manager, Exhaust Systems & Installed Technology Member AIAA The engine cycle aerothermodynamic and operating characteristics also affect the nozzle design (reference Figure 1). For example: high nozzle pressure ratios again mean high area ratios for high performance; large changes in nozzle area ratio to optimize overall performance will more than likely result in nonaxisymmetric nozzles for Mach 4.0 to 6.0 applications; high tailpipe temperatures require more cooling air andlor different cooling schemes; high pressures in the nozzle will require innovative structural design approaches; and installed performance optimization may necessitate auxiliary air systems requiring ejector nozzles for performance optimization as well as for cooling. In addition, integration between the engine and aircraft companies will be absolutely essential in order to optimize the system (airplanelinletlenginelnozzle) installed performance. This paper will discuss some of the issues and design considerations for the Mach 4.0 to 6.0 aircraft with primary emphasis on turboramjet propulsion systems which utilize a common inlet and exhaust system for both operating modes. Some discussion relative to Mach 3f civil transports which utilize turbofan propulsion systems will also be included. Technical Discussion Expanded Airbreathing Propulsion Corridor The operating envelope for future high Mach aircraft is shown on Figure 2 compared to the regime for current aircraft (shaded area). The area of interest for this paper is Mach 2.5 to 6.0 where advanced technology turbofanlturbojet (Mach <4.0) and turboramjet (4.O<M<6.0) propulsion systems are being considered. The aircraft applications include military weapon systems for both defensive and offensive missions and commercial transports.

Most Unmanned Aerial Vehicles (UAV) currently use propulsion systems based on gas turbines or internal combustion engines. However, such systems cannot be used in all sizes of UAVs. To perform a high- endurance mission, a UAV must have an efficient propulsion system and aerodynamic. On the other hand, batteries have a low energy density, which therefore increases the UAV’s weight and imposes penalties on the system. Using fuel cells as the main source of power generation in the UAV propulsion system can increase flight endurance and have reasonable fuel consumption. The aim of the present work is introduce and investigate of a new electric propulsion system based on electric motor, proton exchange membrane fuel cell (PEMFC) and battery for a small UAV. Two different battery discharge strategies are compared for propulsion system design. In addition, static optimization is applied to determine the appropriate size of the various components of the propulsion system. It was found that, the required power at the theoretical endurance speed is about 42.13% lower than that at the actual endurance speed. In addition, the required hydrogen rate and PEMFC area to provide the required power in the proposed hybrid propulsion system are 55.7 g/h and 0.09 m2, respectively. Moreover, the payload weight (with lithium-ion polymer battery) for two battery discharge strategies was 0.71 and 1 kg, respectively. The sensitivity analysis is also used to determine the parameters affecting the final performance of the propulsion system.

The lightweight structures and unconventional configurations being considered for the next generation of aircraft mean that any effort to predict or control the flight dynamics is impacted by the structural dynamics. One of the most severe forms of coupling between aeroelasticity and flight dynamics is an instability called body freedom flutter. The existing tools often assume a relatively weak effect of structural dynamics on the flight dynamics, and are therefore incapable of modeling strong interactions like body freedom flutter. A method of combining different sources of data traditionally used for aeroelasticity and flight dynamics is described by reconciling many of the differences between these models. By building upon past modeling efforts, a level of familiarity in the approach is achieved. Generally the differences from the traditional approaches are subtle but significant. The traditional frequency domain flutter model in a modal coordinate system is converted to a form consistent with a time domain flight dynamics model. The time domain rational function approximation about a non-inertial coordinate system and the unique constraints for the conversion between the inertial and non-inertial coordinate systems are discussed. A consistent transformation of the states of aeroelastic models to flight dynamics models is derived, which enables the integration of data from higher fidelity computational fluid dynamics models or wind-tunnel testing. The present method of integrating multidisciplinary data was used to create models that compare well with X-56A flight-test data, including conditions past the flutter speed.

U.S. Space Station Freedom propulsion system operations

The development of commercial aircraft with hybrid-electric propulsion systems is currently a subject of extensive research in order to improve local air quality and reduce combustion emissions. Among the various types of engines being studied, the two-spool parallel hybrid-electric turbofan engine is particularly challenging due to the low-pressure compressor (LPC). The hybridisation process tends to throttle the LPC, accentuating its significance in the propulsion system. For reliable operation of such systems, accurate predictions of the LPC performance during time-sensitive manoeuvres such as a go-around are important. These manoeuvres are heavily influenced by time-dependent effects that govern the propulsion system’s performance. Often, the aerothermodynamic interplay between these effects is overlooked in propulsion models. In this study, the influence of these aerothermodynamic interdependencies on the modelling results is investigated. To investigate these aerothermodynamic interactions, a dynamic model is developed to simulate the performance of the hybrid-electric turbofan engine. In comparison, a constant mass flow model is used, which is not able to simulate these interdependencies. The results show that the aerothermodynamic interdependencies significantly affect the modelled time-resolved performance, especially for surge margin of the LPC, with this effect becoming more pronounced at higher levels of hybridisation. Therefore, the study recommends the adoption of dynamic simulation methodologies for hybrid-electric engines to guarantee high simulation precision, enhance reliability, and satisfy safety standards.

This paper deals with the modeling and simulation of the ship electric propulsion system. This system is depicted by a non-linear mathematical model. This analysis highlights the operation of a propulsion system of a ship and allows the modeling of its elements. Through a non-linear global model we obtain the performance of a total propulsion system. The numerical simulation is realized through Matlab/Simulink software.