Browse Topic: Electric power
Propeller driven rotors utilize propellers on the main rotor blade to spin the rotor. Past research efforts have highlighted dynamic issues that arise from the rotor-propeller Coriolis interaction. For this paper, a comprehensive multi-body analysis methodology, called Elastic Rotorcraft Analysis (ERA), was applied to various propeller driven rotor datasets. The focus of the modeling effort was on propeller driven rotor twirl phenomenon, which arises from rotor-propeller inertial couplings interacting with rotor blade modes. After describing the phenomenon, the paper is split into two parts: validations and predictions. In Part I of the paper, the ERA propeller driven rotor model was validated using three datasets: (i) a propeller flapping vacuum chamber experiment, (ii) a propeller/rotor loads vacuum chamber experiment, and (iii) a propeller driven rotor hover experiment. The ERA model showed good agreement with the data, and captured the important rotor-propeller Coriolis interaction. In Part II of the paper, predictions for several propeller driven rotor configurations were generated and analyzed. Loads were computed for an isolated propeller and are compared to propeller loads during propeller driven rotor operation. The analysis showed that operating the propeller on the rotor blade introduces significant inertial loads on the propeller. Finally, propeller placement along the main rotor blade span was investigated. The results of the present study agree with earlier research, which showed placing the propeller at the midspan location reduced the electrical power coefficient by nearly half compared to a tip mounted propeller.
Researchers at the National Aeronautics and Space Administration (NASA) have conducted a series of module-level 50-ft dynamic drop tests on electric Vertical Take-off and Landing (eVTOL) Energy Storage Systems (ESS) for the generation of dynamic impact data to support standards developments. The tests were conducted on zero-state-of-charge Electric Power Systems (EPS) Electric Propulsion Ion Core (EPIC) modules at the National Institute for Aviation Research (NIAR), utilizing the NIAR outdoor drop test setup and conducted by NIAR test personnel. Four total tests were conducted on modules oriented in four different orientations. During initial post-test inspections at the drop facility, it was observed that the modules experienced varying amounts of damage in various locations and forms. The damage was quantified to the maximum extent possible via photogrammetric methods such as digital image correlation and marker tracking. Post-test modules were then disassembled, and forensics were conducted, which involved inspections into various containment structures and the cells themselves. A scoring rubric was developed in order to utilize a common methodology to quantitatively assess the damage incurred into each module as way of determining the least and most amount of damage present. Results in the form of post-test inspections, digital image correlation deformations, and scoring rubric values will be all presented, along with proposed paths forward for future work.
Aircraft Certification is a mature and complex bureaucracy that has successfully ensured a very high degree of safety of aircraft design, construction, operation and maintenance. Outside of a very few doing the work, there is a general lack of knowledge of certification details. For novel technologies such as electric power, and innovative configurations such as multi-rotors, the rules are far less mature and still emerging and so also poorly understood. Within the Advanced Air Mobility (AAM) initiative, many new aircraft developments are underway using novel configurations, and the public announcements of regulatory progress toward FAA or EASA Type Certification capitalize on this ignorance by being vague or even misleading. Honeywell conceived the Regulatory Readiness Level (RRL) indicator as an objective measure of certification status to serve the AAM industry and ecosystem, with applicability across aviation. The released RRL Version 1 now enables credible, objective assessment of new aircraft progress toward FAA Type Certification, and Operational Approval for Part 135 operations, to allow consistent apples-to-apples comparisons with other aircraft in development. An emerging complementary version of the rubric for EASA Type Certification is ready for publication to enable RRL determination against the European Union criteria. Future releases will consider other Nation's regulatory authorities, supplemental types certifications (STCs), and risk-based airworthiness assessments such as the Specific Operations Risk Assessments (SORA).
Researchers at the National Aeronautics and Space Administration (NASA) have conducted a series of module-level tests on electric Vertical Take-off and Landing (eVTOL) Energy Storage Systems (ESS) for the generation of dynamic impact data to support standards developments. The tests were conducted on zero-state-of-charge Electric Power Systems (EPS) Electric Propulsion Ion Core (EPIC) modules at the National Institute for Aviation Research (NIAR), utilizing the NIAR outdoor drop test setup and personnel. Four total tests were conducted. For each test, the module was dropped at a specific orientation from a height of 50 feet while connected to a guided trolley in order to assess the effects of a 50-foot drop test on the ESS. The test velocities ranged between 46.9 and 52.8 ft/s with impact angles ranging between a flat, zero-degree impact and 18 degrees. Data were recorded in the form of temperatures, cell-level voltage, module level acceleration and digital image correlation from the tests. Accelerations were in the range of 1,500 g for a few millisecond duration, which were indicative of a shock type loading condition. No modules entered thermal runaway, and post-test inspections revealed a variety of internal deformations and damage present in the various modules tested, with specific damage occurring for specific orientations. The modules were ranked according to a custom developed scoring rubric developed by utilizing the test and post-test inspection results. The results were compiled, reported, and will be used to guide future ESS testing. Part 1 discusses the loading environments in the modules, while Part 2 will discuss the deformation and damage in the modules.
This paper presents an object-oriented, equation-based framework for multi-engineering modeling of a quadrotor UAV, which includes the rigid body dynamics, simplified aerodynamics, gyroscopic effects, electrical power system and battery losses, and DC motor dynamics. An open-source drone modeling library is introduced by explaining the mathematical models and multi-domain components used to model the drone. Animation and visualization techniques for the drone using CAD models are also introduced and explained. The proposed drone model is simulated under different flight scenarios using motor and power system models with different levels of detail, aiming to provide better means for design and understanding, of multi-engineering aspects of UAVs. This model provides a foundation for future UAV open-source model development, electrified power propulsion design, visualization and interaction, and system identification.
This paper analyses the possibility of using hydrogen fuel cells as main energy provider for small to medium-sized eVTOL UAVs. A simplified model for eVTOL UAVs, which covers all relevant areas of aircraft design, including aerodynamics, structural mechanics, propulsion and systems modelling, is presented. Sensitivity studies with various design parameters, including aspect ratio and design cruise speed are performed to show their influence on the configurations’ performance. A comparison between pure battery electric and fuel cell hybrid UAVs is taken. The result of this paper is, that a hydrogen fuel cell hybrid configuration can have a better performance than a battery electric and it can be worth the effort to implement the fuel cell. To achieve this, the mission should require a long endurance and have hover and transition times reduced as far as possible, which both enable the high energy density of the hydrogen system to unfold its full potential. Also, the aircraft needs to be as aerodynamic efficient as possible in order to reduce the fixed weight of the hydrogen fuel cell. Respecting these requirements and accepting a reduced versatility of the aircraft, a fuel cell hybrid eVTOL UAV can easily outperform one with a pure battery electric power supply.
This paper describes the development of performance prediction models for the electric powertrains of group 1 unmanned aerial systems (G1UAS) that use sensorless brushless DC (BLDC) motor architecture consisting of a BLDC motor, electronic speed controller, and a battery. Per US Army definitions, G1UAS are platforms that weigh less than 20 lb. (9 kg). The resulting semi-empirical models for the motor, power electronics, and battery use high-level component specifications to enable pre-conceptual design space exploration and mission-based design optimization of G1UAS without a library of test data. The models also enable tradeoffs analysis between existing and/or conceptual designs without a series of flight tests. To develop, tune, and validate the models, a custom dynamometer test setup was designed and built to measure torque, speed, and electrical power data of small-scale motor drive systems. The validated models reveal that popular claims of high efficiency for electric powertrains are only valid in a narrow band of high speed/low torque operation. This is a critical finding for vehicle designers in the VTOL industry who are increasingly transitioning to electric powertrains in low speed/high torque applications which may decrease the overall system efficiency. A traditional rotor hover test stand was also developed to generate data with a traditional rotor load. The integrated motor and electronic speed controller model was able to predict the total efficiency of the hover stand tests within 5 percent of experimental values. The electrical models presented in this work can be immediately applied to design G1UAS given the torque and speed requirements of the rotors/propellers, the operating voltage of the vehicle system, and certain high-level component specifications for the motor, electronic speed controller, and battery. The models can also be used to BLDC powertrains for small terrestrial or aquatic electric vehicles.
Advances in electric propulsion may provide the opportunity to incorporate multiple distributed propulsors through the use of electric drives and electric power transmission at dramatically lighter weights than mechanical drivetrains. While small-scale UAS (0-20lb) have achieved distributed electric drivetrains, the electro-chemical power-sources do not scale and are weight prohibitive for larger UAS and future eVTOL aircraft with practical range, endurance, and payload requirements. An alternative to the electro-chemical powertrain is the hybrid-electric powertrain. In this paper, a variable-voltage hybrid-electric powertrain is designed, constructed, and tested carefully to examine the effects of variable-voltage generators for variable-RPM electric drives. The paper examines the powertrain component by component, with over 370 test points, from a single electric drive to an integrated powertrain with four distributed propulsors, in a fully instrumented test bed.
Icing of the fuselage and blades may occur when the helicopter is flying in the icing area. If ice accretion occurs in the ADS(Air Data System) of the fuselage, normal speed and altitude information are lost, making it difficult to flight. When windshield icing occurs, the view of pilot is limited and flight is difficult. Also, the ice accretion of the blades deforms the outer shape of the blades (Ref. 1) and makes the dynamic characteristics unstable due to an abnormal weight increase, resulting in deterioration of performance, deterioration of maneuverability, and structural instability. To avoid this, an anti-icing or de-icing system is required. Therefore, if the aircraft is not fitted with a proper anti-icing system, it is not possible to operate under icing conditions. However, it is difficult to design a proper anti-icing system considering the position of anti-icing protection area and icing phenomenon due to limitation of electric power, weight, thermal damage temperature limit, shape and so on. It is essential of understanding of anti-icing mechanism for selection and design of appropriate anti-ice system considering configuration (impingement limit, collection efficiency), material (thermal fatigue limit, heating source) and icing amount (collection efficiency, stagnation point). This paper introduces the preparation procedure (analysis and simulation, design, scaled model test) of icing certification test and the artificial/natural icing flight test method for proving helicopter icing through KUH case, and describes the characteristics of each system for the de/anti-icing system design of the helicopter, the computational analysis method, comparison with the test result, and the improvement method of the de/anti-icing system.
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