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For aircraft electromechanical actuator (EMA) cooling applications using forced air produced by axial fans, the main objective in fan design is to generate high static pressure head, high volumetric flow rate, and high efficiency over a wide operating range of rotational speed (1x∼3x) and ambient pressure (0.2∼1 atm). In this paper, a fan design based on a fan diameter of 86 mm, fan depth (thickness) of 25.4 mm, and hub diameter of 48 mm is presented. The blade setting angle and the chord lengths at the leading and trailing edges are varied in their suitable ranges to determine the optimal blade profiles. The fan static pressure head, volumetric flow rate, and flow velocity are calculated at various ambient pressures and rotational speeds. The optimal blade design in terms of maximum total-to-total pressure ratio and efficiency at the design point is obtained via CFD simulation. A 5-blade configuration yields the best performance in terms of efficiency and total pressure ratio.

The aircraft electromechanical actuator (EMA) cooling fan is a critical component because an EMA failure caused by overheating could lead to a catastrophic failure in aircraft. Fault tree analysis (FTA) is used to access the failure probability of EMA fans with the goal of improving their mean time to failure (MTTF) from ∼O(5×104) to ∼ O(2.5×109) hours without incurring heavy weight penalty and high cost. The dual-winding and dual-bearing approaches are analyzed and a contra rotating dual-fan design is proposed. Fan motors are assumed to be brushless direct current (BLDC) motors. To have a full understanding of fan reliability, all possible failure mechanisms and failure modes are taken into account. After summarizing the possible failure causes and failure modes of BLDC fans by focusing on each failure mechanism, the life expectancy of fan ball bearings based on a major failure mechanism of lubricant deterioration is calculated and compared to that provided in the literature.

The purpose of this study is to set up a laboratory test apparatus to analyze aircraft flight control EMAS' electrical and thermal energy flow under transient and dynamic flight profiles. A hydraulic load frame was used to exert load to the EMA. The actuator was placed within an environmental chamber which simulates ambient temperature as function of altitude. The simulated movement or stroke was carried out by the EMA. The under test EMA's dynamic load, stroke, and ambient temperature were synchronized through a real time Labview DAQ system. Motor drive voltage, current, regenerative current, and motor drive and motor winding temperature were recorded for energy analysis. The EMA under test was subjected to both transient and holding load laid out in a test matrix.

The scaling laws of fans express basic relationships among the variables of fan static pressure head, volume flow rate, air density, rotational speed, fan diameter, and power. These relationships make it possible to compare the performance of geometrically similar fans in dissimilar conditions. The fan laws were derived from dimensionless analysis of the equations for volumetric flow rate, static pressure head, and power as a function of fan diameter, air density and rotational speed. The purpose of this study is to characterize a fan's performance characteristics at various rotational speeds and ambient pressures. The experimental results are compared to the fan scaling laws.

In all-electric aircraft, electromechanical actuators (EMAs) will be used to replace hydraulic actuators. Due to the highly transient mission profiles of the aircraft operation, thermal management of EMAs is a significant issue. In this paper, we study the heat problem of the control and drive units of EMAs, and build a model to calculate and simulate the power loss and heat generation in the driver board. The driver unit consists of a power inverter, a capacitor, a power dissipating resistor and a control circuit. The power loss of each component is studied. The heat loss in the power inverter comes mainly from the power switches: IGBTs. The on-state loss is proportional to the current of the motor, and the switching loss is determined by the switching frequency as well as current.

In the aviation community, there is a high priority to develop all-electric aircraft. Electro-mechanical actuation systems would replace traditional, large, heavy and difficult-to-maintain hydraulic actuation systems. This movement from hydraulic actuation to electrical actuation enhances the flexibility to integrate redundancy and emergency system in future military aircraft. Elimination of the hydraulic fluid removes the possibility of leakage of corrosive hydraulic fluid and the associated fire hazard, as well as environmental concerns. The switch from hydraulic to electrical actuation provides additional benefits in reduced aircraft weight, improved survivability and improved maintainability. The heat load in an electro-mechanical actuation (EMA) is highly transient and localized in nature; therefore a phase change material could be embedded in the heat generating components to store peak heat load.

An Electromechanical Actuation System (EMAS) are an important component for an all electric Aircraft. EMAS would be lighter and require less system maintenance and operational costs than hydraulic actuators, typically used in aircraft systems. Also, hydraulic actuation systems require a constant power load to maintain hydraulic pressure, whereas EMAS only use power when actuation is needed. The technical challenges facing EMAS for aircraft primary flight control includes jam tolerance, thermal management, wide temperature range, high peak electric power draw, regenerative power, installation volume limit for thin wings, etc. This paper focuses on a laboratory test setup to simulate EMAS flight control environment to test and evaluate three important performance parameters of EMAS; thermal management, transient peak power draw, and regenerative power.

The development of electromechanical actuators (EMAs) is the key technology to build an all-electric aircraft. One of the greatest hurdles to replacing all hydraulic actuators on an aircraft with EMAs is the acquisition, transport and rejection of waste heat generated within the EMAs. The absence of hydraulic fluids removes an attractive and effective means of acquiring and transporting the heat. To address thermal management under limited cooling options, accurate spatial and temporal information on heat generation must be obtained and carefully monitored. In military aircraft, the heat loads of EMAs are highly transient and localized. Consequently, a FEA-based thermal model should have high spatial and temporal resolution. This requires tremendous calculation resources if a whole flight mission simulation is needed. A lumped node thermal network is therefore needed which can correctly identify the hot spot locations and can perform the calculations in a much shorter time.

A commercial electromechanical actuator (EMA) is to be dynamically tested with predetermined stroke and load profiles for transient thermal and electric power behavior to validate a numerical model used for aerospace applications. The EMA will follow the stroke profile representative of a real aircraft mission duty cycle. A hydraulic press will exert a corresponding load profile onto the EMA. Specialized hydraulic load control methods must be employed to meet the accuracy requirements. Two of these methods are closed-loop linearization (CLL) and displacement induced disturbance cancellation (DIDC). These methods are implemented along with an external PID compensator, and run in real-time in a series of system identification experiments to observe controller performance.

This paper describes the integrated modeling of a permanent magnet (PM) motor used in an electromechanical actuator (EMA). A nonlinear, lumped-element motor electric model is detailed. The parameters, including nonlinear inductance, rotor flux linkage, and thermal resistances, and capacitances, are tuned using FEM models of a real, commercial motor. The field-oriented control (FOC) scheme and the lumped-element thermal model are also described.

This paper presents the concept of a dual latent heat sink for thermal management of pulse heat generating electronic systems. The focus of this work is to verify the effectiveness of the concept during charging through experimentation. Accordingly, custom components were built and a prototype version of the heat sink was fabricated. Experiments were performed to investigate the implementation feasibility and heat transfer performance. It is shown that this heat sink is practicable and helps in arresting the system temperature rise during charging (period of pulse heat load).

The objective of this paper is to describe the design of an experiment that will examine the effects of elevated acceleration environments on a high-temperature, titanium-water loop heat pipe for actuator cooling. An experimental test setup has been designed for mounting a loop heat pipe on an 8-ft-diameter centrifuge table, which is capable of radial accelerations of up to 12 g's. A high-temperature PAO loop will interface the condenser of the loop heat pipe to simulate the rejection of the transported heat to an elevated temperature. In addition to LHP experimentation, a mathematical model has been developed for aerodynamic heating of highspeed aircraft. A flat plate at zero-incidence, used to model an aircraft wing, was subjected to sub- and supersonic flow to examine whether heat will be rejected or absorbed. The results of this analysis will be used to determine the condenser conditions of the loop heat pipe during centrifuge testing.