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Fluid pressure pulsation in a fluid system is an inherent consideration in applications such as aircraft engine and control systems where mechanical component fatigue life and flow performance are critical. Positive displacement pumps transmitting fluid through hydraulic lines under high pressure impart periodic flow pulses to the fluid which can induce undesirable pressure ripple. Some failures of advanced aircraft prototype hardware were traced to a break in the hydraulic component of the control system due to severe localized responses to periodic pressure pulsations produced by a pump flow-induced ripple at the system resonant frequency. This response is associated with a strong structural fluid resonance that is not sufficiently damped by fluid leakage internal to the aircraft hydraulic system. In the case of pumps or hydraulic motors the main source of pulsation energy is in the flow-induced pressure wave associated with the system plumbing pressure pulsations.

In last few years there has been great research to increase safety of on-road vehicles by providing information of various vehicle parameters to the user/driver while driving on road. Many algorithms have been developed to assess the vehicle run time situations and enable vehicle ECU to take decisions for autonomous driving. These algorithms are derived using data captured from sensors predominantly make use of vehicle dynamic information. The design proposed in this paper discusses capturing of two important and critical vehicle run time parameters i.) Vehicle tire pressure and the ii.) Road gradient. These parameters then help us in determining the effective fuel efficiency of the vehicle and approximate distance that user can drive with the amount of fuel remaining in the tank.

Variable pressure, variable displacement hydraulic pumps are described and their capability to increase the efficiency of advanced aircraft hydraulic systems is discussed. Potential methods of interfacing with the hydraulic system are also discussed and the dynamic response is analyzed.

The V-22 Fly-By-Wire Digital Flight Control System is arguably the most complex yet attempted. In addition to providing control of the aircraft in its multiple flight modes, (helicopter-conversion-airplane) the V-22 FCS is also required to provide an integrated Thrust / Power Management, aircraft maneuver limiting and navigation coupling. The V-22 Flight Control System is also unique in its use of dedicated digital data busses as an integral part of the system architecture, and in its use of inertial reference systems as prime stability augmentation sensors. The FCS also incorporates a 5000 psi hydraulic system. Accordingly, the testing of such a system becomes a very challenging task, especially given the schedule constraints of the V-22 program. This paper will address the means by which the V-22 Flight Control System was tested prior to first flight, with prime focus being the software verification, hardware-software integration and system validation testing.

To the casual airline passenger the appearance of the Boeing 747-400 jetliner is identical to that of the older model 747-300 with the exception of the winglets. However, modifications and improvements include modern digital avionics, a “glass” cockpit and more powerful rudder control surfaces to cope with the growth in engine thrust. The upper rudder is positioned by three parallel actuators and powered by two independent hydraulic systems. Flow control is provided by a triple tandem servovalve which maintains actuator force fight below allowable limits and synchronizes hydraulic flow to the three actuators. This paper describes the requirements, design, development, analysis and iron bird testing of some of the unique features of the upper rudder hydraulic control system.

The United States Space Shuttle Orbiter Auxiliary Power Unit (APU) provides power to the orbiter vehicle hydraulic system. The system operates critical flight functions during a mission including aerodynamic control surface actuation, Space Shuttle main engine thrust vector control, and landing gear deployment, steering, and braking. The APU was designed in the mid-1970s, then flight-certified in support of the approach and landing tests (ALTs) in 1977 and initial orbital flight tests (OFTs) in 1981. During APU design, development, and flight certification, it was necessary to develop a number of technology items representing the state of the art to meet the challenging requirements of space operation. Improvements in the APU are continuing to further enhance life and reliability. More than 1,500 hours of operating time have been accumulated during APU development, certification, and Space Shuttle flights.

Exhaust emissions from a vehicle under road driving condition is affected by the control state of ECU (Engine Control Unit). This control state highly depends on the driving force of the vehicle. The driving force is nearly equal to the driving resistance, which is the sum of the acceleration resistance, the air resistance, the rolling resistance and the gradient resistance. Although it is essential to take an accurate measurement of the road gradient, it is quite difficult to evaluate the gradient resistance in testing on-road driving. In this study, the measurement methods of the road gradient and the altitude with GPS, gyro sensor and height sensor are reported. The road gradient under the on-road driving condition is evaluated by the combination of measuring the pitch angle with the gyro sensor and measuring the vehicle gradient with the two height sensors. Verifying of this method, the altitude of the driving test route is also evaluated.

The pressure losses across the different parts of a regenerative Diesel Particulate Filter (DPF) have been modeled and compared with the measured pressure loss and with the measured changes in the instantaneous weight of the DPF of a commercial automotive diesel engine. The comparisons were made in three operating conditions selected among those included in the transient cycle established in the European Emission Directive. The first one is a low-load mode, with high soot emissions and therefore with high contribution to the DPF charge. The second one is a medium-load mode, in which the balance of soot charge versus spontaneous soot regeneration leads to a slow DPF charging, the temperature at the exhaust manifold being high enough to permit active regeneration. The third one is a high-load mode, in which the spontaneous regeneration leads to a net DPF discharge, the active regeneration becoming useless.

The MD-95 is a 100 passenger twin jet airplane based on the DC-9-30 airplane. This paper describes the changes made to the hydraulic system to improve reliability, reduce maintenance costs, save weight, or simply to update the technology. Changes were also required for interfacing with the flight control system and the status/alerting system.

Development and verification of systems for internal aircraft networks include multiple software layers. These layers are mainly the application-specific components, communication layers, redundancy management and other system services. Verification of these system layers in the early stages of the design process, before a physical network is available, and during the design process has become a critical need in order to reduce design costs and project risks. Time-Triggered Architectures (TTA) and SCADE are both well-established technologies and tools for building safety-critical embedded systems. Both are based on the synchronous paradigm; TTA for the communication infrastructure and distributed embedded computing, and SCADE for simulating and generating code for the application components.

In the filter industry conventional differential pressure indicator devices, which are mounted into the filter housing, generally have only one actuation setting. This setting is based on a predetermined value of differential pressure across the filter element (i.e., the difference in pressure measured upstream and downstream of the filter). When the differential pressure exceeds a certain preload setting, a “pop-up” button and/or electrical signal is activated. Normally the user does not know how long it will take for the filter element to become clogged by the contaminants it continually captures. The user only becomes aware upon actuation that this differential pressure has been reached. The purpose of this paper is to describe a hydraulic fluid monitoring device which has been developed to provide real time, continuous data for the differential pressure developed across a filter within the hydraulic system.

This paper presents the results of an investigation to improve design/analysis techniques for aircraft hydraulic systems. A design/analysis tool was developed by integrating control-surface commands and loads obtained from Aircraft Dynamic Simulator Software (ADSS) with an enhanced version of the HYdraulic TRansient ANalysis (HYTRAN) program. Control-surface commands and loads from an ADSS simulation of a selected maneuver were used as dynamic input to the HYTRAN program so that the hydraulic system response could be predicted throughout the maneuver. Predicted hydraulic system pressures and control-surface positions from the HYTRAN simulation of the maneuver were compared to flight-test data and were found to be in excellent agreement. The successful coalescence of the two independent software programs gives engineers a concurrent design/analysis tool that can be used to optimize hydraulic system designs during the very early stages of design.

This paper describes a dynamic aircraft hydraulic system simulator model. This model uses a flexible approach that accounts for flight control loads, system stiffness, line losses, internal leakage at individual valve packages, pump compensator response, and power input to engine-driven pumps at various engine rpm. The model integrates performance of the flight controls, hydraulics, and power sources allowing accurate simulation of the airplane.

This document presents and discusses the design, development, and demonstration of a 3,000 to 8,000 psi (207 to 552 bar) electronically controlled, overcenter, variable-displacement hydraulic motor (VDHM) in which output speed and position control are effected solely through wobbler position control. This approach contrasts with traditional systems that are normally controlled through a large servovalve which throttles flow across the motor control ports. By connecting the variable-displacement hydraulic motor (VDHM) motor directly to supply and return and controlling through wobbler modulation, the need for a large throttling servovalve with its associated flow losses is eliminated. This can result in significant improvement in peak flow consumption, overall efficiency, and package size.

A passive anti icing system for the engine cowl of the Global Hawk UAV was designed and evaluated under a NASA Phase I SBIR Program. The system utilizes five Loop Heat Pipes to remove 3.8 kW of waste heat from the hydraulic system and deliver it to the engine inlet during icing conditions. During non-icing conditions the heat is bypassed to relief valves on the fuel ullage tank to maintain their temperature above freezing. The system masses 18.5 kg and replaces some 15 lbs of existing equipment. A Phase II program to integrate the system aboard the aircraft is underway with the goal of starting flight tests in early 1999.

Modelling and simulation is of crucial importance for the understanding of system dynamics. In aircraft, simulation has been strong in the area of flight control. Modelling and simulation of the hydraulic systems has also a long tradition. The rapid increase in computational power has now come to a point where complete modelling and simulation of all the sub systems in an aircraft is not far away. This means new challenges in dealing with very complex multi domain systems. Of all the systems in an aircraft fluid power systems is one of the most difficult to handle from a numerical point of view. They are characterised by difficulties such as discontinuities, very strong non-linearities, stiff differential equations and a high degree of complexity. A considerable effort has therefore been made to develop methods suitable for simulation of such systems.

The Smart Crane Ammunition Transfer System (SCATS) is designed to handle, deliver, and reload missiles/ammunitions in the battlefield. The system objectives are as follow: faster missile reloading, decreased utilization of manpower, increased operator safety, reduced missile/munitions damage, reduced operator skills and training, and increased material flow. In this paper, the smart crane ammunition transfer system is introduced and discussed, including the crane, the hydraulic system, the control system, and the development environment. The emphasis is on the design, analysis, and implementation of the SCATS control system. State-of-the-art engineering techniques, such as robotic control, trajectory generation, sensory processing and task planning, are also addressed as to the way in which they will enhance the SCATS system performance and functionality as applied to the automation of complex material handling and resupply actions.

This paper describes how airplane control surface rate limiting can enable a ‘cliff-like’ onset of Pilot-Induced Oscillation, (P.I.O.) and how the danger can be erased by implementation of Conditionally Enabled Phase Compensated Rate Limiters, (PCRLs), in the design of the airplane's flight control system. The application is particularly important for large airplanes where control surface actuator sizing and the associated hydraulic system volumetric flow rate capability cannot be generously over-sized without large weight and cost penalties. It is shown that the PCRL can remain inactive during normal airplane operations where RMS control commands are relatively small thus avoiding adverse control surface response effects that have hindered earlier PCRL acceptance.

Good performance of the Columbus water loop active control system has been demonstrated by several analyses, ground test and is further confirmed by the current flight data. Even so, a comprehensive description of the control within the classical theory is needed, in order to complete the system description, posing also the basis for similar applications to come. Thermal and hydraulic control loops are considered as two separate systems and linear control methods are applied. Loop stability and performance is discussed by computing stability regions of the PI control coefficients at different loop configurations and results compared with available test, flight and simulation data.

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.