Abstract: An aircraft is provided that includes one or more engines, an air-driven generator, and a start-circuit for operably connecting an electrical output of the air driven generator for starting one or more of the engines while the aircraft is in flight. Various embodiments also provide apparatus for cross-ship electric starting of multiple engines while the aircraft is in flight or on the ground using: an electrical output of a generator attached to another engine; on-board batteries; an on board auxiliary power unit; or power received from a ground source. A single start-converter is utilized to supply power to AC starter-motors for electric starting of the aircraft's engines. In some embodiments, the engine is soft-started with the air driven generator, by controlling blade pitch in an air driven turbine driving the air driven generator, without the use of the start converter.

Abstract: Optimal power-limited rendezvous with variable exhaust velocity is investigated for propulsion systems having both upper and lower bounds on thrust magnitude. In this model the spacecraft thrusters have four admissible states, thrusting at the upper saturation level, unsaturated power-limited thrusting, thrusting at the lower saturation level, and unpowered (i.e., engine off.) A fifth chattering state is also possible. The mathematical structure of the solution of the optimal rendezvous problem associated with this propulsion model is found. Computer simulations of rendezvous with a satellite in circular orbit are presented. For rendezvous near circular orbit four classes of chattering solutions of two types can occur. It is assumed that multiple thrusters can be mounted on the spacecraft and that they can operate independently. Applying the efficiency condition presented in a previous paper, a logic for switching individual thrusters on and off for optimal fuel usage and computational efficiency is presented.

Abstract: The vehicle includes a fuselage; a plurality of lifting surfaces attached to the fuselage having control devices attached thereto; and, an articulated propulsion system attached to the fuselage. The propulsion system includes a duct assembly pivotally connected to the fuselage. The duct assembly includes a duct and a propeller assembly mounted within the duct. A motor assembly is connected to the propeller assembly. The duct assembly may be positioned in a substantially vertical position to provide sufficient direct vertical thrust for vertical take-off and landing and may be directed in other positions to provide a varying spectrum of take-off and landing configurations, as well as a substantially horizontal position for high speed horizontal flight. Use of the control surface in the ducted propulsion assembly provides VTOL capability in a very small environment. The environment is not required to be prepared in any special manner. During horizontal flight, the wings provide the lift, which is more efficient than a propeller providing lift. The present invention takes advantage of a center line propulsion, so that there are no asymmetric propulsion loads.

Abstract: Analytical methods were combined with actual thruster data to create a model used to predict the performance of systems based on two types of electric propulsion thrusters, Hall-effect thrusters and ion engines, for several orbit transfer missions. Two missions were trip time constrained: a LEO-GEO transfer and a LEO constellation transfer. Hall thrusters were able to deliver greater payload due to their higher overall specific power. For the power limited orbit topping mission, the choice of thruster is dependent on the user’s need. Ion engines can deliver the greatest payload due to their higher specific impulse, but they do so at the cost of higher trip time. Study of reusable electric orbit transfer vehicle systems indicates that they can offer payload mass gains over chemical systems, but that these gains are less than those offered by other electric propulsion transfer scenarios due to the necessity of carrying propellant for return trips. Additionally, solar array degradation leads to increased trip time for subsequent reusable transfers. * Research Aerospace Engineer, Member AIAA ** Group Leader, USAF Electric Propulsion Lab, Member AIAA This paper is declared a work of the US Government and is not subject to copyright protection in the United States. INTRODUCTION: The US Air Force has recently completed several studies to investigate the potential advantages of advanced space propulsion for several orbit transfer scenarios. The first study investigated advanced propulsion concepts for expendable orbit transfer vehicles and concluded that the potential launch vehicle downsizing that resulted from the use of high specific impulse thrusters provided significant cost savings over base line chemical launch vehicle/upper stage systems. The second study looked at reusable advanced upper stages and preliminary indications are that while there remains the potential for launch vehicle downsizing, it is significantly reduced compared to expendable systems. This difference was largely due to the added propellant required to perform the round trip mission from low-earth orbit to geostationary orbit. Both studies pointed out advantages for advanced electric propulsion systems based on xenon propellant. The objective of this paper is to analyze the tradeoffs between Hall-effect thrusters and ion engines as a high power propulsion system for orbit transfer missions. Both the Hall-effect thruster and the gridded ion engine are classified as electrostatic thrusters and operate on heavy noble gases, primarily xenon. These electric propulsion devices are capable of specific impulses ranging from approximately 1500 to 4000 seconds, compared to chemical systems which typically operate in the range of 300 to 400 seconds. Electric propulsion is a type of rocket propulsion for space vehicles and satellites which utilizes electric and/or magnetic processes to accelerate a propellant at a much higher specific impulse than attainable using classical chemical propulsion. The concomitant reduction in required propellant mass results in increased payload mass capability. The method of analysis used in this study is based on the model developed by Messerole. It has been modified to reflect the most current information on thruster development levels and

Abstract: A method is presented to obtain approximate initial costate values and flight time for the optimal control of a continuous-thrust spacecraft on a coplanar, circle-to-circle transfer. The approximate initial costates are then used as starting values for the associated boundary-value problem to match the desired final states. The exact, nonlinear differential equations are integrated to solve the boundary-value problem with a shooting method. The approximate expressions for the initial costates and flight time are useful when the thrust acceleration is greater than or equal to the change in orbital radius, in canonical units. Numerical examples are provided for a geocentric and an Earth-Mars orbital transfer. EW propulsion technologies have raised interest in the space community for continuous-thrust orbital missions. In military applications, this could mean more responsive deployment of space assets and longer on-orbit lifetimes. Tactical repositioning of a satel- lite using chemical propulsion can consume large amounts of the available fuel mass per maneuver. This is certainly unacceptable for many reasons including lifetime, reliability, and cost. A space- craft propelled by high-efficiency thrusters could accomplish many more maneuvers for the same amount of fuel mass as a chemical propulsion system. Another interesting application of continuous-th rust propulsion is interplanetary space travel. A permanently orbiting space station could serve as a launch point for solar system exploration. Maintain- ing cryogenic fuels for this purpose, however, would be technically difficult and extremely costly. These problems would be less signif- icant if the orbiting space station were used as an assembly point for an exploration vehicle propelled by high-efficien cy continuous thrusters. Such a vehicle could be made reusable much more easily in terms of reliability and cost than a chemically propelled space- craft. Also, a vehicle using continuous thrust with existing technol- ogy could shorten travel times compared to fuel-optimal impulsive maneuvers, then could return to Earth orbit for reconditioning. Although the optimization of impulsive transfers has a direct solution,1 none has been found for the continuous-thrust case. This problem may be solved numerically, and many examples of this are to be found in the literature.2 Optimization of a continuous- thrust trajectory involves the simultaneous solution of an optimal control problem and a boundary-value problem. The initial and final states are normally known, but there is usually no infor- mation available for the initial values of the Lagrange costates. This presents quite a problem, since the optimal control law is often a function of the Lagrange costates that must be initialized for numerical integration. The usual approach is to make an edu- cated guess for the initial values, then update them by solving the boundary-value problem. Trussing,3 Broucke,4 and others2 have re- cast the boundary-value problem in terms of other variables, but the initial values of these must be guessed and refined as well. Prussing3 incorporated the second derivative of the primer vec- tor into a fourth-order dynamics equation, thus eliminating the control variables. Once this is accomplished, four constants of

Abstract: Hall thruster systems based on the SPT-50 and the TAL D-38 were evaluated and mission studies were performed. The 0.3 kilowatt SPT-50 operated with a specific impulse of 1160 seconds and an efficiency of 0.32. The 0.8 kilowatt D-38 provided a specific impulse above 1700 seconds at an efficiency of 0.5. The D-38 system was shown to offer a 56 kilogram propulsion system mass savings over a 101 kilogram hydrazine monopropellant system designed to perform North-South station keeping maneuvers on board a 430 kilogram geostationary satellite. The SPIT-50 system offered a greater than 50% propulsion system mass reduction in comparison to the chemical system on board a 200 kilogram low Earth orbit spacecraft performing two orbit raises and drag makeup over two years. The performance characteristics of the SPF-50 were experimentally evaluated at a number of operating conditions. The ion current density distribution of this engine was measured. The performance and system mass benefits of advanced systems based on both engines were considered.

Abstract: A heating or cooling fluid circuit for an electric vehicle is used to keep the drive or propulsion battery at the right temperature. The fluid is circulated by a pump through a heat exchanger (11) and past the electric driving motor(8). There is a bypass (5) for this main branch(4), controlled by a valve(3). The main branch also cools the battery charger (7) and the control electronics. Upstream from the heat exchanger is a continuous flow heater (9). There is a header tank (13) connected to the circuit by a stub pipe (12) at the highest point.

Abstract: Aerospace Propulsion Systems. Basic Thrust Equations. Propellers, Fans, and Rotors. Aerospace Propulsion Thermodynamic Cycles. The Gas Turbine. Ramjets and Scramjets. Rockets. Environmental Impact.

Abstract: The Aircraft Noise Predication Program (ANOPP) is an industry-wide tool used to predict turbofan engine flyover noise in system noise optimization studies. Its goal is to provide the best currently available methods for source noise prediction. As part of a program to improve the Heidmann fan noise model, models for fan inlet and fan exhaust noise suppression estimation that are based on simple engine and acoustic geometry inputs have been developed. The models can be used to predict sound power level suppression and sound pressure level suppression at a position specified relative to the engine inlet.

Abstract: A system using a stored power unit, such as a battery pack, to supply electric current for driving an electric motor-driven generator, which in turn, drives a propulsion motor drive system. The propulsion motor drive system uses either one, two or several electric propulsion motors to propel a vehicle. The system has a control device that diverts power from the generator output to the stored power unit to recharge and maintain the stored power unit when operating conditions permit. Amorphous energy sources, such as a wind turbine that is mounted on the vehicles roof, for example, can also supply recharging power. By use of strict energy control, the vehicle may be operated over many miles without having to recharge the batteries from a utility power source.

Abstract: A new type of contact-less wafer manipulator, featuring "direct electrostatic levitation and propulsion of silicon wafer" (DELP-SW), has been successfully developed. The novel aspect of this manipulator is that a silicon wafer can be directly levitated and driven via electrostatic forces. In this paper, a brief review of basic principles is presented. This is followed by a description of the structure of a prototype DELP-SW mechanism, including electrode design, position feedback control method, driving principle and the operational procedure. Experimental results which demonstrate completely contact-less transportation of an 8-inch silicon wafer are also presented.

Abstract: A preliminary mixed-compression inlet design concept for potential pulse-detonation engine (PDE) powered supersonic aircraft was defined and analyzed. The objectives of this research were to conceptually design and integrate an inlet/PDE propulsion system into a supersonic aircraft, perform time-dependent CFD analysis of the inlet flowfield, and to estimate the installed PDE cycle performance. The study was baselined to a NASA Mach 5 Waverider study vehicle in which the baseline over/under turboramjet engines were replaced with a single flowpath PDE propulsion system. As much commonality as possible was maintained with the baseline configuration, including the engine location and forebody lines. Modifications were made to the inlet system's external ramp angles and a rotating cowl lip was incorporated to improve off-design inlet operability and performance. Engines were sized to match the baseline vehicle study's ascent trajectory thrust requirement at Mach 1.2. The majority of this study was focused on a flight Mach number of 3.0. The time-dependent Navier Stokes CFD analyses of a two-dimensional approximation of the inlet was conducted for the Mach 3.0 condition. The Lockheed Martin Tactical Aircraft Systems-developed FALCON CFD code with a two equation 'k-1' turbulence model was used. The downstream PDE was simulated by an array of four sonic nozzles in which the flow areas were rapidly varied in various opening/closing combinations. Results of the CFD study indicated that the inlet design concept operated successfully at the Mach 3.0 condition, satisfying mass capture, total pressure recovery, and operability requirements. Time-dependent analysis indicated that pressure and expansion waves from the simulated valve perturbations did not effect the inlet's operability or performance.

Abstract: A unified robust multivariable approach to propulsion control design has been developed at NASA Glenn Research Center. The critical elements of this unified approach are: a robust H/sub /spl infin// control synthesis formulation; a simplified controller scheduling scheme; and a new approach to the synthesis of integrator windup protection gains for multivariable controllers. This paper presents results from an application of these technologies to control design for linear models of an advanced turbofan engine. The objectives of the study were to transfer technology to industry and to identify areas of further development for the technology. The technology elements and industrial development of tools to implement the steps are described with respect to their application to a GE variable-cycle turbofan engine. A set of three-input/three-output three-state linear engine models was used over a range of power levels covering engine operation from idle to maximum unaugmented power. Results from simulation evaluation are discussed and insight is provided into how the design parameter choices affect the results.

Abstract: The focus of propulsion integration technology in the 21st century will be economy. USAF inlet and nozzle technology goals translate into 50% weight reduction and 25% acquisition cost reduction metrics for new aircraft system. Innovative technology to enable these reductions over current state-of-the-art systems in weight and cost is required. For inlet systems, compact diffusers that reduce system volume by 50% will demand fewer parts and improved aerodynamic performance. Exhaust systems will be fixed with fewer parts, requiring a technology like fluidics, for example, to provide area control and thrust vectoring capabilities. Cooperative programs for both inlet and nozzle systems are in place to insure that technologies required to meet weight and cost reduction goals are matured by the year 2000.Copyright © 1996 by ASME

Abstract: This paper gives an overview of the NASA F-18 High Alpha Research Vehicle. The three flight phases of the program are introduced, along with the specific goals and data examples taken during each phase. The aircraft configuration and systems needed to perform the disciplinary and inter-disciplinary research are discussed. The specific disciplines involved with the flight research are introduced, including aerodynamics, controls, propulsion, systems, and structures. Decisions that were made early in the planning of the aircraft project and the results of those decisions are briefly discussed. Each of the three flight phases corresponds to a particular aircraft configuration, and the research dictated the configuration to be flown. The first phase gathered data with the baseline F-18 configuration. The second phase was the thrust-vectoring phase. The third phase used a modified forebody with deployable nose strakes. Aircraft systems supporting these flights included extensive instrumentation systems, integrated research flight controls using flight control hardware and corresponding software, analog interface boxes to control forebody strakes, a thrust-vectoring system using external post-exit vanes around axisymmetric nozzles, a forebody vortex control system with strakes, and backup systems using battery-powered emergency systems and a spin recovery parachute.

Abstract: A propulsion unit, preferably a propulsion unit for a utility vehicle, as well as a method for controlling same are disclosed. The propulsion unit comprises an engine, a transmission and a power take-off unit having a power take-off shaft. A first fraction of the engine output power is directed to the transmission, whereas a second fraction is directed to the power take-off shaft. At least one of the power fractions is detected by means of measuring units. The power fraction transmitted to the transmission is limited to a predetermined limit value. It is, therefore, possible to install a propulsion unit with an engine, the output power of which being higher than the admissible input power of the transmission.

Abstract: The availability of new high-energy permanent magnets at a reasonable cost has opened up the possibility of developing very compact brushless permanent magnet motors with substantially improved overload capability for wheel direct drive applications. Concerning the development of an electric propulsion system for a newly-conceived dual-power city car, this paper discusses the design, construction and laboratory testing of a twin wheel direct drive prototype based on a novel topology of water-cooled axial-flux permanent magnet motor.

Abstract: A marine propulsion engine control system wherein the control includes an arrangement for slowing the speed of the engine by disabling certain cylinders in the event of an abnormal engine running condition. Also, an arrangement is provided for slowing the speed of the engine if a change speed transmission for driving the propulsion shaft by the engine offers more than a predetermined resistance to shifting. The controls are interrelated so that the engine protection control predominates. That is, if the engine is in protection control mode and the operator attempts a shift and more than a predetermined resistance is felt, the shift control routine will not be initiated to effect any additional engine speed reduction. In addition, when the engine speed is reduced, fuel is continued to be supplied by the fuel injectors to avoid backfiring, stalling, and uneven running. When rapid deceleration is called for the spark advance is rapidly retarded but fuel injection amount is gradually decreased.

Abstract: This paper is divided into two major parts. The first part focuses on the YF-22 aircraft and includes an overview of the Advanced Tactical Fighter (ATF) prototype program, a review of the control law development process and a thorough analysis of the YF-22 PIO mishap. This review will include a discussion of the Air Force accident review board conclusions and the corrective action plan that was developed after the accident investigation. The second part of the paper will discuss the development of the F-22 control laws. The F-22 control law design process will be reviewed; including the design philosophy developed for the F-22 and the use of simulation and other analysis tools. Particular emphasis will be placed on the handling qualities and PIO metrics that have been used to shape the longitudinal axis flying qualities of the F-22. Introduction to the ATF Program The YF-22A was the Lockheed, General Dynamics and Boeing entry into the Air Force Advanced Tactical Fighter (ATF) demonstration and validation program. It was a prototype of a single seat, twin engine air superiority fighter that combined stealth, agility and sustained supersonic cruise capabilities. It was designed to optimize the blend of these aircraft characteristics for the demanding threat environment of the 21 century. Aircraft agility was deemed essential across the entire Mach/altrtude envelope of the aircraft. The YF-22 incorporated thrust vectoring nozzles to provide enhanced agility and extend the maneuvering envelope of the aircraft to extreme angles of attack. The external lines of the aircraft Copyright © 1996 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. were very comparable to the F-22, which is shown in Figure 1. The YF-22 was designed to meet the requirements of the ATF program and to reduce the risk for the Engineering Manufacturing Development (EMD) program. An extremely aggressive control law development schedule and flight test program led to the successful demonstration of the outstanding capabilities of the aircraft. As one might expect for a program whose ultimate goal was to demonstrate a number of new technologies, the YF-22 control laws started with a very conventional inner-loop design and added capabilities as aerodynamic and propulsion models matured. The basic command architecture was very similar to that used on the F-16. The initial control law version did not have active thrust vectoring. Many enhanced capabilities were added to the YF-22 control laws as the prototype program matured and flight test approached. The use of thrust vectoring nozzles to augment the aerodynamic pitch control power of the aircraft was incorporated into the control laws and the high angle of attack control laws were also added to the basic structure. For flight test purposes, a switch was provided in the cockpit to engage/disengage thrust vectoring. The flight envelope was initially cleared up to 20° angle of attack with vectoring disengaged. Test points were repeated with vectoring engaged. Envelope expansion into the high angle of attack flight regime with vectoring engaged and disengaged followed the initial buildup. In addition to the incorporation of thrust vectoring, other program objectives, such as the requirement to provide very high pitch rate capability in certain parts of the envelope, were added to the YF-22 control laws. Many of these features were designed to demonstrate capabilities at a specific point in the envelope and, 155 American Institute of Aeronautics and Astronautics due to schedule constraints, did not represent a full envelope, production aircraft design. YF-22 Control Law Development Philosophy The YF-22 control law feedback/feedforward gains were designed using fairly conventional tools. Design goals were based on accepted short period mode, roll mode and dutch-roll mode frequencies, damping ratios and mode shapes. An eigenstructure assignment algorithm was used to calculate the initial feedback gains for the longitudinal axis. However, by combining the pitching moment control power of the horizontal tails and thrust vectoring nozzles into a single, generalized controller, this technique essentially reduced to a pole placement algorithm. The short period mode frequency/damping and the integrator mode frequency were the primary design parameters. The short period mode design goals were also fairly conventional: • Control Anticipation Parameter (CAP)=1.0 • Damping Ratio (Q=0.8 Further refinements to the control law gains were developed from analysis of off-line simulation time histories (step/doublet pitch stick inputs) and comments from piloted evaluations using the fixed-base, YF-22 Handling Qualities Simulator (HQS). Command gradients and gain tweaks to shape rise times and mode characteristics were developed through piloted evaluations in the HQS. Due to the time constraints of the prototype program, much of the analysis of the handling qualities of the YF-22 was performed after the control laws were designed. That analysis generally consisted of stability margin predictions and time history analyses that were incorporated into the Flying Qualities Substantiation report and used to support safety and flight readiness reviews. Demonstration/Validation Flight Test Program The YF-22 development program was extremely successful and, in a very aggressive flight test program, the team was able to demonstrate the exceptional high angle of attack characteristics of the YF-22, as well as the incorporation of a number of other critical technologies. Supersonic cruise and internal weapon carriage/launch capabilities were also demonstrated. The maneuverability enhancements provided by pitch axis thrust vectoring were demonstrated at cruise altitudes (>10K ft). The flight test program commenced in the fall of 1990 and by December 1990 the YF-22 had demonstrated stabilized flight at 60° angle of attack. The brevity of the flight test program required an efficient and coordinated envelope expansion for aircraft loads, flutter and handling qualities. A very rapid response by the Vehicle Management System (VMS) team in Fort Worth allowed several Operational Flight Program (OFP) updates in support of the flight test program. The basic handling qualities of the aircraft were also validated. For all of the maneuvers flown during the initial flight test program the aircraft was very well behaved with predictable flying qualities. The aircraft generally received Level 1 CooperHarper ratings for both Up&Away (UA) and Power Approach (PA) handling qualities tasks. In April of 1991 the Lockheed team was awarded the EMD contract for the ATF program and a followon flight test program with the YF-22 began shortly thereafter. The primary objective of the follow-on test program was to expand the flutter and flying qualities envelopes of the aircraft. This phase of the flight test program ended prematurely when a pilot-induced-oscillation (PIO) occurred during a low approach, forcing a gear-up landing.

Abstract: A new paradigm of propulsion and maneuvering emerges, where unsteadiness in the form of large-scale, closely controlled eddies is the basic mechanism employed. A fish-like actively swimming robotic mechanism exhibits outstanding propulsive and maneuvering capabilities, thus offering exciting possibilities for enhancing the performance of marine craft. But can such technology really work ? This paper provides an outline of recent work, including building machines that move like fish, laboratory data that show the propulsive capabilities of such devices, and then accesses the possibility for transfering this technology to man-made vehicles.

Abstract: The components of a vehicle, in particular of a non-rail vehicle, which is powered by an electric motor traction propulsion system, can be protected against the damaging effects of accelerations which occur during operation by reducing the operating power supplied to a propulsion motor of the vehicle when the accelerations which occur on certain components of the vehicle reach values which pose an immediate danger to these components or reduce their useful life. For this purpose, a traction propulsion system which includes at least one electric motor, a power supply device which supplies the electric motor with operating power and an electronic switching system to control the operating power supplied to the electric motor is equipped with acceleration sensor means and with evaluation means, whereby the acceleration sensor means supply at least one variable which is a function of the acceleration of at least one component of the vehicle, and the evaluation means respond to this variable and control the electronic switching system so that at least in a portion of the range of values of the acceleration-dependent variable, the operating power supplied to the electric motor is reduced in accordance with a specified characteristic.

Abstract: A user powered vehicle which can be propelled upon an ice surface by applying force in repetitive power stroke movements in a rowing manner. The vehicle includes a main framework (10), a forward facing seat (64), a plurality of attached ice runners (14, 16) adapted to support the vehicle for efficient forward movement upon the operating surface, a movable propulsion wheel assembly (18) including at least one propulsion wheel (38) adapted to provide forward thrust when engaged with the operating surface, a rowing assembly (74) and footrest assembly (66) operable for converting rowing motions into propulsion force, and means (86) for transfer of weight from ice runners to propulsion wheel for increased traction during rowing power strokes and for re-transfer of weight to ice runners for improved efficiency between power strokes. The vehicle is adaptable for use on snow or land by replacing ice runners with skis or wheels respectively.

Abstract: A test and evaluation study examining the United Kingdom's 10-cm-diam xenon ion thruster system, known as the UK-10, has been performed at The Aerospace Corporation. The purpose of the study was to determine whether the UK-10 is a suitable replacement on U.S. Air Force satellites for presently employed propulsion devices. The most essential aspects of the UK-10 pertaining to its potential use were evaluated, but a conventional life test was beyond the scope. An examination of performance, lifetime, spacecraft integration issues, mission analysis, and computational modeling and a comparative technology assessment were performed. Emphasis was placed on the creation of a database that will be useful for the integration and operation of satellites using the UK-10 or other gridded electrostatic ion thrusters. Advanced electric-thruster characterization methods produced detailed measurements, in real time, of many fundamental properties of the thruster. The test program concluded that a UK-10 ion propulsion system will be acceptable for use on suitable spacecraft, subject to the considerations discussed, and of considerable potential benefit.

Abstract: In this paper a NafionTM polyelectrolyte ion-exchange membrane (IEM) was used as a propulsion fin for robotic swimming structures such as a boat or fish-like object swimming in water or aqueous medium. The Nafion membrane was chemically plated with platinum. The resulting membrane was cut in a strip to resemble a fish-like caudal fin for propulsion. A small function generator circuit was designed and built to produce approximately plus or minus 2.0 V amplitude square wave at varying frequency up to 50 Hz. The circuit board was mounted on a buoyant styrofoam shaped like a boat or a tadpole. The fin was attached to the rear of the boat. By setting the signal frequency to the desired value and thereby setting the frequency of bending oscillation of the membrane, a proportional forward propulsion speed could be obtained. The speed was then measured using a high speed camera. Several theoretical hydrodynamic models were then presented to characterize speed-frequency of the forward motion using available theories on biological fish motion. The results were compared to experimental data which showed close agreement. It turned out that the forward speed of the object was directly proportional to the product of frequency and amplitude of the fin oscillation as in biological fishes. This relation was further simplified by keeping the voltage constant and therefore amplitude of the oscillation. The proportionality constant could be measured for a known geometry of the fin-boat assembly and reactivity of the Nafion membrane used. The system as a whole presented an autonomous robotic swimming structure with frequency modulated propulsion to investigate application of polyelectrolyte hydrogel membranes and their effect on hydrodynamic behavior of an undulating swimming object. As in fishes the thrust force of the robot was generated by evolution of vortices on the sides of the undulating fin. For a constant forward speed, this thrust is equal to the drag force due to geometry and skin friction of the swimming robot. It was observed that regardless of the laminar or turbulent flow pattern around the robot the relation between speed and frequency holds. This research was a proof of concept for investigating fish propulsion known best for undulatory swimming motion, using polyelectrolyte ion-exchange-metal composite membrane.© (1996) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

Abstract: An improved propulsion unit control system for marine units having at least two propulsion systems. A single lever control controls the speed of both propulsion systems and if a difference in speed occurs, the speed of the faster propulsion unit is reduced. In addition, each propulsion unit includes a respective abnormal condition sensor which outputs a signal to a common control system. This common control system reduces the speed of both propulsion units when an abnormal condition is sensed in either unit.

Abstract: A large, civilian, multi-engine transport MD-11 airplane control system was recently modified to perform as an emergency backup controller using engine thrust only. The emergency backup system, referred to as the propulsion-controlled aircraft (PCA) system, would be used if a major primary flight control system fails. To allow for longitudinal and lateral-directional control, the PCA system requires at least two engines and is implemented through software modifications. A flight-test program was conducted to evaluate the PCA system high-altitude flying characteristics and to demonstrate its capacity to perform safe landings. The cruise flight conditions, several low approaches and one landing without any aerodynamic flight control surface movement, were demonstrated. This paper presents results that show satisfactory performance of the PCA system in the longitudinal axis. Test results indicate that the lateral-directional axis of the system performed well at high attitude but was sluggish and prone to thermal upsets during landing approaches. Flight-test experiences and test techniques are also discussed with emphasis on the lateral-directional axis because of the difficulties encountered in flight test.