Abstract: A hybrid propulsion system (100) for use in road vehicle operations, which propulsion system includes a power splitting mechanical transmission (108), suitably a three shaft epicyclic gearbox (117, 118, 119), for coupling to a tailshaft (115) of the vehicle; a first drive unit (105) arranged for regenerative operation and coupled to the power splitting mechanical transmission (108); a second drive unit (110) arranged for regenerative operation and coupled, independently of said first drive unit, to the power splitting mechanical transmission (108); a third drive unit (113) for coupling, in parallel to said power splitting mechanical transmission, to the tailshaft; and propulsion control means (122) for coordinating operation of the drive units in accordance with a plurality of predetermined modes corresponding to a driving cycle of the vehicle. Two forms of the invention are disclosed, being suited to non-transit and transit operations, respectively. Methods for the optimal control of the hybrid propulsion system of each form of the invention are also disclosed.

Abstract: In this paper we present a model for oscillating-foil propulsion in which springs are used to transmit forces from the actuators to the foil. The expressions for hydrodynamic force and moment on the foil come from classical, linear, unsteady aerodynamics, and these are coupled to linearized rigid-body mechanics to obtain the complete model for swimming. The model is presented as a low-order set of ordinary differential equations, which makes it suitable for the application of techniques from systems and control theory. The springs serve to reduce energy costs, and we derive explicit expressions for spring constants which are optimal in this sense. However, the use of springs can potentially lead to unstable dynamics. Therefore, we also derive a set of necessary and sufficient conditions for stability. A detailed example is presented in which energy costs for one actuator are reduced by 33%.

Abstract: A review of rocket-airbreathing combined-cycle propulsion systems for Earth-to-orbit applications is presented. Rocket-based combined-cycle (RBCC) engines take advantage of the synergistic interactions between the rocket and the airbreathing elements of the engine and the use of high-specie c impulse cycles to yield a mission-averaged specie c impulse that is higher than all-rocket technology can provide. An overview of the multimode operation is given, along with a review of both experimental and modeling work that has been done on this class of engines. Selected issues involved with these engines are discussed. These include engine/vehicle integration, e ow-path design for multimode operation, fuel selection, mixing enhancement and afterburning in rocket-ejector mode, thermal choking, and e ameholding. RBCC propulsion is becoming recognized as a promising technology for achieving a signie cant reduction in the cost of delivering payload to orbit.

Abstract: It has long been known that aircraft flying in close formation can achieve an overall efficiency much greater than is possible for each aircraft flying alone. When coupled with possible near-term advances in solar-electric power storage and aeronautical propulsion systems, it is theoretically possible to create a formation flight system that is capable of truly infinite endurance. This paper discusses a research effort currently underway that addresses the design and implementation of such a system. A brief overview of the system concept is given, and some of the particular problems inherent are more fully discussed. In particular, the difficulties involved in the autonomous control of such systems are presented and discussed.

Abstract: The use of ion propulsion for deep-space missions will become a reality in 1998 with the  ight of the ion-propelled, New Millennium Deep Space 1 (DS1) spacecraft. The anticipation of this event is stimulating the call for improved ion propulsion technologies, a trend that is expected to continue. This paper describes the evaluation of possible advanced solar electric propulsion technologies and their potential beneŽ ts to projected near-term and midterm solar system exploration missions. The advanced technologies include high-performance derivatives of the DS1 ion propulsion technology, scaled-down DS1 systems, and direct-drive Hall-effect thruster systems. The results of this study indicate that signiŽ cant near-term beneŽ ts can be obtained by the development of improved versions of the DS1 ion propulsion system (IPSs) components. In addition, if the current trend to smaller planetary spacecraft continues, then missions  ying these smaller future spacecraft will beneŽ t substantially from the development of scaled-down IPSs that incorporate advanced technologies in the ion engines and the propellant feed systems. The performance of the direct-drive Hall thruster systems is potentially superior to that of all other midterm options, but this technology has the highest development risk. SigniŽ cantly reduced trip times to small bodies and the outer planets may be possible if technology programs work to retire these risks.

Abstract: Minor faults in ship propulsion and their associated automation systems can cause dramatic reduction on ships' ability to propel and manoeuvre, and effective means are needed to prevent simple faults from developing into severe failure. The paper analyses the control system for a propulsion plant on a ferry. It is shown how fault detection, isolation and subsequent reconfiguration can cope with many faults that would otherwise have serious consequences. The paper emphasizes analysis of re-configuration possibilities as a necessary tool to obtain fault tolerance, showing how sensor fusion and control system reconfiguration can be systematically approached. Detector design is also treated and parameter adaptation within fault detectors is shown to be needed to locate non-additive propulsion machinery faults. Test trials with a ferry are used to validate the principles. © 1998 John Wiley & Sons, Ltd.

Abstract: The reduction of Earth-to-orbit launch costs in conjunction with an increase in launcher reliability and operational efe ciency are the key demands on future space transportation systems. Results of various system analyses indicate that these demands can be met with future single-stage-to-orbit (SSTO) vehicles using advanced technologies for both structure and propulsion systems. This paper will provide a classie cation and description of all turbopump feed liquid rocket engine cycles, followed by a combined vehicle/propulsion analysis of the design parameters for the propulsion systems and their associated thermodynamic cycles for the launch of future SSTO vehicles. Existing and projected rocket engine cycles capable of SSTO missions will be presented.

Abstract: An instrumented wheel system for three-dimensional kinetic analysis of upper extremity during wheelchair propulsion has been designed and validated. This system allows the direct measurements of three-dimensional dynamic forces and moments on the handrim during wheelchair propulsion in a laboratory setting as well as in the field. Static loading tests showed a high linearity and little drift (coefficient of determination, r2 > 0.999). Under dynamic loading, the instrumented wheel provided the well-matched measurement forces and moments with the predicted values from the inverse dynamic method using video-based kinematic data (correlation coefficient, p > 0.97). The three-dimensional handrim forces and moments during wheelchair propulsion by a non-disabled subject were demonstrated.

Abstract: The nuclear thermal rocket (NTR) is one of the leading propulsion options for future human missions to Mars because of its high specific impulse (Isp-850-1000 s) capability and its attractive engine thrust-to-weight ratio (approximately equal 3-10). To stay within the available mass and payload volume limits of a "Magnum" heavy lift vehicle, a high performance propulsion system is required for trans-Mars injection (TMI). An expendable TMI stage, powered by three 15 thousand pounds force (klbf) NTR engines is currently under consideration by NASA for its Design Reference Mission (DRM). However, because of the miniscule burnup of enriched uranium-235 during the Earth departure phase (approximately 10 grams out of 33 kilograms in each NTR core), disposal of the TMI stage and its engines after a single use is a costly and inefficient use of this high performance stage. By reconfiguring the engines for both propulsive thrust and modest power generation (referred to as "bimodal" operation), a robust, multiple burn, "power-rich" stage with propulsive Mars capture and reuse capability is possible, A family of modular "bimodal" NTR (BNTR) vehicles are described which utilize a common "core" stage powered by three 15 klbf BNTRs that produce 50 kWe of total electrical power for crew life support, an active refrigeration / reliquification system for long term, "zero-boiloff" liquid hydrogen (LH2) storage, and high data rate communications. An innovative, spine-like "saddle truss" design connects the core stage and payload element and is open underneath to allow supplemental "in-line" propellant tanks and contingency crew consumables to be easily jettisoned to improve vehicle performance. A "modified" DRM using BNTR transfer vehicles requires fewer transportation system elements, reduces IMLEO and mission risk, and simplifies space operations. By taking the next logical step--use of the BNTR for propulsive capture of all payload elements into Mars orbit--the power available in Mars orbit grows to 150 kWe compared to 30 kWe for the DRM. Propulsive capture also eliminates the complex, higher risk aerobraking and capture maneuver which is replaced by a simpler reentry using a standardized, lower mass "aerodescent" shell. The attractiveness of the "all BNTR" option is further increased by the substitution of the lightweight, inflatable "TransHab" module in place of the heavier, hard-shell hab module. Use of TransHab introduces the potential for propulsive recovery and reuse of the BNTR/ERV. It also allows the crew to travel to and from Mar on the same BNTR transfer vehicle thereby cutting the duration of the ERV mission in half--from approximately 4.7 to 2.5 years. Finally, for difficult Mars options, such as Phobos rendezvous and sample return missions, volume (not mass) constraints limit the performance of the "all LH2" BNTR stage. The use of "LOX-augmented" NTR (LANTR) engines, operating at a modest oxygen-to-hydrogen mixutre ratio (MR) of 0.5, helps to increase "bulk" propellant density and total thrust during the TMI burn. On all subsequent burns, the bimodal LANTR engines operate on LH2 only (MR=0) to maximize vehicle performance while staying within the lift capability of two Magnum launches.

Abstract: A pelagic free swimming aquatic vehicle includes a rigid forebody having a predetermined volume; a watertight chamber in the forebody; and a flexible afterbody having a lesser volume than the forebody and including a maneuvering and propulsion propulsion structure and a drive system for driving the structure with a traveling sinusoidal wave motion.

Abstract: Deep Space 1 is a technology demonstration mission scheduled to be launched in October 1998. One of those technologies is the NSTAR 30 cm diameter xenon ion thruster which will provide the primary propulsion. Three Flight-design thrusters were designed and built by Hughes Electron Dynamics Division, with assistance from NASA's Lewis Research Center. The first thruster was a Pathfinder to finalize the fabrication and assembly procedures for the other thrusters. Two flight-worthy thrusters were then fabricated and tested to Protoflight Qualification levels at NASA's Lewis Research Center and Jet Propulsion Laboratory. Each thruster was performance tested before and after Vibration Tests, integrated with different flight power processors and digital control interface units, and underwent Thermal Vacuum Tests with engine starts from -97 °C. Performance tests included neutralizer, discharge chamber, and ion optics characterizations as well as measurements of thruster efficiency over the full 0.5 to 2.3 kW power throttle range. The performance, at both component and thruster levels, was as expected and found to be quite repeatable with negligible dispersion between thrusters. After final functional tests, one thruster was installed on the DS 1 spacecraft while the other was set aside as a flight spare.

Abstract: Since its founding, NASA has been dedicated to the advancement of aeronautics and space science. The NASA Scientific and Technical Information (STI) Program Office plays a key part in helping NASA maintain this important role. The NASA STI Program Office is operated by Langley Research Center, the lead center for NASA's scientific and technical information. The NASA STI Program Office provides access to the NASA STI Database, the largest collection of aeronautical and space science STI in the world. The Program Office is also NASA's institutional mechanism for disseminating the results of its research and development activities. These results are published by NASA in the NASA STI Report Series, which includes the following report types: • TECHNICAL PUBLICATION. Reports of completed research or a major significant phase of research that present the results of NASA programs and include extensive data or theoretical analysis. Includes compilations of significant scientific and technical data and information deemed to be of continuing reference value. NASA counterpart or peer-reviewed formal professional papers, but having less stringent limitations on manuscript length and extent of graphic presentations. • TECHNICAL MEMORANDUM. Scientific and technical findings that are preliminary or of specialized interest, e.g., quick release reports, working papers, and bibliographies that contain minimal annotation. Does not contain extensive analysis. • CONTRACTOR REPORT. Scientific and technical findings by NASA-sponsored contractors and grantees. • CONFERENCE PUBLICATION. Collected papers from scientific and technical conferences, symposia, seminars, or other meetings sponsored or co-sponsored by NASA. • SPECIAL PUBLICATION. Scientific, technical, or historical information from NASA programs, projects, and missions, often concerned with subjects having substantial public interest. • TECHNICAL TRANSLATION. English-language translations of foreign scientific and technical material pertinent to NASA's mission. Specialized services that help round out the STI Program Office's diverse offerings include creating custom thesauri, building customized databases, organizing and publishing research results. .. even providing videos. and addresses the subject of Synergistic Airframe-Propulsion Interactions and Integrations (SnAPII). It is well known that favorable Propulsion Airframe Integration (PAI) is not only possible but Mach number dependent-with the largest (currently utilized) benefit occurring at hypersonic speeds. At the higher speeds the lower surface of the airframe actually serves as an external precompression surface for the inlet flow. At the lower supersonic Mach numbers and for the bulk of the commercial civil transport fleet, the benefits of SnAPII have not been as extensively explored. This is due primarily to the separateness …

Abstract: In order to increase the technology readiness level of the hypersonic scramjet technology, the Air Force Research Laboratory HyTech Program Office funded Aerojet to develop an innovative strut-based dual-mode scramjet engine design under the Storable Fuel Scramjet Flow Path Concepts (SFSFPC) program. Effective system thermal management over the entire flight envelope is critical to the success of the mission. This paper presents a thermal management system operational for a Mach 4 to 8 flight envelope. The design of the thermal system is based (at every Mach number) on delivering fuel to the combustor at the conditions required to support ignition at Mach 4 and high-efficiency combustion at all Mach numbers. The engine thermal management system includes the C/SiC engine shell and struts, a regeneratively fuel-cooled titanium plate between the engine and the vehicle, regeneratively fuel-cooled fins inside the C/SiC struts, a heat exchanger, and a solid propellant gas generator. Due to the radiation cooling, no endothermic reaction is required to cool the engine at full throttle operation at Mach 8. At the Mach 8 cruise condition at the 80 percent power level, the radiation and fuel cooling are augmented by an endothermic reaction requiring less than 40 percent of cracked n-decane fuel. This level of endothermic reaction has been demonstrated under realistic engine conditions. INTRODUCTION Aerojet has been actively pursuing the application of dual-mode ramjet to hypersonic missiles and global reach vehicles (refs. 1-4). To increase the technology readiness level of hypersonic scramjet technology, the Air Force Research Laboratory (AFRL) HyTech Program Office funded Aerojet to develop an innovative, strut-based, dual-mode ramjet engine design under the Storable Fuel Scramjet Flow Path Concepts (SFSFPC) program. A potential application for a Mach 4 to 8 liquid hydrocarbon fueled scramjet propulsion system is a fast-response long-range, hypersonic missile. Mission objectives include a range of at least 750 nautical miles in not more than 12 minutes with carriage from both fighter and bomber aircraft. Effective system thermal management over the entire flight envelope is critical to the success of the mission. This paper presents an efficient, flexible thermal management system for a Mach 4 to 8 flight envelope. The technical challenge is to identify a flexible thermal management system design that will be responsive to the combustor fuel conditioning requirements and to significant variations of structural heat load, fuel temperature, fuel density, and pressure over the entire flight envelope. During Mach 4 operation, the most challenging task is to ensure that the liquid fuel ignites and burns efficiently at ramjet '* Technical Principal, member AIAA, ** Engineering manager, *** Program manager # HyTech SFSFPC Program Manager on the Aerojet Contract This work has been supported by the Air Force Research Laboratory under Contract number F33615-96-C-2693. This paper is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

Abstract: The results of a study performed to evaluate the feasibility and merits of using an electrodynamic tether for propulsion and power generation for a spacecraft in the Jovian system are presented. The environment of the Jovian system has properties which are particularly favorable for utilization of an electrodynamic tether. Specifically, the planet has a strong magnetic field and the mass of the planet dictates high orbital velocities which, when combined with the planet's rapid rotation rate, can produce very large relative velocities between the magnetic field and the spacecraft. In a circular orbit close to the planet, tether propulsive forces are found to be as high as 50 N and power levels as high as 1 MW.

Abstract: As satellite designs shrink, providing maneuvering and control capability falls outside the realm of available propulsion technology. While cold gas has been used on the smallest satellites, hydrogen peroxide propellant is suggested as the next step in performance and cost before hydrazine. Minimal toxicity and a small scale enable bench top propellant preparation and development testing. Progress toward low-cost thrusters and self-pressurizing tank systems is described.

Abstract: This paper presents torque-controlled drives based on machine flux and torque estimation. The theoretical aspects of these methods are discussed and a comparative analysis is provided with emphasis on DSP implementation and experimental results. The problems in the application of these techniques to propulsion systems are also discussed and possible solutions are presented.

Abstract: To attain intensive combustion in M8 flight conditions at which a basic strutless engine failed to initiate intensive combustion, a strut for the inlet contraction ratio of 5 with fuel injectors was installed and the engine was tested in M8 flight conditions, with a total temperature of 2550 K, a total pressure of 10 MPa and an airflow Mach number of 6.7. In the small fuel flow rate regime, pressure increments in the diverging combustor were observed due to the boundary layer combustion. This pressure rise was enhanced with the auxaliry injection from the strut which enhanced the mixing of the main fuel. The intensive combustion within the constant area combustor was attained at higher fuel flow rates, which resulted in transitions to an engine unstart condition. A prediction method was adopted to estimate the limiting fuel flow rate for attainment of the intensive combustion within the constant area combustor and for the transition to the unstart. Introduction Supersonic combustion ramjet (scramjet) is expected to be the most effective propulsion system for Single Stage To Orbit (SSTO) transportation system and the hypersonic transportation system of next generation. Many studies on components of the scramjet engines such as inlet, combustor and nozzle have been carried out. On the other hand, intensive interactions between these components are expected in real engines [1]. Thus, it is necessary to conduct testing of complete engine model to investigate the interactions and the engine performances. However, only limited data on these complete engine models and their performances have been published [2]. Using a blow-down type wind tunnel facility (denoted as RamJet engine Test Facility; RJTF) at NALKRC, we have conducted tests of a scramjet engine at various flight conditions from M4 to M8 [3-8]. This engine had a sidewall compression type inlet section with a contraction ratio of 2.9. The ratio was limited to assure starting capability in the M4 flight conditions. However, this low contraction ratio resulted in low pressure levels at the combustor entrance in the M6 and M8 flight conditions, and no intensive combustion within the engine was observed at both flight conditions [3, 6]. To attain intensive combustion in M6 flight conditions, it was necessary to place a strut within the engine for further compression and low velocity region enlargement [4]. To attain intensive combustion also in M8 flight conditions, a strut for the inlet contraction ratio of 5 with fuel injectors was fabricated and installed. Present paper reports the results of the engine tests in the M8 flight conditions, with a total temperature of 2550 K, a total pressure of 10 MPa and an airflow Mach number of 6.7. Copyright © 1998 by American Institute of Aeronautics and Astronautics, Inc. All right reserved. * Researcher, Ramjet Combustion Section, Ramjet Propulsion Dvision. Member AIAA. t Head, Ramjet System Section. Member AIAA. $ Senior Researcher, Ramjet Aerodynamics Section. § Head, Ramjet Combustion Section. Member AIAA. 1 Head, Ramjet Guidance Section. Member AIAA # Head, Research Coordinate Office. Experimental apparatus Wind tunnel facility and test conditions The RJTF was designed to simulate flight conditions of M4, M6, and M8. The facility is equipped with two heating systems to obtain high enthalpy flow. A storage air heater (denoted as SAH) is used to obtain pure hot air for the M4, M6, and M8 flight condition tests. Two types of vitiation air heaters (VAH) fueled by H2 are used to obtain vitiated hot air for the M6 and M8 flight condition tests, respectively. The M6 flight conditions are achieved with the SAH alone or the VAH alone (denoted as M6S and M6V conditions, respectively), whereas the M8 conditions are achieved with a combination of the SAH and the VAH. Table 1 shows the operational conditions of the RJTF at the simulated Mach 8 flight conditions. The high enthalpy flow obtained at the heaters was accelerated through a rectangular nozzle. The exit cross section of the nozzle was 510 mm x 510 mm. The engine model was set in a low pressure chamber which was connected to a steam ejector system to reduce the pressure within the chamber for simulation of high altitude atmosphere. The engine model was located in such a way that it breathed in the boundary layer on the facility nozzle wall. The displacement thickness of the boundary layer was about 33 mm. Table 1 Flow condition in M8 conditions Stagnant Engine inlet

Abstract: The Hyper-X Program, NASA''s focused hypersonic technology program jointly run by NASA Langley and Dryden, is designed to move hypersonic, air-breathing vehicle technology from the laboratory environment to the flight environment, the last stage preceding prototype development. The Hyper-X research vehicle will provide the first ever opportunity to obtain data on an airframe integrated supersonic combustion ramjet propulsion system in flight, providing the first flight validation of wind tunnel, numerical and analytical methods used for design of these vehicles. A substantial portion of the integrated vehicle/engine flowpath development, engine systems verification and validation and flight test risk reduction efforts are experimentally based, including vehicle aero-propulsive force and moment database generation for flight control law development, and integrated vehicle/engine performance validation. The Mach 7 engine flowpath development tests have been completed, and effort is now shifting to engine controls, systems and performance verification and validation tests, as well as, additional flight test risk reduction tests. The engine wind tunnel tests required for these efforts range from tests of partial width engines in both small and large scramjet test facilities, to tests of the full flight engine on a vehicle simulator and tests of a complete flight vehicle in the Langley 8-Ft. High Temperature Tunnel. These tests will begin in the summer of 1998 and continue through 1999. The first flight test is planned for early 2000.

Abstract: This paper provides an overview of NASA's focused hypersonic technology program, called the Hyper-X Program The Hyper-X Program, a joint NASA Langley and Dryden program, is designed to move hypersonic, air breathing vehicle technology from the laboratory environment to the flight environment, the last stage preceding prototype development The Hyper-X research vehicle will provide the first ever opportunity to obtain data on an airframe integrated scramjet (supersonic combustion ramjet) propulsion system at true flight conditions and the first opportunity for flight validation of experimental wind tunnel, numerical and analytical methods used for design of these vehicles A substantial portion of the program is experimentally based, both for database development and performance validation The program is now concentrating on Mach 7 vehicle development, verification and validation and flight test risk reduction This paper concentrates on the aerodynamic and propulsion experimental programs Wind tunnel testing of the flight engine and complete airframe integrated scramjet configuration flow-path is expected in 1998 and 1999, respectively, and flight test is planned for 2000

Abstract: L OW thrust electric propulsion systems find apromising application field on small-to-medium sizedspacecraft, with mass ranging from a few hundred to about1000 kg and orbital altitudes of 700 km or higher, likethose envisaged for most of the proposed LEO telecom-munication constellations. To evaluate the performanceof FEEP for such missions, a study was carried out on atypical attitude and orbit maintenance case, using currentmodels for perturbation torques and forces. It was assumedthat three-axes attitude control and drag compensation befully performed by means of FEEP thrusters for a missionduration of 5 years, under a maximum thrust constraint of1 mN per thruster. A reference case (900 kg spacecraft in800 km LEO) was analyzed in detail, and a parametricperformance analysis was carried out for the spacecraftmass range 100 - 1000 kg and orbital altitude range 400 -800 km. As a result, the use of FEEP to replace momen-tum and reaction wheels and cold gas or hydrazine thrust-ers was found to be very attractive for satellites of mass inexcess of 400 kg in orbits higher than 400 km about. Out-side this mass and altitude ranges, the use of FEEP maystill lead to significant mass savings, but the AOCS con-figuration must be carefully studied on a case by case ba-sis. For all the cases studied, the mass of a full, 16-thrust-ers propulsion system, including thrusters, neutralizers,propellant, and redundant electronics, is less than 45 kilo-grams.IntroductionElectric thrusters have been traditionally associated withtwo main application domains, that is, in the near term,NSSK of telecommunication satellites and, in the far fu-ture, interplanetary primary propulsion. This picture israpidly changing with the recent trend towards the use ofsmaller spacecraft. Electric propulsion (EP) may play akey role in several classes of missions, spanning from sci-entific drag-free satellites, to earth observation and com-munication microspacecraft, to formation flying and con-stellations. In many cases, the mass savings made pos-sible by the use of EP instead of chemical or cold gaspropulsion is a mission enabling issue.Of course, this new scenario makes a new approach toEP systems necessary. The case of FEEP is a typical ex-ample of the shift in perspective that is occurring for someEP applications. During the 80’s, many efforts were dedi-cated to increasing thrust, in spite of the evident draw-backs associated with the high specific impulse of thisthruster. Those attempts had to face the high power-to-thrust ratio of FEEP (about 60 W/mN, vs. 35 W/mN ofion engines and 20 W/mN of SPT), and development waseventually abandoned. In the 90’s, the low thrust capa-bilities of FEEP have been rediscovered, thanks to thescientific community demand for modulable micronewtonthrusters, and the development efforts have gained newmomentum - this time, in the right direction. Today, theenvisaged application domain of FEEP is the 1 µN - 1 mNrange, were power-to-thrust is not a key factor, but lowsystem mass and low propellant consumption are. Thebasic technology of the 80’s has been fully exploited, andmuch research is underway to refine the thruster design atengineering level and to scale down the subsystems (neu-tralizer and power conditioning and control electronics).For many commercial small satellites applications, FEEPis an attractive choice for such tasks as attitude controland orbit maintenance. Scientific missions may benefit ofthe accurate thrust control capabilities and high bandwidthof FEEP for drag-free control and formation flying

Abstract: A hybrid electric vehicle is propelled with stored energy from a battery or flywheel, plus energy produced by burning fuel in an engine. The cost of the energy consumed, as well as the quantity of air pollutants released, can be reduced by optimizing (1) the ratings of the battery and engine, and (2) the power output that will be delivered by each source under the expected driving conditions. Life-cycle cost can be minimized by running the engine at a constant speed and power, and by avoiding deep discharges of the battery. An on-board "energy manager, which contains an embedded computer, can track the energy content of the battery and optimize the load division between the battery and engine when given a travel time by the driver over a specified route. It can command the engine to deliver more power whenever the alternative is a life-shortening deep discharge of the battery. The vehicle designer needs to perform a system-engineering analysis to optimize the ratings of the engine and battery. For this he needs to understand the power required to move the vehicle at desired speeds over hills on the expected routes, and through headwinds that vary from day to day. In this supporting analysis we modeled the hybrid vehicle by using Lagrange's calculus. We applied the model to a battery-powered electric bicycle which travelled a 200-mile route over a variety of hills. An unexpected result from this model is the energy consumed while travelling over a hilly route. With a specified travel time, no energy could be saved by climbing hills slowly and going down them rapidly. We did not consider the propulsion efficiency in this calculation. The next step will be to incorporate into the model variations in efficiency of the engine and the propulsion motor as speed is varied.

Abstract: This paper describes an evaluation of various propulsion options for robotic interstellar rendezvous missions to stars ranging from 4.5 Light Years (L.Y.) with a 10-year trip time, to 40 L.Y. with a 100-year trip time. Concepts considered included advanced electric propulsion, nuclear (fission, fusion, antimatter) propulsion, beamed energy (e.g., light sails, MagSails) propulsion, electromagnetic catapults, in-situ propellant production concepts (e.g., the interstellar ramjet), and hybrid systems (e.g., antimatter-catalyzed fission/ fusion). The various candidate propulsion options were evaluated using three screening criteria. First, is it possible for the candidate system to achieve the required AV, which can be as much as 0.6 c for a fast, 4.5-L.Y. mission. Second, does the propulsion systems require an extensive, mission-unique supporting infrastructure. Finally, the technology readiness levels of the various subsystem technologies of the propulsion concept are reviewed. This screening process resulted in the selection of beamed energy sail, matter-antimatter, and fusion ramjet concepts as the most promising candidates. Potential mission performance and near-term technology goals of these concepts were then evaluated.

Abstract: A novel lift and propulsion system using alternating-current electromagnets is proposed. The proposed system can simplify the configuration of a non-contacting conveyance system for a steel plate by magnetic force. Both finite element analysis and experiments were carried out. After investigating the basic characteristics of the proposed electromagnetic linear actuator, magnetic levitation experiments with four controlled electromagnets were performed and stable levitation was realized though about 65% of the total mass of the steel plate was supported by uncontrolled alternating-current electromagnets. The results indicate that the proposed system can produce not only the strong levitation force but also the propulsion force on a steel plate.