Advanced launch system (ALS): electrical actuation and power systems improve operability and cost picture

A brief status review of the Advanced Launch System (ALS) is given. The use of electrical actuation integrated with a single-vehicle electrical power system and controls for all actuation and avionic requirements is highlighted. This technology will help attain the ALS primary goals of reduced costs ($300/lb Earth to LEO (low Earth orbit)) and improved operability by significantly reducing the launch operations and servicing requirements relative to current vehicle designs and practices. Six candidate government-supported programs that are involved in the relevant technologies or that could have significant applications for electric actuators and an integrated power-distribution and control system are listed with a brief explanation and description of noteworthy technology.< <ETX xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">&gt;</ETX>

Most of marine electric propulsion ships are propelled by large electric motors. The speed control of the propulsion motor on the electric propulsion vessel is operated by adjusting the frequency. The power system has the inverter installed semiconductor element, such as IGBT or IGCT, for high-speed switching. However, the inverter for power conversion unit of the electric power propulsion on large vessels has fundamentally thermal problems of installed semiconductor, such problems could influence to considerable loss by switching and conducting the inverter in the power system transferring large current with high voltage. In order to calculate power loss and thermal loss of the inverter, this paper sets actual parameters of the electric propulsion system of the large ship in the simulation program considering condition of ship operation. The simulation proves that it is possible to analyze both power loss and thermal loss in the diodes and transistor during switching of the inverter in comparison with field oriented control and direct torque control. As a result, large power loss and thermal loss of switching and conducting the inverter by field oriented control considerably occurs and can also impact damages toward the power electronic devices by the thermal loss generated.

Active development of electric thrustors began 10 years ago. Today, several kinds of thrustors have achieved efficiencies above 90 % and lifetimes of several thousand hours. The following article derives the basic theory of electric thrust production at constant exhaust velocity, and at variable exhaust velocity programmed for optimum vehicle performance. Electrothermal or arcjet; electrostatic or ion; and electrodynamic or plasma thrustors are described. At the present time, ion thrustors of the electron bombardment and of the surface ionization types are the most promising systems. Electric power in space may be generated by solar cells or nuclear-electric generators. It is expected that the incore thermionic converter will eventually be the preferred system. A variety of missions with electric propulsion systems appear feasible and highly desirable, among them orbital station keeping, attitude control, planetary probes, solar and out-of-the-ecliptic probes, deep-space probes, and manned Mars and Venus exploration. For each mission, a careful systems-design study must be made, which will provide the optimum selection of thrustor type, thrust level, exhaust velocity, thrust program, power source, trajectory, and flight plan.

The paper investigates the utility of how nuclear thermal rocket (NTR) upper stage propulsion could substantially reduce the interplanetary transit time while greatly enhancing the economics of performing the manned Mars mission. au ECCV GTO IMLEO Isp kg kg/kw km km/sec kw LEO MEV mw NTR sec !N Nomenclature astronomical unit Earth capture crew vehicle geosynchronous transfer orbit initial mass in low Earth orbit specific impulse kilogram kilograms per kilowatt kilometer kilometers per second kilowatt low Earth orbit Mars excursion vehicle megawatt nuclear thermal rocket seconds delta V 1. Background Initial responses to the launch cost issue have been to suppose the development of heavy lift launch vehicles that are assumed to have fundamental launch cost savings. In general, most studies have isolated launch vehicle systems from the interplanetary vehicle assembled in low Earth orbit. This paper investigates a potential synergism between applying nuclear thermal rockets as both upper stages of the launch vehicle and as departure propulsion for a Mars vehicle departure stage. II. InterPlanetary Vehicle Description The mission to Mars is a propulsion technology dominated mission. The velocity change requirements (llV or delta V) are very large and very sensitive to trip time constraints. Due to this, there has been a heavy emphasis on high specific impulse (Isp) propulsion systems for performing the manned missions to Mars. The fundamental problem of high Isp systems is that they are generally thermally limited by material temperature limits. This issue is eliminated by electric propulsion systems that use electro-magnetic forces for accelerating a working fluid, but are thrust limited due to practical limitations in power supply mass. More exotic propulsion systems tend to use plasma working fluids that can only be contained by such methods as magnetic nozzles. The Space Exploration Initiative (SET) has a long range goal to perform a manned mission to Mars. Preliminary design goals of an interplanetary vehicle for this mission stress extremely high propulsion requirements to minimize transit times while allowing significant stay times on Mars. Related to this key trade between propulsion performance and trip time is the space transportation economics of 20 or more Earth launches required of the current launch vehicle architecture to place adequate vehicle mass in low Earth orbit to perform a single manned mission to Mars. The large initial mass in low Earth orbit (IMLEO) to perform the manned Mars mission is recognized as a dominant mission cost driver using the current launch infrastructure. IMLEO has been used as a preliminary figure of merit to evaluate various design approaches. * Senior Advanced Propulsion Engineer, Member AIAA Most analysis to date has focused on nuclear thermal rockets (Isp ... 1000 sec) and megawatt class electric propulsion systems (Isp ... 5000+ sec). The vehicle strategies have been to split the mission between a one way only cargo mission to Mars and a fast transfer manned vehicle. Mars rendezvous is essential for mission success, and techniques to slow at Earth return have used Earth capture crew vehicles (ECCV) with very high excess velocities (9+ km/sec) that are ballistically braked. Analyzed here is a very simple single vehicle. It uses an NTR driven, high thrust, Earth departure stage that is dropped once the vehicle has entered a fast Mars transfer ellipse. Once the vehicle has exited the deep Earth gravity well, it relies solely on a 5000 second Isp electric propulsion system for capture at Mars, departure from Mars, and final mission capture at Mars with zero excess velocity (a conservative assumption) The payload to Mars is broken into two elements. The ECCV is assumed to weigh 40,000 kg for a crew of four, and is the crew's interplanetary travel habitation module. The Mars excursion vehicle (MEV) is the element used to drop the crew from Mars orbit to the planet surface, be the crew's planetary habitation module, carry all science equipment, and carry the ascent module for returning to Mars orbit. The MEV has a baseline weight of 84,000 kg for a nominal 30 day planetary stay time and is scaled with the squareroot of normalized stay time. The MEV is not returned to Earth. Consumables are estimated at 10 kg per person per mission day. optimized for minimum l1V by adjusting the up leg perihelion and aphelion while maintaining the constraints of Earth rendezvous and travel time. Figure 2 (with the radius in millions of kilometers) shows a transfer ellipse for the up leg for the 320 day travel time, 80 day Mars stay time, 0 degree mean Mars true anomaly case. Figure 3 shows the down leg of that mission. The mission l1V requirement of 29.1 krn/sec is divided into 4.2 krn/sec for Earth departure (in heliocentric space), 6.0 krn/sec for Mars capture, 7.4 krn/sec for Mars departure, and 11.5 krn/sec for Earth capture. Transfer Ellipse Up. III. Interplanetary Mission Modeling The problem of analyzing the interplanetary mission can be driven by very exacting issues such as actual planetary configurations for particular mission opportunities, the requirements for travel time, and the Mars stay time. The issue that needs near term resolution, however, is the general efficacy of various propulsion and/or vehicle configuration options for the Mars mission in general. To answer these type of questions, it was necessary to generalize the Mars transfer problem. To do so, simplifying assumptions were that the Mars and Earth orbits were coplanar, and that the Earth orbit was circular. The nominal planetary configurations for Mars, Earth, and Venus are shown in Figure 1. The radius units are in astronomical units (Earth radius =1 au) and the angular units are the true anomaly of the orbital positions in degrees (Mars perihelion = 0 degrees). tojMJoC:: ~ •• ~: ••• . -

The advantages of using electric propulsion for propulsion are well‐known in the aerospace community. The high specific impulse and, therefore, lower propellant requirements make it a very attractive propulsion option for the Space Exploration Initiative (SEI). Recent studies have shown that nuclear electric propulsion (NEP) is not only attractive for the transport of cargo but that fast piloted missions to Mars are possible as well, with alphas on the order of 7.5 kg/kW. An advanced NEP system with a pecific power (alpha) of 2.5 kg/kW or less would significantly enhance the manned mission option of NEP by reducing the trip time even further. This paper describes an advanced system that combines the PEGASUS Drive with systems of the Rotating Multimegawatt Boiling Liquid Metal (RMBLR) power system that was developed as part of the DOE multimegawatt program and just recently declassified. In its original configuration, the PEGASUS Drive was a 10‐MWe propulsion system. The RMBLR was a 20‐MW electric power system. By combining the two, a second‐generation PEGASUS Drive can be developed with an alpha less than 2.5 kg/kW. This paper will address the technology advancements incorporated into the PEGASUS Drive, the analysis of a fast piloted mission and an unmanned cargo transport Mars mission, and the integration of laser power beaming to provide surface power.

At the initiative of S.P.Korolev, in 1959, Special Design Bureau No.1 (now RSC Energia) established the High-temperature Power Engineering and Electric Propulsion Center which was tasked with development of nuclear electric propulsion for heavy interplanetary vehicles. Selected as the source of electric power was a nuclear power unit based on a thermionic converter reactor, and selected as the engine was a stationary low-voltage magnetoplasmodynamic (MPD) high-power (0.5–1.0 MW) thruster which had thousands of hours of service life. The paper presents the results of extensive efforts in research, development, design, materials science experiments, and tests on the MPD-thruster, including the results of development and 500-hours life tests of an MPD-thruster with a 500-600 kW electric power input that used lithium propellant. The world’s first lithium 17 kW MPD-thruster was built and successfully tested in space. The paper points out that to this day nobody has surpassed the then achievements of RSC Energia neither in thruster output during long steady-state operation, nor in performance and service life. Key words: Martian expeditionary vehicle, nuclear electric rocket propulsion system, electric rocket thruster, magnetoplasmodynamic thruster, lithium, cathode, anode, barium, electric propulsion tests in space.

THE PURPOSE. Consider the issue of the presence of distortions in the form of sinusoidal voltage and current in autonomous electrical power systems with electrical propulsion systems, built on the principle of the unified electric power system. Compare the results of the study of voltage distortion for electrical power systems with electrical propulsion systems of various structures and give recommendations for their application.METHODS. For research, unified electric power systems with DC electrical propulsion systems on the Yeysk ferry and alternating current on the asymmetric icebreaker Baltika are considered. The possibilities of programming with frequency control of modern drives of rudder propellers with AC motors are analyzed.RESULTS. Oscillograms of voltages and currents of generators were obtained using thyristor converters and inverters to control electrical propulsion systems in various modes. Significant ripples and distortions of the sinusoidal voltage and generator currents were noted when thyristor converters are used to power DC propulsion electric motors. Also, small deviations from the sinusoidal form of the ship's network voltage were recorded when inverters were used to control AC propulsion motors as part of rudder propellers.CONCLUSION. The issue of ensuring the quality of electricity produced in autonomous electrical power systems of sea vessels is of great relevance and importance. To ensure the best results, it is advisable to use azimuthing podded drive with AC electric motors, the rotation speed of which is implemented by inverter frequency converters with a DC link.

NASA has completed a preliminary mission and systems study of nuclear electric propulsion (NEP) systems for “split‐sprint” human exploration and related robotic cargo missions to Mars. This paper describes the study, the mission architecture selected, the NEP system and technology development needs, proposed development schedules, and estimated development costs. Since current administration policy makers have delayed funding for key technology development activities that could make Mars exploration missions a reality in the near future, NASA will have time to evaluate various alternate mission options, and it appears prudent to ensure that Mars mission plans focus on astronaut and mission safety, while reducing costs to acceptable levels. The split‐sprint nuclear electric propulsion system offers trip times comparable to nuclear thermal propulsion (NTP) systems, while providing mission abort opportunities that are not possible with “reference” mission architectures. Thus, NEP systems offer short transit times for the astronauts, reducing the exposure of the crew to intergalactic cosmic radiation. The high specific impulse of the NEP system, which leads to very low propellant requirements, results in significantly lower “initial mass in low earth orbit” (IMLEO). Launch vehicle packaging studies show that the NEP system can be launched, assembled, and deployed, with about one less 240‐metric‐ton heavy lift launch vehicle (HLLV) per mission opportunity ‐ a very large cost savings! Technology development cost of the nuclear reactor for an NEP system would be shared with the proposed nuclear surface power systems, since nuclear systems will be required to provide substantial electrical power on the surface of Mars. The NEP development project plan proposed includes evolutionary technology development for nuclear electric propulsion systems that expands upon SP‐100 (Space Power ‐ 100 kw) technology that has been developed for lunar and Mars surface nuclear power, and small NEP systems for interplanetary probes. System upgrades are expected to evolve that will result in even shorter trip times, improved payload capabilities, and enhanced safety and reliability. Non‐nuclear technology development for the NEP system is estimated to cost about $721 M (1993 $). Nuclear technology development costs are not included in the costs in this report, since these costs will be incurred in the nuclear surface power development program. NEP Phase A/B studies are estimated to cost about $154 M. Flight system hardware development (Phase C/D) is estimated to cost about $2.8 B, and fabrication of flight hardware is estimated to be about $7.8 B for four mission opportunities in 2009, 2011, 2014, and 2016.

ABSTRACTThe increasing environmental requirements in the air transport sector pose great challenges to the aviation industry and are key drivers for innovation. Besides various approaches for increasing the efficiency of conventional gas turbine engines, electric propulsion systems have moved into the focus of aviation research. The first electric concepts are already in service in general aviation. This study analyses the potentials of electric and turbo hybrid propulsion systems for commercial aviation. Its purpose is to compare various architectures of electrical powertrains with a conventional turboprop on a regional aircraft, similar to the ATR 72, on engine and flight mission levels. The considered architectures include a turbo-electric (power controlled and direct driven), hybrid-electric (serial and parallel) and a pure electric concept. Their system weights are determined using today’s technology assumptions. With the help of performance models and flight mission calculations the impact on fuel consumption, CO ${}_{2}$ emissions and aircraft performance is evaluated.

The Advanced Launch System (ALS) is envisioned as a family of new generation launch systems, applying both existing and new technology to achieve the desired operability and cost goals, and ultimately providing capability for delivering a range of cargo sizes from approximately 1,000 to 220,000 lbs to low earth orbit. During Phase I of the ALS program, a set of trade studies were initiated to identify the best path for achieving these challenging design goals. Martin Marietta, which was one of seven contractors conducting these trades for Phase I, is one of three prime contractors for Phase I1 who will continue to refine these trades in order to develop a detailed vehicle and operations concept. This paper reviews the requirements of the ALS program, the trends and results of key Martin Marietta Phase I trades, the Phase I concepts that resulted, and finally the trade challenges for the on-going Phase 11 effort. u

Electric propulsion systems are nowadays one of the most important ways to overcome the modern energy paradigm. Among other systems, electric boats can take advantage of using electric power increasing the efficiency of the propulsion chain and reducing the pollutant gas emissions. This paper presents and discusses global aspects of a boat propulsion system investigated using a low scale system. The proposed chain is modeled and analyzed using two different battery types (lead acid and Ni-MH) and a power converter - DC motor set controlled according with the specifications of such systems. Using classical approaches (as for instance, for the DC motor model), the proposed solution is analyzed implementing a speed control system with a sliding mode controller.

This paper examines the role of the TOPAZ II and the 40 kWe thermionic reactors in powering electric propulsion systems for orbital transfer and on‐orbit repositioning applications. Specifically, the feasibility of substituting a TOPAZ II power system and the Russian Stationary Plasma Thruster (SPT) for the conventional power and propulsion systems that might be used on typical Air Force navigation, surveillance, and communications satellites were examined. Additionally, two advanced high‐power missions were analyzed. Four major conclusions were drawn from this top‐level investigation: (1) Stepdown to smaller launch vehicles is possible for both TOPAZ II and 40 kWe systems even though the TOPAZ II at 6 kWe and 1100 kg, is not an extremely effective system for orbital transfer applications; (2) On‐orbit repositioning is an area that profits by the utilization of electric propulsion systems; (3) Advanced high‐power missions are arguably the greatest benefactor of nuclear electric propulsion; and (4) Placing minimum operational altitude requirements on reactor‐propelled systems will significantly diminish their payload capability to the destination orbit.

Application of electrical propulsion for an active debris removal system: a system engineering approach

A logistics model for large space power systems

Electrical propulsion and power systems in naval shipboard applications offer efficient propulsion arrangements for various types of vessels. Adoption has been propelled by the development of semiconductor switching devices since the 1980s. Presently these systems use a rectifier stage to convert the generator side AC system into a regulated DC bus that is distributed to drive the traction motors. The rectification stage typically features either passive diode bridge or thyristorbased line-commutated circuits. These rectifiers interface to the AC distribution using line-frequency transformers which provide galvanic isolation; but are bulky and deteriorate the power density of the overall system. This paper introduces a family of high-frequency isolated and direct AC-DC converters that tap into the emerging technologies of high-frequency high-power transformers and four-quadrant devices. Advantages include increased power density due to elimination of bulky line-frequency transformers, direct AC-DC conversion without multiple power conversion stages, and minimized filtering elements sized for switching frequency harmonics only. The paper also presents broader design and sizing aspects and simulation results of one of the selected candidates.