Spacecraft Propulsion

This technical report presents the Impulse Drive, a propulsion system leveraging Spiral Unified Field Theory (SUFT) principles in five-dimensional (5D) spacetime. By generating micro-portal transitions through the compactified τ -dimension, the Impulse Drive enables high-efficiency or potentially faster-than-light (FTL) travel, requiring ∼ 10 15 Joules for a 1-light-year journey-10 6 times more efficient than chemical rockets. The report details the theoretical foundation, propulsion mechanism, energy source (capacitor bank for prototype, fusion reactor for full-scale), prototype development, experimental validation, and milestones, with a focus on safety and feasibility. The framework is validated and ready for prototype testing, targeting collaboration with SpaceX.

This report outlines current Small Business Innovation Research (SBIR) Post-Phase II opportunity contract award results for the SBIR technology program from 2007 to 2011 for NASA's Aeronautics Research Mission Directorate (ARMD), Human Exploration and Operations Mission Directorate (HEOMD), Science Mission Directorate (SMD), and Space Technology Mission Directorate (STMD). The report provides guidelines for incorporating SBIR technology into NASA programs and projects and provides a quantitative overview of the post-Phase II award patterns that correspond with each mission directorate at NASA Glenn Research Center (GRC). In recent years, one of NASA's goals has been to not only transfer SBIR technologies to commercial industries, but to ensure that NASA mission directorates incorporate SBIR technologies into their program and project activities. Before incorporating technologies into MD programs, it is important to understand each mission directorate structure because each directorate has different objectives and needs. The directorate program structures follow.

Current efforts by the student members of the sole student rocket group at Arizona State University, Daedalus Astronautics, have been focused on development of both a hybrid rocket motor and an aerospike nozzle. Hybrid rocket propulsion has combined properties of both solid and liquid rockets giving it advantages over the individual systems. This is a concept which has gained a more substantial footing with the success of Scaled Composites and SpaceShipOne. Daedalus Astronautics has designed and developed a simple hybrid rocket motor for the use in their high-powered sounding rocket program. The designed motor utilizes nitrous oxide and hydroxyl-terminated polybutadiene (HTPB) for oxidizer and fuel respectively. The key components of the hybrid design and analysis include the oxidizer solenoid valve, injector, pre-combustion chamber, fuel grain, post combustion chamber and nozzle. Initial testing of the hybrid rocket motor have been conducted utilizing a standard "bell" nozzle in order to allow for direct comparison to a more advanced nozzle design. A toroidal aerospike nozzle with an isentropic contour was designed and implemented for use with the hybrid due to its advantages over the standard and widely used "bell" shaped nozzle. An aerospike nozzle has the ability to perfectly expand the exhaust gas at all altitudes increasing the performance of the system. This design includes the use of graphite inserts at the nozzle throat to deter erosion and expansion which has been problems in previous designs. By integrating this nozzle with the hybrid rocket motor a high performance launch system can be achieved. 48th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition

This paper describes a novel magnetic propulsion system designed to generate incremental linear motion using internal magnetic forces. The system employs a central magnet with reversible polarity, a fixed rod with sliding magnets, and switchable stopper magnets to achieve net displacement per cycle. We analyze its mechanics, explore its limitations in a vacuum (space), and evaluate its potential for bolder aspirations (i.e., when tethered to a space station for micro-maneuvers, functioning as a motion storage unit, and eliminating the need to achieve escape velocity to exit a gravitational field). The system is compared to existing propellant-free technologies, such as control moment gyroscopes, solar sails, and the speculative Mach Effect Gravity Assist (MEGA) drive, highlighting its unique self-contained linear thrust mechanism. While promising for terrestrial applications, its scalability for space station movement is constrained by mass, power, and efficiency compared to conventional thrusters.

This pre-university investigative laboratory report examines the effect of propellant gas molar mass on the thrust performance of small-scale ion thrusters. A simple grid-type ion thruster was constructed and tested with gases of varying molar masses, with thrust output measured using a sensitive balance under controlled conditions. Results show that lighter gases generally yield higher exhaust velocities, while heavier gases provide greater impulse per ion at reduced velocities, aligning with predictions from the ideal rocket equation and electric propulsion theory. The study integrates principles from plasma physics, gas dynamics, and aerospace propulsion, offering practical insight into propellant selection for electric propulsion applications.

This study investigates the energy performance of paraffin-based hybrid fuels enhanced with potassium nitrate (KNO 3), ammonium nitrate (NH 4 NO 3), aluminum (Al), titanium (Ti), and stearic acid additives. The fuels were evaluated using thermochemical calculations via ProPEP3 Version 1.0.3.0 software, revealing significant improvements in specific impulse (I sp) and combustion temperature. While formulations with nitrates and aluminum exhibited noticeable increases in combustion efficiency and thermal output, titanium-containing mixtures provided moderate improvements. Stearic acid improved fuel processability and provided a stable burning profile without significant energy penalties. These findings demonstrate that suitable combinations of additives can substantially improve the energy output of paraffin-based hybrid fuels, making them more viable for aerospace propulsion applications.

Any missions to catch 1I/'Oumuamua have the daunting challenge of generating higher hyperbolic excess speeds than the interstellar object's own, i.e. > 26.3 km s -1 with respect to the Sun. To accomplish this task using chemical propulsion, previous papers have investigated a Solar Oberth manoeuvre and alternatively a Jupiter Oberth, these options requiring a thrust from the chemical rocket at perihelion/perijove respectively, points at which the available velocity increment (∆V ) results in maximum augmentation of the kinetic energy of the spacecraft. In this paper we unravel the specifics of a mission requiring a JOM or optionally a passive Jupiter encounter, i.e. with no thrust, the latter having so far not been addressed by Project Lyra. Whereas the previous papers were feasibility studies, this paper delves into what would be achievable with a Jupiter encounter (without any preceding gravitational assists from the inner planets), using currently available off-theshelf solid and liquid rocket stages and assuming the NASA Space Launch System Block 2 can be deployed. Optimal Launch dates are found to lie in the years 2030, 2031 and 2032 with resulting mission durations of around 30-40 years, depending on the particular number and combination of stages exploited. Payload masses to 'Oumuamua of 860kg can readily be accomplished with a duration of 43 or so years, assuming a Centaur D and STAR 48B combination, the latter booster being ignited at perijove. On the other hand if two Centaur Ds are utilised instead of just one, retaining the STAR 48BV for the JOM, 35 years are evinced for the same mass. Furthermore, for lower payloads of ∼ 100kg, the flight duration can be cut accordingly to 31 years, with the additional benefit of requiring no burn at Jupiter.

Planet 9 is an hypothetical object in the outer Solar system, which is as yet undiscovered. It has been speculated that it may be a terrestrial planet or gas/ice giant, or perhaps even a primordial black hole (or dark matter condensate). State-of-the-art models indicate that the semimajor axis of Planet 9 is ∼ 400 AU. If the location of Planet 9 were to be confirmed and pinpointed in the future, this object constitutes an interesting target for a future space mission to characterize it further. In this paper, we describe various mission architectures for reaching Planet 9 based on a combination of chemical propulsion and flyby maneuvers, as well as more advanced options (with a ∼ 100 kg spacecraft payload) such as nuclear thermal propulsion (NTP) and laser sails. The ensuing mission duration for solid chemical propellant ranges from 45 years to 75 years, depending on the distance from the Sun for the Solar Oberth maneuver. NTP can achieve flight times of about 40 years with only a Jupiter Oberth maneuver whereas, in contrast, laser sails might engender timescales as little as 7 years. We conclude that Planet 9 is close to the transition point where chemical propulsion approaches its performance limits, and alternative advanced propulsion systems (e.g., NTP and laser sails) apparently become more attractive.

This paper discusses the current research status of the MET (Microwave Electro-Thermal) thruster. In the MET thruster, an electrode less, vortex stabilized, plasma is produced in a microwave resonator cavity for the purpose of heating gaseous fuel to produce a high temperature rocket exhaust for space propulsion. The higher specific impulse (momentum transfer per unit weight) of these heated gases offers advantages over traditional chemical rockets in terms of reduced fuel mass. In MET devices, dense plasmas have been produced in various possible fuel gases, nitrogen, hydrogen, and ammonia, using 600 to 2200 Watts of microwave power at a frequency of 2.45 GHz. Ammonia has been found to give a specific impulse of 550 sec. It has been found that the plasma is a 98% absorber of microwave * Assoc. Professor of Aerospace Eng.

This paper discusses the current research status of the MET (Microwave Electro-Thermal) thruster. In the MET thruster, an electrode less, vortex stabilized, plasma is produced in a microwave resonator cavity for the purpose of heating gaseous fuel to produce a high temperature rocket exhaust for space propulsion. The higher specific impulse (momentum transfer per unit weight) of these heated gases offers advantages over traditional chemical rockets in terms of reduced fuel mass. In MET devices, dense plasmas have been produced in various possible fuel gases, nitrogen, hydrogen, and ammonia, using 600 to 2200 Watts of microwave power at a frequency of 2.45 GHz. Ammonia has been found to give a specific impulse of 550 sec. It has been found that the plasma is a 98% absorber of microwave * Assoc. Professor of Aerospace Eng.

A rocket is a spacecraft, guided missile, or flying vehicle that boosted by a chemical reaction resulting from the combustion of propellant in the rocket motor. One of the essential parameters in the development of rocket motors is design optimization to improve the propulsion performance of the rocket. Increasing the propulsion performance of the rocket will increase the flight performance of the rocket, in terms of its maximum range or the altitude of the rocket trajectory. This study examined the determination of the design parameter values of a rocket motor by looking at it as an optimization problem with constraints. The problem studied was limited to the case of the second-stage rocket motor. A genetic algorithm was used to solve the resulting optimization problem of propellant grain configuration cases and a characteristic method for designing the bell nozzle. The results obtained indicated an increase in total impulse by 10% compared to the results before optimization.

This chapter briefly introduces hybrid rocket propulsion for general audience. Advantageous of hybrid rockets over solids and liquids are presented. This chapter also explains how to design a test setup for hybrid motor firings. Hybrid propulsion provides sustainable, safe and low cost systems for space missions. Therefore, this chapter proposes hybrid propulsion system for Mars Ascent Vehicles. Paraffin wax is the fuel of the rocket. Propulsion system uses CO 2 =N 2 O mixture as the oxidizer. The goal is to understand the ignition capability of the CO 2 as an in-situ oxidizer on Mars. CO 2 is known as major combustion product in the nature. However, it can only burn with metallic powders. Thus, metallic additives are added in the fuel grain. Results show that CO 2 increase slows down the chemical kinetics thus reduces the adiabatic flame temperature. Maximum flammability limit is achieved at 75% CO 2 by mass in the oxidizer mixture. Flame temperature is 1700 K at 75% CO 2 . Ignition quenches below the 1700 K.

The high cost of launching payloads into orbit e roughly $20,000/kg e continues to deter large-scale exploration and exploitation of space. Ground-based launch systems may radically reduce costs to $200/kg, drastically altering the economics of spaceflight. Low costs will encourage the creation of new markets, including solar-based power satellites and disposal of nuclear waste. The US government should establish a goal of $200/kg by 2020 and provide the resources needed to develop such systems.

The high cost of launching payloads into orbit e roughly $20,000/kg e continues to deter large-scale exploration and exploitation of space. Ground-based launch systems may radically reduce costs to $200/kg, drastically altering the economics of spaceflight. Low costs will encourage the creation of new markets, including solar-based power satellites and disposal of nuclear waste. The US government should establish a goal of $200/kg by 2020 and provide the resources needed to develop such systems.

The Stennis Space Center has been at the forefront of development and application of exhaust plume spectroscopy to rocket engine health monitoring since 1989. Various spectroscopic techniques, such as emission, absorption, FTIR, LIF, and CARS, have been considered for application at the engine test stands. By far the most successful technology h a been exhaust plume emission spectroscopy. In particular, its application to the Space Shuttle Main Engine (SShfE) ground test health monitoring has been invaluable in various engine testing and development activities at SSC since 1989. On several occasions, plume diagnostic methods have successfully detected a problem with one or more components of an engine long before any other sensor indicated a problem. More often, they provide corroboration for a failure mode, if any occurred during an engine test. This paper gives a brief overview of our instrumentation and computational systems for rocket engine plume diagnostics at SSC. Some examples of successful application of exhaust plume spectroscopy (emission as well as absorption) to the SSME testing are presented. Our on-going plume diagnostics technology development projects and future requirements are discussed. I.

The Stennis Space Center has been at the forefront of development and application of exhaust plume spectroscopy to rocket engine health monitoring since 1989. Various spectroscopic techniques, such as emission, absorption, FTIR, LIF, and CARS, have been considered for application at the engine test stands. By far the most successful technology h a been exhaust plume emission spectroscopy. In particular, its application to the Space Shuttle Main Engine (SShfE) ground test health monitoring has been invaluable in various engine testing and development activities at SSC since 1989. On several occasions, plume diagnostic methods have successfully detected a problem with one or more components of an engine long before any other sensor indicated a problem. More often, they provide corroboration for a failure mode, if any occurred during an engine test. This paper gives a brief overview of our instrumentation and computational systems for rocket engine plume diagnostics at SSC. Some examples of successful application of exhaust plume spectroscopy (emission as well as absorption) to the SSME testing are presented. Our on-going plume diagnostics technology development projects and future requirements are discussed. I.

-kilowatt electric propulsion systems are well developed and have been used on commercial and military satellites in Earth orbit for several years. Ion and Hall thrusters have also propelled robotic spacecraft to encounters with asteroids, the Moon, and minor planetary bodies within the solar system. High power electric propulsion systems are currently being considered to support piloted missions to near earth asteroids, as cargo transport for sustained lunar or Mars exploration, and for very high-power piloted missions to Mars and the outer planets. Using NASA Mars Design Architecture 5.0 as a reference, a preliminary parametric analysis was performed to determine the suitability of a nuclear powered, MW-class electric propulsion system for Mars cargo transport. For this initial analysis, high power 100-kW Hall thrusters and 250-kW VASIMR engines were separately evaluated to determine optimum vehicle architecture and estimated performance. The DRA 5.0 cargo mission closed for both propulsion options, delivering a 100 t payload to Mars orbit and reducing the number of heavy lift launch vehicles from five in the baseline DRA 5.0 architecture to two using electric propulsion. Under an imposed single engine-out mission success criteria, the VASIMR system took longer to reach Mars than did the Hall system, arising from the need to operate the VASIMR thrusters in pairs during the spiral out from low Earth orbit.

This is chiefly a review paper with additional considerations and suggestions about aspects of matter-antimatter propulsion for space exploration. The paper is divided into six main sections. In the first section the fundamentals of the low-energy antiproton-nucleon annihilation physics are presented. In the second section geocentric, interplanetary and interstellar flights possible by means of antimatter engines are reviewed. In the third section the big problems related to the antimatter handling (production, storage, control) are discussed. In the fourth section proposed matter-antimatter thruster design concepts for Earth-space, interplanetary and out-of-Solar-System missions are examined and some results from computer simulations are discussed. In the fifth section a comparative cost between antimatter-based engines and conventional engines is made. Finally, aspects of a world-wide cooperation to try to make antimatter propulsion a reality are pointed out.

The MSFC/UAH team has been developing of a novel power management and distribution system called a Linear Transformer Driver (LTD). LTD's hold the promise of dramatically reducing the required mass to drive a z-pinch by replacing the capacitor banks which constitute half the mass of the entire system. The MSFC?UAH tea, is developing this technology in hope of integrating it with the Pulsed Fission Fusion (PuFF) propulsion concept.

The MSFC/UAH team has been developing of a novel power management and distribution system called a Linear Transformer Driver (LTD). LTD’s hold the promise of dramatically reducing the required mass to drive a z-pinch by replacing the capacitor banks which constitute half the mass of the entire system. The MSFC?UAH tea, is developing this technology in hope of integrating it with the Pulsed Fission Fusion (PuFF) propulsion concept. High-Voltage pulsed power systems used for Z-Pinch experimentation have in the past largely been based on Marx Generators. Marx generators deliver the voltage and current required for the Z-Pinch, but suffer from two significant drawbacks when applied to a flight system: they are very massive, consisting of high-voltage capacitor banks insulated in oil-filled tanks and they do not lend themselves to rapid pulsing. The overall goals of Phase 1 are to demonstrate the feasibility of the LTD concept for a Z-Pinch propulsion system and to learn techniques for d...

Deep-space exploration faces significant challenges in propulsion, spacecraft manoeuvrability, and communication systems, impacting mission duration, operational flexibility, and information exchange with Earth. To address these challenges and enhance deep-space mission capabilities, this paper investigates the viability of manoeuvrable solarsailed small satellites for propulsion and potential use as communication relays. Solar sails, relying on sunlight pressure for thrust, can theoretically operate indefinitely, limited only by material durability and electronic systems. However, achieving high performance depends on optimizing sail design for manoeuvrability within small satellite constraints. This research examines existing solar sail technologies and their application in small satellites, focusing on a novel dual-sail approach that combines a large monolithic sail for primary propulsion with a heliogyro system for manoeuvring and attitude control. The study presents a detailed analysis of solar radiation pressure models, including quantum and electromagnetic descriptions, and explores the optical force model for realistic sail behaviour. A conceptual design for manoeuvrable solar-sailed satellites is presented, investigating specific details such as sail configuration, material selection, and deployment mechanisms. The design incorporates a 100 m² monolithic square sail made of aluminized Kapton for primary propulsion, complemented by a heliogyro system with long, thin blades for enhanced manoeuvrability. The paper also explores the integration of flexible solar panels on the heliogyro blades for additional power generation. Finite Element Analysis (FEA) is employed to assess the structural integrity and thermal performance of the sail design under various space conditions. The analysis provides insights into sail deformation and stress distribution, crucial for ensuring the design's feasibility and durability in the harsh space environment. This research, conducted by members of the Space Generation Advisory Council (SGAC) within the Small Satellites Project Group (SSPG), aims to balance the benefits of spacecraft manoeuvrability through solar sails with the constraints of small satellites, potentially advancing future deep-space exploration capabilities.

This paper presents results of a thermal analysis performed on a gaseous oxygen/Polymethylmethacrilate hybrid rocket motor for a CubeSat application. This work was focused on deriving temperatures inside the combustion chamber and verifying the selected materials can withstand the high temperatures and stresses. Results show that the fuel starts to break-up before the end of the burn. The addition of a graphite post-combustion chamber was also evaluated but it is not recommended for long burn times due to high temperature reached in the steel. This work is an important intermediate step in the design process to determine system performance.

Virtual Cathode lifetimes and floating potential measurements in a Polywell T M fusion device MATTHEW CARR, JOE KHACHAN, School of Physics, University of Sydney -The Polywell T M is a spherically convergent ion focus concept first developed by R Bussard in the 1980s as a possible device for controlled thermonuclear fusion. The device aims to magnetically confine electrons with a quasi-spherical-cusp magnetic field, forming a deep potential well in the centre of the device, which can attract and maintain a high density of local energetic ions passing through the potential well. Careful design of the magnetic field and coil formers might help to substantially reduce the former/grid collision losses that plague other IEC devices. Floating potential measurements in the core of a Polywell have shown that a virtual cathode is established, with floating potentials of up to -250V obtained for milliseconds. The lifetime of the virtual cathode was determined only by the shape and duration of the magnetic coil current. This implies that currents of increasing duration will increase the lifetime of the virtual cathode. Further measurements reveal that virtual cathode formation could only be established within a narrow magnitude range of coil currents. We find that the floating potential increases with decreasing gas pressure.

In this study, we are aiming to develop an electric propulsion system suitable for miniature satellites especially CubeSats. Electron Cyclotron Resonance Ion Thruster (ECR Ion Thruster), combining the advantages of high density plasma jet and high specific impulse, could be the promising candidate for satellite's propulsion subsystem. In the phase of preliminary design, SIMION software was used to define the configuration of ion thruster depending on the simulation results of ion behaviors and ion optics. Currently, the prototype system is capable to produce nearly 0.9 milli-newtons thrust and around 1025~3076s specific impulse with the calculation of plasma diagnostics.

The International Space Station (ISS) will require periodic reboost due to atmospheric aerodynamic drag. This is nominally achieved through the use of thruster firings by the attached Progress M spacecraft. Many Progress flights to the ISS are required annually. Electrodynamic tethers provide an attractive alternative in that they can provide periodic reboost or continuous drag cancellation using no consumables, propellant, nor conventional propulsion elements. The system could also serve as an emergency backup reboost system used only in the event resupply and reboost are delayed for some reason. The system also has direct application to spacecraft and upper stage propulsion. Electrodynamic tethers have been demonstrated in space previously with the Plasma Motor Generator (PMG) experiment and the Tethered Satellite System (TSS-IR). The advanced electrodynamic tether proposed for this application has significant advantages over previous systems in that higher thrust is achievable with significantly shorter tethers and without the need for an active current collection device, hence making the system simpler and much less expensive. (Author)

The first interstellar object to be discovered, 1I/'Oumuamua, exhibited various unusual properties as it was tracked on its passage through the inner solar system in 2017/2018. In terms of the potential scientific return, a spacecraft mission to intercept and study it in situ would be invaluable. As an extension to previous Project Lyra studies, this paper elaborates an alternative mission to 1I/'Oumuamua, this time also requiring a Jupiter Oberth Manoeuvre (JOM) to accelerate the spacecraft towards its destination. The difference is in the combination of planetary flybys exploited to get to Jupiter, which includes a Mars encounter before proceeding to Jupiter. The trajectory identified is inferior to previous finds in terms of higher ∆V requirement (15.6 km s-1), longer flight duration (29 years) and less mission preparation time (launch 2026), however it benefits from a feature absent from previous JOM candidates, in that there is little or no ∆V en route to Jupiter (i.e. a free ride) which means the spacecraft need not carry a liquid propellant stage. This is marginally offset by the higher ∆V needed at Jupiter, requiring either 2 or 3 staged solid rocket motors. As an example, a Falcon Heavy Expendable with a CASTOR 30B booster followed by a STAR 48B can deliver 102kg to 1I/'Oumuamua by the year 2059. Other scenarios with shorter flight durations and higher payload masses are possible.

The first definite interstellar object observed in our solar system was discovered in October of 2017 and was subsequently designated 1I/'Oumuamua. In addition to its extrasolar origin, observations and analysis of this object indicate some unusual features which can only be explained by in-situ exploration.For this purpose, various spacecraft intercept missions have been proposed. Theirpropulsion schemes have been chemical, exploiting a Jupiter and Solar Oberth Maneuver (mission duration of 22 years) and also using Earth-based lasers to propel laser sails (1-2 years), both with launch dates in 2030. For the former, mission durations are quite prolonged and for the latter, the necessary laser infrastructure may not be in place by 2030. In this study Nuclear Thermal Propulsion (NTP) is examined which has yet to materialise as far as real missions are concerned, but due to its research and development in the US government sponsored Rover/NERVA programs, actually has a higher TRL than laser propulsion. Various solid reactor core options are studied, using either engines directly derived from these programs, or more advanced options, like a proposed particle bed NTR. With specific impulses at least twice those of chemical rockets, NTP opens the opportunity for much higher ΔV budgets, allowing simpler and more direct, time-saving trajectories to be exploited. For example a spacecraft with an upgraded NERVA/Pewee-class NTR travelling along an Earth-Jupiter-1I trajectory, would reach 1I/'Oumuamua within 14 years of a launch in 2031. The payload mass to 1I/'Oumuamua would be around 2.5metric tonnes, but even larger masses and shorter mssion durations can be achieved with some of the more advanced NTR options studied. In all 4 different proposed NTRs and 5 different trajectory scenarios are examined.

This paper will examine the development and test philosophy of a new, miniature pulsed plasma thruster destined for flight on the UW Dawgstar nanosatellite, and the development and test philosophy of small hydroxyl ammonium nitrate (HAN) thrusters intended for satellite use. The key to cost effective development is to know when both a rigorous analysis and a full test program is needed, and when the use of heritage hardware in a similar application may permit more latitude. The discussion will include a review of system challenges and considerations. Techniques used to design, build, and test new propulsion systems will be compared with techniques used to adapt mature hydrazine technology to new applications. As a specific example, long-term, full temperature range, materials compatibility testing for "green" monopropellants such as HAN is critical to developing suitable tanks, lines and components. In contrast, system level thermal analysis (without testing) will usually be adequate because HAN propulsion systems will be designed to take advantage of heritage hydrazine systems "lessons

Engineering details are presented for a magnetized target fusion (MTF) propulsion system designed to support crewed missions to the outer solar system. Structural, thermal and radiation-management design details are presented. Propellant storage and supply options are also discussed and a propulsion system mass estimate is given.

The Advanced Concepts Office at the Marshall Space Flight Center is an organization that can trace its roots to the days just before the launch of the Saturn launch vehicles and has distinguished itself for systems engineering and systems analysis necessary to define the future of the center and Agency. Systems analysis has evolved considerably in recent history and as computing power and knowledge has increased, the processes by which concept trade decisions are made have evolved comparably. This paper will describe the processes and tools used at Marshall Space Flight Center in the Advance Concepts Office and the Engineering Directorate Vehicle Integrated Performance Analysis (VIPA) team, the products generated using these tools and processes and future plans for further development of capabilities for advance transportation systems for the exploration of near Earth space and deep space; both human and robotic.

Electric propulsion is now a successful method for primary propulsion of deep space long duration missions and for geosyncronous satellite attitude control. Closed Drift Plasma Thruster, so called Hall Thruster or SPT (stationary plasma thruster) was primarily conceived in USSR (the ancient Soviet Union) and since then, it has been developed by space agencies, space research institutes and industries in several countries such as France, USA, Israel, Russian Federation and Brazil. In this article, we present the main features of the Permanent Magnet Hall Thruster (PMHT) developed at the Plasma laboratory of the University of Brasilia. The idea of using an array of permanent magnets, instead of an electromagnet, to produce a radial magnetic field inside the cylindrical plasma drift channel of the thruster is very significant, specially because of the possibility of developing a Hall Thruster with power consumption low enough to be used in small and medium size satellites. Descriptions such as the plasma density, temperature space profiles inside and outside the thruster channel, ion temperature measurements based on Doppler broadening of spectral lines and ion energy measurements are shown for different plasma production regimes, the space plasma potential, the measured propulsion and power consumption are shown through the paper. We also compare our data with the results reached by other types of thrusters. Finally, some of the most significant events on the evolution of the electric propulsion and their contributions to the progress on the space missions are reviewed.

Satellites have become an almost indispensable part of our lives today. All
communication systems, smart devices, the Internet of Things (IoT), data reception or
transmission, and many other scientific endeavors are carried out with the assistance
of satellites. Upon separation from the launch vehicle and during all mission lifetimes,
a satellite needs to perform many specific maneuvers. Although maneuver needs may
vary depending on the mission's orbit, they are mainly as follows: orbit raising to
increase its altitude from the launcher's initial orbit to the mission orbit; station
acquisition to capture the mission window precisely; station keeping maintaining its
position as desired against perturbations during mission life; and transferring to the
graveyard orbit for de-orbiting at the end of mission life. The propulsion systems are
an irreplaceable requirement for all satellites that aim to perform maneuvers.
In this thesis, first, electrical-based propulsion systems that have significant
advantages for the satellite sector are to be examined. Currently, they are less preferred
in the Turkish satellite sector due to their complex and unknown details. According to
literature research, it is predicted that electric propulsion systems, which have been the
trend of recent years and are widely used globally and increasing dramatically, will
have a greater adoption in the Turkish satellite sector in the next decade.
The knowledge and experiences gained in national and international projects in the
satellite sector are aimed at contributing to the literature within the scope of the near
future of the Turkish satellite sector, considering the needs of the global market.
In this thesis, it is aimed to first determine which satellite platforms for the electric
propulsion system are cost-effective and design-advantageous, and then to introduce
the design solution that would emerge if it were applied to a sample communication
satellite platform. Subsequently, it is aimed to focus on the development of the
mathematical model of the electric propulsion system in the MATLAB-SIMULINK
environment, which is rare (or even non-existent) in the Turkish industry. Finally, the
developed mathematical model is to be verified in a sample maneuver scenario.

A Helicon Plasma Thruster prototype in the 0.5–1.5 kW range is being designed and built by EP2 and SENER. The prototype is proposed as a flexible testing platform for different geometries, magnetic configurations, antennas, frequencies and gases. The goal of the prototype is to verify the prediction of current models, understand the power loss mechanisms in the helicon plasma thruster, and investigate the optimization of its performance. The lessons learned from the first test campaign at ESA-ESTEC-EPL will help to improve the design and scale up to higher powers in the future. This paper presents the preliminary analysis and design results and discusses the main aspects that are expected to drive performance of the device.

We perform one-dimensional particle-in-cell (PIC) simulation of external electromagnetic field excitation into magnetized plasmas. We consider two models for the electromagnetic field excitation: electrostatic excitation by electrodes and electromagnetic excitation by current antenna. Here, the external electrodes are placed outside plasma region, background magnetic field is perpendicular to the one-dimensional direction, and the externally applied field frequency is chosen in a range below the lower-hybrid frequency. For both models, we will discuss the electromagnetic field excitation processes by varying the externally applied field frequency and the plasma radius.

Characteristics of High-Density Helicon Plasma Sources and Their Application to Electrodeless Electric Propulsion 1 S. SHINOHARA, H. NISHIDA, T. NAKAMURA, A. MISHIO, H. ISHII, N. TESHIGAHARA, H. FU-JITSUKA, S. WASEDA, TUAT, T. TANIKAWA, Tokai U., T. HADA, F. OTSUKA, Kyushu U., I. FUNAKI, T. MATSUOKA, JAXA, K. SHAMRAI, T. RUDENKO, INR -High-density but low temperature helicon plasmas have been proved to be very useful for fundamental research as well as for various applications. First, we introduce our very large helicon sources [1] with a diameter up to 74 cm. For the industrial and propulsion applications, we have reduced the aspect ratio (axial length-to-diameter) down to 0.075, and examined the discharge performance and wave characteristics. Then, we discuss our small helicon sources [1] for developing new electrodeless acceleration schemes. Some experimental and theoretical results [2] by applying the rotating magnetic (or electric) fields to the helicon plasma under the divergent magnetic field will be presented, along with other propulsion schemes. In addition, an initial plasma production experiment with very small diameter will be described.

The transition from OLD SPACE to NEW SPACE along with increasing commercialization has a major impact on space flight, in general, and on electric propulsion (EP) by ion thrusters, in particular. Ion thrusters are nowadays used as primary propulsion systems in space. This article describes how these changes related to NEW SPACE affect various aspects that are important for the development of EP systems. Starting with a historical overview of the development of space flight and of the technology of EP systems, a number of important missions with EP and the underlying technologies are presented. The focus of our discussion is the technology of the radio frequency ion thruster as a prominent member of the gridded ion engine family. Based on this discussion, we give an overview of important research topics such as the search for alternative propellants, the development of reliable neutralizer concepts based on novel insert materials, as well as promising neutralizer-free propulsion concepts. In addition, aspects of thruster modeling and requirements for test facilities are discussed. Furthermore, we address aspects of space electronics with regard to the development of highly efficient electronic components as well as aspects of electromagnetic compatibility and radiation hardness. This article concludes with a presentation of the interaction of EP systems with the spacecraft.

A model of the maximum achievable exhaust velocity of a conical theta pinch pulsed inductive thruster is presented. A semi-empirical formula relating coil inductance to both axial and radial current sheet location is developed and incorporated into a circuit model coupled to a momentum equation to evaluate the effect of coil geometry on the axial directed kinetic energy of the exhaust. Inductance measurements as a function of the axial and radial displacement of simulated current sheets from four coils of different geometries are fit to a two-dimensional expression to allow the calculation of the Lorentz force at any relevant averaged current sheet location. This relation for two-dimensional inductance, along with an estimate of the maximum possible change in gas-dynamic pressure as the current sheet accelerates into downstream propellant, enables the expansion of a one-dimensional circuit model to two dimensions. The results of this two-dimensional model indicate that radial curren...

The China-Brazil Earth Resources Satellite (CBERS-FM2) was designed to be three axes stabilized, in other words, an inertial stabilized satellite. As part of the qualification tests, the spacecraft hydrazine propulsion system had to be completely checked against any possible significant leakage which would dangerously reduce its operational lifetime while in orbit. These tests included the Local Leak Tests, where each part of the propulsion system was individually leak tested using Helium Leak Detectors, and the Global Leak Tests, where the whole spacecraft Service Module (SM) was installed in a Space Simulation Chamber, a large thermal vacuum system, where an external Helium Leak Detector was attached for leakage rate measurements. This paper describes the development of the employed methodologies and does an analysis of the obtained results.

Cylindrical disks composed of HfB 2 -26 vol%SiC f are successfully fabricated by combining the Self-propagating High temperature Synthesis (SHS) and Spark Plasma Sintering (SPS) processes. Specifically, HfB 2 powders synthesized by SHS are mixed with SiC fibers and the resulting mixture is consolidated by SPS. The presence of SiC fibers was retained in the nearly full dense materials sintered at 1800 °C. The measured values of Vickers hardness, fracture toughness, flexural strength and elastic modulus at room temperature are 21.6±0.8 GPa, 6.2±0.5 MPa•m 0.5 , 429±45 MPa and 312±17 GPa, respectively. The obtained products are found to well resist to ablation, with weight losses below 0.25% when exposed to the most aggressive heat flux conditions (1250 W/cm 2 ). A nozzle component of the desired shape and size is finally obtained for space propulsion applications after electrical discharge machining the SPS material.