Showing papers on "Spacecraft propulsion published in 2012"

Nuclear Thermal Propulsion (NTP): A proven growth technology for human NEO/Mars exploration missions

Abstract: The nuclear thermal rocket (NTR) represents the next “evolutionary step” in high performance rocket propulsion. Unlike conventional chemical rockets that produce their energy through combustion, the NTR derives its energy from fission of Uranium-235 atoms contained within fuel elements that comprise the engine's reactor core. Using an “expander” cycle for turbopump drive power, hydrogen propellant is raised to a high pressure and pumped through coolant channels in the fuel elements where it is superheated then expanded out a supersonic nozzle to generate high thrust. By using hydrogen for both the reactor coolant and propellant, the NTR can achieve specific impulse (I sp ) values of ∼900 seconds (s) or more — twice that of today's best chemical rockets. From 1955–1972, twenty rocket reactors were designed, built and ground tested in the Rover and NERVA (Nuclear Engine for Rocket Vehicle Applications) programs. These programs demonstrated: (1) high temperature carbide-based nuclear fuels; (2) a wide range of thrust levels; (3) sustained engine operation; (4) accumulated lifetime at full power; and (5) restart capability — all the requirements needed for a human Mars mission. Ceramic metal “cermet” fuel was pursued as well, as a backup option. The NTR also has significant “evolution and growth” capability. Configured as a “bimodal” system, it can generate its own electrical power to support spacecraft operational needs. Adding an oxygen “afterburner” nozzle introduces a variable thrust and I sp capability and allows bipropellant operation. In NASA's recent Mars Design Reference Architecture (DRA) 5.0 study, the NTR was selected as the preferred propulsion option because of its proven technology, higher performance, lower launch mass, versatile vehicle design, simple assembly, and growth potential. In contrast to other advanced propulsion options, no large technology scale-ups are required for NTP either. In fact, the smallest engine tested during the Rover program — the 25,000 lb f (25 klb f ) “Pewee” engine is sufficient when used in a clustered engine arrangement. The “Copernicus” crewed spacecraft design developed in DRA 5.0 has significant capability and a human exploration strategy is outlined here that uses Copernicus and its key components for precursor near Earth object (NEO) and Mars orbital missions prior to a Mars landing mission. The paper also discusses NASA's current activities and future plans for NTP development that include system-level Technology Demonstrations — specifically ground testing a small, scalable NTR by 2020, with a flight test shortly thereafter.

Solar Sails: Technology and demonstration status

Abstract: Solar Sail propulsion has been validated in space (IKAROS, 2010) and soon several more solar-sail propelled spacecraft will be flown. Using sunlight for spacecraft propulsion is not a new idea. First proposed by Frederick Tsander and Konstantin Tsiolkovsky in the 1920’s, NASA’s Echo 1 balloon, launched in 1960, was the first spacecraft for which the effects of solar photon pressure were measured. Solar sails reflect sunlight to achieve thrust, thus eliminating the need for costly and often very-heavy fuel. Such “propellantless” propulsion will enable whole new classes of space science and exploration missions previously not considered possible due to the propulsive-intense maneuvers and operations required.

Nammo Hybrid Rocket Propulsion TRL Improvement Program

Abstract: In 2010, Nammo initiated an ESA-funded hybrid propulsion Technology Readiness Level (TRL) improvement program based on ideas proposed in a study performed for Ariane 5 ME a year before. In a program lasting less than one year, Nammo executed 29 hybrid rocket firing tests using hydrogen peroxide as the oxidizer. All the tests were successful, demonstrating hybrid rocket propulsion exhibiting high performance under different conditions.

Cold Gas Propulsion System - An Ideal Choice for Remote Sensing Small Satellites

Abstract: Cold gas propulsion systems play an ideal role while considering small satellites for a wide range of earth orbit and even interplanetary missions. These systems have been used quite frequently in small satellites since 1960’s. It has proven to be the most suitable and successful low thrust space propulsion for LEO maneuvers, due to its low complexity, efficient use of propellant which presents no contamination and thermal emission besides its low cost and power consumed. The major benefits obtained from this system are low budget, mass, and volume. The system mainly consists of a propellant tank, solenoid valves, thrusters, tubing and fittings (fig. 1). The propellant tank stores the fuel required for attitude control of satellite during its operation in an orbit. The fuel used in cold gas systems is compressed gas. Thrusters provide sufficient amount of force to provide stabilization in pitch, yaw and roll movement of satellite. From design point of view, three important components of cold gas propulsion systems play an important role i.e. mission design, propellant tank and cold gas thrusters. These components are discussed in detail in section 3. Selection of suitable propellant for cold gas systems is as important as above three components. This part is discussed in section 2 of this chapter. Section 4 describes the case study of cold gas propulsion system which is practically implemented in Pakistan’s first prototype remote sensing satellite PRSS.

Peregrine Hybrid Rocket Motor Ground Test Results

Abstract: To further develop and demonstrate the applicability of liquefying-fuel hybrid rocket technology to low-cost launch applications, a small team of engineers is developing a medium-scale liquefying-fuel hybrid sounding rocket using storable propellants (paraffin wax and N2O) that will carry a 5 kg payload to the edge of space. This rocket, known as Peregrine, is being developed by engineers from NASA Ames, Stanford University, Space Propulsion Group Inc. (SPG, Sunnyvale, CA) and NASA Wallops, with a launch from Wallops anticipated at some point in the future. This paper focuses on the propulsion ground test results obtained to date.

Solar Electric Propulsion Vehicle Demonstration to Support Future Space Exploration Missions

Abstract: Human and robotic exploration beyond Low Earth Orbit (LEO) will require enabling capabilities that are efficient, affordable, and reliable Solar Electric Propulsion (SEP) is highly advantageous because of its favorable in-space mass transfer efficiency compared to traditional chemical propulsion systems The NASA studies have demonstrated that this advantage becomes highly significant as missions progress beyond Earth orbit Recent studies of human exploration missions and architectures evaluated the capabilities needed to perform a variety of human exploration missions including missions to Near Earth Objects (NEOs) The studies demonstrated that SEP stages have potential to be the most cost effective solution to perform beyond LEO transfers of high mass cargoes for human missions Recognizing that these missions require power levels more than 10X greater than current electric propulsion systems, NASA embarked upon a progressive pathway to identify critical technologies needed and a plan for an incremental demonstration mission The NASA studies identified a 30kW class demonstration mission that can serve as a meaningful demonstration of the technologies, operational challenges, and provide the appropriate scaling and modularity required This paper describes the planning options for a representative demonstration 30kW class SEP mission

Analysis of Thrust Characteristics of a Magnetic Sail in a Magnetized Solar Wind

Abstract: Themagnetic sail is an advanced space propulsion concept that uses an artificial magnetosphere for capturing the solar wind energy. In this study, the interaction of the solar wind with themagnetosphere of a magnetic sail has been simulated based on the resistive magnetohydrodynamics model in two-dimensional space and the plasmadynamic characteristics of magnetic sail were evaluated. When the solar wind is not magnetized by the interplanetary magnetic field, the attitude of the magnetic sail spacecraft is static stable when the magnetic moment vector is perpendicular to the solar wind flow direction. The interplanetary magnetic field not only enhances a drag force in the direction leaving the sun (i.e., thrust) but also acts on the pitching moment; the pitching moment due to the interplanetary magnetic field rotates the magnetic sail spacecraft so as to align the magnetic moment vector parallel to the interplanetarymagnetic field. Despite the weak interplanetarymagnetic field adopted in the simulation, which is 1 order of magnitude lower than the typical value, the pitching moment coefficient is significant. The attitude stability of the magnetic sail is hence strongly affected by the interplanetary magnetic field.

Hypergolic Systems: A Review in Patents

Abstract: Hypergolic reactions may be useful in civil and military applications. In the area of rocket propulsion, they constitute a potential oeld due to the reduced Zeight and comple(ity of fuel inMection systems, alloZing yet controlled use of the propulsors. This manuscript aimed at presenting different hypergolic systems and their particularities, comparing them Zith chemical propulsion systems, Zhich are most commonly employed in rocket motors, for e(ample.

Permanent magnet synchronous motor to drive propellant pump

Abstract: The paper presents problems investigated by Mikroma in the project “In Space Propulsion”, realized within 7th Framework Programme financed by the European Union. The investigation concerns elaboration of an electric motor to drive a fuel pump in the new concept of cryogenic rocket propulsion with small thrust force. The problems related to elaboration of the motor suited to work in cryogenic conditions are discussed. Chosen results of investigation of the influence of cryogenic temperature on the properties of manufacturing materials of the motor are shown. Moreover, results of laboratory tests of constructed prototype motors are presented.

Nuclear Thermal Rocket (Ntr) Propulsion: A Proven Game-Changing Technology for Future Human Exploration Missions

Abstract: The NTR represents the next evolutionary step in high performance rocket propulsion. It generates high thrust and has a specific impulse (Isp) of approx.900 seconds (s) or more V twice that of today s best chemical rockets. The technology is also proven. During the previous Rover and NERVA (Nuclear Engine for Rocket Vehicle Applications) nuclear rocket programs, 20 rocket reactors were designed, built and ground tested. These tests demonstrated: (1) a wide range of thrust; (2) high temperature carbide-based nuclear fuel; (3) sustained engine operation; (4) accumulated lifetime; and (5) restart capability V all the requirements needed for a human mission to Mars. Ceramic metal cermet fuel was also pursued, as a backup option. The NTR also has significant growth and evolution potential. Configured as a bimodal system, it can generate electrical power for the spacecraft. Adding an oxygen afterburner nozzle introduces a variable thrust and Isp capability and allows bipropellant operation. In NASA s recent Mars Design Reference Architecture (DRA) 5.0 study, the NTR was selected as the preferred propulsion option because of its proven technology, higher performance, lower launch mass, simple assembly and mission operations. In contrast to other advanced propulsion options, NTP requires no large technology scale-ups. In fact, the smallest engine tested during the Rover program V the 25,000 lbf (25 klbf) Pewee engine is sufficient for human Mars missions when used in a clustered engine arrangement. The Copernicus crewed spacecraft design developed in DRA 5.0 has significant capability and a human exploration strategy is outlined here that uses Copernicus and its key components for precursor near Earth asteroid (NEA) and Mars orbital missions prior to a Mars landing mission. Initially, the basic Copernicus vehicle can enable reusable 1-year round trip human missions to candidate NEAs like 1991 JW and Apophis in the late 2020 s to check out vehicle systems. Afterwards, the Copernicus spacecraft and its 2 key components, now configured as an Earth Return Vehicle / propellant tanker, would be used for a short round trip (approx.18 - 20 months)/short orbital stay (60 days) Mars / Phobos survey mission in 2033 using a split mission approach. The paper also discusses NASA s current Foundational Technology Development activities and its pre-decisional plans for future system-level Technology Demonstrations that include ground testing a small (approx.7.5 klbf) scalable NTR before the decade is out with a flight test shortly thereafter.

Nuclear Cryogenic Propulsion Stage

Abstract: The fundamental capability of Nuclear Thermal Propulsion (NTP) is game changing for space exploration. A first generation Nuclear Cryogenic Propulsion Stage (NCPS) based on NTP could provide high thrust at a specific impulse above 900 s, roughly double that of state of the art chemical engines. Characteristics of fission and NTP indicate that useful first generation systems will provide a foundation for future systems with extremely high performance. The role of the NCPS in the development of advanced nuclear propulsion systems could be analogous to the role of the DC-3 in the development of advanced aviation. Progress made under the NCPS project could help enable both advanced NTP and advanced NEP.

Testing of the NASA Hypersonics Project Combined Cycle Engine Large Scale Inlet Mode Transition Experiment (CCE LlMX)

Abstract: Status on an effort to develop Turbine Based Combined Cycle (TBCC) propulsion is described. This propulsion technology can enable reliable and reusable space launch systems. TBCC propulsion offers improved performance and safety over rocket propulsion. The potential to realize aircraft-like operations and reduced maintenance are additional benefits. Among most the critical TBCC enabling technologies are: 1) mode transition from turbine to scramjet propulsion, 2) high Mach turbine engines and 3) TBCC integration. To address these TBCC challenges, the effort is centered on a propulsion mode transition experiment and includes analytical research. The test program, the Combined-Cycle Engine Large Scale Inlet Mode Transition Experiment (CCE LIMX), was conceived to integrate TBCC propulsion with proposed hypersonic vehicles. The goals address: (1) dual inlet operability and performance, (2) mode-transition sequences enabling a switch between turbine and scramjet flow paths, and (3) turbine engine transients during transition. Four test phases are planned from which a database can be used to both validate design and analysis codes and characterize operability and integration issues for TBCC propulsion. In this paper we discuss the research objectives, features of the CCE hardware and test plans, and status of the parametric inlet characterization testing which began in 2011. This effort is sponsored by the NASA Fundamental Aeronautics Hypersonics project

Ascent trajectory optimisation for a single-stage-to-orbit vehicle with hybrid propulsion

Abstract: This paper addresses the design of ascent trajectories for a hybrid-engine, high performance, unmanned, single-stage-to-orbit vehicle for payload deployment into low Earth orbit A hybrid optimisation technique that couples a population-based, stochastic algorithm with a deterministic, gradient-based technique is used to maximize the nal vehicle mass in low Earth orbit after accounting for operational constraints on the dynamic pressure, Mach number and maximum axial and normal accelerations The control search space is first explored by the population-based algorithm, which uses a single shooting method to evaluate the performance of candidate solutions The resultant optimal control law and corresponding trajectory are then further refined by a direct collocation method based on finite elements in time Two distinct operational phases, one using an air-breathing propulsion mode and the second using rocket propulsion, are considered The presence of uncertainties in the atmospheric and vehicle aerodynamic models are considered in order to quantify their effect on the performance of the vehicle Firstly, the deterministic optimal control law is re-integrated after introducing uncertainties into the models The proximity of the final solutions to the target states are analysed statistically A second analysis is then performed, aimed at determining the best performance of the vehicle when these uncertainties are included directly in the optimisation The statistical analysis of the results obtained are summarized by an expectancy curve which represents the probable vehicle performance as a function of the uncertain system parameters This analysis can be used during the preliminary phase of design to yield valuable insights into the robustness of the performance of the vehicle to uncertainties in the specification of its parameters

Thermal Characterization of a Hall Effect Thruster

Abstract: : The thermal characteristics of a Hall thruster directly influence thruster and spacecraft design. High temperatures affect the magnetic coil capabilities and cause higher insulator erosion rates, influencing both thruster performance and lifetime. The Hall thruster transfers heat through both radiation and conduction, and the spacecraft must handle this additional thermal energy. An infrared camera provides a non-intrusive method to analyze the thermal characteristics of an operational Hall thruster. This thesis contains the thermal analysis of a Busek Co. Inc. 200 W Hall thruster, using a FLIR ThermaCAM SC640 infrared camera. The Space Propulsion Analysis and System Simulator Laboratory at the Air Force Institute of Technology on Wright-Patterson Air Force Base provided the location for thruster set up and operation. The infrared camera furnishes the surface temperatures for the entire thruster, and approximates the transient heating behavior during start up, steady state, and shut down. Thermocouples verify and correct the camera data. Experimentally determined emissivities characterize the materials of the thruster. In addition, a view factor analysis between the camera pixels and the alumina sprayed portion of the cathode determines the exchange of radiation between the pixels and cathode surface. This process develops a technique to map surface temperatures of complex geometries with confidence in the actual values. Accurately mapping the surface temperatures of a Hall Effect thruster will improve both thruster efficiency and lifetime, and predict the thruster's thermal load on a satellite.

NOFBX™ single stage to Orbit Mars Ascent Vehicle

Abstract: Nitrous Oxide Fuel Blend (NOFBX™) is a nitrous-oxide-based mono-propulsion technology developed specifically for a Mars Ascent Vehicle (MAV) application by Firestar Technologies, largely funded by the NASA Mars Program Office, NASA Small Business Innovation Research (SBIR), and NASA International Space Station (ISS). The resulting propellant and engine technology offers significant benefits over previously contemplated solid and storable liquid propulsion solutions that have been proposed for planetary and in-space missions including a MAV, enabling a Single State to Orbit (SSTO) ascent vehicle for Mars Sample Return (MSR). Important characteristics of the technology include specific impulse performance that equals or exceeds state-of-the-art nitrogen textroxide/ monomethyl hydrazine (NTO/MMH) propulsion systems, restartability via spark ignition, deep throttleability (100:1) and high thrust-to-weight ratio. The propellant is completely non-toxic, non-contaminating to the planetary environment and storable over a very wide range of temperatures ( +70°C). Together these attributes offer the potential for dramatic simplification to the MSR program architecture as well as to the MAV itself, increasing margins, eliminating failure modes, improving reliability and decreasing cost. Firestar has demonstrated NOFBX™ thrusters in the 0.4 N to 445 N (0.1–100 lbf) thrust range, with measured vacuum specific impulse (vIsp) performance of up to ∼325 seconds. Innovative Space Propulsion Systems (ISPS) has been formed by Firestar and other strategic partners to commercialize NOFBX™ technology and will be demonstrating the 445 N (100 lbf) engine and associated feed system components on the ISS in 2013 under a National Lab award from NASA. Also, a current Phase 2 SBIR from NASA will demonstrate the proposed SSTO MAV propulsion module in a series of ground tests. This paper will describe NOFBX™ characteristics and applications in greater detail and status of the ISS flight demonstration. The proposed Single Stage to Orbit MAV architecture will then be described with associated performance estimates and benefits to both the MAV itself and the surrounding program architecture. We will show how NOFBX™ directly addresses and mitigates many of the key technology challenges identified in recent NASA MAV architecture studies.

Plasma propulsion for geostationary satellites for telecommunication and interplanetary missions

Abstract: The advantages of electric propulsion for the orbit maintenance of geostationary satellites for telecommunications are described. Different types of plasma sources for space propulsion are presented. Due to its large performances, one of them, named Hall effect thruster is described in detail and two recent missions in space (Stentor and Smart1) using French Hall thrusters are briefly presented.

Design of an Oxygen Turbopump for a Dual Expander Cycle Rocket Engine

Abstract: : The design of a pump intended for use with a dual expander cycle (LOX/H2) engine is presented. This arrangement offers a number of advantages over hydrogen expander cycles; among these are the elimination of gearboxes and inter-propellant purges and seals, an extended throttling range, and higher engine operating pressures and performance. The target engine has been designed to meet the needs of Phase III of the Integrated High Payoff Rocket Propulsion Technology (IHPRPT) program; thus, this pump must meet the program's reliability, maintainability, and service life goals. In addition, this pump will be driven by warm gaseous oxygen. In order to meet the needs of this engine, the pump will need to be capable of delivering 106 lbm/s (48.1 kg/s) at 4500 psi (31 MPa); this will necessitate a turbine capable of supplying at least 2215 hp (1652 kW). The pump and turbine were designed with the aid of an industry standard design program; the design methodology and justification for design choices are presented. Appropriate materials of construction and bearings for this pump are discussed.

Chemical rocket propulsion

Abstract: 1Department of Aerospace Engineering, Ryerson University, Toronto, ON, Canada M5B 2K3 2Aerophysics Branch, Air Force Research Laboratory, Edwards AFB, CA 93524, USA 3Dipartimento di Ingegneria Meccanica e Aerospaziale, Politecnico di Torino, 10129 Torino, Italy 4Department of Aeronautical Engineering, Kumaraguru College of Technology, Coimbatore, Tamil Nadu 641006, India 5 School of Mechanical and Mining Engineering, The University of Queensland, Brisbane, QLD 4072, Australia

A Study of Gas-Centered Swirl Coaxial Injectors using X-ray Radiography

Abstract: : Gas-centered swirl coaxial injectors, a specific type of airblast atomizer, are of interest in rocket propulsion applications. These applications require good mixing of the liquid and gas to ensure complete combustion within the engine. While strides are being made on the computational front, predictions of the mass distributions achieved with this type of injector remain too costly or too inaccurate for engineering design. There has been, therefore, a reliance on experimental results. Unfortunately, the mass flow rates and the strong gas phase typically encountered in rocket engines create sprays with high optical densities and render the vast majority of optical and laser techniques ineffective. Data has been obtainable through mechanical patternation, but the technique has limitations. Time-gated ballistic imaging has also shown promise in rocket injectors but produces only qualitative information. An x-ray radiographic technique with a high-brilliance x-ray source (Advanced Photon Source) has been applied to these high-optical-density sprays. To achieve this testing a new, mobile flow facility was constructed; this facility simulates the rocket flows using water and nitrogen instead of fuel and oxidizer. The x-ray radiography technique has been able to measure equilvent path length in gas-centered swirl coaxial injectors at a range of typical operating conditions. These results and their implications for gas- centered swirl coaxial injector performance in liquid rocket engines are discussed.

Multiple NEO Rendezvous Using Solar Sail Propulsion

Abstract: The NASA Marshall Space Flight Center (MSFC) Advanced Concepts Office performed an assessment of the feasibility of using a near-term solar sail propulsion system to enable a single spacecraft to perform serial rendezvous operations at multiple Near Earth Objects (NEOs) within six years of launch on a small-to-moderate launch vehicle. The study baselined the use of the sail technology demonstrated in the mid-2000 s by the NASA In-Space Propulsion Technology Project and is scheduled to be demonstrated in space by 2014 as part of the NASA Technology Demonstration Mission Program. The study ground rules required that the solar sail be the only new technology on the flight; all other spacecraft systems and instruments must have had previous space test and qualification. The resulting mission concept uses an 80-m X 80-m 3-axis stabilized solar sail launched by an Athena-II rocket in 2017 to rendezvous with 1999 AO10, Apophis and 2001 QJ142. In each rendezvous, the spacecraft will perform proximity operations for approximately 30 days. The spacecraft science payload is simple and lightweight; it will consist of only the multispectral imager flown on the Near Earth Asteroid Rendezvous (NEAR) mission to 433 Eros and 253 Mathilde. Most non-sail spacecraft systems are based on the Messenger mission spacecraft. This paper will describe the objectives of the proposed mission, the solar sail technology to be employed, the spacecraft system and subsystems, as well as the overall mission profile.

Nuclear Thermal Propulsion for Advanced Space Exploration

Abstract: The fundamental capability of Nuclear Thermal Propulsion (NTP) is game changing for space exploration. A first generation Nuclear Cryogenic Propulsion Stage (NCPS) based on NTP could provide high thrust at a specific impulse above 900 s, roughly double that of state of the art chemical engines. Characteristics of fission and NTP indicate that useful first generation systems will provide a foundation for future systems with extremely high performance. The role of the NCPS in the development of advanced nuclear propulsion systems could be analogous to the role of the DC-3 in the development of advanced aviation. Progress made under the NCPS project could help enable both advanced NTP and advanced Nuclear Electric Propulsion (NEP).

Ignition and Combustion of Magnesium Particles in Carbon Dioxide

Abstract: Metal-CO2 propulsion is less known than in-situ resource utilization (ISRU) technologies. This concept, based on using Martian carbon dioxide as an oxidizer in jet or rocket engines, offers the advantage of no chemical processing for CO2 and thus requires less power consumption than ISRU alternatives. In this paper, we study the burning behavior of the Mg in a CO2 atmosphere to assess the feasibility of using Mg/CO2 reactions as an in situ resource utilization technology for rocket propulsion and energy generation on other planets. From the experimental results, we can see that the critical ignition temperature increases with increasing the particle size and decreases with increasing the ambient pressure. In the CO2 atmosphere, we found the complicated sequence of interaction modes including pulsating combustion in a wide range of ambient temperatures. The pulsation frequency is determined by the sample temperature at the phases of slow heterogeneous combustion between the flashes. The combustion mechanisms are discussed with consideration of processes in both a surface film and gas phase.

Technological and programmatic development of Zefiro 40 Solid Rocket Motor

Abstract: In the frame of VEGA Evolution Program the development of a technological demonstrator, named Zefiro 40 Solid Rocket Motor, will allow to introduce the materials and technologies improvements able to increase the propulsive performance of VEGA Launcher solid rocket motors reaching a capability of greater payloads insertion in higher orbits. System analyses performed by ELV, Prime Contractor of VEGA Launcher, showed an initial possible step foreseeing the first stage propellant increase from 80 to about 120 tons and second stage propellant increase from 23 to about 40 tons. Zefiro 40 will represent therefore the means for Avio S.p.A., leader in solid space propulsion and responsible of design and production of Zefiro 23 and Zefiro 9 SRMs, as well as P80 Loaded Motor Case, to develop innovative solutions with the primary target to increase the motor mass fraction, as an extremely low density thermal protection rubber, a high performance carbon-epoxy own pre-preg production, a self-protected nozzle flexible joint, an undercut casting design. Moreover further technology targets were identified in order to improve the industrial processes of Avio facilities, in particular to reach higher values of nozzle divergent diameters and greater loaded propellant masses, but also to avoid defects occurrence during nozzle and case fabrication, to reduce the propellant compounds class risk and to prepare the advent of larger scale solid rocket motors, which will require, for example, segmented composite cases and improved casting facilities. In this paper an overview is presented of planned development and qualification program of Zefiro 40 solid rocket motor.

Chapter 14 – Space Propulsion

Abstract: Publisher Summary This chapter discusses propulsion of a rocket in outer space. The thrust or specific impulse is measured by the total impulse delivered by the engine per unit weight of propellant consumed. Electric propulsion systems have also been looked into. Jhan (1968) defines electric propulsion as the acceleration of gases for propulsion by electric heating and/or by electric and magnetic body forces. Thus electro-thermal, electrostatic, and electromagnetic propulsion devices can be considered for propulsion in space. Electro-thermal devices like arcjets are popular and find space in the study. Propellants for arcjets include hydrazine (N2H4), ammonia (NH3), and hydrogen (H2). An arcjet using ammonia as the propellant produced a thrust of 2N at a specific impulse of 800 s. One-dimensional electrostatic thruster, ion stream speeds are also discussed. Hall thrusters are also studied. Under electromagnetic propulsion devices, the magneto-plasma-dynamic accelerator (MPD) is discussed in details.

Investigation of Singly Ionized Iodine Spectroscopy in Support of Electrostatic Propulsion Diagnostics Development

Abstract: : This effort examines the spectroscopy of the second spectrum of the iodine atom (I II) in order to determine one, or more, useful transitions for laser-induced fluorescence of an accelerated atomic iodine singly charged ion (I+). While the second spectrum of iodine has been analyzed, it is not particularly well characterized. Nor has it been studied substantially within a plasma such as those of interest to the spacecraft propulsion community. Our goal is to examine the spectral data available in the literature and determine transitions suitable for development into diagnostics tools, such as laser- induced fluorescence (LIF), to examine the plasma acceleration within an electrostatic plasma propulsion thruster. While xenon remains the preferred propellant for electrostatic spacecraft propulsion, a number of alternative propellants are being analyzed in various laboratories. Some of the propellants that have been investigated in the recent literature include krypton, bismuth, and iodine. Of these alternative propellant candidates, iodine is the least well investigated. However, due to its close mass (127 versus 131 amu) compared to xenon, it has strong potential for use as an electrostatic propulsion propellant. Iodine's benefis include a solid density of 4.9 g/cc, a low boiling point of 183 degrees C. Compared to xenon storage density of 1.2 g/cc at 2,000 psi, or the bismuth boiling point of 1,564 degrees C, there appear to be system level advantages to iodine fueled electrostatic spacecraft propulsion. This effort focuses on the development of a laser-induced fluorescence diagnostic tool for the iodine ion.

Beamed-Energy Propulsion (BEP) Study

Abstract: The scope of this study was to (1) review and analyze the state-of-art in beamed-energy propulsion (BEP) by identifying potential game-changing applications, (2) formulate a roadmap of technology development, and (3) identify key near-term technology demonstrations to rapidly advance elements of BEP technology to Technology Readiness Level (TRL) 6. The two major areas of interest were launching payloads and space propulsion. More generally, the study was requested and structured to address basic mission feasibility. The attraction of beamed-energy propulsion (BEP) is the potential for high specific impulse while removing the power-generation mass. The rapid advancements in high-energy beamed-power systems and optics over the past 20 years warranted a fresh look at the technology. For launching payloads, the study concluded that using BEP to propel vehicles into space is technically feasible if a commitment to develop new technologies and large investments can be made over long periods of time. From a commercial competitive standpoint, if an advantage of beamed energy for Earth-to-orbit (ETO) is to be found, it will rest with smaller, frequently launched payloads. For space propulsion, the study concluded that using beamed energy to propel vehicles from low Earth orbit to geosynchronous Earth orbit (LEO-GEO) and into deep space is definitely feasible and showed distinct advantages and greater potential over current propulsion technologies. However, this conclusion also assumes that upfront infrastructure investments and commitments to critical technologies will be made over long periods of time. The chief issue, similar to that for payloads, is high infrastructure costs.

Progress Toward Demonstration of Remote, Autonomous Attitude Control of a CubeSat Using Ion Electrospray Propulsion Systems

Abstract: As a solution to the problem of scalable propulsion in small satellite architectures, the microfabricated ion Electrospray Propulsion System (iEPS) developed by MIT’s Space Propulsion Laboratory has progressed to a point where it is ready to be characterized on a realistic testbed. In this paper, developments in the iEPS thruster design and testing equipment are outlined. Design changes address the performance and testing issues encountered with the rst version. These changes include features to mitigate the formation of liquid current paths and ease grid-to-tip alignment and grid removal. Additionally, a 1-DOF free-oating CubeSat testbed with an integrated autonomous remote control system and high-voltage PPU has been developed and tested for use as a thrust balance and attitude control demonstrator.

Program AHRES and its Contribution to Assess Features and Current Limitations of Hybrid Rocket Propulsion

Abstract: In the year 2011 the German Aerospace Center (DLR) has started the program AHRES (Advanced Hybrid Rocket Engine Simulation). The aim is the development of software engineering tools and CFD tools for conceptual design and optimisation of hybrid rocket engines (HRE). To provide high accuracy software tools, the AHRES program includes complex experiments with hybrid rocket engines in laboratory and prototype scale, for pressures up to 70 bars. For this purpose an existing test facility at the DLR site Trauen (former EUROPA 2 upper stage test range) was enhanced, in order to conduct tests with a HRE. In this paper a short description of the operational test bed is given, which is equipped with sophisticated measurement techniques. The HRE test bed is currently equipped with oxidiser supply systems for high-test peroxide (HTP) and gaseous oxygen. The solid fuel grains are based on hydroxyl-terminated polybutadiene (HTPB) mixed with different metallic additives to improve the performance. Currently conducted tests are aimed to characterise the HRE’s combustion processes and to generate an experimental database to develop the above mentioned simulation tools. Based on experimental test data, calculations on developed simulation tools and literature sources a comparison of the HRE with solid and storable liquid rocket engines is made. To illustrate one possible application of a HRE in near-future a HRE which could be applied as an upper stage is designed and compared with existing solid rocket engines. For the comparison the third stage “Zefiro 9A” of the VEGA launcher is taken. The preliminary results and conclusions are presented, including specific impulse, combustion and engine efficiency, as well as the achievable regression rate dependent on fuel composition. The last factor is of tremendous importance to determine the achievable thrust level of a hybrid rocket engine. This again, determines whether a HRE might be promising for the application as an upper stage. The results are presented graphically, in table form, analysed, and commented.