Showing papers on "Propellant published in 2005"

Burning of Nano-Aluminized Composite Rocket Propellants

Abstract: Several aluminum nanopowders were examined and compared with the final goal to evaluate their application in solid rocket propulsion. A detailed investigation of pre-burning properties by the Brunauer-Emmet-Teller method, electron microscopy, X-ray diffraction, and X-ray photoelectron spectroscopy was carried out. Ballistic properties and the combustion mechanism of several aluminized propellant formulations were investigated. In particular, aggregation and agglomeration of metal particles at and near the burning surface were analyzed by high-speed high-resolution color digital video recordings. All tested nano-powders are of Russian production; their physical characterization was carried out at the Istituto Donegani (Novara, Italy); ballistic studies were performed at the Solid Propulsion Laboratory (Milano, Italy) using laboratory and, for comparison, industrial composite propellants based on ammonium perchlorate as an oxidizer. Results obtained under a fair variety of operating conditions typical of rocket propulsion indicate, for increasing nano-Al mass fraction or decreasing nano-Al size, larger steady burning rates with essentially the same pressure sensitivity. While aggregation and agglomeration phenomena still occur, their significance may be reduced by using nano-Al instead of micro-Al.

Transcritical oxygen/transcritical or supercritical methane combustion

Abstract: Injection of liquid fluid initially at subcritical temperature into an environment in which the temperature and pressure exceed the thermodynamic critical conditions is an important phenomenon in many high performance devices like liquid propellant rocket engines. This is found, for example, in the Space Shuttle main engines or in the Ariane 5 Vulcain engine both operating with liquid oxygen (LOx) and gaseous hydrogen (GH2). This article is concerned with the less standard situation where both reactants are in a transcritical state. One case of current interest in propulsion, that of combustion of cryogenic oxygen and methane injected at high pressure, is investigated experimentally. A coaxial injector delivers oxygen at a temperature of 85 K and methane at 120 or 288 K. The pressure in the chamber takes values between 4.5 and 6 MPa. Emission images from excited state OH (A2Σ, denoted OH*) and CH (A2Δ, denoted CH*) are recorded and averaged. The Abel transform is used to determine the mean flame structure from these average images. Data indicate that the flame is stabilized in the vicinity of the injector. When both propellants are transcritical, the flame features two conical regions of light emission, one spreading close to the liquid oxygen boundary and the other located further away from the axis near the liquid methane boundary. The outer flame boundary is also conical with a relatively large expansion angle. This flame structure notably differs from that observed when one of the propellants is injected in a subcritical or transcritical state while the other is gaseous. An analysis of the relevant characteristic times suggests that under transcritical conditions the rate of combustion is mainly controlled by turbulent energy transfer to the propellants. This determines the mass fluxes from the dense regions to the lighter gaseous streams governing the rate of conversion into products.

A MEMS-based solid propellant microthruster with Au/Ti igniter

Abstract: A solid propellant microthruster with Au/Ti igniter is demonstrated as an improved micropropulsion system for microspacecraft. The new design provides the microthruster with a high degree of flexibility, maneuverability and integration. Single microthruster and microthruster arrays have been successfully fabricated using standard microfabrication technologies. The propellant combustion process in the micro-chamber of the microthruster has been visually observed. Initial tests employing gunpowder-based solid propellants, have produced 2.11 × 10−5 to 1.15 × 10−4 N s of total impulse and 2.68–14.65 s of specific impulse at sea level, and 3.52 × 10−5 to 2.22 × 10−4 N s of total impulse and 4.48–28.29 s of specific impulse in vacuum. The performance of the solid propellant microthruster with Au/Ti igniter is also compared with that of a solid propellant microthruster having a wire igniter.

Catalyzed combustion of hydrogen–oxygen in platinum tubes for micro-propulsion applications

Abstract: The present investigation addresses the need to understand the physics and chemistry involved in propellant combustion processes in micro-scale combustors for propulsion systems on micro-spacecraft. These spacecraft are planned to have a mass less than 50 kg with attitude control estimated to be in the 1–10 mN thrust class. Micro-propulsion devices behave differently than macro-scale devices because of the differences in magnitude of flow rates and heat transfer. Reducing the combustor size increases the relative surface area, increasing the heat loss, and as combustors are continuously reduced in size, they approach the quenching dimensions of the propellants. Combustors of this size are expected to significantly benefit from surface catalysis processes. A miniature flame tube apparatus is chosen for this study because microtubes can be easily fabricated from known catalyst materials, and their simplicity in geometry can be used in fundamental simulations for validation purposes. Experimentally, we investigated the role of catalytically active surfaces within 0.4 and 0.8 mm internal diameter microtubes, with special emphases on ignition processes in fuel rich gaseous hydrogen and gaseous oxygen. Calculations of flame thickness and reaction zone thickness predict that the diameters of our test apparatus are below the quenching diameter of the propellants in most atmospheric test conditions. The temperature and pressure rise in resistively heated platinum microtubes and the exit hydrogen concentration were used as an indication of exothermic reactions. Data on imposed heat flux/preheat temperature required to achieve ignition versus mass flow rate are presented. With a plug flow model, the experimental conditions were simulated with detailed gas-phase chemistry and surface kinetics. Computational results, in general, support the experimental findings.

New Kinetics for a Model of Heterogeneous Propellant Combustion

Abstract: In earlier work we describe an unsteady, three-dimensional, phase-coupled combustion code which, with the use of a random packing algorithm to construct model propellants, and the use of a homogenization strategy to account for unresolvably small propellant particles, can be used for the simulation of heterogeneous propellant combustion. This work uses a simple two-step kinetic model for ammonium perchlorate (AP)/hydroxyl-terminated polybutadiene (HTPB) combustion which fails to accurately predict variations in the burning rate with AP concentration for a homogenized AP/HTPB blend supporting a one-dimensional flame. Here we describe a three-step model, one which captures the three flames of the Beckstead-Derr-Price (BDP) combustion model, and show that kinetic parameters can be adopted so that one-dimensional AP burning rates and one-dimensional AP/HTBP blend burning rates can be correctly predicted. We discuss the stability of the underlying flame structures and highlight a difficulty that arises in these instability-prone systems when simple kinetic models are used to describe them. The combustion model, with the new kinetics, is used to reexamine the burning of random packs, and improved agreement with the experimental burning rates of Miller packs is demonstrated. We also reexamine the problem of sandwich-propellant combustion and investigate the trend in surface shape and burning-rate variations with pressure and binder width. These trends are compared with experimental results of Price. The sandwich configuration is used to measure the importance of the primary diffusion flame of the BDP model.

New Aluminum Agglomeration Models and Their Use in Solid-Propellant-Rocket Simularions

Abstract: Random packs of ammonium perchlorate and aluminum particles in fuel binder, of the kind used to mimic the morphology of heterogeneous propellants, define distributions of aluminum particles that can be used as the starting point of agglomeration studies. The goal is to predict the fraction of aluminum that agglomerates and the size distribution of the agglomerates. Three phenomenological models are described, each with one or two parameters that can be adjusted to fit experimental data, and a number of such fits are attempted. It is shown that the agglomeration models can be calibrated to match a wide variety of propellant outputs, as needed for the numerical simulation of rocket chamber flows with aluminum injection. Results for such flows are presented and provide information about the distribution of the aluminum droplets and of the alumina smoke particles that arise from its presence.

Application of Supercritical Carbon Dioxide in Energetic Materials Processes: A Review

Abstract: The preparation of micro- and nanostructured energetic materials has recently drawn considerable attention as a potential method that can be used to obtain energy release more rapidly than conventional materials. Formation of solid particles with well-defined properties (e.g., particle size, particle size distribution, particle shape) and free of solvent inclusions for production of energetic materials using compressed gases were studied. It is possible to process moderately solids, like energetic materials that are difficult to comminute due to their sensitivity to mechanical or thermal stress. The characteristics of compressed gases allow the variation of morphology of solid particles in a wide range. It is possible to produce crystalline particles with a small size and narrow size distribution without defects (i.e., free of solvent inclusions). In this paper, the different processes to form very fine particles of explosives and propellant with supercritical fluids are reviewed. The first method is the ...

Benchmark Wall Heat Flux Data for a GO2/GH2 Single Element Combustor

Abstract: Wall heat flux measurements in a 1.5 in. diameter circular cross-section rocket chamber for a uni-element shear coaxial injector element operating on gaseous oxygen (GOz)/gaseous hydrogen (GH,) propellants are presented. The wall heat flux measurements were made using arrays of Gardon type heat flux gauges and coaxial thermocouple instrumentation. Wall heat flux measurements were made for two cases. For the first case, GOZ/GHz oxidizer-rich (O/F=l65) and fuel-rich preburners (O/F=1.09) integrated with the main chamber were utilized to provide vitiated hot fuel and oxidizer to the study shear coaxial injector element. For the second case, the preburners were removed and ambient temperature gaseous oxygen/gaseous hydrogen propellants were supplied to the study injector. Experiments were conducted at four chamber pressures of 750, 600, 450 and 300psia for each case. The overall mixture ratio for the preburner case was 6.6, whereas for the ambient propellant case, the mixture ratio was 6.0. Total propellant flow was nominally 0.27-0.29 Ibm/s for the 750 psia case with flowrates scaled down linearly for lower chamber pressures. The axial heat flux profile results for both the preburner and ambient propellant cases show peak heat flux levels a t axial locations between 2.0 and 3.0 in. from the injector face. The maximum heat flux level was about two times greater for the preburner case. This is attributed to the higher injector fuel-to-oxidizer momentum flux ratio that promotes mixing and higher initial propellant temperature for the preburner case which results in a shorter reaction zone. The axial heat flux profiles were also scaled with respect to the chamber pressure to the power 0.8. The results at the four chamber pressures for both cases collapsed to a single profile indicating that at least to first approximation, the basic fluid dynamic structures in the flow field are pressure independent as long as the chamber/njector/nozzle geometry and injection velocities remain the same.

A parallel solution-adaptive scheme for multi-phase core flows in solid propellant rocket motors

Abstract: The development of a parallel adaptive mesh refinement (AMR) scheme is described for solving the governing equations for multi-phase (gas–particle) core flows in solid propellant rocket motors (SRMs) An Eulerian formulation is used to describe the coupled motion between the gas and particle phases A cell-centred upwind finite-volume discretization and the use of limited linear reconstruction, Riemann solver based flux functions for the gas and particle phases, and explicit multi-stage time-stepping allows for high solution accuracy and computational robustness A Riemann problem is formulated for prescribing boundary data at the burning surface and a mesh adjustment algorithm has been implemented to adjust the multi-block quadrilateral mesh to the combustion interface A flexible block-based hierarchical data structure is used to facilitate automatic solution-directed mesh adaptation according to physics-based refinement criteria Efficient and scalable parallel implementations are achieved with domain

Development of a low-temperature co-fired ceramic solid propellant microthruster

Abstract: Low-temperature co-fired ceramic (LTCC) technology has been first successfully employed for the realization of a solid propellant microthruster. The microthruster has potential applications in microspacecraft as an excellent micropropulsion system for high-accuracy station keeping, attitude control, drag compensation and orbit adjust. The design, fabrication and experimental investigation of LTCC microthrusters are reported. Results from experiments on microcombustion, and thrust and impulse measurements both at sea level and in vacuum are presented. Initial tests employing gunpowder-based solid propellant have produced 3.81 × 10−5–1.27 × 10−4 N s of total impulse and 5.55–14.41 s of specific impulse at sea level, and 1.31 × 10−4–2.79 × 10−4 N s of total impulse and 19.05–31.55 s of specific impulse in vacuum. The performance of the LTCC solid propellant microthruster is also compared with that of a silicon-based solid propellant microthruster.

Numerical Modeling of Single-Pulse Gasdynamics and Performance of Pulse Detonation Rocket Engines

Abstract: Pulse detonation rocket engines (PDREs) offer potential performance improvements over conventional designs bu tr epresent a challenging modeling task. A quasi-one-dimensional, finite rate chemistry computational fluid dynamics model for PDREs is described and implemented. Four different PDRE geometries are evaluated in this work: a baseline detonation tube, a detonation tube with a straight extension, and a detonation tube with two types of converging/diverging (CD) nozzles. The effect of extension length and CD nozzle area ratio on the single-pulse gasdynamics and performance of a PDRE is studied over a wide range of blowdown pressure ratios (1‐1000). The results indicate that a CD nozzle is generally more effective than a straight extension in improving PDRE performance, particularly at higher pressure ratios. Additionally, the results show that the blowdown process of the CD nozzle systems could be beneficially cut off well before the pressure at the endwall reaches the ambient value. The single-pulse performance results are also compared to some recent experimental measurements as well as a steady-state rocket system using similar modeling assumptions.

Burning Mechanism of Aluminized Solid Rocket Propellants Based on Energetic Binders

Abstract: This paper reports results obtained from an experimental study of the combustion mechanism of aluminized propellants based on an energetic binder. The techniques used in this investigation include:

The Noble-Abel Equation of State: Thermodynamic Derivations for Ballistics Modelling

Abstract: : Accurate modeling of gun interior ballistics promotes more efficient gun and propelling charge design. In order to simulate interior ballistic flowfields, such models require a description of the thermodynamic behaviour of the propellant gas. The Noble-Abel equation provides a simple and reasonably accurate equation of state for propellant gases at the high densities and temperatures experienced in guns. Most computational fluid dynamics-based ballistics models, however, require additional thermodynamic functions which must be derived from the equation of state. This note presents the derivation of such thermodynamic functions for Noble-Abel gases. Although the derivations are geared toward the functional requirements of the commercial Fluent code, the results are equally applicable to all computational fluid dynamics solvers. Also presented is a brief numerical example for a typical propellant; highlighting the different thermodynamics of the Noble-Abel and ideal gas equations.

High Pressure Testing Of Composite Solid Rocket Propellant Mixtures: Burner Facility Characterization

Abstract: Much Research on composite solid propellants has been performed over the past few decades and much progress has been made, yet many of the fundamental processes are still unknown, and the development of new propellants remains highly empirical. Ways to enhance the performance of solid propellants for rocket and other applications continue to be explored experimentally, including the effects of various additives and the impact of fuel and oxidizer particle sizes on burning behavior. One established method to measure the burn rate of composite propellant mixtures in a controlled laboratory setting is to use a constant-volume bomb, or strand burner. To provide high-pressure burn rate data at pressures up to 360 atm, the authors have installed and characterized a new strand burner facility at the University of Central Florida. Details on the new facility and the measurement procedures are summarized. Repeatability between different batches of the same mixture has been demonstrated to be very good, and two common HTPB/Ammonium Perchlorate (AP) propellant mixtures containing 7/3 and 5/5 bimodal AP distributions were tested in the burner. The resulting burn rates are compared to data from the literature with good agreement in burn rate exponent and the impact of changing from a 7/3 to a 5/5 AP split.

Flight Research of an Aerospike Nozzle Using High Power Solid Rockets

Abstract: Flight research has been conducted on an aerospike rocket nozzle using high power solid rockets. Two aerospike rockets and one conventional rocket were flown successfully to supersonic speeds, providing the first known set of transonic flight performance data for aerospike rockets. This paper describes the rockets, solid rocket motors, nozzles, and rocket instrumentation system. Flight test results are also discussed and compared with ground test results. Flight data show that all of the rockets successfully reached supersonic speeds with a maximum Mach number of 1.6 and a peak pressure altitude of nearly 30,000 ft. The aerospike nozzle efficiency was determined to be 0.96 from computational fluid dynamics (CFD) analysis. The rocket chamber pressures and thrusts of the aerospike rocket motors were lower than the conventional rocket motors. Because the same propellant formulation was used in all of the rocket motors, the discrepancy in pressure and thrust was most likely caused by a larger actual aerospike nozzle throat area than the designed throat area. Potential causes for the larger aerospike nozzle throat area are also discussed.

Flow Separation in Rocket Nozzles, a Simple Criteria

Abstract: Cold and hot flow tests were conducted to investigate the flow separation in rocket nozzles. The results are presented. A separatio n data base including a wide range of literature data is established to evaluate the influence of propellant combination and nozzle design on flow separation. As a result a simple separation criteria is suggested. Nomenclature p = pressure cf = friction coefficient Ma = wall Mach number Mades = design Mach number κ = adiabatic exponent θ = deflection angle σ = oblique shock angle u,U = velocity δ = boundary layer thickness δ *

Thrust Oscillations in Reduced Scale Solid Rocket Motors Part I : Experimental Investigations

Abstract: The subscale motors are very relevant tools to study pressure oscillations and to understand the vortex shedding induced by instabilities observed in large segmented solid rocket motors such as the Ariane 5 P230 motor. The present work describes an experimental firing test campaign with different set-ups (LP6, LP9, LP10), carried out at the Fauga-Mauzac Propulsion Laboratory. It focuses on the effects of different propellant geometries and of the nozzle position. Comparisons between full-scale and different subscale motors tests are presented to conclude the paper.

Energetics of Propellant Options for High-Power Hall Thrusters

Abstract: Krypton, cadmium, iodine, cesium, mercury, and bismuth are compared with xenon in several areas of performance, including thrust, specific impulse, probability of ionization, maximum theoretical efficiency, and sputter yield. The lighter propellants such as krypton and cadmium are favorable for high-Isp, low-thrust applications, whereas heavier propellants such as mercury and bismuth are preferable for low-Isp, high-thrust missions. Calculations of the ionizing collision rate show that cesium had the highest ionization probability (and lowest ionization energy), krypton had the lowest probability (and highest ionization energy), with the other propellant falling between cesium and krypton. Sputter erosion calculations show that for a carbon surface, heavier atoms will sputter less at low ion energies (less than 2000 eV) than light atoms, and will sputter much less on a kilograms-per-kilogram basis.

The prediction of thermal stability of self-reactive chemicals

Abstract: An advanced study on the thermal behaviour of double base (boost and sustain propellant) rocket motor used in a ground to air missile has been carried out by differential scanning calorimetry (DSC). The presence of two propellants as well as the different experimental conditions (open vs. closed crucibles) influence the relative thermal stability of the energetic materials. Several methods have been presented for predictions of the reaction progress of exothermic reactions under adiabatic conditions. However, because decomposition reactions usually have a multi-step nature, the accurate determination of the kinetic characteristics strongly influences the ability to correctly describe the progress of the reaction. For self-heating reactions, incorrect kinetic description of the process is usually the main source of serious errors for the determination of the time to maximum rate under adiabatic conditions (TMRad). It is hazardous to develop safety predictive models that are based on simplified kinetics determined by thermoanalytical methods. Applications of finite element analysis (FEA) and accurate kinetic description allow determination of the effect of scale, geometry, heat transfer, thermal conductivity and ambient temperature on the heat accumulation conditions. Due to limited thermal conductivity, a progressive temperature increase in the sample can easily take place resulting in a thermal explosion. Use of both, kinetics and FEA [1], enables the determination of the reaction progress and temperature profiles in storage containers. The reaction progress and temperature can be determined quantitatively at every point in time and in space. This information is essential for the design of containers of self-reactive chemicals, cooling systems and the measures to be taken in the event of a cooling failure.

Stable pharmaceutical solution formulations for pressurized metered dose inhalers

Abstract: An aerosol solution composition for use in an aerosol inhaler comprises an active material, a propellant containing a hydrofluoroalkane, a cosolvent and optionally a low volatility component to increase the mass median aerodynamic diameter (MMAD) of the aerosol particles on actuation of the inhaler. The composition is stabilized by using a small amount of mineral acid and a suitable can having part or all of its internal metallic surfaces made of stainless steel, anodized aluminium or lined with an inert organic coating.

Modeling of Propellant Tank Pressurization

Abstract: Accurate modeling of propellant tank pressurization is an essential element of the prediction of rocket performance. This is the case even more so for hybrid rockets that use a self-pressurizing oxidizer because the thrust produced by the motor is dependent on oxidizer tank pressure. Described herein is a real fluid, quasi-phase equilibrium thermodynamic propellant tank model that is applicable to a broad range of propellants including those that are highly volatile and/or cryogenic. The model has been validated by comparison with the measured results of cold-flow blow-down tests of a self-pressurizing nitrous oxide system. Data from one of the cold-flow tests showed evidence of stratification within the oxidizer tank during the blow down. It was found that accurate modeling of the propellant evaporation rate is a crucial element in the prediction of the propellant expulsion process.

The microwave electro-thermal (MET) thruster using water vapor propellant

Abstract: The research to develop the microwave electro-thermal (MET) thruster at Research Support Instruments, Inc. (RSI) using a variety of gases as fuel is described. The MET has undergone dramatic evolution since its first inception, and it is now moving toward flight development. The MET uses an electrodeless, vortex-stabilized microwave discharge to superheat gas for propulsion. In its simplest design, the MET uses a directly driven resonant cavity empty of anything except gaseous propellant and the microwave fields that heat it. It is a robust, simple, inexpensive thruster with high efficiency, and has been scaled successfully to operate at 100 W, 1 kW, and 50 kW using 7.5-, 2.45-, and 0.915-GHz microwaves respectively. The 50-KW, 0.915-GHz test was perhaps the highest power demonstration of any steady-state Electric thruster. The MET can use a variety of gases for fuel but the use of water vapor has been shown to give superior performance, with a measured specific impulse (I/sub sp/) of greater than 800 s. When this added to the safety, ease of storage and transfer, and wide availability of water in space, the potential exists for using a water-fueled MET as the core propulsion system for refuelable space platforms.

Optimal Design and Control of Hybrid Rockets for Access to Space

Abstract: A nested direct/indirect method is used to find the optimal design of hybrid rockets for suborbital missions. The direct optimization of the parameters which affect the engine design is coupled with the trajectory indirect optimization. Different propellant combinations are analyzed. First, the simplest blowdown feed system is considered. A more complex pressurization system with an additional gas tank, which allows a phase with constant propellant tank pressure, is then analyzed. The optimization procedure provides the engine design and the trajectory, which maximize the mission performance index.

Space depot for spacecraft resupply

Abstract: A propellant depot ( 40, 150 ) includes a utility box ( 42, 42 ′) that has space flight equipment. A propellant cartridge adaptor ( 95 ) is coupled to the utility box ( 42, 42 ′) and to an exchangeable propellant cartridge system ( 41 ). The propellant depot ( 40, 150 ) also includes a docking adaptor ( 44 ) for coupling to an approaching spacecraft ( 24 ). A controller ( 66 ) controls the transfer of propellant from within the exchangeable propellant cartridge system ( 41 ) to the spacecraft ( 24 ). A method of providing propellant to a spacecraft in space includes launching an orbital propellant depot ( 40, 150 ) into space. The spacecraft is docked to the orbital propellant depot ( 40, 150 ) in space. Propellant is transferred to the spacecraft. The spacecraft is separated from the orbital propellant depot ( 40, 150 ).

Device for applying cryogenic composition and method of using same

Abstract: A device of the present invention for applying a cryogenic composition includes a machining tool (20) or tool holder having a channel (38) positioned therethrough and a capillary tube (44) positioned within the channel (38). A dense cryogenic fluid is passed through the capillary tube (44) while a diluent or propellant fluid is passed through the channel. The diluent or propellant fluid flows within the channel (38) and about the capillary tube (44). Upon exiting the capillary tube, the dense fluid admixes with the diluent or propellant fluid to form a cryogenic composite fluid or spray (24). The cryogenic composite fluid or spray is selectively directed onto a substrate (22) for cooling, cleaning or lubrication purposes, or onto the machining tool (20) for cooling purposes.

Atmospheric Mining in the Outer Solar System

Abstract: Atmospheric mining in the outer solar system has been investigated as a means of fuel production for high energy propulsion and power. Fusion fuels such as Helium 3 (3He) and hydrogen can be wrested from the atmospheres of Uranus and Neptune and either returned to Earth or used in-situ for energy production. Helium 3 and hydrogen (deuterium, etc.) were the primary gases of interest with hydrogen being the primary propellant for nuclear thermal solid core and gas core rocket-based atmospheric flight. A series of analyses were undertaken to investigate resource capturing aspects of atmospheric mining in the outer solar system. This included the gas capturing rate, storage options, and different methods of direct use of the captured gases. Additional supporting analyses were conducted to illuminate vehicle sizing and orbital transportation issues. While capturing 3He, large amounts of hydrogen and 4He are produced. With these two additional gases, the potential for fueling small and large fleets of additional exploration and exploitation vehicles exists. Additional aerospacecraft or other aerial vehicles (UAVs, balloons, rockets, etc.) could fly through the outer planet atmospheres, for global weather observations, localized storm or other disturbance investigations, wind speed measurements, polar observations, etc. Deep-diving aircraft (built with the strength to withstand many atmospheres of pressure) powered by the excess hydrogen or helium 4 may be designed to probe the higher density regions of the gas giants. Outer planet atmospheric properties, atmospheric storm data, and mission planning for future outer planet UAVs are presented.

Engine heat exchanger with thermoelectric generation

Abstract: A vehicle has a body and a source of a propellant. An engine is carried by the body. The engine reacts the propellant to produce thrust. The engine has a heat exchanger transferring heat from the reaction to at least a component of the propellant and generating electricity thermoelectrically.