Patent Number: 048256471
Section: summary

FIELD OF THE INVENTION This invention relates to thrusters intended to be used for low-power applications such as in the orbital positioning of spacecraft. The thruster disclosed is of the type using propellants such as hydrazine or hydrogen in which the propellant is heated to a desired temperature prior to exiting through a rocket propulsion nozzle. The heating provides a high specific impulse and facilitates decomposition and/or combustion. These thrusters are normally used during the lifetime of a three axis stabilized or spin stabilized satellite (presently 8-10 years for synchronous orbit satellites) in order to place in, to change or to maintain orbit station. Satellite on-board propulsion is frequently required to finalize and, in some instances, make major corrections to achieve final orbit circularization and/or orbit station. When this is accomplished with a typical hydrazine-fueled engine, large quantities of propellant may be expended. Use of a performance augmented engine (using electric energy to extend the nominal chemical reaction performance level) for this function would conserve and retain more fuel for on-orbit functions. Typically, excess electric power is available on a spacecraft even during orbit/station insertion maneuvers. Correction firings are time-spaced, with off periods between firings so as to permit battery recharge for subsequent firings. By this augmentation process, fuel usage can be reduced by as much as 30 percent or more. The thruster may also be used for correcting a satellite orbit which has decayed, or to maintain the orbit of a satellite which experiences some significant atmosphere drag or for repositioning the satellite to another location or station. Such thrusters can also be used for propelling satellites which follow other satellites or for evading tracking satellites. Another application of the performance-enhanced engine is to change the orbital path of a satellite in order to evade ground tracking or to make ground tracking more difficult. An application of this would be in satellite maneuvering for the sole purpose of decoying or saturating would-be tracking capabilities. In usage, this engine could be ground-controlled by the spacecraft operating agency, or in some instances of covert operation, may be preprogrammed for on-orbit automatic control. BRIEF DESCRIPTION OF THE PRIOR ART Liquid propellant fueled spacecraft engines operate at performance levels limited by the chemical reaction energy of the propellant. Performance is generally maximum for steady state operating periods of more than a minute and reduced for pulsing operation. For a monopropellant fueled engine using a propellant such as hydrazine, either a catalyst bed or a thermal decomposer is used to initiate the exothermic reaction process. Of these processes, the catalyst bed is the most common. The thermal decomposer is typically brought to operating temperature by means of an electrical resistance heater. There decomposers serve only to initiate the chemical reaction, but do not add to or augment the chemical performance level. To extend the performance level, investigators have suggested use of electrical resistance heaters to boost the chemical performance level by exposing the chemically reacted or reacting propellant to a high temperature heater, thereby increasing the propellant temperature prior to expansion of the propellant through a nozzle. The inherent problem of such a device is the fact that direct contact between the heater and the flow requires that the propellant be as free as possible of contaminants (standards in excess of those typical of most rocket engine usage specification levels). This also precludes use of the more complex propellant formulations such as any that would contain carbon or oxygen, due to possible chemical interactions with the heating element. In the prior art, the heaters were designed for maximum output during propellant contact. Without heat removal by the propellant the heater would attain excessive temperature and heater burn-out could occur. Accordingly, the power could only be switched on when propellant was flowing, and this meant that the successful transfer of energy from the heating filament to the propellant could be accomplished only when a sufficient amount of electric power was available for heating the propellant at the rate that the propellant was being utilized for thrust. This not only prevented operation during times when battery power was substantially low, but also precluded preheating the thruster with the heater before propellant flow was initiated. This also places limitations upon attainable temperatures. Such an engine cannot be off-flow modulated or pulsed with off periods greater than a few milliseconds and as such is limited in its usefulness. A further characteristic of flow-coupled devices is that the heater is subjected to all pressure fluctuations of the propellant supply and reaction process. Gas dynamic forces from any propellant reaction instability will be transmitted to the heater and may cause a heater distortion failure. Since the flow is circulating through the heater region there is constant flow impacting and washing of the heater. Further, the heat transfer area is limited by the finite surface area of the heater. This prior art design also requires use of high temperature sealed electrical feed-through(s) into the chamber. This places restrictions on the overall engine design as to operating temperature and power. The prior art thrusters used an outer shield having a low emissivity surface in order to reflect as much heat as possible back to within the heating portion of the thruster. This minimized energy losses by maintaining a higher temperature within the heating section of the thruster. Because the minimization of heat loss was accomplished in part by minimizing exterior surface cooling, the exterior tended to remain hot causing heat to be transferred through the thruster's supporting structure to the satellite proper. A further disadvantage of having a high exterior temperature was that infrared sensors could easily distinguish a warm satellite's components from the surrounding space. The rocket nozzle section of thrusters also presented a source of high temperature emissions. This resulted from the high temperatures present at the nozzle's throat and internal expansion chamber areas, which high temperatures were conducted as heat to the outer portions of the rocket nozzle. In prior augmentation designs, the liquid propellant is first decomposed, vaporized and reacted in an uninsulated, thermally separate chamber allowing some of the reaction energy to be lost. Nozzle expansion area ratios of most rocket engines including prior augmentation thrusters are characteristically several hundred or less. Rocket engine test facilities characteristically have limited capability to simulate a space environment, therefore actual spaceflight is typically required to fully evaluate nozzle performance for large expansion ratio nozzles of engines at thrust levels greater than a few tenths of pounds thrust. Ground test data as it exists for these engines due primarily to the inadequacy of space simulation in the test facilities suggests ineffectiveness for expansion ratio nozzles of greater than several hundred. Further, state-of-the-art analytical projections are not definitive as "universal agreement does not exist regarding the correct procedures and assumptions for calculating the propulsive performance for nozzles with these high area ratios," as stated in the following reference: Cooper, L. P. (NASA LeRC), Advanced Propulsion Concepts for Orbital Transfer Vehicles, AIAA paper-83-1243, June 1983. BRIEF DISCUSSION OF THE STATE OF THE ART IMMEDIATELY PRECEDING THE PRESENT INVENTION Use of the electrical power supply of a space vehicle to augment and/or to induce propellant dissociation can result in achievement of more thrust per unit mass of gas as the gas temperature is raised to increasingly higher values. Since satellite launch capabilities limit the mass of material that can be carried as propellant, the higher the temperature of the propellant outflow, generally the longer the useful lifetime of the space vehicle. For communication, navigation, weather or surveillance satellites, space stations, space platforms and space probes, great gains can usually be obtained by increasing the stagnation temperature of the propellant flow. In one aspect of the invention disclosed in parent application Ser. No. 517,265, a thruster is provided which permits propellant to be heated without directly contacting a heater filament. The heater may have single or multiple elements to permit operation at different power levels and/or to have element redundancy. It is a further objective to increase the efficiency of heat transfer from the heater element by increasing the ratio of transfer of thermal energy to the propellant over thermal energy loss. In a further aspect thereof, an increase in propulsion performance is provided by permitting significant amounts of thermal energy to be stored within a heat exchanger for fractions of seconds or for considerable periods of time, such as several minutes. This enables the thruster to operate with reduced amounts of electrical power when necessary or advantageous. This thermal capacity of the heat exchanger also permits the engine to operate in a periodically modulated (interrupted) flow mode with constant heater power to accommodate specific flight operations or circumstances for either balancing or unbalancing the thrust vector. Typically, control engines are operated as matched thrust level pairs mounted on a vehicle to provide parallel thrust vectors which when summed together provide a resultant vector that would generally extend through the center of mass of the vehicle. In the event of an unbalanced disturbance torque on the vehicle or of inadvertent single engine performance degradation, i.e., a fractional reduction of thrust pressure in one engine, the opposite engine could be flow modulated at a rate that would maintain a thrust vector sum that would have the desired orientation. The option is also available to input a torque into the vehicle by this means. For a spinning spacecraft with engines located on either side of the spin axis if one engine were to fail, the second, if flow modulated during each revolution, could maintain a desired thrust vector. In a further aspect of the invention disclosed in the parent application, a thruster is provided which may be operated at maximum efficiency by off-modulating either its heater filament or the propellant as necessary to match thruster pairs and/or to achieve an optimum performance balance under typical spacecraft conditions of a reducing propellant flow rate due to blow down of the propellant tank pressurant over mission life and changing spacecraft power supply due to a let down of battery voltage during a firing sequence and/or power supply capability degradation over mission life. In a further aspect of the invention disclosed in the parent application, a thruster assembly is provided which is efficient in transfer of heat energy from an electrical heating element by means of effective utilization of heat shielding. It is a further object to provide such a thruster which, despite an effective means of maintaining heat in the thruster, maintains a relatively cool exterior surface and thereby presents a cool attachment point for a supporting structure. In a yet further aspect of the invention disclosed in the parent application, a thruster is provided which, despite maintaining heat within the thruster, presents a cool exterior surface which is difficult to track with infrared scanning devices prior to actual ignition of the thruster, thereby decreasing the possibility that an enemy could detect an intended ignition of the thruster, even though the thruster may have a pre-heat capability. In a further aspect, a thruster is provided which has as an option to transfer as little heat as possible from its nozzle throat section to the outside portions of its nozzle. In a further aspect, a thruster is provided which operates with a cool exterior surface so as to make the thruster more difficult to track with infrared sensing devices when the thruster is expelling propellant into space. In one further more specific aspect of the invention disclosed in the parent application, a thruster assembly is provided with a heating filament located within a heater cavity that communicates to space and which has a propellant guiding structure surrounding the heater cavity. This propellant guiding structure provides for the propellant inflow (injection) to occur at a location near the engine supporting structure, by design the coolest zone of the heat exchanger, and the propellant is then guided or channeled to flow through the inlet to the heat exchanger to the hottest zone which is adjacent to the nozzle throat. In this manner, the flow is heated by acquiring some of the heat that would otherwise be lost from the heat exchanger due to conduction into the supporting structure. This regenerative heating of the propellant both increases the efficiency of the heat exchanger and helps achieve higher propulsive performances for a given amount of available electric power. The heater located within the heater cavity of the invention disclosed in the parent application can be assembled and tested separately from the heat exchanger. This modular construction feature also permits the heat exchanger to be assembled and tested separately from the heater, using a test heater or heater simulator. For flight applications, this permits operational testing of a flight heater in a non-degrading environment typical of the used for preflight checkout of rocket engines or spacecraft. A flight heater may be joined with the flight heat exchanger for a preflight vibration test, then removed and replaced with a ground-test-only heater (1) for preflight validation of the heat exchanger-nozzle and engine operating characteristics and (2) calibration. Subsequently the flight heater which has been checked out and calibrated in a separate test series in the non-degrading environment is reinstalled into the heat exchanger/engine in readiness for flight. Another feature of this arrangement permits a heater replacement, if desired, subsequent to engine installation on a vehicle or in a test facility without needing to remove the heat exchanger/engine. The heater disclosed in the parent application may be comprised of one or more radiating heating elements or a combination of heaters and/or a thermionic emitter. A preferred configuration if a thermionic emitter is used would be to energize a heater to bring the cathode emitter to emission temperature. An embodiment featuring use of a cathode emitter to transfer energy to the heat exchanger requires the heater assembly to be electrically insulated from the heat exchanger and the heat exchanger would then function as an anode to receive the electron transfer from the cathode. The heater filament disclosed in the parent application may have a number of configuration options as to shape, spacing and material selection. The heater may include one or more heating elements. Multiple elements may provide heater redundancy and/or the capability to operate at one or more power levels. The heaters may be free standing (self-supporting) or may be provided with additional electrically insulated mechanical supporting structures. The heater material and size are selected to provide an energy transfer capability to match or nearly as is feasible the spacecraft power supply capability making minimum use of additional power controllers and/or voltage-current regulators. The radiating heater materials will be made primarily of tungsten. Additional materials and processing are used with the tungsten to obtain specific predetermined characteristics. Three percent rhenium is added thereto to create the alloy W3Re to provide (1) ease in forming the element and (2) high vibration resistance. Selected trace elements and processing with the tungsten (without 3% rhenium) are used to make a high temperature resistant (in excess of 1925 degrees K.) wire more "sag" or droop resistant in the presence of gravitational and/or centrifugal force fields than would otherwise be attainable. This type of material combination and processing is typical of that used for filaments in aircraft landing lights. Application of this same type of filament material for the radiating heater(s) in the thruster provides a heater that can operate in a gravitational and/or centrifugal force field with less "sag" or deformation than would typically occur with a W3Re filament. This "sag" resistant wire permits extended periods, in excess of 100 hours, of high-g flight time or ground (one g) test time without resorting to heater rotation at rates of one rotation or more per minute; (such rotation rates are required to prevent "sag" of a high temperature self-supporting W3Re heater filament in a typical thruster configuration.) This "sag" resistant wire makes it possible to use a radiating high temperature filament on a spin stabilized spacecraft (characteristically having a rotation rate of 40 to 80 revolutions pr minute) with the engine being mounted away from the spin axis and exposed to centrifugal forces of 2 to 6 g's. A further aspect of the invention disclosed in the parent application is the option of sealing the heater cavity containing a non-reactive gas, such as nitrogen, to enable gas pressurization of the filament. This pressurization will reduce heater filament vaporization rates. Conduction through the gas and gas convection induced by a "g" field will also transfer significant amounts of power from the heater element to the heat exchanger, resulting in a lowered temperature (as much as 220 degrees K. lower) of the coil for the transfer of a given amount of power. This combination of a reduced evaporation rate and a lower coil temperature to transfer a given power from the coil to the heat exchanger can increase the lifetime of a coil by over one order of magnitude, e.g. from 60 hours to over 600 hours. Pressurant gas dynamic forces in the heater cavity may also be used to counteract distorting g forces. That is, the heater filament may be configured in relationship to the cavity so as to interact with the pressurant gas to cause a gas convection force to oppose the "sag" forces. The heating filament may be switched "on" for significant periods of time when propellant is not flowing through the passageway and a heat-sinking capability of the propellant guiding structure permits heat to be transferred to propellant when the filament is switched "off". The propellant guiding structure may be formed in multiple layers to provide plural thermal zones of increasing temperature for the propellant as the propellant is passed through the structure. In order to retain heat within the structure, the shields will be separated by means of physical indention or preformed to specific configurations with thermal processing. Multiple radiation shields may be used internally within the heat exchanger, surrounding and at the base of the heater assembly and external of the heat exchanger. While interior shields have low emissivity in order to reflect and hold heat inwardly, the exterior surface of the thruster may have a coating having a high emissivity in order to present as cool an exterior surface as possible. The tendency of the exterior surface to remain cool by emitting heat enables operation of the thruster with higher internal temperatures, hence more efficiently. The heat exchanger/engine supporting structure, typically designated in the art as a barrier tube, connects and mechanically couples the engine to the spacecraft mount. The preferred embodiment of this barrier tube as disclosed in the parent application, uses a thin tube of extended length, with material cut-outs, formed of a low thermal conductivity material such a titanium to minimize the heat loss through this thermal conductivity path. This extended length barrier tube may be configured as concentric cylinders connected at alternating ends to minimize packaging volume with acceptable engine structural support to meet typical spacecraft launch vibration load requirements. The heating filament and the interior surface of the heater cavity as disclosed in the parent application, may also be provided with high emissivity coatings by means of surface treatment and/or coatings in order to promote a rapid transfer of energy from the heating filament to the materials surrounding the heater cavity with minimum temperature differentials between the wire and the cavity. The heating filament may operate in either a vacuum environment or may be pressurized with an inert or non-reactive gas or with reacted propellant gases in order to prolong the life of the filament. Reacted and/or energized propellant gases may be introduced into the heater cavity directly from a heat exchanger bleed for moderate level pressurization, (40 to 150 psia), or from a bleed from the expansion nozzle wall for less than one atmosphere (as low as 10.sup.-3 psia) pressurization. In that the cavity would be moderately well sealed (low leak of several cubic centimeters of gas per hour or less permitted), the gas is essentially stagnant in the absence of a "g" field. No significant measure of propellant is lost during this pressurization process. The filament itself may be provided with a bifilar helix configuration. In this mode, electromagnetic forces resulting from current flowing through each filament half will cause the filament to maintain a desired central position relative to the other, thereby axially stabilizing the filament when it is hot. The construction is such that the fuel passageways are formed as helix threads or as grooved passageways extending in one or more plural layers along the length of the thruster housing coaxial with the heater cavity. The concentric relationship of the fuel passageways and associated structure, including the shield, permits the thruster to be assembled with a minimum of weldments or other fastening devices. The thruster assembly as disclosed in the parent application may be provided with an injection passage such that the propellant can be introduced as a liquid and heat from the performance augmentation section will thermally decompose it without the use of a dissociation catalyst. The fluid passageways may be coated or plated with a material that is resistant to chemical interaction or, when desired, to enhance the dissociation process of the propellant, permitting use of less costly materials for the passageway such as TZM molybdenum alloy. In a further aspect of the invention as disclosed in the parent application, a thruster assembly such as described above can be formed with a nozzle having a nozzle throat insert. The nozzle throat insert has a high temperature capability, whereas the remainder of the expansion area of the nozzle is not required to have the same high temperature properties. The insert construction also provides a means to reduce thermal emissions from the thruster's nozzle expansion portion. While the thruster assembly disclosed in parent application Ser. No. 517,265 has met with commercial success, potential for improvement exists in its structure and operation in the following particulars: (a) Firstly, the thruster assembly disclosed in the parent application decomposes hydrazine in a typical rocket engine catalytic bed in a separate chamber and feeds the reaction products into the thruster through a feed tube. This technique, however, is found to result in the loss of a significant percentage (up to 30%) of the hydrazine decomposition chemical energy from the gas flow while retaining high fractions of undissociated ammonia. For separate reasons then, these approaches to propellant (hydrazine) decomposition and injection have disadvantages. (b) The heat exchanger surrounding the heater coil will be typically made from Mo4ORe bar stock. This material is presently available in diameters of 11/8 inch or less, and at the maximum size there is some porosity in the material within the inner 1/2 inch core. This material costs about $2,700/kg (1983) and most of the purchased weight has to be machined out and discarded. This present limit at to maximum diameter available of Mo4ORe imposes restrictions on engine thrust level due to heat exchange surface area limitations. This material size restriction also limits the amount of power that can be transferred by radiation to the limited diameter cavity walls, as there are packaging, operating condition and lifetime limits of current heater technology, as demonstrated in FIG. 22. The power radiated from a heater coil with a fixed major helix diameter can be increased somewhat higher than 750 watts and still maintain adequate life by any or all of the following techniques: (i) Enhance the emissivity of the coil by surface roughing or by using a surface coating such as hafnium carbide (maximum increase of power transfer is about a factor of four). (ii) Pressurize the cavity to reduce the surface loss rate and permit the coil to operate at a higher temperature and thus radiate more power-per-unit surface area (see FIG. 27) (maximum increase of power transfer is about a factor of two and one-half). (This feature would also permit some power to be conducted from the heater to the heat exchanger wall by convection through the gas so that the total power increase factor by means of pressurization could be three or three and one-half.) (iii) Lengthen the cavity and simultaneously lengthen the coil. This results in longer unsupported heater coil lengths and weights, and thereby approaches a maximum length limit where the coil will eventually fail either in vibration or in "creep" or "squirm". An estimate on the potential power increase available by lengthening is a factor of about two. (c) Stresses in the heat exchanger section surrounding the inner heater coil, caused by the internal gas pressure, lead to creep rupture distortion and/or failure in lifetimes up to about 500 hours. If wall thicknesses are increased to extend lifetimes, undesirable weight is added. To help minimize the material stresses and wall thickness, the internal pressure is reduced by external pressure drop mechanisms from a typical propellant or fluid supply pressure range of 350-100 psia to values ranging from 100-40 psia over life. This reduces the Reynold's number of the flow through the nozzle and consequently results in a lower thrust coefficient of the nozzle. Hence, the specific impulse of the augmenter is lower than what it could be by utilizing the full pressure available. (d) In order to reduce energy loss out the open end of the heat exchanger cavity, three short (one-inch length) radiation shields (FIG. 16) are placed between the coil and the cavity surface. All of the cavity can be emissivity-enhanced by roughening or by thermo-chemical treatment of or plating the surface. Such processing can increase the effective emissivity by several factors. These arrangements maximize the amount of heater radiation absorbed by the cavity. However, as the surface area of the cavity and the surface emissivity (enhanced or not) is uniform over the cavity length, the radiative heat transfer into the wall is relatively uniform. This arrangement did not permit any significant concentration of radiation transfer near the engine nozzle throat inlet section thereby limiting the peak specific impulse (propellant temperature) operating conditions. (e) Thermal shielding of the heat exchanger components can effectively be accomplished only by using radiation shielding. In existing designs, scrolls of molybdenum foil are used, together with some discs. Both are usually spaced or separated by surface indentations and preforming or by wires. If the ratio of the contact area to the shield areas exceeds 10.sup.-5, the effectiveness of the shields is significantly reduced. It is difficult to obtain an area ratio lower than 10.sup.-4 by using surface indentations and preforming or wire separations. Good shielding at the nozzle end is extremely important to ensure that the heat exchanger in this region attains its highest possible temperature value. If gas from the nozzle exit is permitted to expand and enter the outer shield cover and surrounding the shields, the shields' effectiveness is reduced. In this regard, reference is made to Tables 1A and 1B. If the shield cover is not effectively sealed to the nozzle exit, then the radiation shields will have degraded performance. TABLE 1A ______________________________________ POWER LOST IN WATTS FROM RADIATION SHIELDED HEAT EXCHANGER AS FUNCTION OF AMBIENT PRESSURE AND SHIELD CONTACT SURFACE RATIO A.sub.c /A P(Torr) 10.sup.-3 10.sup.-4 10.sup.-5 10.sup.-6 10.sup.-7 ______________________________________ 1.0E 01 296.86 269.29 266.16 265.84 265.88 16 1.0E 00 227.41 174.28 167.47 166.70 100.73 SHIELDS 1.0E-01 189.87 108.53 95.52 94.08 93.96 T.c = 1900 1.0E-02 184.06 95.45 78.63 76.47 76.25 deg K. = 1.0E-03 183.46 93.94 76.32 73.96 73.71 (2961 1.0E-04 183.40 93.77 76.07 73.69 73.44 deg F.) 1.0E-05 183.40 93.76 76.05 73.66 73.42 .010 in. Spacing 1.0E 01 291.03 259.50 255.88 255.52 255.53 22 1.0E 00 215.77 152.50 144.07 143.20 143.14 SHIELDS 1.0E-01 183.98 92.66 77.45 75.74 75.59 T.c = 1900 1.0E-02 179.52 82.11 63.48 61.00 60.74 deg K. = 1.0E-03 179.05 80.93 61.62 58.92 58.63 (2961 1.0E-04 179.00 80.81 61.43 58.70 58.41 deg F.) .008 in. Spacing 1.0E 01 305.64 282.26 279.64 279.39 279.4 10 1.0E 00 249.37 209.85 205.06 204.58 204.56 SHIELDS 1.0E-01 203.30 138.12 128.38 127.33 127.25 T.c = 1900 1.0E-02 194.83 120.10 106.45 104.83 104.65 deg K. = 1.0E-03 193.87 117.83 103.33 101.55 101.37 (2961 1.0E-04 193.77 117.60 103.00 101.20 101.02 deg F.) 1.0E-05 193.76 117.57 102.97 101.17 100.99 .018 in. 1.0E-06 193.76 117.57 102.96 101.17 100.99 Spacing 1.0E-07 193.76 117.57 102.96 101.17 100.99 1.0E-08 193.76 117.57 102.96 101.17 100.99 ______________________________________ O.D. 1.5 L = 2.75 TABLE 1B ______________________________________ POWER LOST IN WATTS FROM RADIATION SHIELDED HEAT EXCHANGER AS FUNCTION OF AMBIENT PRES- SURE AND SHIELD CONTACT EFFECT OF COATING SHIELD WITH RHODIUM RADIATION SHIELD PERFORMANCE COMPARISON 16 Shields T.c = 1900 deg K. = (2961 deg F.) .010 in. Spacing A.sub.c /A P(Torr) 10.sup.-3 10.sup.-4 10.sup.-5 10.sup.-6 10.sup.-7 ______________________________________ No Coating 1.0E 01 300.90 274.05 271.00 270.70 270.73 1.0E 00 233.06 181.24 174.62 173.93 173.89 1.0E-01 196.54 117.93 105.68 104.34 104.23 1.0E-02 190.91 105.75 90.12 88.09 87.88 1.0E-03 190.32 104.36 87.98 85.75 85.52 1.0E-04 190.26 104.21 87.75 85.50 85.27 Rh Coating 1.0E 01 291.86 263.58 260.28 259.95 259.98 1.0E 00 219.31 162.27 154.90 154.14 154.09 1.0E-01 179.05 90.91 77.15 75.66 75.53 1.0E-02 172.80 77.12 60.64 58.71 58.51 1.0E-03 172.13 75.56 58.60 56.56 56.35 1.0E-04 172.07 75.40 58.38 56.34 56.12 ______________________________________ SUMMARY OF THE INVENTION The present invention improves upon the prior art as discussed hereinabove in the following particulars: (a) The inner and outer heat exchanger components are brazed together along the lands of the threads or the tops of the grooves thereof (FIG. 16). Braze material could be a material such as, for example, molybdenum or iridium. One method of several options to apply the braze material may be by vapor-deposition on the inner diameter of the outer heat exchanger component, the outer diameter of the inner heat exchanger component, or both. Another method of putting the braze in position is by layering between the parts a thin foil of the braze material. A third method is to provide a channel in the meshing parts which contains the braze material, i.e. in the form of a wire (see FIG. 46). The braze can be accomplished by any one of several methods: (i) Heating the whole structure in a vacuum furnace. (ii) Placing a heater assembly in the heat exchange cavity and heating in a vacuum environment (preferably ion-pumped to under 10.sup.-4 Torr). This is a preferential brazing technique, since the inner component will be somewhat hotter than the outer component. The extra thermal expansion of the inner component will close any gaps that may exist between the components over the surface that is to be brazed. Effecting this braze reduces the tension load on the outer component and the compression load on the inner component to negligible values so that the wall thickness (and weight) can be substantially reduced. Also, the creep-rupture problem is almost completely eliminated, since each flow passage is now equivalent to a tube and all of the metal becomes structural. This improvement will permit lower specific weight for the heat exchanger while giving simultaneously a life extension of over 500 hours as compared to an unbrazed structure. (b) Further, a coiled tube replaces the outer heat exchanger passage of the prior art configuration discussed hereinabove. This tube can be made of molybdenum-rhenium, rhenium or other high-temperature materials. Use of centrifugal flow passages for either the inner or outer heat exchanger passage and the resulting effects of the induced secondary flow which is generated into the flowing propellant tends to greatly increase the heat transfer rates over those usually obtainable in the straight flow channel. Also, the gas is heated more uniformly because of the mixing induced by the secondary flows. For better performance, the coiled tube can also have an inside coating of iridium, tungsten or rhenium. Iridium is an advantageous coating material since it would enhance catalyzation of the reaction of converting hydrazine (N.sub.2 H.sub.4) into only hydrogen (H.sub.2) and nitrogen (N.sub.2). It would also help transform any ammonia (NH.sub.3) that might be injected into the tube at the inlet, into hydrogen and nitrogen molecules and thus ensure that this would not occur in the hotter components where some damage to the material could result from the intermediate reaction products (e.g., N or N.sup.+) reacting with some component of the material. Useful thicknesses of the coating will be discussed hereinafter. Some of the advantages of using this injection tube heat exchanger component to duct the propellant (typically hydrazine products) into the inner heat exchanger are as follows: (i) The outer diameter of the "effective" heat exchanger is now not limited by the present manufacturing size limits of molybdenum-rhenium; (ii) The weight and cost of this component are much lower than they would be by machining it from bar stock, as compared even to present available bar stock; and (iii) The stresses are comparatively low in the tube, thus resulting in an extension of typical lifetimes. The further feature of coating the inner diameter of the tube will provide additional lifetime in excess of 500 hours. Several methods may be used to coat the inner diameter of the tube. One method is to vapor-deposit the coating by typical chemical vapor deposition (CVD) techniques. A second method is to vapor-deposit the coating from a wire strung on the axis through the tube. This concept and apparatus are considered to be an integral part of the invention and are described in detail hereinafter. (c) The power radiated from the heater coil is reflected from the inner shield into the preferentially placed optimum (high absorbent) energy absorber structure. In order to maximize this effect, the heater enclosing radiation shields internal to the cavity are lengthened to extend the full distance of the cavity instead of covering only the first approximately one-third as in prior designs. An energy absorber component is brazed or welded into the nozzle end of the cavity wall in order to maximize the power transfer into this region. Power is transferred to this energy absorber component by direct radiation from the heater coil and by reflection from the inner surface of one of the shields. By making the ratio of the gap length to gap spacing in the energy absorber component high enough, impinging photons will undergo enough multiple reflections in the energy absorber component to be absorbed, thus giving the energy absorber component an effective emissivity of unity. The cross-sectional area of each cylinder of the energy absorber component must be adequate to ensure that the temperature difference between the braze joint and the far edge of each cycle is less than 20.degree. C. with maximum energy flux. The energy absorber component may be fabricated from any of the following materials: tungsten, tungsten-rhenium, molybdenum, rhenium, or molybdenum-rhenium. It can be fabricated as a series of cylinders brazed to a disc, or as a scroll brazed to a disc. The combined advantage of the full length shields and the energy absorber component is to have over 50% of the power radiated from the heat coil transferred to the energy absorber component. In this way, the peak temperature of the nozzle inlet end of the heat exchanger can be raised over 200.degree. C. above what it might be without these improvements. (d) A nozzle inlet area heat exchanger component may also be brazed or welded into the cavity wall face opposite the energy absorber component. This nozzle inlet area component can be fabricated as a spiral, attached at both ends to a housing, or as a series of cylinders with gaps at alternate ends. The cross-sectional area of the metal needs to be large enough to ensure that axial temperature differences between the cavity end and the nozzle inlet area end are less than 20.degree. C. This component can be fabricated from any of the following materials: tungsten, rhenium, tungsten-rhenium, molybdenum or molybdenum-rhenium. The advantage of this component is that it will ensure that the gas temperature is heated to within 20.degree. C. of the peak temperature attained anywhere in the heat exchanger and that this peak heat exchanger temperature is adjacent to the inlet sonic orifice of the nozzle. Over one-half of the electrical power input can be transferred to the propellant on this nozzle heat exchanger component. (e) In order to obtain the best performance of the radiation shields, wherever possible they are made as disc-cylinder combinations. They may, if desired, be made of 0.001" thick tungsten foil formed and annealed into cylinders and welded to discs that are 0.005" or 0.010" thick tungsten. These shields can then be nested and held together with a means that mimimizes the contact area between successive shields. The advantage of this method of shield fabrication and support is that the contact-area to shield-area ratio can be kept to values of under 10.sup.-5, making the shield effectiveness close to the theoretical maximum as will be described hereinafter. Where appropriate, these shields can be coated with a low-emissivity metal to enhance their performance. The case and the nozzle exit can be joined together in such a manner as to eliminate propellant leakage or back flow to ensure that the pressure inside of the enclosure which holds the shields is always under 10.sup.-3 Torr at operational conditions. In order to help maintain this internal enclosure pressure low, the enclosure can also be vented at the end where the heater assembly is attached, at a location remote from the nozzle. (f) In order to additionally concentrate the energy flux into the absorber and nozzle heat exchanger, the heat exchanger components may be replaced by a support tube and a second coiled tube heat exchanger, as will be described in greater detail hereinafter. Since the support tube now only serves to support the structure of the energy absorber component and nozzle heat exchanger component, its cross-sectional area for heat conduction away from these components can be much smaller than that of the components described in (a) above (see FIG. 20). This will permit a further increase in the temperature of these components for a given input power and mass flow rate, thus further increasing the specific impulse. It will also reduce the weight and cost of the thruster. This tube may also be internally coated with tungsten or rhenium in the same manner as the coiled tube described hereinabove with reference to FIGS. 16 and 20. With this dual-coiled tube configuration, if the nozzle heat exchanger components, as well as the pressure vessel are made out of rhenium or tungsten-rhenium, operating temperatures can be increased to over 2475.degree. K. (4000.degree. F.). For hydrazine propellant this could allow the mission average specific impulse (I.sub.sp) to approach 340 seconds as explained by the graph in FIG. 21. In order to get this value of I.sub.sp, it may also be necessary to increase the number of shields in the cavity and also around the nozzle heat exchanger. (g) If more of the chemical power of hydrazine decomposition could be retained (and/or used for ammonia dissociation), considerable savings of electrical power would result, especially at the higher thrust levels. For example, at a thrust level of 0.50 lb.sub.f and a flow rate of 5.8 lb.sub.m /hr (I.sub.sp =310 sec), the loss of chemical power could be as high as 230 watts (with a gas temperature of 900.degree. F. at the injection point of fully decomposed hydrazine into the augmenter). The methods of saving most of this power involve adequately insulating the catalytic bed chamber with radiation shields to reduce the heat loss and provide a longer bed (more propellant dwell time and catalyst contact) to maximize ammonia dissociation. Radiation shielding may be accomplished using ultra-light, low-emissivity disc-cylinder radiation shields; preferably micro-arc welded thereto using, for example, the process and equipment described in U.S. Pat. No. 4,404,456 to Cann. This shielding may also be coated to achieve optimum emissivity properties in manner to be described hereinafter. As the augmenter design is improved to permit both higher operating power levels and increased propellant flow rate, an additional advantage can be realized. This higher flow rate through the injector into the decomposer will result in a higher level of absorption of energy before boiling. For example, if the flow rate is 7 lb/hr, by heating all of the liquid from 298.degree. K. to 398.degree. K., the liquid would absorb 206 watts. Additionally, at this flow rate, the liquid flow is likely to be turbulent, thus promoting mixing and better cooling capabilities by turbulent heat transfer from the tube to the fluid (see Table 2). TABLE 2 __________________________________________________________________________ HEAT TRANSFER AND HEAT ABSORPTION PROPERTIES OF LIQUID HYDRAZINE FLOWING THRU THE INJECTION TUBE UPSTREAM OF THE DECOMPOSER T .degree.C. T .degree.F. p.sup.psi .mu. c.sub.p K .rho. Pr ##STR1## ##STR2## w.sub.Lm/sec. w.sub.Lm/sec. P.sub.liqWatts P.sub.liqWatts __________________________________________________________________________ 0 32 5.80.sup.-2 1.28.sup.-3 3,650 .921 1,080 4.95 308 925 3.60 10.8 -5.6 -16.8 50 122 3.48 0.65.sup.-3 3,560 .850 980 2.72 607 1,822 3.96 11.9 5.6 16.8 100 212 9.28 0.40.sup.-3 3,770 .783 925 1.93 987 2,961 4.20 12.6 17.1 51.4 150 302 40.6 0.31.sup.-3 4,390 .703 882 1.94 1,270 3,820 4.41 13.2 30.0 90.0 200 392 149.0 0.24.sup.-3 6,490 .624 818 2.50 1,650 4,930 4.75 14.3 47.1 141.0 250 482 387.0 0.18.sup.-3 10,900 .544 755 3.61 2,193 6,580 5.15 15.4 74.5 228.0 300 572 890.0 0.12.sup.-3 24,300 .452 675 6.45 3,290 9,870 5.76 17.3 350 662 1,335.0 0.08.sup.-3 .368 535 4,930 14,800 7.26 21.8 __________________________________________________________________________ This feature makes the utilization of propellant thermal decomposition eminently feasible with minimum problem from NVR build-up due to boiling and/or three-phase flow in the liquid injection tube. In order to ensure good fluid mixing and minimum power input in the injection tube near the outlet, a mixer and shield assembly may be incorporated into the fluid injection system downstream of the valve, and just upstream of the injection point. This incorporation of a thermal decomposer made integral with the augmenter will permit retention of close to 100% of the chemical power in the gas flow. (h) Using a further configuration, the limitation on power imposed by the maximum presently available material stock diameter of molybdenum-rhenium is removed and the diameter can now be increased to any optimum value. This makes it possible to design radiation transfer augmenters, using combinations of the power enhancing features of this disclosure, to operate over a broad range of power levels, even in excess of 20 kw. Removal of these size limitations makes it possible to optimize performance with H.sub.2 propellant as compared to the performance available with devices of the prior art. Hydrogen typically requires longer dwell time or higher heat transfer rate to achieve a given temperature increase as compared to a hydrazine augmenter. (i) As the power input to the augmenter increases, the coil size and weight increases. On a spinning spacecraft, this coil weight will almost certainly lead to unacceptably high sag rates of the coil when it is hot, thus reducing the lifetime below desired values. To overcome this problem, a technique of combining radiative and thermionic heating may be employed. A thermionic emitter element is mounted in the cavity of the heat exchanger with cylinders interleaving those of the thermal absorber. This emitter element is electrically insulated from the heat exchanger and is connected electrically to the negative terminal of the power bus. The heat exchanger is electrically connected to the positive terminal of the power bus and the operational sequence for this heater is as follows: (1) Power is supplied to the coil by closing a switch. (2) The power radiated from the coil heats the thermionic element and the heat exchanger. (3) Once the temperature of the emitter element gets above about 1650.degree. K. (2500.degree. F.), great numbers of electrons are emitted by the thermionic emitting material, such as thoriated tungsten, from which the emitter is fabricated. The electric field between the emitter and the heat exchanger accelerates these electrons toward the heat exchanger where they impact and are absorbed. (4) This electron flow constitutes an electric current which flows across a potential drop equal to that of the power source, such as a battery, delivering energy to the heat exchanger at a rate given by: EQU P=IV (1) where: P=power; PA1 I=current; PA1 V=potential drop. PA1 I=current; PA1 A=active emitter area; PA1 .epsilon..sub.o =capacitivity of vacuum; PA1 .vertline.e.vertline.=charge on the electron; PA1 m.sub.e =mass of the electron; PA1 x=gap; PA1 V=potential drop between the emitter and the heat exchanger. PA1 1. Propellant characteristics PA1 2. Stagnation pressure and throat area PA1 3. Nozzle contour PA1 4. Nozzle exit area PA1 5. Nozzle wall temperature distribution PA1 thermal diffusion (species diffusion in a temperature gradient) PA1 pressure diffusion (species diffusion in a pressure gradient) PA1 interspecies energy transfer PA1 T.sub.c =stagnation temperature in chamber PA1 .gamma.=ratio of specific heats of gas PA1 M=free stream Mach number in flow PA1 P.sub.r =Prandtl number PA1 r.sub.c =radius of curvature of throat PA1 r*=throat radius PA1 f(.gamma.)=a function of the specific heat ratio, .gamma..perspectiveto.0.97+0.86 PA1 Re*=flow Reynold's number based on throat diameter PA1 .gamma.=ratio of specific heats of the gas PA1 .DELTA.C.sub.F =increment to thrust coefficient from supersonic portion of the nozzle PA1 p.sub.c =chamber pressure PA1 A*=throat flow area PA1 p*=gas pressure at the throat PA1 .rho.*=gas density at the throat PA1 w*=gas velocity at the throat PA1 p.sub.w =gas pressure at the wall PA1 .tau..sub.w =shear stress at the gas-wall interface ##EQU5## .mu..sub.w =viscosity of the gas at the wall W.sub.f.s. =free stream velocity of the gas PA1 .delta.=momentum thickness of the boundary layer PA1 r=radial coordinate in cylindrical coordinate system PA1 R=radial coordinate in spherical coordinate system PA1 .theta.=half angle of the nozzle PA1 r.sub.e =radius at nozzle exit at the nozzle exit, or PA1 F=thrust PA1 p=gas pressure PA1 .rho.=gas density PA1 w=gas velocity PA1 r=radial variable PA1 r.sub.e =radius at nozzle exit PA1 .theta.=half angle of the conical nozzle PA1 M.sub.e =Mach number at the exit of the nozzle PA1 .gamma.=ratio of specific heats of the gas PA1 .theta.=half angle of conical nozzles PA1 Re*=Reynold's number based on the throat diameter PA1 .varies.=a "variable" constant Almost all of this energy is deposited in the thermal absorber, adjacent to the nozzle. (5) This power (P) heats the thermal absorber to temperatures above that of the emitter. Most of this power is transferred to the gas in the heat exchanger near the nozzle. However, some small fraction is radiated back to the emitter, supplying the work function energy to maintain the electron emission and temperature of the emitter, which in turn keeps the electric current flowing. (6) The gaps between the emitter and the absorber are designed to values which control the level of current at the space-charge-limited level, given by the Child-Langmuir equation: ##EQU1## where: (7) Once the design current is flowing and the steady-state operational temperature with propellant flowing is established, a switch can be opened, permitting the coil to cool down to the heat exchanger temperature. Since the coil is used essentially as an initiator, being hot for only a few minutes each firing, the coil lifetime can be many thousands of augmenter operational hours before it would fail due to factors of evaporation and/or distortion or sag. (8) When the firing is to be terminated (after 40-60 minutes), a switch is opened and the propellant flow rate is stopped by closing the valve. The advantages of this type of heating over pure radiative heating are as follows: (a) Unlimited power can be transferred by increasing the voltage, , the area of the emitter, , or by decreasing the gap, . (b) The energy is deposited exactly where it is most useful, near the nozzle throat entrance, thus permitting this final and most thermally isolated section of the heat exchanger to operate at the highest temperature possible and with the highest thermal efficiency. (c) The structure of the emitter can be very rigid and thus not move or deform appreciably under the "g" loading in a spinning spacecraft. (d) The cylindrical structure of the emitting and absorbing elements give great rigidity to the surfaces so that electro-thermo-mechanical instability leading to hot-spot development cannot readily occur. (e) The emission characteristics of thoriated tungsten match exceptionally well to the temperature range to which the heat exchanger (absorber) and propellant gases need to be heated to get specific impulses of 300 to 340 sec. (FIG. 32 is explanatory in this regard). By designing for current densities of under 0.2 amps/cm.sup.2, the emitter temperature can vary from 1650.degree. to 2200.degree. K. (2500.degree. to 3500.degree. F.) before the current flow becomes "emission" limited, rather than "space charge" limited. With the collector part of the heat exchanger operating 38.degree. to 93.degree. C. (100.degree. to 200.degree. F.) hotter than the emitter, an ideal temperature is produced for operating a high-performance augmenter using hydrazine decomposition products, hydrogen or ammonia, as the propellant gases. (f) The intermeshed emitter-collector structure disclosed hereinafter with reference to FIGS. 30-31 represents a near-ideal configuration for implementation in a thruster. The large surface area of both components permits low current density operation. This has the following advantages: (i) A wide operating temperature range at constant power input is available. (ii) The large active emitting surface area permits gaps of over 0.020 inches between the emitter and the collector with conventional spacecraft voltages of about 40 volts (see FIG. 32). These size gaps can be maintained at an adequately constant spacing, even with significant differential thermal expansion between the components. (iii) With low current density, the power density to the collector is low (.perspectiveto.8 watts/cm.sup.2). Thermal conduction in the metal can hence overpower any tendency for hot-spot development due to small changes that may occur locally in the gap spacing. (g) During the lifetime of the spacecraft on which the engine is being used, the average voltage of a typical space vehicle power source during a firing decreases. For most current communication spacecraft this is usually from 41 to 36 volts. Simultaneously, the storage pressure feeding the propellant will decay over life from, typically, 300 psia to 100 psia. Since the power output varies as the voltage to the 5/2 power, thermionic heating will have less change in specific impulse over life than will pure radiation heating, where the power input varies approximately as the 8/5 power of the voltage. In order to further improve upon the performance of thrusters such as that which is disclosed in the parent application, modifications in the nozzle assembly are desirable and the present invention includes aspects of nozzle design and analysis resulting in improved thruster performance. The objective of the nozzle analysis and design optimization is to determine the nozzle configuration which will deliver the maximum I.sub.sp at the specified thrust and to predict the off-design performance over a range of operational parameters likely to be encountered during operations. A further objective is to investigate options for minimizing the back flow contamination of the space vehicle by the exhaust plume, plume interaction with vehicle structure that could be located in the flow path such as solar panel array and plume loss mechanisms in satellite configurations where the exhaust gases must pass through a long large-diameter duct. This last problem area is presently postulated as limiting the usefulness of augmented engines on some spinning spacecraft, the coil "sag" problem in the g-field having been successfully solved recently by the applicant. Nozzle and engine operating and configuration variables to be considered are: The importance of working for a high thrust coefficient is described in greater detail hereinafter and indications are that at a specific impulse of 310 sec., the stagnation gas temperature can be reduced from 2077.degree. K. (3280.degree. F.) to 1950.degree. K. (3050.degree. F.) if the thrust coefficient is increased from 1.60 to 1.65. Alternatively, the specific impulse could be raised to 320 sec. at a gas temperature of 2077.degree. K. (3280.degree. F.) by increasing the thrust coefficient from 1.60 to 1.65 (see in this regard FIG. 33). Since the procedures for analytically and experimentally determining the thrust coefficient were generated for high thrust rockets exhausting to an ambient atmosphere, some of the physical processes that can affect the thrust coefficient in low thrust rockets, whose laminar boundary layers may encompass a significant fraction of the flow, exhausting to the vacuum of space, have not been given due consideration. A curious anomaly appears to have occurred in the design of nozzles for rockets used to control space vehicles. Although rocket nozzle designers realize that they operate only in the vacuum of space, the area ratios of most nozzles are designed to give best performance in the steady-state vacuum achievable in the ground test facility where the engine will be tested. No significant attempt has been made in the prior art to utilize the thrust available, through proper design, from the low pressure expansion region. Nozzles also have traditionally been cooled to minimize chemical and physical erosion, to reduce radiative power loss and to minimize nozzle weight by taking advantage of the higher strength of the material at low temperatures. This design approach needs to be reexamined for low thrust rockets that operate only in a space environment. This is especially true for the high performance augmented engines in which new mechanisms occur when the gas consists of a mixture of atoms or molecules of very different molecular weight, such as is found in the decomposition products, hydrogen and nitrogen, of hydrazine at high temperature. These new mechanisms can be identified by the processes of: It may be possible to utilize one or all of these effects to enhance the thrust coefficient over what it would be for a lower operating temperature engine or as compared to a propellant of a single species gas of the same temperature and molecular weight. The phenomena to be considered are: 1. Thermal diffusion tends to separate the species when a temperature gradient exists. For a mixture of hydrogen and nitrogen, the hydrogen is concentrated in the higher temperature regions. The computed and measured thermal diffusion coefficient is plotted as a function of the species number density ratio in FIG. 34. The amount of species separation that could be achieved is plotted in FIG. 35 as a function of the temperature ratio, T.sub.hydrogen /T.sub.nitrogen. If the nozzle were run hot, then the gas in the boundary layer would have a higher percentage of hydrogen than average (i.e., greater than 67%). This reduces the viscosity in the boundary layer as well as the mass flux through the boundary layer, both of which can reduce viscous momentum losses and may increase the thrust coefficient. 2. Pressure diffusion permits the lighter molecular weight component (in this case hydrogen) to have a higher velocity than the nitrogen as the gas expands through the pressure gradient in the nozzle. This effect becomes more pronounced as the static pressure drops below one Torr. This is shown graphically in FIGS. 36 and 37 where the velocity separation achievable with a fixed pressure gradient is plotted as a function of the pressure. In the "core" flow one wants to discourage this from occurring by having the hydrogen molecules accelerate the nitrogen molecules by way of collisions; this leads to the highest values of the thrust coefficient. A design criterion becomes: The stagnation pressure should be as high as possible so that when the two species do "uncouple", at low pressure, the static temperature of the gas is as low as possible. 3. Interspecies Energy Transfer. For a given mass of gas in the chamber, two-thirds of the energy resides in the hydrogen and one-third in the nitrogen. If this gas is now expanded through a perfect nozzle with no pressure diffusion so that the hydrogen and nitrogen molecules form a univelocity beam, one-eighth of the energy is carried by the hydrogen, and seven-eighths is carried by the nitrogen. For this "massive" energy transfer to occur (more than one-half of the energy in the gas), the temperature of the hydrogen, as well as its directed velocity, must be higher than that of the nitrogen through the nozzle. An estimate of the temperature difference needed to keep the two species moving at the same velocity may be determined as a function of pressure. Again, to ensure that this energy transfer takes place, the pressure should be as high as possible. An enhancement mechanism may be available by considering the following model: (a) Operate the nozzle wall, both upstream and downstream of the throat, above the recovery temperature of the hydrogen. (b) The high thermal conductivity of the hydrogen may transfer energy into the hydrogen in the core, where collisions with the nitrogen will transfer this added energy to the nitrogen. (c) This addition to the stagnation enthalpy of the gas in the nozzle will tend to increase the thrust coefficient by a factor that is proportional to the square root of the ratio of the enhanced stagnation temperature over the unenhanced stagnation temperature. As a result of the above considerations, several important design criteria result: (a) The stagnation pressure in the heat exchanger should be as high as possible. (b) The nozzle wall temperature should be operated at a temperature equal to or greater than the recovery temperature in the gas. This must be "optimized" by including considerations of radiation power loss from the nozzle. The recovery temperature, T.sub.r, is defined as: ##EQU2## where T.sub.r =recovery temperature The curvature of the nozzle at the throat, r.sub.c, is another important parameter of the nozzle design. It appears prominently in the expression for the discharge coefficients C.sub.d in the following form: ##EQU3## where C.sub.d =discharge coefficient How the value of the discharge coefficient affects the thrust coefficient is not immediately obvious. This is investigated by developing a novel method of computing the thrust coefficient. The thrust of a rocket (F), operating in a vacuum, can be computed by two methods: 1. Evaluating the integral: ##EQU4## where C.sub.F =thrust coefficient for subsonic portion of the nozzle (2) Integrating the stress tensor over the axial projection of all interior and exterior surfaces. The approach adopted here is to compute the thrust that is generated up to the nozzle throat using method 1 above, and then to compute the additional thrust in the expanding section using method 2 above. The two components of the thrust coefficient, C.sub.F, are identified as follows: ##EQU6## where EQU C.sub.F =C.sub.F *+.DELTA.C.sub.F and, Since the velocity w* is purely axial at the throat, cylindrical coordinates are used in computing C.sub.F *. For invisid gas, accelerated at constant enthalpy and with a conical diverging nozzle, the integrals can be evaluated. The results are: ##EQU7## where Note: On the above calculations, the flow at the throat has been made spherically symmetric for convenience. Equations 9 through 11 represent the results of the classical approach to computing the thrust coefficient. Assuming that the pressure is independent of the radius at the throat, the viscous effect on C.sub.F * can be computed. The result is: ##EQU8## where EQU C.sub.D =discharge coefficient The expression indicates that the radius of curvature at the throat should be small so that C.sub.D is kept as high as possible. This conclusion may be somewhat modified by the desire to continue heating the gas as it accelerates through the throat. Some indication of the optimum nozzle shape can be determined by using equation 7. Immediately downstream of the throat there will be a negative increment to C.sub.F since p.sub.w sin .theta.&lt;.tau..sub.w cos .theta.. Once the expansion angle is increased to make the expression in brackets positive, the angle .theta. must be adjusted throughout the expansion to ensure that: EQU p.sub.w sin .theta.-.tau..sub.w cos .theta.&gt;0 (13) Eventually the nozzle half-angle will approach 90.degree., becoming a disc perpendicular to the axis of the throat. When the disc is extended out sufficiently far radially, such that substantially no collisions are occurring between propellant particles at the periphery thereof, at that circumferential location, a conical end piece may be provided having an angle with respect to the longitudinal axis of the nozzle designed to maximize deflection of propellant particles in the direction parallel with the nozzle axis. Since the disc part and conical end piece of the nozzle can be fabricated from extremely thin sheet material, the conical end piece should extend to the maximum diameter permitted by space vehicle constraints. The pressure at the wall, p.sub.w, is a strong function of .theta. and .tau..sub.w, a weak function of .theta.. Both decrease as R is increased. An analysis of the nozzle in accordance with the teachings of the present invention should permit an optimization of the nozzle contour and determine the exit area for the range of operational Reynold's numbers. In order to obtain the most accurate results for the nozzle design downstream of the throat, equation 7 or 35 may be calculated for spaced nozzle wall increments as low as one millimeter or less. Such calculations may be done by computer for greater efficiency and accuracy. Taking into consideration the test data and the design implications of the mechanisms discussed earlier, nozzle configurations that have analytical interest are sketched in FIGS. 38, 39 and 40. The rationale for each design feature is indicated on the figures. Some predicted and test data is available for estimating the values of the thrust coefficient for various nozzle shapes and gases (Murch, C. K., Broadwell, J. E., Silver, A. H. and Marcisz, T. J. "Performance Losses in Low-Reynolds-Number Nozzles", J. Spacecraft, Vol. 5, #9; Potter, J. Leith and Carden, William H., "Design of Axisymmetric Contoured Nozzles for Laminar Hypersonic Flow", J. Spacecraft, Vol. 5, #9; and Kinslow, Max and Miller, John T., "Nonequilibrium Expansion of a Diatomic Gas Through a Convergent-Divergent Nozzle", The Physics of Fluids, Vol. 9, #9. The first reference gives throat Reynold's numbers in data for nitrogen and hydrogen flows which are comparable throat Reynold's numbers to those found in "EPAT" and "ACT" in FIG. 41. This data has been examined so that comparisons with the existing test data from the two augmented engines, "EPAT" and "ACT", could be made, and also to determine the feasibility of scaling to higher throat Reynold's numbers. In this comparison a surprising difference in performance between nitrogen and hydrogen is seen, FIG. 41. This may be because of the different rates of freezing the vibrational and perhaps even the rotational energy as the pressure drops during the expansion. Using the definition of nozzle efficiency contained in the first reference, the efficiency values can be converted to thrust coefficients by multiplying the ordinate by 1.7498, the thrust coefficient for a perfect gas expanding through an area ratio of 100 and turned so that the velocity is only in the thrust direction. Scaling curves may be drawn based on the following relation: ##EQU9## where C.sub.F (.gamma., A/A*)=the functional dependence of the thrust coefficient on the ratio of specific heats, and the area ratio Since most of the data for various area ratios and for the two gases passes through the point ##EQU10## where .eta..sub.n =nozzle efficiency Extrapolated performance predictions using this relation are plotted in FIG. 41 out to Reynold's numbers of 20,000 using a value of .gamma.=1.40 with data from the first reference also being shown on this plot. When a higher gas pressure is used it may be possible to recover more of the energy from vibrational and rotational de-excitation. The maximum available thrust coefficient from a gas with a value of .gamma.=1.31 has been calculated assuming an area ratio of 100, and that profile losses scale in the same manner as computed previously. This curve is also displayed in FIG. 41. Finally, the best estimates of the thrust coefficients from several thruster implementations are shown in FIG. 41. These implementation are described in FIG. 41 with abbreviations defined as follows: ______________________________________ HiPEHT a TRW electrothermal hydrazine augmented disclosed in U.S. Pat. No. 4,322,946 ACT a Rocket Research Corporation implementation of a catalytic hydrazine augmenter as disclosed in parent application serial number 517,265 filed July 26, 1983 EPAT TRW test data from applicant's implementation of the catalytic hydrazine augmenter disclosed in the parent application Lord(a) resistojet disclosed in a publication by J. A. Donovan, W. T. Lord and P. J. Sherwood entitled "Fabrication and Preliminary Testing of a 3 KW Hydrogen Resistojet" given at the AIAA 9th Electric Propulsion Conference, April 17-19, 1972. Lord(b) resistojet disclosed in a publication by J. A. Donovan and W. T. Lord entitled "Performance Testing of a 3 KW Hydrogen Resistojet" Yoshida resistojet disclosed in a publication by R. Y. Yoshida, C. R. Halbach and C. R. Hill entitled "Life Test Summary and High Vaccum Tests of 10 MLB Resistojets" ______________________________________ It should be emphasized that the predicted performance, as well as the measured performance from the first reference is based on an area ratio of 100. HiPEHT and ACT have area ratios of 250 to 300, but were tested with ambient pressure of 0.1 to 0.5 Torr. EPAT has an area ratio of 700 and a flat plate nozzle continuation out to a diameter of 2 inches. The test was also conducted with an ambient pressure of 10.sup.-5 to 10.sup.-2 Torr during the test. The large area ratio and the low background pressure in the test chamber accounts for the higher thrust coefficients shown for EPAT. Further data on the effect of the ambient pressure on thrust has been found in papers (by Lord and Yoshida) presenting results from a 3KW hydrogen resistojet and a 10 MLB resistojet. In two cases, thrust was measured as a function of pressure and the thrust coefficient increased from 1.14 to 1.40 in one case as the pressure was decreased from 20 microns to 2.6 microns (1000 microns=1 Torr) and in the other case, the thrust coefficient increased from 1.25 to 1.51 as the pressure was decreased from 1000 micron to 1 micron. This and other data from these papers are plotted in FIG. 41 and are identified as data points by Lord and Yoshida.