Patent Number: 048256471
Section: description

Before discussing the present invention, the subject of matter of FIGS. 1-13 as disclosed in the patent application will be set forth. Referring to FIGS. 1-10, a thruster assembly 11, according to the present invention, is comprised of a heater assembly 13 and heat exchanger assembly 15. The heat exchanger 15 contains fuel passages 17 through which propellant is caused to enter. Energy is transferred to the propellant for any or all of the following purposes: to be vaporized, decomposed, reacted and/or heated to a higher temperature. The heat exchanger fuel passages 17 may be coated or plated at 14 as shown in FIG. 7 for the purpose of chemically isolating the fuel passages 17 from the propellant or to enhance decomposition or reaction of the propellant. One example of a material which enhances catalytic decomposition of the propellant hydrazine is platinum. The heated propellant will then pass on its way to a rocket propellant expansion nozzle 19 which includes a nozzle throat 21. The nozzle throat 21 opens into an expansion section 23. The propellant may be any one of, for example, hydrazine, N.sub.2, NH.sub.3, H.sub.2, etc. Referring to FIGS. 2, 4 and 6, the heater assembly 13 comprises a heater element 31 which is formed as a coil 33 surrounding a center conductor and/or supporting post 35. The coil 33 is formed of wire which is first tightly wound into a small diameter helix (about 0.1 inches in diameter) which is then wound into a large diameter helix (of about 1/2 inch to 3/4 inch outside diameter). The coil 33 is connected to an electric power source 200 (FIG. 6) by way of a pair of power conductors 35, 37 and 39. The power conductors 35, 37 and 39 are attached to the heater filament 33 at end 41 of a heating cavity 43 which houses the heater filament 31. The heater assembly is an elongate structure including in this embodiment the coil 33 formed of two segments that are bifilar wound as a double helix. Each lead conductor is made of several sections corresponding with temperature transition from the high temperature radiating heater or emitting cathode to the cold, less than 100.degree. C., connection to a power supply 200. For simplicity, only one section of the lead system is shown in FIGS. 1 and 6. The radiating coil wire 33 forming in the helixes includes a portion that is not coiled that extends in straight line paths away from the coiled portion until it reaches a lower temperature zone at 31 where the temperature is less than 1000.degree. C. This straight wire is a section of the power lead conductor. The coil 33 is centered along a center axis 47 (FIG. 2) of the thruster 11 from the end 41 of the heater assembly 13 to the lower temperature zone at 31. The center conductor 35 extends along the center axis 47 and is attached to the coil 33 at the end 41 of the coil. When power is provided to the power conductors 37 and 39 the heater filament 31 elevates in temperature and becomes less rigid than it would be with the power switched off. With current passing through the heater element 33, electro-magnetic forces are formed which cause the coil 33 to be biased in such a way as to center about the center post and/or conductor 35. The bifilar heating filament provides the primary source of heat from the heater assembly 13. In order to enhance the transfer of heat form the coil filament 33, the coil filament 33 can be coated (FIG. 8) or surface processed (FIG. 9) to produce a high-emissivity surface, thereby maximizing the transfer of thermal energy from the coil filament 33 to the heat exchanger cavity wall 38. Additionally, the heat exchanger 11 cavity wall 38 can have a high emissivity coating 25B as shown in FIG. 11 or surface processing as shown in FIG. 10. The use of a high emissivity surface permits a greater transfer of power for any given temperature or, alternatively, reduces the temperature required to obtain a certain amount of power transfer. In operation, for a given amount of current, a faster warm-up time is also obtained by the use of the high emissivity surface. Advantageously, the emissivity with an optimum coating or surface treatment is expected to be 0.85 or greater, with 1.0 being perfect emissivity and 0 being perfect reflectivity. In operation, the heat exchanger cavity wall 38 has a large geometric view as compared to the heater coil 35 or center post 35 and, therefore, intercepts a high percentage of radiation emitted by the heater coil 33. To reduce the amount of radiant energy that would be lost out of the open end of the heat exchanger 15 (an area defined as the plane P--P of an opening (FIG. 4) in the heat exchanger cavity wall 38), radiation shields 82, 84 and 88 are located to intercept and reflect this energy back into the heater cavity 43, where most of the energy will be intercepted subsequently by the heat exchanger cavity wall 38. This embodiment of the invention illustrates an arrangement of a number of reflective discs, shown in FIG. 4, spaced along the power lead-heater support channel. Each metallic disc 82 has cut-outs 201 to allow passage therethrough without contact by the lead connectors 37 and 39. To reduce radial outflow of radiation, reflecting cylinders 84 are attached to several of the discs 82. The coil 33 illustrated is configured as a single element with a center tap 41 and support conductor or post 35. This element may be operated either as a single total element with the post 35, in this mode, merely performing a support function, or with the post 35 being connected in common with both filaments 33, or as two distinct heater elements. Additional elements might be also enclosed within the heating cavity 43 for the purpose of providing (1) additional step levels of operating power, (2) a non-harmful ground test circuit, or (3) redundant heating elements for greater reliability and/or extension of operating capability. The heater assembly 13 and heat exchanger assembly 15 are configured such that they can be fabricated and independently tested as separate entities, and substitute or test heater assemblies may be interchanged with the flight heater assembly 13 (see FIGS. 2 and 4). The heater assembly 13 may be attached to the heater exchanger assembly 15 so as to maintain an opening gap 18 as shown in FIGS. 2 and 4 which permits pressure within the cavity 43 to reach equilibrium with ambient pressure outside of the thruster 11. Since the thruster 11 is designed to operate under extraterrestrial conditions, the ambient pressure will be quite low. Thus, the pressure inside the cavity 43 will be nearly a vacuum and energy from the element 31 will be transferred to the heat-transfer structure 15 and the nozzle 53 primarily by radiation. An alternate embodiment (not shown) would provide a complete closing or sealing of the heater cavity opening gap 18 so as to entrap and/or enclose a cavity 43 pressurant such as an inert gas which may be placed in the cavity during assembly or be permitted to bleed into the cavity from a heat exchanger flow passage 12 to the heater cavity 43 bleed 20 (FIG. 4) during engine/heat exchanger operation or be pressurized from the nozzle 19 through a nozzle flow to heater cavity bleed line 30 as shown in FIG. 6. The heater assembly 13 may also be configured to provide radiant heating and/or thermionic emission energy transfer. An exemplary embodiment 213 illustrative of this emphasis is shown as FIG. 12. Here, the heater coils 233 are shown schematically as simple coils supported by a center post cathode lead connector 235. If desired, the coil 233 could be formed with helixes like those of FIGS. 2 and 4. The heater leads 237 and 239 are similar to those illustrated previously in FIGS. 1, 2 and 4. This heater embodiment allows operation in relatively high force fields, that is, 5 "g"s without detrimental sag. In this illustration the center post cathode lead connector 235 supports a cylinder-disc cathode 236. The cathode 236 is fabricated to conform with the shape of the heat exchanger anode cavity wall 238 and a separation gap 243. FIG. 6 is a schematic illustration of the inventive propellant flow control valve 198 and power leads 35, 37, 39 and associated power system 199. Also shown are power/voltage converters 197 and 197' if required as well as power switched 196 and 196'. As shown, separate power supplies 200 and 200' may be used for the radiation heater 233 and the thermionic converter 236 (FIG. 12) respectively, as well as for the two parallel coils 33 shown in FIG. 4. The separate power supplies and separate controls therefor enable a large variety of adjustments in heater intensity to be made. If desired, the power supplies 200 and 200' may be operated in a pulsed mode with "on" condition thereof corresponding to opening of valve 198 and "off" condition thereof corresponding to closing of valve 198. FIG. 12 illustrates an embodiment wherein a relatively large surface area cathode emitter 236 would be used for the primary mode of energy transfer and the radiation heater 233 would be used to heat the cathode 236 and anode 238 to emission temperature conditions. The radiation heater 233 can also be configured to augment or serve as a backup device to transfer energy by radiation to heat exchanger wall 238. In a typical embodiment, the emitter cathode 236 will be adequately supported by a bracket or brackets 241 to maintain separation gap 243 with a supporting distance insulator 242 separating the cathode 236 and the anode 238. An alternative embodiment to provide maximum spacing between the heating coil 233' and the heat exchanger wall 238' is shown in FIG. 13. This configuration is useful for full operation life of a radiation coil 233' in a gravitational and/or centrifugal force field where all energy transfer is to be accomplished with a radiative heater 233' as contrasted to the embodiment illustrated in FIG. 12 where the principal use of the radiation heater 233 is to preheat an emitter 236 and the emitter performed the principal amount of energy transfer. Since the transfer of energy from the filament 33 to the heat exchanger assembly 15 is primarily by radiation, the outer surface of the heater filament 33 and the inner surface of the heat exchanger assembly 15 are preferably provided with high emissivity coatings or are surface treated to effect a higher than normal emissivity. Referring to FIG. 8, a cross-section of a heater filament 33 is shown to have a high emissivity coating or plating 250 formed thereon. This coating or plating may be, for example, hafnium carbide. FIG. 11 shows a similar coating or plating 256 on the inner surface of heat exchanger assembly wall 38. Referring now to FIG. 9, a blown up side view of a heater filament 33 is shown to include surface treatment 252 for the purpose of the increasing the surface area there to enhance and increase heat transfer therefrom. This surface treatment may be accomplished through mechanical or thermo-chemical means. FIG. 10 shows a similar treatment 254 on the inner surface of heat exchanger wall 38. This surface treatment may increase the filament surface area and cavity surface area by at least 20% and possibly by as much as 100% or more. For the purpose of this disclosure, extraterrestrial conditions mean the conditions normally present where orbital satellites are located. This normally includes the ionosphere and above provided that the satellite is within planetary orbit about the earth. Emissivity, .epsilon., is a property of a surface which permits the surface to radiate heat across the surface. It is given a dimensionless value, with a pure reflector having an emissivity of .epsilon.=0 and pure black body having an emissivity of .epsilon.=1. Typical values of emissivity are: ______________________________________ Gold = 0.05 Molybdenum = 0.15 Tungsten = 0.2 Hafnium Carbide (HfC) = 0.8-0.9 Tungsten Carbide (WC) = 0.5 ______________________________________ For the purpose of this patent application, high emissivity means .epsilon.&gt;0.4. The high emissivity materials should have an emissivity as high as is practical, considering the thermal stresses to which the material is exposed. Advantageously, the emissivity value of the high emissivity material should be .epsilon.&gt;0.5 and preferrably .epsilon.&gt;0.75. If possible, the emissivity of these materials should have a higher value, such as .epsilon.&gt;0.85. In the preferred embodiment, hafnium carbide is used for its high emissivity and ability to withstand high temperatures. However, tungsten may be thermally and chemically treated to modify the surface to increase the nominal .epsilon. of 0.2 to 0.5 or higher. The heat exchanger assembly 15, as mentioned above, contains fuel passages 17 which are provided in layers about the heater cavity 43. Propellant enters the fuel passages 17 either as a gas or as a liquid through a propellant inlet line 61 which directs propellant to an intermediate temperature propellant passageway 63 which connects with an elevated temperature propellant passageway 65. Passage of propellant from the intermediate temperature propellant passageway 63 to the elevated temperature passageway 65 is by way of two propellant flow passageways, to with, fore and aft conduits 67 and 69 located at fore and aft ends of the helixes, respectively. The elevated temperature propellant passageway 65 communicates with a short expansion nozzle structure conduit 71 which, in turn, communicates with a propellant expansion chamber 19. The propellant inlet line 61, the intermediate temperature propellant passageway 63, the elevated temperature propellant passageway 65 and conduits 67-71 are all considered a part of the fuel passages 17. The elevated temperature propellant passageway 65 and the intermediate propellant passageway 63 are each cut as a helix within the heat exchanger assembly 15. A series of laminations 81 arranged concentrically about the propellant passageways 63, 65 and provide a means to retain as much heat as possible within the heat exchanger assembly 15. The laminations provide a thermal insulting function within the heat exchanger assembly 15. A set of laminations 83 between the intermediate temperature and elevated temperature propellant passageways 63, 65 forms a thermal shield. A second set of laminations located concentrically outside the intermediate temperature propellant passageway 63 and forms a second shield 85. Beyond the second shield 85 are additional laminations 94 and 97. The helixes defining the intermediate and elevated temperature passageways 63, 65 are formed as thread-like cuts 87 in thermally conductive material which is defined as propellant passageway material 89. There are, or course, no mating threads for the thread-like cuts, as propellant passes through these cuts 87 instead. The thread cut arrangement facilitates fabrication because, prior to assembly, the cuts 87 are on the outside of their respective portions of the propellant passageway material 89. Due to the concentric relationship of the propellant passageways 63, 65, they are able to be assembled by merely nesting concentric layers. As best shown in FIG. 3, the propellant passageway material 89 forming the propellant passageways 63, 65 extends to a first concentric tube 91. A second concentric tube 92 is located concentrically outside of the first concentric tube 91. A third concentric tube 93 is located concentrically outside of the second concentric tube 92. The concentric tubes are separated from each other by laminations 81 which, together with the concentric tubes 91-93, form outer thermal shields. These outer thermal shields comprise the second thermal shield 85 and third thermal shield 94 and an external shield 97. The third concentric tube 93 is made continuous with a foreplate 95. The third concentric tube 93 and foreplate 95 form an exterior layer of the external shield 97. The exterior surface of the external shield 97 is coated with a low emissivity coating. A portion of the propellant passageway material 89 extends outwardly the first concentric tube as a first connecting ring portion 99. A second connecting ring portion 101 extends between the first and second concentric tubes 91, 93, at fore ends of the first and second concentric tubes 91, 92. Laminations separate the fore ends of the first and second concentric tubes 91, 92, as well as the second concentric ring portion 101 from the foreplate 95. An exterior connecting ring portion 103 extends between the second concentric tube 92 and the third concentric tube 93 at aft portions of the second and third concentric tubes 92, 93. The connecting ring portions 99-103 and the concentric tubes 91-93, as well as the foreplate 95, form a supporting structure for the propellant passageway material 89 and that part of the thruster 91 located aft of the fore plate 95. The locations of the connecting ring portions 99-103 cooperate with the concentric tubes 91-93 to form a folded configuration for the supporting structure. Thus, direct heat conduction through the supporting structure must take a tortuous path from the propellant passageway material 89 to the foreplate 95. The foreplate 95 is attached to a thruster mount 105 which is a part of the satellite designed to support the thruster. Because of the folded arrangement achieved by the concentric tubes 91-93, as connected by the connecting ring portions 99-103, the foreplate 95 is kept relatively cool, thus presenting a minimum of thermal heat conduction to the satellite via the thruster mount 105. In order to further reduce the temperatures to which the thruster mount 105 is exposed, the exterior surface of the thruster 11, particularly exterior surface 107 of the third concentric tube 93, is coated or surface conditioned to obtain a high emissivity coating. A preferred high emissivity coating would be hafnium carbide (HfC). The use of the high emissivity coating on exterior surface 107 increases radiation heat loss from the third concentric tube 93, thereby conducting less heat to the thruster mount 105. The reduction in temperature of the third concentric tube 93 is believed to also affect the infrared radiation by causing emission to occur at longer wavelengths. This not only makes it difficult for an outside observer to determine when the thruster 11 is being heated, but also makes the thruster more difficult to trap using infrared sensors. By providing the high emissivity coating on exterior surface 107, the operation temperature of the thruster 11 at the propellant passageway material 89 can be increased even though it may be necessary to maintain a low temperature at exterior surface 107. This enables the thruster 11 to operate at high efficiencies because of the use of the high emissivity coating on exterior surface 107. It should be further noted that without the use of the folded arrangement of the concentric tubes 91-93, separated by the laminations 81, it would be necessary to reduce heat loss at the exterior surface and a low emissivity coating on the exterior surface would be less practical. In addition to the thruster mount 105, various controls are attached to the thruster 11. The reduction of temperature accomplished by the use of the high emissivity coating at exterior surface 107 (similar to that shown in FIG. 11) reduces the maximum temperature to which external components of these controls are exposed. Heat conduction through the supporting structure 107, 92 and 91 may be reduced by having material cut-outs 40 as illustrated in FIGS. 1 and 3. When propellant is being expelled from the expansion nozzle 19 (FIGS. 2 and 5) to produce thrust, high temperatures are created at the nozzle throat 21. To withstand these high temperatures, it is necessary to use high temperature or refractory materials at that location. An option as shown in FIG. 5 is to use a separate insert 109 for the nozzle throat area 21. The insert extends to connect the heat exchanger flow passage 117 with the expansion nozzle structure 111. By using the separate insert 109, costly materials are only required for the hottest portions of the heat exchanger chamber 15. The expansion section of the nozzle 111 is exposed to a lower temperature. This part of the nozzle can be made as a separate section or shell which covers laminations of external radiation shields 112 which are located in that area. It can be seen that, because of the uncoupled heat-exchange relationship of the insert 109 and the expansion portion 111, the amount of heat loss through the expansion nozzle due to conduction and radiation during the operation of thruster 11 is reduced. In the preferred embodiment, the insert 109 is made of thoriated tungsten. Further, the preferred material for the interior walls of the heat exchanger 138 and the nozzle inlet 118 is molybdenum/rhenium, and the expansion nozzle 111 is preferably made of TZM (a moderate cost molybdenum alloy) or titanium. By separating insert 109 from the expansion portion 111, a means is provided to reduce radiation losses from the nozzle 119. The insert 109 is not mechanically joined directly to the expansion portion 111 and a blocking effect is accomplished by a thin diaphragm 258 between insert 109 and the expansion portion 111. As shown, the diaphragm 258 is located on the expansion side of the nozzle throat 109 and acts as a "block" to prevent propellant flow from entering the radiation shield area 112 through the gap that would otherwise be present between insert 109 and expansion portion 111. Thus, the diaphragm 258 acts as a layer of metal blocking the gap from flow-through. In the preferred embodiment, this diaphragm 258 will be made of tungsten foil. Therefore, less power is transferred from the hot insert 109 to the expansion portion 111. It is expected that temperatures at the insert 109 will reach a range of 1700.degree. to 1900.degree. K., whereas temperatures at the intermediate divergent portion 111 will reach a range of 1100.degree. to 1400.degree. K. Without a separation of the nozzle portions 109-113, energy losses would be represented by: EQU P=.sigma..epsilon..sub.n A.sub.n T.sub.N.sup.4 (16) where P=power-energy/unit time PA1 .sigma.=(Stefan-Boltzman) constant PA1 .epsilon..sub.n =integrated emissivity of the nozzle 111 PA1 A.sub.n =effective area of the nozzle 111 PA1 T.sub.N =temperature of the nozzle 111 PA1 .epsilon..sub.i =emissivity of the insert 109 PA1 A.sub.i =area of the insert 109 PA1 T.sub.i =temperature of the insert 109 PA1 T.sub.p =temperature of the expansion portion 111. PA1 d.sub.equ =equivalent or hydraulic diameter PA1 =4Ac/Pc PA1 A.sub.c =cross-sectional PA1 P.sub.c =perimeter of wetter surface PA1 Nu=Nusselt number PA1 K=thermal conductivity of the gas PA1 m=mass flow rate flowing through the channel PA1 T.sub.w =temperature of the wall PA1 T.sub.r =recovery temperature in the gas PA1 Pr=Prandtl number PA1 d1=element of length along the flow channel PA1 r=radius of curvature PA1 u=flow velocity PA1 .upsilon.=kinematic viscosity PA1 Re=Reynold's number of the flow PA1 (1) Power is supplied to the radiative heating element 601 by closing the switch 613 with the switch 615 being in the open position. PA1 .gamma.=ratio of specific heats of the gas PA1 p.sub.w =gas pressure at the wall PA1 M=Mach number at the inner edge of the boundary layer PA1 r.sub.c =radius of curvature of throat PA1 r*=throat radius PA1 f(.gamma.)=a function of the specific heat ratio, EQU .gamma..perspectiveto.0.97+0.86.gamma. PA1 Re*=flow Reynold's number based on throat diameter PA1 .gamma.=ratio of specific heats of the gas PA1 .rho.=gas density PA1 w=gas axial velocity PA1 r.sub.e =radius at nozzle exit PA1 r=radial variable PA1 dr=differential of radial variable PA1 F=thrust PA1 p.sub.c =chamber gas pressure PA1 A*=throat flow area with the separated structure of the preferred embodiment, energy losses would be represented by: EQU P=.sigma.{.epsilon..sub.i A.sub.i T+.epsilon..sub.n (A.sub.n -A.sub.i)T.multidot.p.sup.4 } (17) where These equations are approximate models because of such factors as thermal conductivity and direction of thermal radiation. Using a 0.1 pound thrust engine for an illustrative example, the typical power loss values for an integral nozzle without a diaphragm would be: ##EQU11## For the nozzle with a diaphragm, approximate values would be: ##EQU12## These examples indicate that the radiative power loss from the nozzle can be reduced by more than a factor of 3 by using the diaphragm and thermal uncoupling. To provide for flow modulated operation, the inflow of propellant through inlet 61, FIGS. 1, 2 and 3, can be shut on and off by a flow control valve 198 shown in FIG. 6. OPERATING PROCEDURES Prior to operation, a warm-up procedure is normally followed. First, non-stored electrical energy, if available, is applied to the heater filament 33 in order to gradually increase the internal temperature of the thruster 11. Typically, such non-stored energy would be provided by solar cells or by a reactor power supply on the space vehicle and would provide an initial warm-up without taxing the vehicle's battery storage system. If the power available from such a non-stored energy source is fairly low, it may be desired to use the center conductor 35 in combination with one or both of the power conductors 37, 39, thus reducing the optimum operating voltage of the filament 33 in half. Such an initial phase of warm-up may last typically from several minutes to a couple of hours and is not essential to the successful operation of the device. Warm-up may also be accomplished by flowing reacted propellant through the device. A full warm-up procedure is then initiated. During the full warm-up procedure, current is applied to the coil filament 33, normally through the power conductors 37, 39 in order to bring the temperature of the elevated temperature propellant passageway 65 to a temperature at which the thruster 11 is ready for thrusting operations. When the temperature of the elevated temperature propellant passageway 65 is elevated in such a manner, the intermediate temperature propellant passageway also warms, with temperature gradually decreasing toward the third concentric tube 93. The coil filament 33 may be off-modulated when the overall temperature of the heater assemblies 13 is at a maximum limit or when the temperature at the elevated temperature propellant passageway 65 and the expansion nozzle 19 is sufficiently high for operation. Obviously, a number of control programs can be designed in accordance with reduced energy consumption and a necessary degree of readiness. At this time, the high emissivity coating on the exterior surface 107 and the folded structure of the concentric tubes 91-93 causes the exterior surface 107 to remain at a fairly low temperature. The low temperature operation, as stated above, prevents excess thermal conditions from occurrence at the thruster mount 105 and reduces the possibility that a warm-up of the thruster 11 can be readily detected. Typically this stage of the warm-up takes between a couple of minutes and a half hour. In the event of a lower power supply or when conditions otherwise require reduction of electrical consumption, a longer warm-up is employed. When the temperature occurring at the elevated temperature propellant passageway 65 and at the expansion nozzle 19 is sufficiently high, the thruster 11 is throttled on by causing propellant to enter the propellant inlet line 61. This causes the fuel passages 17 to cool, thus requiring additional heat from the heater assembly 13. The thruster 11 has a heat-sinking capability which permits the heater to be controlled by off-modulation, rather than by partial attenuation of current. This not only enables the heater coil filament 33 to operate at maximum efficiency, but also increases the efficiency of DC electrical power supply in that voltage-changing devices or resistor banks are not required for attenuation. The heater coil filament 33 is thus switched "on" and "off" by switch 196 in order to provide a desired minimum temperature for the propellant without greatly exceeding that temperature, in order to provide optimum and safe operation. When the cooling effect of the propellant is greater than the heat able to be produced by both the exchange of heat from the expansion nozzle 19 and the heat produced by the heater assembly 13, it is possible to off-modulate the propellant supply. The heater assembly 13 can then provide enough heat to heat the propellant passageway material 89 and the expansion nozzle 19 until the propellant can be caused to flow at an optimum rate. The ability of the thruster 11 to operate in such an intermittent manner enables an increased efficiency of operation, thereby reducing the requirement for electrical power consumption and conserving propellant fuel. An additional advantage of (1) the ability to off-modulate the heater assembly 13 and (2) the ability to operate the heater assembly 13 in a way which brings the internal temperatures of the thruster to proper levels without propellant passing through the fuel passageways 17, is the fact that the propellant can be selectively throttled, with the thruster being constantly ready for thrusting operations. This gives the engineers controlling the thruster a great deal of flexibility in the operation of the satellite and permits them to rapidly change the position of the satellite as circumstances require. Referring now to FIGS. 16-20, several aspects of the present invention will be discussed. Firstly, note the first opening 298 and second opening 299. The heater assembly 311 is mounted into the first opening 298 and the nozzle 321 opens to the second opening 299. As discussed hereinabove with regard to FIG. 2, reference number 89 refers to a pair of concentric members defined as propellant passageway materials, each of which has cut therein thread-like cuts 87 which define passageways for the propellant which are connected to one another via conduits 67. With reference back to FIG. 16, it is seen that the optimized performance augmenter 300 includes propellant passageway material 389 having cut therein screw thread-like passageway means 387. Between the threads of the thread-like passageway means 387, a plurality of lands are formed which are designated by reference number 388. These lands 388 define the interface between the passageway material 389 and outside wall 317 at the inner pass heat exchanger assembly 313. The inner pass heat exchange assembly 313 includes an interface surface 316 which faces and engages the lands 388 of the propellant passageway material 389. In one aspect of the present invention, the surface 316 of the inner pass heat exchanger assembly 313 is brazed directly to the lands 388 of the propellant passageway material 389 of the inside wall of the inner pass heat exchanger component which also serves as the enclosing cavity of the heater assembly 311. As discussed hereinabove, the braze material could be a material such as, for example, vanadium, molybdenum or iridium. Several methods are available for use in brazing the inner pass wall 388 of heat exchanger assembly 313 to the outside wall 316 of the inner pass heat exchanger assembly 313 via the lands 388 and the surface 316. One such method may comprise vapor deposition on the inner diameter surface 316 of the outside wall of the heat exchanger assembly 313 as well as on the lands 388 of the inside wall 389 of the inner heat exchanger assembly 313. A further method of brazing may comprise putting the braze material in position by layering between the parts of thin foil of the base material. Another method would be to locate grooves or channels 351 in the lands 388 of the inner pass heat exchanger 313 and placement of the braze material 353 in this channel 351 as seen in FIG. 46. After this is completed, the braze may be accomplished by any one of several methods including (1) heating the whole structure in a vacuum furnace, (2) placing a heater assembly in the heat exchange cavity defined between the inner wall 389 and outer wall 317 of the inner pass heat exchanger assembly 313 and heating in a vacuum environment preferably ion-pumped to under 10.sup.-4 Torr. This technique is preferred since the inner component 389 will thereby become somewhat hotter than the outer component 317 and the extra thermal expansion of the inner component 389 caused by this extra heating will close any gaps which may exist between the components over the surface which is to be brazed. As described hereinabove, affecting this braze reduces the tension load on the outer component 317 and the compression load on the inner component 389 to negligible value so that the wall thickness and thereby the weight thereof may be substantially reduced. Also, the problem of creep and rupture of these components is almost completely eliminated since each flow passage is now equivalent to a tube and all of the metal becomes structurally involved in the assembly. This improvement alone will permit lower specific weight for the heat exchanger, while simultaneously extending the life thereof to over 500 hours of use. With reference back to FIG. 2, as discussed hereinabove, there are two concentric structures of propellant passageway material 89. As shown in FIG. 16, the outermost propellant passageway material 89 from FIG. 2 is now designated as the outer pass heat exchanger 315 and now takes the form of a continuous coiled tube 391 which extends from the propellant inlet tube 361 to an exit point 392 which opens into a intermediate passageway 367 which is equivalent to the connecting passageway 67 of FIG. 2. The intermediate connecting passageway 367 communicates the coiled tube 391 with the spiral passageway 387 within the propellant passageway material 389 to thereby allow a continuous flow of propellant therethrough. If desired, the coiled tube 391 may be made of molybdenum-rhenium, rhenium alone, or other high-temperature materials. As described hereinabove, for better performance the coiled tube 391 may 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 into separate hydrogen and nitrogen. Such a coating would also help transform any ammonia (an intermediate decomposition product of hydrazine) that might be injected into the tube 391 at the inlet thereof into hydrogen and nitrogen molecules, to thus ensure that such ammonia would not reach the inner spiral passageway 387, which is much hotter than the coiled tube 391 where some damage to the material could result from intermediate reaction products such as N or N.sup.+ reacting with some component of the material. As shown in FIG. 22, useful thicknesses for this inner coating may easily be determined. Applicant has discovered through research that when propellant is flowing in a laminar manner, greater heat exchange results through a coiled tube than through a linear tube. Programs have been developed by applicant which compute the distribution of temperature and of heat flux rate through all components of a heat exchanger. These calculations must be iterated with calculations for the gas properties at each point in the heat exchanger. The expression used to describe the heating is: ##EQU13## where h.sub.o =stagnation enthalpy of gas with ##EQU14## T.sub.g =local static temperature of the gas M=gas Mach number Once values for the Nusselt number are established, the integration can precede. Standard texts on heat transfer give the Nusselt number for straight pipes and ducts. However, as will be seen later, the effect of curvature on the skin friction is very pronounced. By Reynold's analogy, a similar effect can be expected with the heat transfer. Hence, the same curvature corrections to the friction factor will be used to correct the Nusselt number. These corrections are significant in the helical passages proposed (over a factor of 2). Helical, rather than straight flow passages can hence increase the heat transfer rate to the gas very significantly, all other factors being equal. This enhanced heat transfer rate and friction factor occurs because a secondary flow is induced in the gas, as shown in FIGS. 43 and 44. This secondary flow will also mix the gas, giving a much more uniform enthalpy to the flow at any given cross-section. The enhancing effect of curvature also permits the flow to stay laminar to higher Reynold's numbers, while at the same time giving higher heat flux rates and skin friction coefficients than would be obtained from turbulent flow at these Reynold's numbers. Using these expressions and the procedure outlined hereinbelow, reasonable agreement has been obtained between the calculated and measured pressure drops in applicant's heat exchanger designs. The influence of curvature is stronger in laminar than in turbulent flow. The characteristic dimensionless variable, which determines the influence of curvature in the laminar case, is the Dean number D: ##EQU15## where R=radius of the cross-section The measurements carried out by M. Adler for the values: r/R=50, 100, and 200, demonstrated the existence of a large increase in the resistance to flow caused by the curvature for Re.sqroot.R/r&gt;101/2. According to his calculations the resistance coefficient, .lambda., for laminar flow in a curved pipe is given by ##EQU16## where Re=Reynold's number of flow and where .lambda..sub.0 denotes the coefficient of resistance of a straight pipe. Measurements indicate, however, that the above equation only has asymptotic validity, and may be used for values of the parameter .sqroot.R/r exceeding about 10.sup.2.8. The results of measurements are approximated with a higher degree of precision by the following empirical equation, first given by L. Prandtl. ##EQU17## This equation gives good agreement with experimental results in the range EQU 10.sup.1.6 &lt;Re(R/r).sup.1/2 &lt;10.sup.3.0 (23) C. M. White has found that the resistance coefficient for turbulent flow in a curved pipe can be represented by the equation ##EQU18## These differ somewhat from, but are in general agreement with C. M. White's equation above. On the basis of calculations and data evaluation to date, a number of further design criteria can be defined: (i) The flow channels should be designed so that the Mach number increases monotonically from the injection point up to the nozzle outlet. (ii) The flow passages should be as large as possible throughout the heat exchanger to ensure the maximum possible pressure at the throat. A comparison of FIGS. 2 and 16 reveals that the invention illustrated in FIG. 16 includes further structure not contemplated by the invention shown in FIG. 2. In particular, the terminus of the inner heat exchanger assembly 313 shown in FIG. 16 includes the provision of an energy absorber structure 309 and a pre-nozzle entrance heat exchanger 310. The heat exchanger 310 may, if desired, be brazed into the wall 302 on a side thereof opposite to the side to which the energy absorber component 309 is brazed. The nozzle heat exchanger 310 may be fabricated either as a spiral brazed at both ends to a housing, FIG. 48, or as a series of cylinders with gaps at alternative ends, FIG. 49. It may be fabricated from tungsten, rhenium, tungsten-rhenium, molybdenum or molybdenum-rhenium. There are also a trio of radiation shields 306, 307 and 308 which are located within the heat exchange cavity 303 extending the entire length of the cavity with the open ends facing the energy absorber structure 309. Each of the radiation shields 306, 307 and 308 is comprised of a disc 312 extending transverse to the longitudinal axis of the heating coil 305 and some radiation shields have the further provision of a cylindrical member 304 extending along this longitudinal axis. In FIG. 16, only the cylindrical member 304 associated with the disc 312 of the shield 306 is shown, however it may be seen from FIG. 20 that these cylindrical components may extend the full distance of the heat exchange cavity 303 or 303'. Each metallic disc 312 has cut-outs 338 to allow passage therethrough without contact by the lead connectors 340, 342, 344. These discs 312 are supported in the structure by a plurality of rods 346 (FIGS. 17 and 18) anchored in insulator segment discs similar to those illustrated in FIGS. 4, 12 and 13. Short tungsten springs 348 are mounted over the four support rods 346 and placed between the discs 312. These springs accurately position the discs axially and permit thermal expansion without inducing excessive stresses in any component. In order to accurately position the discs axially a predetermined compression is induced in all springs during assembly. The length of these springs can vary along the channel in which the discs are mounted in order to improve the efficiency of the radiation shielding. To reduce radial outflow of radiation, the reflecting cylinders 304 are attached to several of the discs 312. A plurality of smaller discs similar to the discs 86 in FIG. 4 are included on the lead connector to block radiation leakage through the cut-outs 338 in the larger radiation discs 40 wherein the gap is provided for noncontact passage of the lead connectors 340, 342 and 344. This embodiment of radiation shields permits a radiation transfer efficiency to the energy absorber component 309 of 90 to 95 percent. Similar structure to the disc-cylinder shielding structure described hereinabove may be used in the thruster housing between the opening for the heating element and the outer skin to reduce power losses. As shown in FIGS. 1-3, holes 40 are formed in the outer skin and inner support structure so as to expose the housing interior to the vacuum of outer space. Referring to FIG. 16, it is seen that the thruster 300 has a wall 350 structurally connecting the nozzle 321 with the outer walls of the housing. It is important to keep the wall 350 free of holes 40 so as to prevent flow of propellant leaving the nozzle 321 from entering into the housing. FIG. 19 shows an expanded view of a portion of FIG. 16 explaining the manner of installation of space optimized radiation shielding in the housing. As shown in FIG. 16, the shielding consists of a plurality of discs 370 having respective cylinders 371 preferably micro-arc welded thereto at 372 much in the manner disclosed in U.S. Pat. No. 4,404,956 to applicant herein. As better seen in FIG. 19 rods 373 are assembled through holes 374 formed in the discs 370, which holes are comprised of opposed annular beveled surface 375 to thereby define a circular line 376 which contacts the rod 373 so as to minimize the surface area of contact therebetween. In order to separate the discs 370 between respective disc pairs are inserted wire rings 377 which contact each of the adjacent discs 370 as well as the rod 373 in a circular line contact only which minimizes the surface area of contact therebetween. Applicant has discovered through research that the smaller the ratio of surface area of contact between shields to surface area shields, the smaller the power loss through the shields. Similarly, as the pressure of the atmosphere in which the shields are mounted is reduced, reduction in power loss is evidenced. These discoveries are demonstrated in Table 1A as set forth hereinabove. For example, referring to Table 1A, it is seen that with 16 shields being used at pressure of 10.sup.-2 Torr, as this ratio is reduced from 10.sup.-3 to 10.sup.-7, the power loss is reduced from 184.06 watts to 73.71 watts. Further, with reference to Table 1A, at a constant ratio of 10.sup.-5, as the pressure is reduced from 10 Torr to 10.sup.-5 Torr, the power loss is reduced from 266.16 watts to 76.05 watts. The energy absorber component 309 is brazed into the nozzle end of the cavity wall 302 so as to maximize power transfer from the coil 305 to the fuel which has made its way to the pre-nozzle entrance heat exchanger component 310. Power is transferred to the energy absorber component 309 from two sources, firstly directly through radiation from the coil 305, and secondly by reflection from the inner surface of the shield 306. As described hereinabove, by making the ratio of the gap length to gap spacing in the energy absorber component 309 high enough, photons impinging therein will undergo enough multiple reflections so as to be absorbed thereby, thus giving the energy absorber component 309 an effective emissivity of substantially unity. If desired, the energy absorber component 309 may be fabricated from any one of tungsten, tungsten-rhenium, molybdenum, rhenium, or molybdenum-rhenium. If desired, it may be fabricated as a series of cylinders brazed to a disc, FIG. 50, or as a scroll brazed to the disc, FIG. 51. By utilizing the full length cylindrical members exemplified by the cylindrical member 304 in conjunction with discs so as to comprise the radiation shields 306, 307 and 308 in conjunction with the energy absorber component 309, a great increase in the power radiated from the coil 305 which is transferred to the energy absorber component 309 is realized, thus enabling the peak temperature of the nozzle end of the heat exchange to be raised 200.degree. C. above what it might be able to be raised to without these improvements. Referring again to FIG. 2, it is seen that the propellant supply passages 87 extend in two concentric rows of thread-like passages. As discussed hereinabove with reference to FIG. 16, the outer passages 87 may be replaced with a continuous coiled tube 391 and modifications to the inner passages may be made, including brazing operations, to reduce tension loads on the outer component 317 and for other purposes specifically set forth hereinabove. Further improvement may be made in the FIG. 2 thruster by completely replacing the inner threads 87 thereof with a coiled tube like the tube 391 of FIG. 16. In this light, reference is made to FIG. 20 which shows a thruster 300' wherein the inner thread-like passageway 87 of FIG. 2 has been replaced with an inner coiled tube 322 to convey propellant from outer tube 391 to the nozzle heat exchanger component 310. As may be seen from a comparison of FIGS. 16 and 20, the support tube 320 of the FIG. 20 does not include the thread structure of the corresponding component of FIG. 16, designated by reference number 313, and further the support tube 320 with its cavity wall 302' is only required to (1) isolate the heater assembly 311 from propellant supply, (2) support the energy absorber component 309 and (3) support the nozzle heat exchanger 310. Accordingly, the tube 320 may be made much thinner than the tube 313 of FIG. 16, which results in: (1) saving in weight, and (2) higher levels of heat transfer from the heater assembly 311 to the propellant. As discussed hereinabove with regard to the tube 391, the tube 322 may similarly be internally coated with tungsten or rhenium. With the dual-coiled tube 322, 391 configuration shown in FIG. 20, if the nozzle heat exchanger 310 and the pressure vessel are made of rhenium or tungsten-rhenium, operating temperatures may be increased to over 4000.degree. F., and with hydrazine as a propellant this could allow the mission average specific impulse (I.sub.sp) to approach 340 seconds as shown in the FIG. 21 graph. For hydrogen as a propellant, this operating temperature would result in an average specific (I.sub.sp) approaching 850 seconds. As stated hereinabove, the inside surfaces of the coiled tubes 322 and 391 may advantageously be coated with iridium, tungsten or rhenium. Iridium is an advantageous coating material for use with hydrazine 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). Referring now to FIGS. 23 and 25, applicant has devised a device for coating the inside surfaces of a tube T before it is coiled to form tubes 322 and 391. As shown in FIG. 23, the coating apparatus 400 includes a top plate 401 and a bottom plate 403 which close openings formed in a high temperature glass tube 405 to form an enclosed chamber 407. An opening 409 is formed in the bottom plate 403 and a conduit 411 connects this opening 409 with a vacuum pump 412 which is designed to maintain the chamber 407 at a pressure of 10.sup.-4 Torr or below. Rigidly mounted to the bottom plate 403 is a bracket assembly 413 which includes an upstanding rod-like support 415. Extending outwardly from the support 415 are an upper arm 417 and a lower arm 419. The upper arm 417 has attached thereto at its extreme end a bracket 418, while the lower arm 419 has similarly attached thereto a bracket 420. Referring now to FIG. 24, it is seen that the bracket 420 has attached thereto a tube supporting member 427. The bracket 420 includes a radially inwardly directed shoulder 421 and a side wall 423 which slidably accommodate therein the member 427. An insulating disc 429 is slidably mounted in the tube supporting member 427, by virtue of the inner longitudinal wall 431 and radially inwardly projecting surface 433 thereof. The insulating disc 429 is preferably made of boron nitride and includes an opening 435 therethrough sized to slidably receive the high temperature lead 437. A flexible power lead 439 is suitably electrically attached to the high temperature lead 437. The high temperature lead 437 has an opening 441 therein designed to slidably accommodate one end of coating wire 443. A threaded opening 445 is formed transverse to and intersecting with opening 441 and a set screw is threaded into the opening 445 so as to forcibly engage the side of the wire 443 to thereby retain it in mounted configuration within the opening 441. In a similar fashion, the tube supporting member 427 includes an opening 449 in which is slidably inserted the tube T. The tube supporting member 427 further includes a transverse threaded bore 451 in which is threaded a set screw 453 which is provided so as to bearingly engage the tube T and thereby retain it within the opening 449. The high temperature lead 437 has welded thereto a vapor shield 438 which is provided so as to keep vapors created by the operation of this device from welding the tube T to the tube supporting member 427 and further prevents any short circuits involving the high temperature lead 437. A flexible power lead 440 is attached in a manner well know to those skilled in the art to the tube supporting member 427. With further reference to FIG. 24, it is seen that the bracket 418 has attached thereto a further tube supporting member 455 which includes an opening 457 for receipt of the top of the tube T and a transverse threaded bore 459 which receives therein a threaded set screw 461 which bearingly engages the surface of the tube T to thereby retain the tube within the bore 457. Further, the tube supporting member 455 includes a plurality of transverse bores 463 which are provided for the same purpose as the transverse bores 452 in the tube supporting member 427, to wit, to enable the venting from the tube T of vapors formed by the heating of the wire 443. The tube supporting member 455 further includes a further bore 465 which is provided to slidably receive therein the wire 443 and the tube supporting member 455 further includes a further transverse threaded bore 467 which threadingly receives a threaded set screw 469 which bears against the wire 465 to thereby retain it within the bore 465. As should be evident from FIG. 24, a flow path for electrical current is created by the structure shown therein from the flexible power lead 439, through the high temperature lead 437, through the wire 443, through the tube supporting member 455, through the tube T, through the tube supporting member 427 and to the flexible power lead 440. It is seen that the insulator 429 acts to electrically insulate the high temperature lead 437 from the tube supporting member 427 so as to prevent any short circuits. The vapor shield 438 prevents any vapors from impinging upon the insulator 429 so as to prevent any completion of circuitry between the high temperature lead 437 and the tube supporting member 427. As disclosed hereinabove, advantageous coating materials for the inside surfaces of the tube T include iridium, tungsten or rhenium. Accordingly, the wire 443 is made of whichever one of the hereinabove listed materials is desired to be used for the coating of the interior surfaces of the tube T. With this wire suitably attached to the high temperature lead 437 and the tube supporting member 455, a source of current of approximately 10 to 15 amps is placed across the flexible power leads 439 and 440. The wire 443 has an inherent resistance which increases substantially linearly with the temperature thereof. Accordingly, the resistance of the circuit may be measured and from the circuit resistance, a good estimate of the temperature of the wire 443 may be determined. In this way, the temperature of the wire 443 may be controlled. It is anticipated that in order to properly coat the inner surfaces of the tube T, the wire 443 will have to be electrically heated via the flexible power leads 439 and 440 for a time period of approximately one to four hours. It is noted that the current in the above described circuit flows through the tube T in the opposite direction to the flow of current through the wire 443. As a result, the electrons flowing in the tube T and the wire 443 tend to repel one another and thereby the wire 443 is maintained in a centered position within the tube T. It is further noted that as the wire 443 is heated, it will tend to expand. It is for this reason that the high temperature lead 437 is slidably mounted within the bore 435 of the insulator 429. In this way, as the wire 443 expands, the high temperature lead 437 will slide downwardly due to the force of gravity through the bore 435 to maintain the wire 443 in a taut configuration within the tube T. As stated hereinabove with regard to FIGS. 16 and 20, the shields 306, 307 and 308 may be coated with a low emissivity metal to thereby enhance their performance. Low emissivity metals include, for example, gold, silver and rhodium. With reference now to FIGS. 25 and 26, an apparatus will be described which has been devised so as to enable the coating of the shields 306, 307 and 308. As shown in FIGS. 25 and 26, the shield coating apparatus 500 is seen to include a first elongated tube 501 having an elongated slit 503 therein which extends approximately half of its longitudinal extent as best seen in FIG. 25. In surrounding relation to this tube 501, a further tube 505 is provided which includes a section removed therefrom defining faces 507 and 509 as best seen in FIG. 26. A slotted cylinder 511 is welded to the tube 501 and includes a slit 513 which is aligned with the slit 503 in the tube 501 prior to welding the tube 501 and the slotted cylinder 511 together. As may be seen through a comparison of FIGS. 25 and 26, the slotted cylinder 511 includes a first portion 515 which includes no slit therein and is completely cylindrical in nature so as to provide a guide mechanism for a slot control rod 517. A further portion 519 of the slotted cylinder 511 includes not only the slit 513, but has 180.degree. of its circumferential extent removed, as best seen in FIG. 26. Accordingly, the rod 517 may be reciprocated in guiding relation with the portion 515 and the end 518 of the rod 517 will overlie adjustable portions of the slit 513 to thereby control the opening thereof. A plate 521 is provided at one end of the apparatus 500 and has connected thereto the portion 519 of the slotted cylinder 511, one end of the tube 501 and one end of the tube 505. A hole 523 is provided through the end plate 521 for the purpose of inserting therethrough in mounting relation a pyrometer (not shown) which is utilized to measure the radiation within the tube 501 and therefrom to determine the temperature within the tube 501. The pyrometer may be utilized to sense the temperature therein and from this sensed temperature to control the current supplied to the apparatus 500 to thereby control the temperature within the tube 501. An insulator 525, which may, if desired, be made of boron nitride is installed between the tubes 501 and 505 at the end thereof opposite the end plate 521. An electrically conducting plug 527 is inserted into that end of the tube 501 adjacent the insulator 525 and a power lead 529 is electrically attached thereto. Further, an electrically conducting device 531 is attached to the outer tube 505 and a power lead 533 is connected thereto. Accordingly, an electrical circuit is created between the device 531, the tube 505, the end plate 521, the coating material contained within the tube 501 and the electrically conducting plug 527. A mounting bracket 535 is provided which enables the mounting of the apparatus in a suitable location. Further, the end plate 521 has mounted thereto shields 537. As shown in FIG. 26, in order to operate the apparatus 500, the material 539 which is to be coated on the radiation shields is inserted into the tube 501 in powdered form. In order to ensure proper operation of the device, a continuous line of powder 539 must extend from the end plate 521 to the electrically conducting plug 527 so as to complete the circuit. Thus, the device must be maintained in a level orientation so as to ensure that this electrical circuit is maintained in a complete condition. The slotted cylinder 511 acts as a nozzle with its slit 513 to control the direction of conduction of vapors caused by the evaporation of the powder 539 due to its heating by the electrical current which is supplied across the power leads 529 and 533, and which may be at a current level of approximately 100 amps. The apparatus 500 is specifically designed so as to enable the coating of the inner surfaces of these cylindrical shields, as well as the outer surfaces thereof. As shown in FIG. 26, a dashed line 543 is intended to be indicative of the shield with its interior surfaces being coated, whereas the dashed line 545 is intended to be indicative of a shield with the outer surfaces thereof being coated by the apparatus 500. In the operation of the apparatus 500, a source of current of approximately 100 amps is placed across the power leads 529 and 533. The slit controlling rod 517 is adjusted in a lateral fashion so as to enable the exposure of a predetermined longitudinal extend of the slit 513, with this adjustment depending upon the longitudinal extent of the shield which is to be coated by the apparatus 500. The pyrometer (not shown) is inserted into the hole 523 and as the current flows across the circuit melting the powder into a liquid extending between the end plate 521 and the plug 527, and being further heated to form vapors, the pyrometer senses the radiation in the vapors to thereby enable the temperature within the tube 501 to be determined, and control means (not shown) may be utilized to thereby control the current to control the temperature. Further radiation shields 538 are provided in a circumferential direction about the outer tube 505 so as to concentrate the heat formed by the electrical circuit within the tube 501. As the vapors are formed by the melting and evaporating of the powder 539, these vapors escape through the slits 503 and 513 and are guided by the portion 519 of the slotted cylinder 511 onto the surface of the cylindrical shield which is being coated thereby. The surface which is being coated is attached to a device (not shown) which enables the shield to be slowly rotated with respect to the slit 513 to thereby ensure a uniform coating thereof. In a similar manner, the disc portions of the shields may be coated by suspending them over the portion 519 of the slotted cylinder and rotating them with respect to the slit 513 to thereby ensure uniform coating thereof. With reference now to FIGS. 30 and 31, a further modification of the concepts taught in the parent file is set forth. A comparison of FIGS. 2 and 12 reveals that the heater depicted in FIG. 2 supplies only radiant heat, whereas the heating element of FIG. 12 provides both radiation and emission. In this vein, the embodiments described herein with regard to FIGS. 16 and 20 may also be modified so as to provide both radiation and emission-type heating. FIG. 30 shows the thruster 600 as including fuel supply passages similar to those shown in FIG. 16. Of course, as desired, the fuel supply passages as depicted in FIG. 20, may be utilized with this particular embodiment. As shown, the thruster 600 includes a radiation heating element 601 and an emissive heating element 603. As shown in FIG. 30, the energy absorber component 609 is comprised of a plurality of substantially concentric cylinders and the emissive heating element 603 is designed to comprise a plurality of concentric cylindrical members which interleave with the cylindrical members in the energy absorber component 609. As shown in FIG. 31, one possible power supply scheme for the heaters of the embodiment of FIG. 30, comprises a common power supply 611 which supplies the radiative heating element 601 via the switch 613 and the emissive heating element 603 via the switch 615. In this manner, easy control of the heating elements is possible. If desired, separate power supplies for each of the radiative heating element 601 and the emissive heating element 603 may be provided. A preferred mode for operating the embodiment illustrated in FIGS. 30 and 31 is as follows: (2) The power radiated from the coil 601 heats the thermionic element 603 and the energy absorber component 609 which is included in the circuit for the emissive heating element 603. (3) Once the temperature of the emissive heating element 603 is above approximately 1650.degree. K. (approximately 2500.degree. F.), the switch 615 may be closed so that great numbers of electrons are emitted by the thermionic emitting material of the emissive heating element 603 which may comprise, for example, thoriated tungsten. The electric field between the emitter 603 and energy absorber component 609 accelerates these electrons toward the energy absorber component 609 where they impact and are absorbed. (4) This electron flow constitutes an electrical current I which flows across a potential drop V equal to that of the power source, such as a battery, thereby delivering energy to the energy absorber component 609 at at rate given by P=VI; virtually all of this energy is deposited in the thermal absorber adjacent to the nozzle. (5) This power P heats the energy absorber component 609 to temperatures above that of the emissive heating elements 603. 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 emissive heating element 603 thereby supplying the work function energy to maintain the electron emission and temperature of the emissive heating element, which in turn keeps the electric current flowing. (6) The gaps between the emissive heating elements 603 and the energy absorber component 609 which comprise the interstices between the respective concentric cylindrical portions thereof, are specifically designed to values which control the level of current at the space-charge limited level given by the Child-Langmuir equation as described hereinabove. (7) Once the design current is flowing and the steady state operational temperature with propellant flowing is established, the switch 613 may be opened to thereby permit the coil to cool down to the temperature of the heat exchanger. Since the coil is used essentially as an initiator, being elevated in temperature for only a few minutes each firing, the coil lifetime may be many hundreds of thruster operational hours before it begins to sag and then touches another thruster component to thereby fail. (8) When the firing of the thruster is to be terminated, after for example, 40 to 60 minutes, the switch 615 may be opened and the flow of propellant may be stopped by closing a propellant supply valve. With reference now to FIGS. 28 and 29, a further aspect of the present invention will be described. In prior art thrusters, in order to supply the thruster with vaporized, preheated and/or decomposed fuel, a separate pre-heater or decomposer assembly was necessary upstream of the fuel supply conduit extending through the outer housing of the thruster. A principal problem of the prior art of such assemblies is solved by the device set forth in FIGS. 28 and 29. It is noted that although great care is exercised in the manufacturing and handling of propellants and oxydizers to be used for long missions on spacecraft or satellites, some impurities inevitably are found in the propellant. Some of these are in the form of metallic oxides, carbonates, and/or other compounds that, when deposited on feed tube surfaces, adhere thereto and having a comparatively low vapor pressure, cannot be vaporized off these surfaces or readily removed by other means. These deposited impurity compounds have been known by the term non-volatile residues. If the supply tube temperatures or the injector orifice temperature rises above the boiling point of the propellant and/or oxydizer, then nucleate boiling of the liquid adjacent to the wall will occur. Experience has shown that when nucleate boiling occurs, these non-volatile residues deposit on the hot walls and if the surface area where nucleate boiling occurs is small, such as at the location of a fuel supply injection orifice, then the non-volatile residues will build up and partially or wholly block the propellant or oxydizer supply tube. This process of adherence of non-volatile residues to the supply tubes has been identified by those skilled in the art as the probable cause of blockages which have been observed in the feed tubes of low thrust hydrazine engines of the prior art. The invention shown in FIGS. 28 and 29 achieves the object of (1) ensuring that the temperature of the injection tubes during propellant flow never exceeds the boiling point of the propellant and/or oxydizer and (2) ensures that a high percentage of the power conducted, convected or radiated to the feed tube is regeneratively returned to the decomposition or reaction chamber with the injected fluid. In this vein, the wires 713 which are welded into the mixing chamber 709, serve several purposes: (1) The wires intersect all parts of the flow of fuel, thus permitting energy transfer of the total flow of fluid. The wires 713 will also accomplish some mixing of the fluid tending to give the heated fluid a relatively uniform temperature. (2) The wires 713 increase significantly the surface area available for heat transfer from the metal components to the fluid. (3) The combination of aspects discussed above in (1) and (2) permit the fluid flow to absorb significant quantities of power, even as much as 20 watts at the lowest flow rates, to thereby reduce the temperature of the metal components of the mixing chamber to values below the boiling point. With no nucleate boiling, there will be little or no non-volatile residue build-up on the walls of the mixing chamber 709 or on the surfaces of the wires 713. One further feature is noted, to wit, the internal diameter of the injection orifice 705 is specifically sized to get the desired injection velocity of the fluid. Internal radiation shields 721 preferably made of tungsten act to reduce or prevent convective energy transfer from the hot decomposition products or reacted propellant gases to the injection orifice 705. The liquid mixing and injecting device 700 shown in FIGS. 28 and 29 includes a fluid inlet 701 connected to an outlet 703 via an injection orifice 705. The fluid inlet 701 terminates at a diverging flow passage 707 which leads to a mixing chamber 709 which connects to the injection orifice 705 via a converging flow passage 711. Within the mixing chamber 709, a plurality of wires 713 are welded so as to enhance the heat transfer therein as well as fluid mixing. If desired, the wires 713 may be made of tungsten-rhenium. As best shown in FIG. 29, the wires 713 are oriented in circumferentially staggered relationship with respect to one another so as to provide a tortuous path for fuel flowing therethrough. The outlet 703 of the injector leads to a preheater decomposition chamber 715 which feeds the fuel to a screen pack or other decomposition and/or heat transfer structure 717 which may, if desired, be surrounded by a heater source 719. The screen pack or other decomposition and/or heat transfer structure 717 comprises a decomposition and/or heating structure and the fuel flows through the screen pack or structure 717 and thence into the fuel inlet of the thruster, for example, denoted by reference number 361 in FIG. 16. Accordingly, the device shown in FIGS. 28 and 29 has been developed for attachment to the inlet pipe of the preheater and/or decomposers 715 or directly into the thruster housing which is shown in FIG. 16 with reference numeral 361, in FIG. 20 with reference numeral 361' and in FIG. 25 with reference numeral 661. See, in this regard, FIG. 47. In the prior art there was little attempt to thermally isolate and achieve optimum energy efficiency from the preheater/decomposer and/or to provide vaporized propellant to the thruster at ideal decomposition or thermal state. This was due to the concern for the blockage problem just discussed and now solved with the feature of the present invention set forth hereinabove. With this solution, optimum designing calls for the preheater to operate with minimum thermal loss from within its self-contained heater source or from the chemical energies released from an exothermic decomposing propellant. This is achieved by thermally isolating the preheater/decomposer 700, FIG. 28 with radiation shielding 721 as shown and by using other standard thermal isolating techniques for this objective. Further, it is advantageous for optimum augmenter performance to provide as near as possible fully reacted (dissociated) propellant (in the case of hydrazine to have most of the intermediate reaction ammonia dissociated) out of the decomposer. To achieve this, the screen pack or other decomposition element 717 should be sufficiently long and operated at adequate temperatures to achieve this end. This is contrary to operation of typical decomposing thrusters which function without electrical enhancement of performance. Such thrusters minimize ammonia decomposition. This assembly 200 can also be utilized to provide additional, auxiliary preheating of a propellant such as hydrogen and/or to effect desirable chemical reactions prior to entrance into the high temperature thruster. NOZZLE DESIGN--PRIOR ART Most rocket nozzles have been designed for maximum thrust when exhausting to an ambient pressure higher than one Torr. At one Torr the mean free path is of the order of 2.times.10.sup.-3 cm, several orders of magnitude smaller than the throat diameter, or the boundary layer thickness of the gas near the nozzle exit. When these conditions prevail, the nozzle shape that results in the highest performance is one that has the invisid or "core" flow flowing parallel to the axis of the nozzle at the exit of the nozzle. This, then, results in a "bell-shaped" nozzle as the optimum configuration provided that near the exit: EQU p.sub.w sin .theta.&gt;.tau..sub.w cos .theta. (26) For example let: ##EQU19## This indicates that if the expansion half angle of the nozzle, .theta., is less than 30.degree. near the exit, then the thrust of the rocket is decreasing as the gas expands further, due to the preponderant effect of the shear stress term ##EQU20## Most, if not all, nozzles on rockets tend to terminate when some such condition is reached. Also, the angle .theta., at the nozzle exit is usually considerably smaller than 30.degree. in order to straighten the "core" flow. SPACE OPTIMIZED NOZZLE DESIGN A typical nozzle designed in this manner has a contour similar to that shown in FIG. 45. Also, in the figure, the pressures and area ratios at various positions along the nozzle are indicated. Also, drawn in phantom in FIG. 45 is an example of a nozzle designed in accordance with the present invention. If the ambient gas is much lower than one Torr and the gas in the boundary layer is expanded further to have a mean free path that is comparable with the boundary layer thickness, then a further condition on the angle .theta. can be calculated. In this region, the shear stress .tau..sub.w can be written a ##EQU21## where: .tau..sub.w =shear stress at the wall When the gas is expanded to this extent, then: ##EQU22## Even if M is only unity and .gamma.=1.40 EQU tan .theta.&gt;1.89 EQU .theta.&gt;62.degree. In general, the Mach number M at the edge of the boundary layer will be greater than unity hence the expansion half angle will be greater than 62.degree. in order to increase the thrust by expanding to high area rations. The above considerations indicate that the nozzle design procedure shold be as follows: Step I--use conventional design procedure to establish the contour between the throat and the area ratio at which ##EQU23## i.e., a point at which the thrust gain is twice the shear stress loss (the reason for the "2" on the formula). This point is chosen because shortly beyond this point, thrust loss due to shear stress will dominate over thrust gain due to pressure. This point is shown in FIG. 45. Step II--continue increasing the angle .theta., as needed, to maintain the same ratio between the thrust gain and the shear stress loss, i.e., a ratio of 2/1. Step III--it may be possible to "fine-tune" the angle .theta. as a function of the area ratio in order to get even more thrust. Step IV--at very large area ratios (greater than 2000) let .theta.=90.degree. and continue this flat plate nozzle out to the largest practical diameter. As discussed hereinabove, several improvements to the design of nozzles in the thruster art would be helpful in increasing the efficiency and life expectancy of thrusters. In this vein, theoretical aspects of nozzle design and analysis were discussed hereinabove in the Summary of the Invention. With reference now to FIGS. 38, 39 and 40 a few applications of theory with regard to nozzle design for thrusters will be set forth in greater detail. The nozzle configuration shown in FIG. 38 is specifically designed for continuous heating through M.perspectiveto.2. The design criteria for the nozzles shown in FIG. 38 were discussed hereinabove in the Summary of the Invention, and these design criteria are repeated here for convenience as follows: (1) The stagnation pressure in the heat exchanger should be as high as possible. (2) 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.revreaction. by including considerations of radiation power loss from the nozzle. 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 D.sub.d in the following form: ##EQU24## where C.sub.d =discharge coefficient How the value of the discharge coefficient affects the thrust coefficient is not immediately obvious. This will be investigated by developing a novel method of computing the thrust coefficient. The thrust on a rocket, F, operating in a vacuum, can be computed by two methods: 1. Evaluating the integral: ##EQU25## where p=gas pressure at the nozzle exit or, 2. Integrating the stress tensor over the axial projection of all interior and exterior surfaces. The approach adopted here will be to compute the thrust that is generated up to the 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 are identified as follows: EQU C.sub.F =F/p.sub.c A* (33) where where ##EQU26## Since the velocity w* is purely axial at the throat, cylindrical coordinates are used in computing C.sub.F. In practice .DELTA.C.sub.F can be analytically maximized by choosing various configurations (.theta. as a function of R) and nozzle surface temperature distributions (.tau..sub.w) as a function of R) and then using numerical procedures to solve the Navier-Stokes equations in the nozzle. Since the different gases have different thermo-dynamic and transport properties, the nozzle shape may change from gas to gas. Also, since relaxation effects in the gas (e.g, atom-atom recombination) depend upon pressure and residence time of the gas in the nozzle, the best nozzle shape may change for any given gas with the chamber pressure of the gas. As pointed out elsewhere, best nozzle performance will be obtained by making the chamber pressure of the gas as high as possible. In order to obtain the most accurate results for the nozzle design downstream of the throat, equation 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. For invisid gas, accelerated at constant enthalpy and with a conical diverging nozzle, the integrals can be evaluated. The results are: ##EQU27## Equations 36 through 38 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: ##EQU28## The expression indicates that the radius of curvatures 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 35. Immediately downstream of the throat there will be a negative increment to C.sub.F since p.sub.w sin .theta.-.tau..sub.w cos .theta.&lt;0. 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 (40) Eventually the nozzle 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 of the nozzle axis. Since the disc part and conical end piece of the nozzle can be made from extremely thin sheet material, the weight thereof can be kept low and the conical end piece should extend to the maximum diameter permitted. 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 Reybold's number. In light of the discussion hereinabove with regard to test data and design implications of nozzles, nozzle configurations worthy of analytical investigation are shown in FIGS. 38, 39 and 40. TEST PROCEDURE Inasmuch as the thruster 11, according to the present invention is designed to operate in an outer space environment, a special test facility is provided for ground testing, as shown in FIG. 42. The thruster 11 is placed in a vacuum enclosure 121, with an outside exhaust duct 123 provided in communication with expansion nozzle 19. It is recognized that the provision of the outside exhaust duct 123 would create a sea-level ambient pressure condition at the expansion nozzle 19 and the fuel passages 17, particularly at times when propellant is not being supplied to the thruster 11. The remainder of the thruster 11 is exposed to a vacuum created by sorption pumps 127 and finally by ion pumps 127 which evacuate the vacuum enclosure 121. Because of the terrestrial gravity environment to which the heater may be rotated at about 30 RPM or greater, repositioned 180 degrees as needed to compensate for any sag of coil 33 or a magnetic coil 129 may be implemented in surrounding relation to the thruster 11. The magnetic coil arrangement 129 serves to support the heater coil 33 to the extent necessary to counteract the force of gravity. Force supplied by the magnetic coil arrangement 129 is calculated to provide a force equal and opposite to that of the acceleration of gravity on the material of the heater coil 33 when it is at appropriate operating temperatures. The magnetic coil arrangement 129 is also modulated to an extent necessary to reflect changes in forces in the coil 33 during conditions of acceleration caused by the thrust of the thruster 11. Other provisions for testing in the vacuum enclosure 121 include special power features 133 and a propellant inlet supply 135. Testing will occur at a pressure of less than 10.sup.-5 Torr. What has been described are preferred embodiments of the invention. It should be noted that it is possible to provide various other arrangements. For example, while an air vent opening for the heater cavity 18 has been described, it is also possible to seal the heater assembly 13 with the heat exchanger assembly 15 in a vacuum, with a small amount of pressurant being permitted to remain within the heater assembly 13. This pressurant would affect the vaporization rate of material from the heater coil 33. It is also possible to provide various arrangements for the expansion nozzle 19 in accordance with the specific needs and application of the thruster 11. Further, if the fuel is injected into the thruster 11 in the unreacted liquid state, the heater assembly 13 is used to pre-heat the thruster 11 to a safe temperature above the thermal decomposition temperature, about 1000.degree. K. This internal coupling of the exothermic decomposition and the electrical performance augmentation eliminates the heat losses from an externally mounted decomposition chamber and from the connecting injection tube. The present invention may be utilized with such propellants as H.sub.2, N.sub.2, N.sub.2 H.sub.2, NH.sub.3, CO.sub.2, CO, CH.sub.3, H.sub.2 O, etc. Accordingly, the above description is not intended to be limiting, but is, instead, intended to be exemplary in nature.