Patent Publication Number: US-6216097-B1

Title: Power measuring cooling plant system and method

Description:
BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     The present invention relates to a thermal balancing system, and in particular to a thermal balancing system and method for terrestrial thermal vacuum testing of spacecraft to simulate on orbit conditions by the replication of the heat transport functions of heat pipes which are dysfunctional due to gravitational forces. The present invention also relates to a system and method for preheating a spacecraft prior to terrestrial thermal vacuum testing, and a method for design of a cooling plate utilized by the above systems. 
     2. Background of Related Art 
     Unmanned space vehicles, in particular satellites, are now used for everything from telecommunications to weather reporting to missile defense. With this growing demand has come increased size, complexity, and capability resulting in greater cost. As a result, ever increasing amounts of terrestrial, ground-based, testing are required to ensure that such spacecraft will function properly throughout their effective life. 
     In its basic elements, a spacecraft can be broken up into six functional groups or subsystems: the spacecraft structure; the power system; the attitude control system; the telemetry, tracking, command and communications system; the propulsion system; and the thermal control system. The thermal control system is an important element as its purpose is to achieve thermal balance of the spacecraft, permitting proper performance of all subsystems during the spacecraft&#39;s proposed operational modes. As one might expect, through the various phases of a spacecraft&#39;s mission there are variations in heat generation and dissipation due to internal component operation and surface fluxes caused primarily by solar heating and eclipse. As each on board system has a thermal operating range, it is the responsibility of the thermal control system to maintain the temperature of subsystems within their best operating range. 
     During the initial design phase of a spacecraft, analytical thermal system models are developed. The models not only predict a thermal system balance temperature but also define heat transfer paths for which heat dissipation devices can be strategically placed to discard the waste heat. While these models are adequate for design, the predictions must be verified by actual system testing. 
     As there is no or negligible air at spacecraft orbital altitudes, the primary heat transfer mechanisms are conduction and radiation. Conductive heat transfer is achieved through material and surface-to-surface contact, while radiative heat transfer is primarily achieved through the use of radiator panels which radiate the heat into deep space. Excess heat generated by internal components or solar heating is conducted from the source to the radiator surfaces, which then rejects the heat to space. While there are a variety of convective heat transfer methods to move the excess heat from the source to the radiative panels, one preferred method is a heat pipe. The advantage of heat pipes is that they typically exhibit thermal conductivity greatly in excess of most thermally conductive metals and transport thermal energy at efficiencies greater then 90%. A heat pipe is a closed loop system which contains an evaporator section, a condenser section, and a working fluid. Should external geometric requirements make it necessary, an adiabatic transport section can be included to further separate the evaporator and condenser portions. Internal to the heat pipe are typically a series of axial grooves which provide a transport path for condensed fluid to flow from the condenser section to the evaporator section. 
     The heat pipe is mounted such that the evaporator section is attached to a heat source whose temperature is to be maintained within a predetermined range. The condenser section is generally attached to a heat sink such as a space heat radiator for radiation of the heat to space. It is not uncommon for a spacecraft to have several heat pipes, or banks of heat pipes, located throughout the vehicle and oriented in a variety of positions depending upon heat sources and heat sinks. 
     In operation, the heat generated by a heat source is absorbed by the working fluid in the evaporator section of the heat pipe, vaporizing the working fluid. The vaporization causes an increase vapor pressure difference resulting in the vaporized fluid traveling, inside the heat pipe, from the evaporator section to the condenser section. Since the condenser section is attached to a heat sink, the vaporized fluid reaching the condenser section is cooled, causing it to condense and release the latent heat of vaporization. The condensed working fluid is transferred back to the evaporator section by capillary action, where it is again evaporated, absorbing the heat from the evaporator section. 
     In a substantially gravity-free space environment, the transfer of the working fluid over the length of the heat pipe is not a problem. However, on Earth gravity may inhibit the return of the working fluid, especially if the condenser section is located below the evaporator section. This is because capillary action of the working fluid in a one-G environment is generally limited to a vertical rise of about 13 millimeters or less. As a result, the functionality of a heat pipe in a terrestrial environment is directly affected by its orientation. 
     Terrestrial thermal system testing of spacecraft, otherwise known as thermal vacuum testing, is typically done in a large vacuum chamber capable of housing an entire satellite at vacuum levels lower than of 1×10 −5  torr. As a result of the limited capillary rise of a working fluid in a one-G environment, thermal vacuum testing can be quite difficult and give incomplete results. This is due to the fact that often one or more of the heat pipes are oriented in such a manner that they are functionally inoperable because of the gravitational force. While data can be acquired by those systems that are functional, such testing cannot develop a complete thermal balance for the entire vehicle. As a result, either multiple tests run must be conducted with the spacecraft manually repositioned to different orientations, or the spacecraft must be mechanically rotated during test. While these approaches provide additional data with respect to the functionality of the thermal control system, such testing can still not provide a complete thermal balance since not all systems are functional at the same time. In addition to the inability to provide a thermal balance, such alternate approaches are costly both in material and man-hours. Each manual repositioning of the satellite requires the test chamber to be vented to the atmosphere, the satellite rotated, a full pump-down cycle completed, and the spacecraft thermally stabilized. It can take from 12 to 48 hours between successive tests. If the satellite is rotated within the chamber during test, special tooling must be designed and manufactured. This tooling must not only be capable of supporting and rotating the satellite during test, but also withstanding the thermal and vacuum environment. 
     There is thus a need for method that would permit full thermal system testing of spacecraft in a single test cycle and provide a system thermal balance. Additionally, such a system should be able to provide a full system balance within a margin of error of 10% or less, preferably 5% or less. 
     SUMMARY OF THE INVENTION 
     The present invention is directed to a system and method for terrestrial thermal vacuum testing of spacecraft which is capable of developing a system thermal balance within a 5% error. This is accomplished by attaching a cooling plate to a heat source on a spacecraft which extracts heat from the source, conducts it through a series of working fluid lines and discards the heat at a cooling cart. Heat flow is calculated by measuring the temperature of the working fluid both before and after entering the cooling plate and measuring the fluid flow rate for the system. Based on the initial heat flow calculation an equivalent amount of heat is added back to the spacecraft at an appropriate location thus simulating on-orbit conditions. A second measurement is read for the system and the corresponding heat load delivered. This process is repeated until thermal balance is achieved, i.e. the measured heat flow values do not change. The system is preferably used in situations in which heat pipes onboard the spacecraft are disfunctional due to Earth&#39;s gravitational force. 
     The present invention also addresses the design and selection of the cooling plate which is based on both the physical constraints of the spacecraft and the proposed heat load. Based on the data derived from analytical system thermal models, the physical constraints of the cooling plate, the working fluid chosen and the pressure limitations of the system, a determination is made as to the number and size of through fluid passageways needed in the cooling plate. The cooling plate is fabricated based on the results of this analysis. 
     The present invention is also directed to preheating a spacecraft prior to terrestrial thermal vacuum testing. In such a case, the above-described system is operated prior to spacecraft activation. A heat extraction system in the cooling cart is deactivated and a heating unit, also located within the cooling cart, is operated to heat the working fluid, which is pumped into the cooling plate passing the heat to the spacecraft. The cooled working fluid is then returned to the cooling cart to be reheated, continuing the cycle. The use of the system to preheat the spacecraft reduces the time required for the spacecraft to reach its specific thermal state before commencing thermal vacuum testing, resulting in shorter test cycles and less cost. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     These and other features, aspects, and advantages of the present invention will be apparent to those skilled in the art from the following detailed description together with the accompanying drawings in which: 
     FIG. 1 is a block diagram of the power measuring cooling plate system; 
     FIG. 2 is a flow chart depicting the steps for the design and selection of a cooling plate; 
     FIG. 3 is an elevation view of a heat pipe system; 
     FIG. 4 is a broken perspective view of the cooling plate; 
     FIG. 5 is a broken sectional view of the cooling plate taken along section lines  5 — 5  of FIG. 4; 
     FIG. 6 is a schematic diagram of the power measuring cooling plate system; 
     FIG. 7 is a flow chart depicting the steps for conducting a thermal balance for a spacecraft; 
     FIG. 8 is a partially perspective view of a single bayonet cooling plate; 
     FIG. 9 is a broken sectional view of the cooling plate taken along section lines  9 — 9  of FIG. 8; and 
     FIG. 10 is a partial perspective view of an alternate bayonet cooling plate configuration. 
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The present invention is for a system and a method for conducting system thermal balancing of a spacecraft housed in a thermal vacuum chamber. The present invention is typically used in situations in which heat transfer systems on a spacecraft are dysfunctional due to a combination of their orientation and gravity. The system of the present invention extracts heat from a heat generating surface on a spacecraft, determines the heat flow from the surface and supplies an equivalent amount of heat back to the spacecraft at a specified location. This process is repeated until a point at which the measured heat flow remains constant, this is the thermal balance point. 
     The system of the present invention, as depicted in FIG. 1, is used for terrestrial thermal testing of a spacecraft  20  located within a thermal vacuum chamber  22 . Attached to a heat generating surface on spacecraft  20  is a cooling plate  24  which in turn is connected to a cooling cart  26  via a set of working fluid lines  28 . A working fluid  30  is pumped, by a magnetic gear drive fluid pump  32 , between cooling plate  24  and cooling cart  26  receiving heat from spacecraft  20  and discarding the added heat at a heat exchanger  34 . If necessary, working fluid  30  can be heated by an immersion heater  35  to compensate for excess heat dissipated by heat exchanger  34 . Heat flow from spacecraft  20  is determined by a data acquisition and control computer  36  based on the temperature data received from a plurality of temperature sensors  38  and fluid flow rate data received from a turbine flow meter  40 . An amount of heat equivalent to that extracted by cooling plate  24  is reintroduced back to spacecraft  20  at a secondary location by spacecraft heater  42  simulating on orbit conditions. The above process is repeated until a point at which measured heat flow values remain constant, at this point thermal balance is achieved. 
     The use of the power measuring cooling plate system begins with the design and selection of the cooling plate  24  as depicted in FIG.  2 . During the initial stages of a spacecraft&#39;s design several computer models are developed. These models provide the cooling plate designer with predicted heat load requirements and dimensional constraints for spacecraft  20  (step  200 ). The heat load requirements consist of the predicted heat flow ({dot over (Q)}) for spacecraft  20  and the predicted temperature of the cooling plate mounting surface on the spacecraft (T sc ) to which the heat is to be conductively transferred. The dimensional constraints of spacecraft  20  determine the maximum size of cooling plate  24 , in particular its cooling passageway length (L 1 ). Based on the proposed heat load, an initial working fluid  30  is selected (step  202 ). Once initial working fluid  30  is selected and the physical constraints of the cooling plate  24  are defined, an initial set of assumptions are made, including: ΔT (temperature of working fluid  30  measured at the supply end of cooling plate  24 —temperature of working fluid  30  measured at the return end of cooling plate  24 ; N 1  (number of fluid passageways  43  within the cooling plate  24 ); and D 1  (the diameter of base plate through fluid passageways  43 ) (step  204 ). The accuracy of temperature sensing device  38  governs the ΔT value, which is chosen to maximize the measurement accuracy. ΔT out-in  is chosen such that the measurement accuracy error between ΔT and {dot over (V)} (volumetric flow rate of working fluid  30 ) is balanced to minimize root mean squared error factor. Based on the above defined constants ({dot over (Q)},L 1 ), the physical properties of the working fluid  30 , and the assumed variables (ΔT, N 1 , D 1 ), the equations 1, 2, 3, and 4 are simultaneously solved (step  206 ):                V   .     =       Q   .       ρ                   C   p        Δ                   T     OUT   -   IN                   (   1   )                         
     where:                  Δ                   T     OUT   -   IN         =         T     f   ,   out       -     T     f   ,   in         =       Q   .       ρ                   C   p          V   .                               Re   1     =           V   1          D   1       ν     =       (     4   π     )          (       V   .         N   1          D   1        ν       )                   (   2   )                       Δ                 p     =                    1   2        ρ                       V   1   2          [         f   1          (       L   1       D   1       )       +       ∑     j   =   1     m            (     K   LOSS     )     j         ]         Within     Cooling   Plate           +                                1   2        ρ                       V   2   2          [         f   2          (       L   2       D   2       )       +       ∑     l   =   1     q            (     K   LOSS     )     l         ]           Outside                 of       Cooling                 Plate                       ≈                    1   2              ρ        (       4        V   .         π                   N   1          D   1   2         )       2          [         (     0.0813   -     0.01216                     log   10          [     Re   1     ]           )          (       L   1       D   1       )       +       K   LOSS          [     2   ×       L   1       L   ref         ]         ]         +                                1   2              ρ        (       4        V   .         π                   D   2   2         )       2          [         (     0.0813   -     0.012161                     log   10          [     Re   2     ]           )          (       L   2       D   2       )       +       K   LOSS          (     N   2     )         ]                       (   3   )                         
     where:          Re   2     =           V   2          D   2       ν     =       (     4   π     )          (       V   .         D   2        ν       )                        f≅0.0813−0.01216log 10 (Re) 
     
       
         
           
             
               
                 
                   
                     
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         {dot over (Q)}=ηh f A fs (T s −T f,avg )=h f A fs (T fs −T f,avg ) 
       
     
     for turbulent flow, where Re 1 &gt;2300 and where            νρ                   c   p       k     &gt;   0.5                             h   j          A   fs       =     0.023            4   0.8          π   0.2          k   0.67          V   .          ρ   0.33          c   p   0.33          N   1   0.2          L   1        1.0         ν   0.47          D   1   0.8                     T     f   ,   avg       =         T     f   ,   out       -     T     f   ,   in         2                     h   f     ≈                  k     D   1            [     0.023          (         V   1          D   1       ν     )     0.8            (       νρ                   c   p       k     )     0.33       ]                   =                  k     D   1            [     0.023          (       4        V   .         π                   N   1          D   1        ν       )     0.8            (       νρ                   c   p       k     )     0.33       ]                   =                0.023        [         k   0.67       D   1   1.8              (     4   π     )     0.8            (     νρ                   c   p       )     0.33       ]                   =                0.023          (     4   π     )     0.8              k   0.67            V   .     0.8          ρ   0.33          c   p   0.33           ν   0.47          N   1   0.8          D   1   1.8                                 
     Where: 
     {dot over (V)}=volumetric flow rate of working fluid  30  (m 3 /sec); 
     {dot over (Q)}=heat flow (watts); 
     ρ=density of working fluid  30  (kg/m 3 ); 
     c p =heat capacity of working fluid  30  (J/kg ° k); 
     T f,out =temperature of working fluid  30  at the outlet of cooling plate  24  (° k); 
     T f,in =temperature of working fluid  30  at the inlet of cooling plate  24  (° k); 
     Re 1 =average Reynolds number of a fluid passageway  43  within the cooling plate  24  (unitless); 
     Re 2 =average Reynolds number of a fluid passageway outside of the cooling plate  24  (unitless); 
     f=friction factor within the fluid passageway (unitless); 
     V 1 =average fluid velocity within the cooling plate flow passageway  43  (m/sec); 
     V 2 =average fluid velocity within the cooling line outside of cooling plate  24  (m/sec); 
     ν=kinematic viscosity of working fluid  30  (m 2 /sec); 
     D 1 =diameter of a fluid passage way  43  (m); 
     N 1 =number of fluid passageways  43  (unitless); 
     Ac 1 =cross-sectional area of fluid passageway  43  within cooling plate  24  (m 2 ); 
     ΔP=total pressure drop across the power measuring cooling plate system (N/m 2 ); 
     L 1 =combined length of fluid passageways  43  (m); 
     L 1,ref =length of a single passageway  43  of fluid within the cooling plate  24  (m);            2        [       L   1       L   ref       ]       ≅     number                 of                 90      °                 bends                   (   unitless   )         ;                   
     number of 90° bends (unitless); 
     K LOSS =head loss coefficient (unitless); 
     L 2 =length of working fluid line  28  and bayonets  44  (m); 
     D 2 =average inside diameter of working fluid line  28  and bayonets  44  (m); 
     Ac 2 =cross sectional area of fluid passageway outside of cooling plate  24  (m 2 ); 
     N 2 =number of 90° bends outside of cooling plate  24  (unitless); 
     h f =heat transfer coefficient of working fluid  43  within the cooling plate fluid passageway  43           (     W     °                   Km   2         )     ;                   
     A fs =surface area of cooling fluid passageway  43  within the cooling plate  24  (m 2 ); 
     T fs =fluid passageway surface temperature within the cooling plate  24  (° k); 
     T f,avg =average fluid temperature within the cooling plate  24  (° k); 
     η=thermal efficiency of cooling plate  24  (unitless); 
     T s =mounting surface temperature on the cooling plate where it will be mounted to spacecraft  20  (° k); 
     T sc =mounting surface temperature on the spacecraft  20  where cooling plate will be mounted (° k); and 
     G sc-s =interface thermal conductance between the cooling plate  24  and the spacecraft mounting surface (w/° k). 
     The above equations 1, 2, 3 and 4 are iterated by changing the variables D 1  and N 1 , and if necessary ΔT until; the Reynolds number is greater then 2300 (step  208 ) to assure that the flow is in the turbulent region to maximize heat transfer, ΔP value is less then maximum operating pressure for the system (step  210 ), and T sc  is less then the predicted temperature of the transfer surface (step  212 ). When these conditions have been satisfied the corresponding D 1  and N 1  then define the physical structure of base plate  45  for fabricating cooling plate  24  (step  214 ). 
     Once cooling plate  24  has been fabricated, the face surface  46  is attached, adjacent to, or directly upon, the spacecraft&#39;s heat generating surface  47 . Cooling plate  24  can be attached to a variety of surfaces on spacecraft  20 , commonly to a spacecrafts heat pipe  49  or towards the evaporator section  50  of heat pipe  48 . At, or adjacent to, the condenser end  52  of the heat pipe  48  is commonly located spacecraft heater  42  used to introduce heat back to the radiative surface  54  of spacecraft  20 , see FIG.  3 . 
     The method of attachment of face surface  46  to surface  47  can vary depending on the expected ΔT interface  which is equal to T sc −T s . If ΔT interface  is large, the sensitivity with respect to the measurement of heat transfer is not as critical and as such, thermal bonding between the surfaces ( 46 / 47 ) is also not as critical and thus a variety of techniques know to those skilled in the art can be used, for example bolting or adhesives. It is desired to have ΔT interface  small, so the bond interface between surfaces  46  and  47  must be a good thermal interface, for example a solder joint. If the method of attachment chosen is soldering, then depending upon the material selected for base plate  45 , special surface preparation might be required for surface  46 . 
     After attachment of base plate  45  to surface  47 , spacecraft  20  is placed in a secured test structure in thermal vacuum chamber  22 . A plurality of Bayonets  44 , namely a supply and return bayonet, are inserted into base plate  45  sealing the bore with an O-ring  56  against the base plate bore  58  and then rotated to lock bayonet locking flange  60  into base plate locking slot  62 . A locking fastener  64  is passed through holes  66  and  68  to prevent bayonet rotation. A temperature sensor  38  is installed in a first aft end opening  70  of both bayonets  44  and connected with a temperature sensor wire  72  to the inside surface of a chamber feed through  74 . A second aft end opening  76  on both bayonets  44  is connected to feed through  74  by a working fluid line  28  and leak checked, see FIGS. 4 and 5. 
     Once cooling plate  24  has been designed, fabricated and attached to the heat generating surface  47  of spacecraft  20 , it is connected to cooling cart  26  as shown in FIG.  6  and depicted in flow chart  7 . In FIG. 6, electrical connectors  78  and  80  are fastened to the outside surface of chamber feed through  74  and the data acquisition and control computer system  36 . A working fluid line  28  is also attached to the outside surface of chamber feed through  74  for both supply and return bayonet connecting them to cooling cart  26  (step  218 ). 
     Upon connection of cooling plate  24  to cooling cart  26 , working fluid  30  is added in sufficient quantity to fill the system and the working fluid reservoir  82 . Thermal vacuum chamber  22  is then sealed and the system pumped down to the required temperature and pressure levels (step  220 ). Upon achieving the required environmental conditions, several spacecraft systems are activated allowing the spacecraft to thermally stabilize at which point both spacecraft  20  and the power measuring cooling plate system of the present invention are fully activated and commence thermal vacuum testing (step  222 ). 
     During operation, magnetic gear drive fluid pump  32  circulates working fluid  30  throughout the system. Working fluid  30  flows from pump  32  through a pressure relief valve  84  which is used to protect the system from over-pressurization. The working fluid  30  flows over a set of thermal couple probes  86  immersed within the flow stream which measure the temperature of the working fluid  30 . The probes  86  send this information, via electrical connector  88 , to a proportional integral derivative (PID) controller  90  which among other things controls an immersion heater  35 . The working fluid  30  passes a manual by-pass ball valve  94  which diverts a portion of working fluid  30  reducing the flow rate to the rest of the system. The diverted working fluid  30  is reintroduced to the system at a point in front of single pass counter flow heat exchanger  34 . 
     The remaining portion of fluid  30  flows into a motorized flow control ball valve  96 , which, like valve  94 , is used to further regulate the flow rate of working fluid  30  entering the remainder of the system. Ball valve  96  is connected to and receives control signals from PID controller  88  via electrical connector  98 . The working fluid  30  exiting ball valve  96  flows through turbine flow meter  40  which measures the flow rate of working fluid  30 . Flow rate information is sent, via electrical connector  100 , to a flow indicator  102  which in turn sends the data, via electrical connectors  104  and  106 , to both the data acquisition and control computer system  36  and the PID controller  90  which are in turn connected by an electrical connector  105 . 
     The working fluid  30  exits cooling cart  26  and enters a working fluid line  28 , which is connected to chamber feed through  74 , passes through feed through  74  and into a second working fluid line  28  connected to second aft end opening  76  of bayonet  44 . The working fluid  30  enters the second end opening  76  where it flows over a temperature sensor  38  housing a Resistance Temperature Device (RTD) sensor  108 . RTD sensor  108  is connected and sends signals to a Hewlett Packard 3852 data acquisition unit (not shown) which in turn sends signals to an IBM compatible personal computer (not shown). 
     The working fluid  30  flows into base plate  45  where it receives heat conducted from spacecraft  20 , exits base plate  45  and enters return bayonet  44  where it flows over a second temperature sensor  38  housing a second RTD temperature sensor  108 . Like the first, the second RTD temperature sensor  108  sends a signal to the Hewlett Packard 3852 data acquisition unit which in turn sends a signal to the personal computer. Heated working fluid  30  exits return bayonet  44  entering a working fluid line  28  which is connected to chamber feed through  74 , passes through chamber feed through  74  and into another working fluid flow line  28  which is connected to the entrance of cooling cart  26 . Heat exchanger  34  cools the working fluid  30  by conductively passing the heat to a separate cooling fluid system  110 . 
     Cooling fluid system  110  supplies heat exchanger  34  with a cooling fluid, commonly liquid nitrogen, which is provided by a cooling fluid supply  112  passing through a cooling system shut-off manual ball valve  114  and through a cooling system cryogenic solenoid valve  116 . Ball valve  114  is used to shut off the flow of the cooling fluid to heat exchanger  34  and solenoid valve  116  is used to regulate the flow rate, which in turn regulates the amount of heat extracted from the working fluid  30 . Additionally, solenoid valve  116  is connected to, via electrical connector  118 , and controlled by, PID controller  90 . The cooled working fluid  30  exits heat exchanger  34  and enters an immersion heater  35  where, if necessary, is heated to make up for excess heat extraction caused by heat exchanger  34 . Immersion heater  35 , like solenoid valve  116 , is connected to, via electrical connector  120 , and controlled by, PID controller  90 . From immersion heater  35  working fluid  30  flows back through the fluid pump  32  completing the cycle (step  224 ). 
     The temperature of working fluid  30  is measured at both supply and return bayonets  44  and sent to the data acquisition unit, which in turn sends the data to the personal computer (step  226 ). The data acquisition unit also receives flow rate signals from flow indicator  102 , previously received from flow meter  40  (step  228 ), which are in turn sent to the personal computer. The personal computer calculates an initial heat flow ({dot over (Q)}) from spacecraft  20  based on above equation (1) and the extrapolated curvilinear equations for ρ and C p  of working fluid  30  (step  230 ). Based on the initial heat flow calculation, spacecraft heater  42  adds an equivalent amount of heat back to the spacecraft  20  (step  232 ), commonly at condenser section  52  of heat pipe  48 . A second measurement is then read for the system and the heat input adjusted to reflect the measured change (step  234 ). This process is iterated (step  236 ) until the thermal balance is achieved (step  238 ), i.e. the point at which an increase in heat supplied back to spacecraft  20  has no effect on the heat load determined by the power measuring cooling plate system. With a proper design of cooling plate  24  and selection of an appropriate turbine flow meter  40  and RTD temperature sensors  108 , a thermal balance can be achieved with an error of less then 5%. 
     In an alternate embodiment, as shown in FIGS. 8 and 9, the present invention is directed to cooling plate  24  consisting of a single bayonet  44  with two or more through fluid passageways for passing working fluid  30 . Base plate  45  contains one or more base plate through fluid passageways  43  forming a variety of shapes such as a loop. In a single bayonet system, bayonet  44  contains a first through fluid passageway  122  for suppling working fluid  30  to base plate  45  and a second through fluid passageway  124  for the return of heated working fluid  30  after passing through base plate  45 . Located at the forward end of the bayonet  44 , between the bayonets through fluid passageways, is a bayonet sealing device  126  such as a rubber gasket. The device seals against the bottom of base plate bore  58  creating a fluid tight seal and separating the inlet and outlet through fluid passageways. 
     In a second alternate embodiment, as shown in FIG. 10, bayonet  44  consists of two or more section, namely a first bayonet section  128  and second bayonet section  130 . The forward end of first bayonet section  128  is attached to base plate back surface  132  in the same manner as discussed above, or permanently attached by an attachment means such as welding or soldering. The aft end of first bayonet section  128  contain one half of a leak-free bayonet connector  134  with the second half attached to the forward end of bayonet section  130 . The aft end of the second bayonet section  130  contain two or more first openings  70  for attachment of temperature sensor  38  and two or more second openings  76  for attachment of working fluid line  28  as described above. 
     The present invention is also directed to a system for preheating a spacecraft prior to terrestrial thermal vacuum testing. In such a case, the power measuring cooling plate system of the present invention is operated prior to spacecraft activation with heat exchanger  34  deactivated, by electromagnetically closing valve  116 , and immersion heater  35  activated to heat working fluid  30 . Heated working fluid  30  is pumped, as described above, through the system entering cooling plate  24 , passing heat to spacecraft  20  and in return cooling working fluid  30 . The cooled working fluid  30  is then returned through the immersion heater  35  where it is reheated, completing the cycle. 
     The power measuring cooling plate system of the present invention has many advantages over presently available methods for thermal vacuum testing including the fact that the present system permits not only the generation of a system thermal balance, but also provides a method for developing a system thermal balance within an error of 5% or less. In addition, the above system and cooling plate provide a method for preheating a spacecraft situated in a thermal vacuum chamber, thus reducing the time and cost required for a spacecraft to reach its specific thermal state before commencement of thermal vacuum testing. 
     Although the present invention has been described in considerable detail with reference to certain preferred versions thereof, other versions are possible. Therefore, the spirit and scope of the appended claims should not be limited to the description of the preferred version contained herein.