Abstract:
The invention relates to a field of feeding reaction engines, and in particular a turbopump  9  for feeding at least one combustion chamber  2  with a first propellant, the turbopump comprising a pump  9   a , a turbine  9   b  to the pump  9   a  in order to drive it, and a variable and/or disengageable braking device  11.

Description:
BACKGROUND OF THE INVENTION 
     The present invention relates to the field of feeding at least one combustion chamber with at least one propellant. 
     In the description below, the terms “upstream” and “downstream” are defined relative to the normal flow direction of a propellant in a feed circuit. 
     In reaction engines, and in particular in rocket engines, thrust is typically generated by hot combustion gas that is produced by an exothermal chemical reaction that has taken place within a combustion chamber and that expands in a propulsive nozzle. Consequently, high pressures normally exist in the combustion chamber while it is in operation. In order to be able to continue to feed the combustion chamber in spite of those high pressures, propellants need to be introduced at pressures that are even higher. Various means are known in the prior art for achieving this. 
     First means that have been proposed comprise pressurizing the tank containing the propellants. Nevertheless, that approach greatly restricts the maximum pressure that can be reached in the combustion chamber and thus restricts the specific impulse of the reaction engine. Consequently, in order to reach higher specific impulses, the use of feed pumps has become common practice. Various means have been proposed for actuating such pumps, and most frequently they are driven by at least one turbine. In such a turbopump, the turbine itself may be actuated in various different ways. For example, the turbine may be actuated by combustion gas produced by a gas generator. Nevertheless, in so-called “expander cycle” rocket engines, the turbine is actuated by one of the propellants after it has passed through a heat exchanger in which it is heated by the heat produced in the combustion chamber. Thus, this transfer of heat can contribute simultaneously to cooling the walls of the combustion chamber and/or of the propulsive nozzle, while also actuating at least one feed pump. 
     Typically, propellant feed circuits are arranged to reach an operating equilibrium in which a specific flow rate of each propellant is delivered to the combustion chamber. Consequently, a rocket engine fed by such feed circuits reaches a stable level of thrust. Nevertheless, under certain circumstances, it may be desirable to be able to select between a plurality of stable levels of thrust. In particular, it is now desired for the rocket engines of the final stages of satellite launchers to have not only a function of putting the payload into orbit, but also a function of de-orbiting the final stage. In order to perform such de-orbiting, and in particular in order to ensure that the final stage falls at an accurate point, it is preferable to make use of a level of thrust that is substantially smaller than the level of thrust used while putting the payload into orbit. 
     In the survey “Design and analysis report of the RL 10-IIB breadboard low thrust engine”, FR-18046-3, written for NASA on Dec. 12, 1984, a system for feeding propellant to a combustion chamber is proposed that is capable of obtaining a low-thrust mode by opening a passage for bypassing the turbine that drives the pumps for the two propellants. Nevertheless, that solution requires additional complication in the propellant feed circuit, in particular to the detriment of their reliability. 
     OBJECT AND SUMMARY OF THE INVENTION 
     The present invention seeks to remedy those drawbacks. In particular, the invention seeks to propose a feed device that enables the rate at which propellant is delivered to a combustion chamber to be controlled without leading to additional complication of the propellant feed circuits. 
     In at least one embodiment, this object is achieved by the fact that, in a turbopump for feeding at least one combustion chamber with a first propellant, and including at least a first pump and a first turbine coupled to be driven by the first pump, there is also a variable and/or disengageable braking device. 
     By means of the braking device, it is possible to apply a braking torque to the turbopump while its speed is rising in order to restrict the flow rate at which propellant is delivered to the combustion chamber. The speed of the engine can thus be stabilized at a low level of thrust. 
     In particular, the braking device may be an electromagnetic braking device, a friction braking device, or a hydrodynamic braking device. If the braking device is electromagnetic, it may for example be a device of the type that makes use of eddy currents induced in a rotor in order to generate a braking torque. If the braking device is a friction braking device, it may for example be of the type having a friction pad actuated by an actuator or a device having mechanical jaws. A hydrodynamic braking device may for example comprise an inverse Pelton turbine. 
     The present invention also provides a feed circuit for feeding at least one combustion chamber with a first propellant, and in particular a circuit having at least a first turbopump comprising at least a first pump and a first turbine coupled to the first pump in order to drive it, a variable and/or disengageable braking device, and a heat exchanger situated downstream from the first pump and upstream from the first turbine and suitable for heating a flow of the first propellant between the first pump and the first turbine by using heat generated in said combustion chamber. In particular, the second turbine may be situated downstream from the first turbine. Thus, the feed circuit is a circuit of the so-called “expander” type making use of this transfer of heat to the first propellant simultaneously for cooling the walls of the combustion chamber and/or of the propulsive nozzle, and also for actuating at least the first turbopump. 
     In order to contribute to controlling the first turbopump, the circuit may also include a passage for bypassing at least the first turbine, which passage is fitted with a first bypass valve, thus enabling the first turbopump to be bypassed in selective manner. 
     The present invention also provides a feed system for feeding at least one combustion chamber with propellants, the system comprising at least a first circuit for feeding the at least one combustion chamber with a first propellant, a second circuit for feeding the at least one combustion chamber with a second propellant, and a second turbopump comprising at least a second pump situated in the second circuit and a second turbine situated in the first circuit and coupled to the second pump in order to drive it. In this way, the feed system can deliver two propellants under pressure to the combustion chamber, the flow rates of the two propellants being variable simultaneously as a function of the braking torque applied to the first turbopump by its braking device. 
     Nevertheless, in order to contribute to controlling the second turbopump, the first circuit may also include a passage bypassing the second turbine and fitted with a second bypass valve, thus enabling the second turbopump to be bypassed selectively. 
     The invention also provides a method of controlling a combustion chamber. In particular, in a method in at least one implementation, the combustion chamber is fed with a first propellant by a first turbopump comprising at least a first pump, a first turbine coupled to the first pump in order to drive it, and a variable and/or disengageable braking device, and a braking torque is applied to the first turbopump by its braking device in order to restrict the speed of the turbopump and thus restrict the flow rate delivered by the turbopump to the combustion chamber. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWING 
       The invention can be well understood and its advantages appear better on reading the following detailed description of an embodiment given by way of non-limiting example. The description refers to the accompanying FIGURE, which is a diagram showing a rocket engine with a propellant feed system in this embodiment of the present invention. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The rocket engine  1  shown in the accompanying FIGURE has a combustion chamber  2  with a diverging nozzle  3 , tanks  4 ,  5 , and a system  6  for feeding the combustion chamber  2  with propellants coming from the tanks  4 ,  5 . The tank  4  contains a first propellant and the tank  5  contains a second propellant. In particular, in the embodiment shown, the tanks  4 ,  5  may be cryogenic tanks containing respectively liquid hydrogen and liquid oxygen. 
     The feed system  6  has a first circuit  7  for the first propellant and a second circuit  8  for the second propellant. The first circuit  7  is connected to the tank  4  via a valve  27  and includes a first turbopump  9  and a regenerative heat exchanger  10  incorporated in the walls of the combustion chamber  2 . The first turbopump  9  comprises a first pump  9   a  and a first turbine  9   b  coupled to the first pump  9   a  in order to actuate it, together with an electromagnetic, friction, and/or hydrodynamic braking device  11  for applying a braking torque T brake  to the turbopump  9 . The first circuit  7  is configured in such a manner that the heat exchanger  10  is situated downstream from the first pump  9   a  and upstream from the first turbine  9   b . A second turbine  12   b  is also situated downstream from the first turbine  9   b  in this first circuit  7 . This second turbine  12   b  is coupled to a second pump  12   a  in order to actuate it, said second pump  12   a  being situated in the second circuit  8  for pumping the second propellant. Together, the second pump  12   a  and the second turbine  12   b  form a second turbopump  12 . The first circuit  7  also includes a passage  13  bypassing both turbines  9   b  and  12   b , which passage is fitted with a first bypass valve  14 , and the first circuit  7  also has a passage  15  bypassing the second turbine  12   b  and fitted with a second bypass valve  16 . Directly downstream from the first pump  9   a , the first circuit  7  also has a branch connection leading to a blow-off line  17  for the first propellant and including a first propellant blow-off valve  18 . Directly upstream from the injectors  19  for injecting the first propellant into the combustion chamber  2 , the first circuit  7  also has a valve  20  for admitting the first propellant into the combustion chamber  2 . 
     The second circuit  8 , which is connected to the tank  5  via a valve  28 , also includes a branch connection downstream from the second pump  12   a  to a line  21  for blowing off the second propellant, with a second propellant blow-off valve  22 . The second circuit  8  leads to the injectors  23  for injecting the second propellant into the combustion chamber  2  via a dome  24  covering the combustion chamber  2 . Directly upstream from the dome  24 , the second circuit  8  also has a valve  25  for admitting the second propellant into the combustion chamber  2 . The combustion chamber  2  also has an ignitor  26 . The braking device  11 , the valves  14 ,  16 ,  18 ,  20 ,  22 ,  25 ,  27 , and  28 , and the ignitor  26  are all connected to a control unit (not shown) in order to control the operation of the rocket engine  1 . 
     Prior to igniting the rocket engine  1 , the valves  27  and  28  are initially opened to admit the propellants into the circuits  7  and  8  and to cool the circuits down. During this cooling, the blow-off valves  18  and  20  remain open, as do the bypass valves  14  and  16 . Once the circuits  7  and  8  have been cooled, the valves  20  and  25  are opened to admit the two propellants into the combustion chamber  2 . The ignitor  26  is then activated in order to ignite the mixture of propellants in the combustion chamber  2 . Once ignition has occurred, the heat exchanger  10  begins to heat the flow of the first propellant that passes therethrough. The blow-off valves  18  and  22  and the bypass valves  14  and  16  can then be closed progressively in order to enable the speed of the turbopumps  9  and  12  to increase. During this increase in speed, an increasing flow of the first propellant, as heated in the heat exchanger  10 , actuates the turbines  9   b  and  12   b  before being injected into the combustion chamber  2  via the injectors  19 . In turn, the turbines  9   b  and  12   b  activate the pumps  9   a  and  12   a  respectively, thereby increasing the flow rates of both propellants during this period of increasing speed. 
     The rise in speed of the first turbopump  9  is governed by the following equation: 
     
       
         
           
             
               I 
               ⁢ 
               
                 
                   ⅆ 
                   ω 
                 
                 
                   ⅆ 
                   t 
                 
               
             
             = 
             
               
                 T 
                 turbine 
               
               - 
               
                 T 
                 pump 
               
               - 
               
                 T 
                 brake 
               
             
           
         
       
     
     where I represents the inertia of the turbopump  9 , ω represents its speed of rotation, T turbine  represents the torque generated by the expansion of the first propellant in the first turbine  9   b , and T pump  represents the torque consumed by the first pump  9   a in order to pump the first propellant. At the end of this rise in speed, the first turbopump  9  reaches equilibrium in which the torque T turbine  generated by the first turbine  9   b  is equal to the sum of the torque T pump  consumed by the first pump  9 a plus the braking torque T brake  from the braking device  11 . Since the braking device  11  is variable and/or disengageable, it is thus possible to control the speed at which the first turbopump  9  reaches its operating equilibrium. Indirectly, this also affects the operating equilibrium of the second turbopump  12 , with the torque generated by the second turbine  12   b  depending on the flow rate of the first propellant pumped by the first pump  9   a . By means of the braking device  11 , it is thus possible to control the rocket engine  1  so as to obtain different levels of thrust. 
     Although the present invention is described above with reference to a specific embodiment, it is clear that various modifications and changes may be applied thereto without going beyond the general scope of the invention as defined by the claims. In addition, the individual characteristics of the various embodiments mentioned may be combined in additional embodiments. Consequently, the description and the drawings should be considered as being illustrative rather than restrictive.