Abstract:
The invention provides a turbomachine including a device for supplying pressurized gas, the gas being bled from the compressor of the turbomachine via a bleed orifice and cooled in a cooling cavity disposed within the compressor rotor upstream of the bleed orifice. This arrangement has the effect of cooling the bled gas by heat transfer through the outer wall of the compressor rotor to the gas stream which is being compressed by the compressor upstream of the bleed orifice, and has the result of reducing the drop in the performance of the compressor as a result of the gas being bled off.

Description:
BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     The invention relates to turbomachines, particularly but not exclusively to turbomachines for aircraft, and more specifically to turbomachines including at least one device which bleeds gas from the compressor of the turbomachine to provide a supply of pressurized gas. The invention also relates to the application of this device to the cooling of the rotor of the turbine of the turbomachine. 
     2. Summary of the Prior Art 
     Turbomachines are well known machines which essentially comprise a rotary compressor with vanes, a combustion chamber and a turbine, the stream of gas which passes through them from the upstream end to the downstream end being subjected to an appropriate thermodynamic cycle. Turbomachines are used in particular in aeronautics for aircraft propulsion, and are also used in industry. They often include at least one device for supplying pressurized and somewhat cold gas intended for various uses, such as cooling parts such as the turbine, controlling operating clearances, heating the fuel or the lubricant, and aircraft auxiliaries. 
     Such a device for supplying pressurized gas customarily includes means for bleeding gas from the compressor at a desired pressure, means for cooling the gas thus bled to a desired temperature, and means for conveying this gas to the items which will use it. 
     The means for bleeding the gas from the compressor comprises at least one bleed orifice opening into the gas stream and arranged at a point in the compressor where the pressure has reached a sufficient level. 
     The means for cooling the bled gas essentially comprises a heat exchanger connected to the bleed orifice by a pipe. The heat exchanger consists of a cooling cavity which has an inlet and an outlet and which is bordered over an adequate area by a cooling surface in contact with a coolant fluid, the bled gas brushing against the cooling surface while passing through the cooling cavity from the inlet to the outlet. The coolant fluid may be ambient air, blown if necessary, but may equally be the lubricant or the fuel, which then acts as a heat-transfer fluid which has to be cooled itself, this most often being done using other coolers in contact with the ambient air. 
     The gas is then led to the items which are to use it by pipes. Cooling the turbine rotor is more tricky. The rotor disks usually have a radial cavity extending into each blade. The duct conveying the gas to the radial cavity passes through the center of the turbomachine near its axis of rotation and may consist of the turbine shaft itself. 
     Such devices for supplying pressurized gas exhibit numerous drawbacks. Firstly, bleeding gas from the compressor reduces the performance of the compressor, which has a direct impact on the thermodynamic performance of the turbomachine. Secondly, cooling the bled gas requires coolers which are placed outside the machine, these coolers generating drag in the case of turbine engines for aircraft. When use is made of an intermediate heat-transfer fluid; such as the lubricant or the fuel, effective regulation needs to be in place in order not to run the risk of carbonizing this heat-transfer fluid as a result of excessive heating. Finally, cooling the turbine rotor requires a complex and cumbersome circuit, part of which has to pass through the center of the turbomachine and interfere with the turbine shaft. 
     SUMMARY OF THE INVENTION 
     The invention addresses two problems, the first being that of reducing the fall in compressor performance, and the second being that of simplifying and lightening the cooling of the turbine rotor. 
     As a solution to the first problem the invention provides a turbomachine including a rotary compressor having an upstream end, a downstream end and means defining a flow path therebetween through which a stream of gas is compressed as it passes from said upstream end to said downstream end, said compressor including vanes disposed in said flow path and a compressor rotor having an outer wall which is contacted by said stream of gas as it flows along said flow path, and a pressurized gas supply device comprising a gas bleed orifice disposed in said rotary compressor and opening into said flow path thereof, and a cooling cavity disposed within said compressor rotor for cooling the pressurized gas bled from said gas stream via said gas bleed orifice, said cooling cavity comprising an inlet connected to said gas bleed orifice, an outlet, and a cooling surface for contact by said pressurized gas bled from said gas stream as said pressurized gas passes from said inlet to said outlet, said cooling surface being formed at least partly by at least a portion of the outer wall of said compressor rotor upstream of said gas bleed orifice whereby said bled pressurized gas is cooled by heat transfer through said outer wall of said compressor rotor to said gas stream upstream of said gas bleed orifice. 
     This arrangement has the result of returning to the gas stream, in the compressor, some of the energy which is bled off, thus reducing the drop in compressor performance resulting from the gas being bled off. 
     It will be understood that the transfer of heat from the bled gas back into the gas stream is made possible by the fact that this heat exchange takes place upstream of the gas bleed orifice, that is to say in a zone of the compressor where the gas stream is at a lower temperature than the temperature it has in the region of the bleed orifice. This exchange is made efficient by the presence of the rotor blades, which receive, by thermal conduction, some of the heat which is transferred to the outer wall of the compressor rotor, these blades returning the heat to the stream of gas by virtue of their large surface area which is swept at high speed by the gas flowing in the stream. 
     It will be understood that the cooling surface has to be large enough to allow the bled gas to be cooled, the person skilled in the art determining said heat-exchanger surface area according to the characteristics of the turbomachine, the point at which the gas is bled off, and the flow rate and temperature of the pressurized gas to be supplied. 
     The invention has the advantage of being simple to implement and of mainly using existing means, thus making it possible to reduce the mass and cost of the turbomachine. 
     Preferably, the cooling surface extends overall from downstream to upstream, that is to say that the bled gas sweeps the cooling surface from downstream to upstream. This arrangement has the effect of bringing the bled gas against zones of the cooling surface which, on the whole, are increasingly colder, and has the result of achieving a greater drop in the temperature of the bled gas. This arrangement also has the advantage of returning the heat to the gas stream uniformly along the compressor, which allows the operation of the compressor not to be disturbed. 
     It will be understood that all that is required in order to obtain the effect sought by the invention is for the flow of the bled gas in contact with the cooling surface to be generally in the direction from downstream to upstream. This effect will be maintained in spite of limited returns of bled gas in the downstream direction, it being possible for such returns to result from mechanical or aerodynamic constraints. 
     Preferably, the bleeding of gas is centripetal, the bleed orifice being located in the outer wall of the compressor rotor. This arrangement has the effect of shortening the path that the bled gas has to follow from the bleed orifice to the inlet of the cooling cavity, and has the result of allowing a short connection. In a preferred embodiment, the bleed orifice also constitutes the inlet to the cooling cavity, that is to say the bleed orifice opens directly into said cavity. 
     The turbomachine may have a plurality of cooling cavities in the compressor rotor, so as to be able to supply bled gas to different receivers and more or less independently of each other. In other words, the flow rate of pressurized gas supplied by one cooling cavity will have only a small influence on the temperature of the pressurized gas supplied by another cooling cavity. 
     The cooling cavities will preferably be arranged in the upstream-downstream direction, so as simultaneously to supply bled gas under different pressure and temperature conditions. Thus, a cooling cavity located toward the upstream end of the compressor will supply pressurized gas at a temperature and at a pressure which are lower than those for gas supplied by a cooling cavity nearer the downstream end of the compressor. 
     In the case of a turbomachine including a turbine having a rotor which is coaxial with the compressor and located downstream of said compressor and which is arranged to be cooled by the centrifugal passage of a flow of gas, the turbomachine is preferably provided with a tube disposed coaxially with the compressor and with the turbine, the upstream end of this tube being connected to the outlet of the cooling cavity and the downstream end of the tube being connected to the turbine rotor that is to be cooled. The function of this tube is to bring the bled pressurized gas directly to the turbine rotor from the cooling cavity, passing from upstream to downstream through the center of the turbomachine, that is to say near its geometric axis. This very simple arrangement makes it possible to make use of the central region of the turbine engine which is usually underused. It also allows the cooling gas to be conveyed through a straight short pipe of large cross section, which will minimize pressure drops. 
     Preferably, in the case of a turbine rotor with at least two stages, first and second stages A and B will be connected each to a respective one of first and second cooling cavities A and B by a respective one of first and second tubes A and B, the first turbine stage A being upstream of the second stage B, the first cooling cavity A being downstream of the second cooling cavity B, and the second tube B passing through the inside of the first tube A. An arrangement of this kind has the effect of conveying of the bled gases from the cooling cavities to the respective turbine stages that are to be cooled in the manner of concentric flows which do not cross, the flow of bled gas from the second cooling cavity B being conducted along the inside of the flow of bled gas from the first cooling cavity A. This arrangement thus allows the various flows of bled gas to be conveyed simply, and without crossing, from the cooling cavities which produced them to the turbine stages that are to be cooled. The arrangement also makes it possible to supply the various turbine stages with bled gas under temperature conditions suited to each stage: a stage located nearer the downstream end of the turbine, which is therefore not as hot, receiving bled gas from a cooling cavity located nearer the upstream end of the compressor, which is therefore also less hot. Finally, the arrangement makes it possible to maintain large passage cross sections which do not introduce pressure drops in the flow of bled gas. 
     The present invention is particularly effective when the compressor rotor is of the disk type, because the disks penetrate into the cooling cavity and increase the cooling surface area, which has the effect of increasing the ability of the cooling cavity to cool the bled gas. 
    
    
     BRIEF DESCRIPTION OF THE DRAWING 
     FIG. 1 shows a partial view, in cross section, of a twin-spool turbine engine for an aircraft, said engine having two devices A and B for supplying pressurized gas in accordance with the invention. 
    
    
     DESCRIPTION OF THE PREFERRED EMBODIMENT 
     The drawing shows part of a turbomachine  1  which, on the whole, is generally symmetrical about a longitudinal axis  2  and through which a gas stream  5  passes from the upstream end  3  to the downstream end  4 . The gas stream  5  passes first of all through a rotary compressor  10  including a stator  11  which externally bounds the flow path for the gas stream  5  though the compressor and which supports a plurality of stationary vanes  12 . The compressor  10  also includes a compressor rotor  15  which rotates about the longitudinal axis  2  and has an outer wall  16  which internally bounds the gas stream flow path, the outer wall  16  supporting a plurality of turning vanes  17 , often referred to as blades, which extend radially across the gas stream  5  and compress it in collaboration with the stationary vanes and the particular shape of the gas stream flow path. In order to withstand the centrifugal force produced by the rotation of the compressor rotor  15 , a disk  18  secured at its periphery to the outer wall  16  of the compressor rotor  15  is arranged under each stage of turning vanes  17 . Each disk  18  is disposed roughly in a plane which is radial with respect to the longitudinal axis  2 , and defines two oppositely facing lateral flanks  19  and  20  and a central aperture  21 . 
     The gas stream  5  then passes through a combustion chamber  25  where it experiences a significant rise in temperature, before entering and passing through a turbine  30 . 
     The turbine  30  comprises a stator  31  externally bounding the flow path for the gas stream  5  through the turbine, the stator  31  supporting stationary vanes  32 . The turbine  30  also comprises a rotor  35  which rotates about the axis  2  and, at its periphery, carries a plurality of turning vanes or blades  36  which extend radially across the gas stream flow path and receive some of the energy built up in said gas stream  5 . The “mobile” turning vanes or blades  36  are distributed in two stages  37 , namely an upstream stage  37   a  and a downstream stage  37   b.  Each stage  37  has a turbine disk  38  for withstanding the centrifugal force of rotation, and each turbine disk  38  has a radial cavity  39  allowing the centrifugal flow of a cooling gas  40  which is conducted through the center  41  of the disk  38  and passes radially through the turbine disk  38  and the turning vanes or blades  36 . 
     The turbine  30  is coaxial with the compressor  10 , and the turbine rotor  35  is connected to the compressor rotor  15  by a drive shaft  45 , usually of large diameter. 
     In this example, the turbomachine  1  has two devices A and B for supplying pressurized gas, each being arranged inside the compressor rotor  15 . Each pressurized gas supply device comprises a centripetal gas bleed orifice  51  arranged in the outer wall  16  of the rotor  15 , and the gas bled off is referenced  52 . The supply device also comprises a cooling cavity  53  inside the compressor rotor  15 . The cooling cavity  53  uses the internal volume of the compressor rotor  15  and has an inlet  54  which, in this example, coincides with the gas bleed orifice  51 . The cooling cavity  53  also has a cooling surface  55  formed at least partly by at least part of the outer wall  16  of the compressor rotor  15 , this cooling surface  55  being located upstream  3  of the gas bleed orifice  51 . The cooling cavity  53  is, in this example, bounded on the upstream  3  side by a disk  61  and internally by a tube which extends almost to the disk  61 , the gap  58  left between the disk  61  and the tube  57  constituting the outlet from the cooling cavity  53 . The cooling cavity  53  may have a varying number of disks  18  extending radially across it, the tube  57  passing through the central aperture  21  of each of these disks  18  with sufficient clearance  59  to allow the bled gas  52  to pass from the disk from downstream to upstream. The tube  57  extends in the downstream direction as far as the radial cavity  39  of a turbine disk  38  to which it is connected. In this example, the cooling cavity  53  is bounded at the downstream end by a downstream partition  60  of the compressor rotor  15 , the downstream partition  60  being attached at its periphery to the outer wall  16  of the rotor downstream of the gas bleed orifice  51 . 
     The way in which the assembly works is as follows. The stream of gas  5  passes through the compressor  10  from upstream to downstream, undergoing a compression which increases its temperature and pressure. Compressed, and therefore hot, gas  52  is bled from the gas stream  5  through the bleed orifice  51 , passes through the cooling cavity  53  from its inlet  54  to its outlet  58  where it enters the tube  57 , passes along the tube  57  to the radial cavity  39  of the turbine disk  38 , moves out as far as the turning vanes or blades associated with the turbine disk  38 , and finally returns to the gas stream  5 , usually through cooling orifices, not shown, formed in the turning vanes or blades  36 . As it passes through the cooling cavity  53 , the bled gas  52  passes through the disks  18  via the clearances  59  between the central apertures  21  and the tube  57 . During this passage, the bled gas  52  is swept in a complex swirling movement across the cooling surface  55  and the flanks  19  and  20  of the disks  18 . Because the cooling surface  55  is upstream  3  of the bleed orifice  51 , this cooling surface  55  and the turning vanes  17  attached to it are in contact with the gas stream  5  which is colder at this point than it is toward the bleed orifice  51 . This allows heat to be transferred from the bled gas  52  to the gas stream  5 , this heat passing through the disks  18  which extend across the cooling cavity  53 , through the cooling surface  55 , and at least partially through the turning vanes  17 . 
     The relative arrangements of the two gas supply devices A and B will now be described. To do this, the letters a, b will be used in conjunction with the reference numerals used to indicate components of the devices as described above. The first gas supply device A comprises a cooling cavity  53   a  toward the downstream end of the compressor rotor  15 , this cooling cavity  53   a  being connected by a tube  57   a  to the turbine disk  38   a  located at the upstream end of the turbine rotor  35 . The second gas supply device B comprises a cooling cavity  53   b  connected by a tube  57   b  to the turbine disk  38   b.  The cooling cavity  53   b  is upstream of the cooling cavity  53   a,  the tube  57   b  passes through the inside of the tube  57   a,  and the turbine disk  38   b  is downstream of the turbine disk  38   a.    
     The downstream cooling cavity  53   a  is bounded at the downstream end by a downstream partition  60  of the compressor rotor  15 , this downstream partition  60  being connected at its outer periphery to the downstream end of the outer wall  16  of the compressor rotor  15 , and being connected at its inner periphery to the tube  57   a.  The downstream cooling cavity  53   a  is bounded at its upstream end by a common disk  61  which has the edge of its central aperture  21  attached to the tube  57   b.  The attachments described may be achieved by any means which offers sufficient gastightness to the bled gas, having regard to the pressures reached inside the turbomachine. 
     The upstream cooling cavity  53   b  is bounded at its downstream end by the aforementioned common disk  61 , and at its upstream end by the first disk  62  of the compressor  10 . 
     The tubes  57   a  and  57   b  may be thin in order to reduce their mass, and preferably they will be connected together by spacer pieces  63  to make the assembly more rigid, said spacer pieces  63  obviously allowing the bled gas  52  to pass. 
     In this example, the invention is applied to the so-called “high-pressure” spool of a “twin-spool” turbomachine. The other, so-called “low-pressure”, spool is not shown, but has a shaft  65  passing coaxially through the tube  57   b  to connect the “low-pressure” turbine to the “low-pressure” compressor. 
     The advantages of the present invention applied to the cooling of the turbine rotor should now be clearly apparent. 
     1) In comparison with turbine rotor cooling performed directly with uncooled bled gas, the invention allows the same cooling of the turbine rotor to be performed with a lower flow rate of gas, thus increasing the efficiency of the overall cycle of the turbomachine and thereby reducing its fuel consumption. 
     2) The method employed for cooling the bled gas leads to only a very small increase in the mass of the turbomachine, represented, in practice, by the mass of the tubes  57 , in contrast to prior systems with separate and therefore heavier coolers. This advantage is significant in the field of aeronautics. 
     3) The cooler is intrinsic to the compressor rotor and therefore causes no aerodynamic braking of a turbomachine used for the propulsion of aircraft, unlike conventional coolers which are usually positioned in the turbomachine bypass flow and therefore degrade the total pressure of the flow and the thermodynamic efficiency of the turbomachine. 
     The present invention is not restricted to the particular example which has just been described, and is intended to cover all variations which may be conceived without departing either from its scope or spirit. 
     The turbomachine may be a single-spool or a multi-spool turbomachine, and the invention may be applied equally to any of the spools. 
     In this example, the turbomachine comprises a so-called “in-line” compressor. The present invention is also applicable to a turbomachine which has a centrifugal compressor or a hybrid compressor. 
     In this example, the bleed orifices  51  open directly into the cooling cavities  53  and are thus coincident with the inlets  54  of said cavities. However, the bleed orifices  51  could just as well be distinct from said inlets  54 . 
     In this example, the turbomachine  1  comprises a compressor rotor  15  of the type having disks  18 . The invention may also be applied to a turbomachine including a compressor rotor  15  without disks, and in this case partitions will be provided to axially bound the cooling cavities  53 . 
     Finally, in this example, the term “stage”  37  is used in the strict sense and corresponds to a turbine disk  38  with its turning vanes or blades  36 . In application of the invention, however, a stage  37  may encompass a number of turbine disks  38  and their associated vanes or blades  36 , which will all be cooled together by one and the same flow of bled gas  52 . A solution of this type is preferable because of its simplicity towards the downstream end of the turbine, where the temperature is lower.