Abstract:
A method and apparatus for controllable distribution of power from a turbine of a gas turbine engine between two rotatable loads of the gas turbine engine, comprises transferring a shaft power of the turbine to the respective rotatable loads using differential gearing operatively coupled with the turbine and the rotatable loads, respectively; and controlling the power transfer using machines operatively coupled with the respective rotatable loads, operable as a generator or a motor for selectively taking power from one of the rotatable loads to drive the other of the rotatable loads, or the reverse.

Description:
FIELD OF THE INVENTION 
   The present invention relates to gas turbine engines, and more particularly, to gas turbine engines in which the power extraction from the compressor and other rotatable loads can be modulated, without affecting turbine operative condition. 
   BACKGROUND OF THE INVENTION 
   A gas turbine engine generally includes in serial flow communication, one or more compressors followed in turn by a combustor and high and low pressure turbines disposed about a longitudinal axis centerline within an annular outer casing. During operation, the compressors are driven by the respective turbines and compressor air which is mixed with fuel and ignited in the combustor for generating hot combustion gases. The combustion gases flow downstream through the high and low pressure turbines which extract energy therefrom, for driving the compressors, and for producing other output power either as shaft power or thrust for powering an aircraft in flight. For example, in other rotatable loads, such as a fan rotor in a by-pass turbo fan engine, or propellers in a gas turbine propeller engine, power is extracted from the high and low pressure turbines for driving the respective fan rotor and the propellers. 
   It is well understood that individual components in operation require different power parameters. For example, the fan rotational speed is limited to a degree by the tip velocity and, since the fan diameter is very large, rotational speed must be very low. The core compressor, on the other hand, because of its much smaller tip diameter, can be driven at a higher rotational speed. Therefore, separator high and low turbines with independent power transmitting devices are necessary for the fan and core compressor in prior art aircraft gas turbine engines. Furthermore since a turbine is most efficient at higher rotational speeds, the lower speed turbine driving the fan requires additional stages to extract the necessary power. These additional stages and the separate power transmitting devices result in weight penalties which are undesirable in aircraft applications. 
   Efforts have been made to minimize turbine weight of aircraft gas turbine engines by for example, a differential gearing system which distributes power from a single turbine to at least two different components, such as a core compressor and a fan rotor. This is known in the prior art, as described in U.S. Pat. No. 4,251,987, issued to Adamson on Feb. 24, 1981. In a differential geared turbine engine, the compressor, the fan and the turbine are all mechanically linked, and therefore modulating means are necessary to modulate the rotational speed and torque in order to optimize the individual component performances under various engine operation conditions. Various means are available within the existing technology for modulating the torque requirements of various components. Adamson suggests the use of known torque and flow varying techniques, such as variable pitch fans, variable core compressor stators, bleed air extraction, etc., which selectively vary the engine flow passage-defining geometry in order to modulate the torque versus speed characteristics of the individual components. However, a variable engine flow passage-defining geometry increases the complexity and therefore reduces the reliability of an aircraft gas turbine engine. 
   Use of machines operable as either generators or motors for shaft power transfer in gas turbine engines is known in the art. Hield et al. in their U.S. Pat. No. 5,694,765 which issued Dec. 9, 1997, describe a multi-spool gas turbine engine for an aircraft application, which includes a transmission system operated to transfer power between relatively rotatable engine spools. In a number of embodiments, each shaft is associated with a flow displacement machine operable as a pump or a motor, and in other embodiments, permanent magnet or electromagnetic induction type machines operable as motors or generators, are used. However, Hield et al.&#39;s shaft power transfer system does not offer, disclose or teach differential geared gas turbine engines, because they direct themselves to the transfer shaft power between two independently rotatable (i.e. not differentially-geared) engine spools. 
   Therefore, it is desirable to provide an aircraft gas turbine engine configuration in which the turbine weight is minimized without compromising the engine flow passage-defining geometry thereof. 
   SUMMARY OF THE INVENTION 
   One object of the present invention is to provide a gas turbine engine adapted to modulate engine power distribution between different rotatable loads in order to meet with various engine operation requirements, but without affecting an optimum turbine operation condition. 
   In accordance with one aspect of the present invention, there is provided a gas turbine engine including a compressor and a turbine in serial fluid communication, a rotatable load, and a differential gearing system for receiving power from the turbine and transmitting power to the respective compressor and the rotatable load. A first motor/generator mechanism is coupled to the compressor for operating either as a motor to drive the compressor, or as a generator to take power from the compressor. A second motor/generator mechanism is coupled to the rotatable load for operating either as a motor to drive the rotatable load, or as a generator to take power from the rotatable load. The first and second motor/generator mechanisms are controlled for selectively modulating the torque versus speed characteristics of the compressor and the rotatable load, and for modulating the rotational speed relationship between the turbine, the compressor and the rotatable load. 
   In one embodiment of the present invention, the differential gearing system comprises a first sun gear driven by the turbine at turbine rotational speed, and planet gearing engaging the sun gear and operatively connected to the compressor for rotationally driving the compressor at a first output rotational speed with respect to the turbine. A planet carrier is provided for operatively supporting the planet gearing and is rotatable together with the planet gearing. The planet carrier is operatively connected to the rotatable load for driving the rotatable load in a rotational motion at a second output rotational speed with respect to the turbine. The first and second motor/generator mechanisms are preferably permanent magnet motor/generators. 
   In accordance with another aspect of the present invention there is a method provided for controllably distributing power from a turbine of a gas turbine engine between two rotatable loads of the gas turbine engine, which comprises transferring a shaft power of the turbine to the respective rotatable loads, using differential gearing operatively coupled with the turbine and the rotatable loads, respectively; and controlling the power transfer using machines operatively coupled with the respective compressor and rotatable load, operable as a generator or a motor for selectively taking power from one of the rotatable loads to drive the other of the rotatable loads, or the reverse. 
   The differential-geared gas turbine engine with motor/generator regulating mechanisms according to the present invention, advantageously provides a high overall efficiency of performance and requires a minimum number of gears, compressor stages and turbine stages. No bleed valves or variable geometry of the engine fluid path is required. Higher speeds are achievable for turbines, compressors and other power output shafts at off-design points which usually occur under take-off conditions. 
   Other advantages and features of the present invention will be better understood with reference to a preferred embodiment described hereinafter. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
     Having thus generally described the nature of the present invention, reference will now be made to the accompanying drawings, showing by way of illustration the preferred embodiments thereof, in which: 
       FIG. 1  is a partial cross-sectional view of a portion of a gas turbine engine incorporating one embodiment of the present invention, showing the structural configuration of a differential gearing system and the motor/generator regulating mechanisms; and 
       FIG. 2  is a schematic illustration of a gas turbine turbopropeller engine, illustrating an embodiment of the present invention as applied in different engines. 
   

   DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
   Referring to the drawings, particularly  FIG. 1 , an exemplary gas turbine engine  10  includes in serial flow communication about a longitudinal central axis  12 , a fan  13  having a plurality of circumferentially spaced apart fan or rotor blades  14 , a compressor  16  having a plurality of circumferentially spaced apart compressor blades  17 , an annular combustor (not shown), and a turbine (not shown). The turbine includes a rotating shaft  18  extending along the longitudinal central axis  12 , and is operatively coupled with a differential gearing system  20 . The compressor  16  is coupled to the differential gearing system  20  by a rotor shaft  22  which is rotatably supported by bearing  23  on a stationary structure of the engine  10  and which extends co-axially with respect to the turbine rotating shaft  18 . The fan  13  is coupled to the differential gearing system  20  by a rotor shaft  24  extending along the longitudinal central axis  12 . Conventional annular combustor (not shown) and fuel injecting means (not shown) are also provided for selectively injecting fuel into the combustor, to generate combustion gases for powering the engine  10 . 
   A conventional annular casing  26  surrounds the engine  10  from the compressor  16  to the turbine, and defines with the compressor  16 , a compressor inlet  28  for receiving a portion of ambient air  30 . The downstream end of the casing  26  defines an exhaust outlet (not shown). A plurality of stator vanes  32  which are circumferentially spaced apart and are positioned downstream of the fan blades  14  and are provided for supporting the engine  10  within a nacelle  34 . A second group of stator vanes  36  which are circumferentially spaced apart and positioned further downstream of the stator vanes  32  are also provided for supporting the engine  10  within the nacelle  34 . 
   A portion of the air  30  compressed by the fan blades  14  adjacent to fan blade roots  38 , passes through the stator vanes  32  and  36 , and is further compressed by compressor blades  17 . The compressed portion of the air  30  is mixed with fuel to generate combustion gases which rotate the turbine and the turbine rotating shaft  18 , and are then discharged through the outlet of the casing  26 , thereby providing thrust. 
   The turbine rotating shaft  18  inputs a portion of the engine power to the differential gearing system  20  in order to further distribute this portion of engine power which is required during engine operation, and which will be further described hereinafter. 
   The nacelle  34  which surrounds the fan blades  14  and at least the upstream portion of the casing  26 , is spaced radially outwardly from the casing  26 , to define with the casing  26 , an annular duct  40  for permitting the radially outer portion of the air  30  compressed by the fan blades  14  to bypass the engine  10 . The nacelle  34  includes an inlet  42  at its upstream end for receiving the ambient air  30 , and an outlet (not shown) at its downstream end for discharging the portion of air  30  which has been compressed by the fan blades  14  and passes through the annular duct  40 , in order to provide a portion of the thrust. 
   The differential gearing system  20  includes a sun gear  44  affixed to the forward end of the turbine rotating shaft  18  at the turbine speed. A plurality of first planet gears  46  are disposed about and engage the sun gear  44 . The first planet gears  46  are each affixed to an associated shaft  48  which is rotatably secured at one end thereof to a carrier disk  50 . Second planet gears  52  are each affixed to the associated shaft  48  so that the second planet gears  52  are rotatable together with the corresponding first planet gears  46  and their respective, associated shafts  48 . Thus, the first and second gears  46  and  52  surround the longitudinal central axis  12  and the individual pairs of the first and second gears  46 ,  52  are positioned around the longitudinal axes  54  of the respective associated shafts  48 . The second planet gears  52  also surround and engage a gear  56  which is affixed on the forward end of the compressor rotor shaft  22  for transmitting rotary motion and torque to the compressor  16 . 
   Alternatively, each first planet gear  46  can be integrated with the corresponding one of the second planet gears  52  such that each integrated unit of first and second planet gears  46 ,  52  is rotatably mounted on the corresponding associated shaft  48  which is in turn affixed at one end thereof to the carrier disk  50 . This alternative configuration will perform the same function as the planet gear configuration described in the preceding paragraph. 
   The carrier disk  50  is rotatably supported by a bearing  58  positioned around the turbine rotating shaft  18  at the forward end thereof. Therefore, each pair of the first and second gears  46 ,  52  are rotatable together about both the longitudinal central axis  12  and the respective longitudinal axes  54  which are in turn rotatable about the longitudinal central axis  12 . The carrier disk  50  includes cylindrical section  60  extending axially and forwardly therefrom, which is coaxially coupled with the fan rotor shaft  24  by, for example engaging inner and outer gears or keys  62  disposed therebetween, in order to transmit rotary motion and torque to the fan rotor shaft  24 . 
   It should be noted that the diameters of first and second planet gears  46  and  52  differ, and the relative sizes are a function of the desired gear ratio and therefore provide the desired speed and torque relationship between the compressor rotor  16  and the fan assembly  13 . 
   Thus, in operation, turbine power is transferred through the turbine rotating shaft  18  to rotate sun gear  44 , which in turn drives first planet gears  46 . As a result of the coupled relationship between the first planet gears  46  and the second planet gears  52 , and between the first planet gears  46  and the carrier disk  50 , the first planet gears  46  drive the respective second planet gears  52  and the carrier disk  50 , which both in turn drive the respective compressor rotor shaft  22  and fan rotor shaft  24  rotationally, but not necessarily at the same speed about the turbine rotating shaft  18 . 
   The differential gear characteristics are well known, that is,
 
 N   T   =k   1   N   F   +k   2   N   C  
 
where
         N T =turbine speed (RPM)   N F =fan speed (RPM)   N C =compressor speed (RPM)   k 1 , k 2 =constants dependent upon the associated gear ratios.       

   Therefore, fan speed can increase, compromising the compressor speed and vice versa, through equilibrium of the gear elements. The constants k 1  and k 2  set the torque relationship between the fan  13  and compressor  16 . Power is merely the product of the torque and speed, and thus the power ratio between the fan  13  and the compressor  16  can be varied to accommodate any particular operating condition where the total turbine power is known. Therefore, the speed relationship between the fan  13 , compressor  16  and the turbine, can be adjusted by modulating the torque versus speed characteristics of either the fan  13  or compressor  16 . The torque versus speed relationship will naturally set at values determined by the torque characteristics of the fan  13  and compressor  16 . Changes in speed will vary the aerodynamic characteristics of the components. For example, during periods of relatively low aircraft velocity the fan bypass flow rate may be increased at the expense of the core engine by adjusting the torque of the compressor, thereby increasing fan speed and decreasing compressor speed. Conversely, during high speed operation, the torque of the fan may be modulated to decrease the bypass ratio and fan speed, and increase the compressor speed and core engine high velocity exhaust gas flow. 
   Conventionally, the torque modulation of the fan or compressor has been achieved by varying the geometry of a particular region of the relative fluid path, for example by setting the adjustment of variable compressor stator vanes, or adjusting compressor air bleed valves. According to the embodiment of the present invention shown in  FIG. 1 , the torque modulation of the fan and compressor is achieved by use of machines operable as motors or as generators. 
   In  FIG. 1  a first machine which is preferably a permanent magnet motor/generator, indicated by numeral  64 , includes a permanent magnet rotor  66  and a stator  68 . 
   The permanent magnet rotor  66  is mounted on the compressor rotor shaft  22  and is rotatable together with the compressor rotor shaft  22 . The stator  68  has a cylindrical configuration and includes electric windings installed therein. The cylindrical stator  68  is supported within a stationary structure of the engine  10 , and surrounds the permanent magnetic rotor  66  in a radially spaced apart but very close relationship therewith. A second machine  70  which is similar to the permanent magnet motor/generator  64  is also provided. The machine  70  has a permanent magnet rotor  72  mounted on the cylindrical section  60  of the carrier disk  50  and is rotatable together with the carrier disk  50 . Alternatively, the permanent magnet rotor  72  can also be mounted directly on the fan rotor shaft  24  and be rotatable together therewith. A cylindrical stator  74  which includes electric windings is secured to a stationary structure of the engine  10  and closely surrounds the permanent magnet rotor  72 . The electric windings of the respective stators  68  and  74  are electrically connected to a controller which can be a part of the engine or aircraft control system and is therefore not shown in FIG.  1 . 
   In operation, turbine power is input from the turbine rotating shaft  18  into the differential gearing system  20  and is transferred to the respective compressor rotor shaft  22  and fan rotor shaft  24 , to meet the respective torque versus. speed characteristic requirements of the fan and the compressor, depending on the gear ratios k 1  and k 2  which are independent and variable in the design of the engine and are chosen by the designer as a result of due consideration of the projected operating environment and the aerodynamic characteristics of the individual components. When the engine  10  is operated under an off-design points condition, one of the machines  64  and  70  can be controlled to operate as a generator for taking power to produce electric current which is then delivered to the other of the machines  64  and  70 , causing it to operate as a motor for driving the shaft coupled therewith. Therefore, the machines  64  and  70  can be controlled for selectively modulating the torque versus speed characteristic of the compressor  16  and the fan  13 , and for modulating the rotational speed relationship between the turbine, compressor  16  and the fan  13 . The machines  64  and  70  can either or both be advantageously used as electric starters. In such an operation, either or both of the compressor rotor shaft  22  and the fan rotor shaft  24  can be rotated by machines  64  and  70  which in this case receive electrical power and operate as motors to electrically start the engine. 
   The differential gearing system  20  can be used to distribute turbine power, not only between the compressor and the fan, but also among other rotatable loads of the engine. As an example, the differential gearing system  20  can be used with a gas turbine propeller engine as is schematically illustrated in FIG.  2 . 
   In  FIG. 2  the differential gearing system and the machines operable as motor, generator and starter are similar to their equivalent parts illustrated in  FIG. 1 , and are indicated by the respective identical numerals. The gas turbine propeller engine which is generally indicated as  110 , has a longitudinal central axis  112 . The gas turbine propeller engine  110 , similarly to gas turbine engine  10  of  FIG. 1 , includes compressor  116 , and turbine  119  which includes two stages as shown in FIG.  2 . The turbine  119  is connected to the differential gearing system  20  by turbine rotating shaft  118 , and the compressor  116  is connected to the differential gearing system  20  by compressor rotor shaft  122 . Both shafts  118  and  122  are operatively supported within the engine  110  and are rotatable about the longitudinal central axis  112 . A propeller  113  is provided at the front end of the engine  110  and is coupled to the differential gearing system  20  by a propeller shaft  123 , which is operatively mounted within the engine  110  and which is rotatable about the longitudinal central axis  112 . The gas turbine propeller engine  110  is also provided with a combustor  125  and fuel injection means (not shown). According to the embodiment shown in  FIG. 2 , the gas turbine propeller engine  110  is further equipped with a heat recuperator  127 . 
   The permanent magnetic rotor  66  of the machine  64  is mounted on the compressor rotor shaft  122  and is rotatable together with same. The cylindrical stator  68  is secured to a stationary structure of the engine  110 , and the electrical windings thereof are electrically connected to a controller  129 . The permanent magnet rotor  72  of machine  70  is mounted on the propeller shaft  123  and is rotatable together with same. The cylindrical rotor  74  of machine  70  is secured to a stationary structure of the engine  110 , and the electrical windings thereof are electrically connected to the controller  129 . 
   In operation the machines  64  and  70 , can either or both be used as electrical starters to electrically start the engine  110 . When the engine  110 , has started, ambient air  130  entering the engine  110  from the front end thereof is compressed by the compressor  116 . The compressed air  130  then passes through the heat recuperator  127  where the compressed air  130  is heated. The heated compressed air  130  then exits the heat recuperator  127  and is mixed with fuel for combustion in a combustor  125 , thereby producing combustion gases  133 . The combustion gases  133  rotate the turbine  119  to power the engine  110  and then enter the heat recuperator  127  for heat exchange. In the heat recuperator  127  the remaining heat energy carried by combustion gases  133  exhausted from the turbine  119 , is transferred to the compressed air  130  therein to increase the compressor air temperature, thereby improves combustion efficiency. Combustion gases  133  are then discharged from the heat recuperator  127  into the surrounding air. 
   In contrast to the gas turbine engine  10  in  FIG. 1 , combustion gases  133  generated in gas turbine propeller engine  110  do not directly provide thrust to the aircraft which carries the engine  110 . Therefore, combustion gases  133  deliver a substantial amount of power and energy carried thereby, to turbine  119 . The turbine power is then distributed by the differential gearing system  20  to the compressor  116 , and to the propeller  113  which produces the entire amount of thrust required to fly the aircraft. 
   The working status of the machines  64  and  70 , as a motor or a generator are controlled by the controller  127  according to the different requirements for torque versus speed characteristics of the compressor  116  and the propeller  113 , and in order to adjust the speed relationship between the propeller  113 , compressor  166  and the turbine  119 . 
   In essence, the advantage of this arrangement according to the present invention, is that the differential gearing ratios can be chosen so that the three shafts rotate, each at the optimum speed for their individual components, and the three components can each be adapted to maintain an optimum speed matching at different power settings. Thus, the turbine can be designed for the most efficient performance without compromising in consideration of compressor or other rotatable load speeds. 
   The machines operable as a generator or a motor in the embodiments described above are permanent magnet motor/generators. However, other types of machines such as electromagnetic induction motor/generator or hydraulic motor/generator (pump) machines can be alternatively used for this purpose. The gas turbine engine and the gas turbine propeller engine illustrated in  FIGS. 1 and 2  are exemplary only, and therefore the present invention can be used with other types of gas turbine engines, for distributing turbine power between various rotatable loads of the engines. Also, though a planetary epicyclic gear system is disclosed, any suitable epicyclic or other type of gear system may be used. 
   Modifications and improvements to the above-described embodiments of the present invention may become apparent to those skilled in the art. The foregoing description is intended to be exemplary rather than limiting. The scope of the invention is therefore intended to be limited solely by the scope of the appended claims.