Abstract:
A method of powering a rotary-wing aircraft includes selectively coupling and uncoupling a first power turbine to change a power distribution between the rotor system and the secondary propulsion system.

Description:
REFERENCE TO RELATED APPLICATIONS 
       [0001]    The present application is a Continuation Application of patent application Ser. No. 11/998,248, filed Nov. 29, 2007. 
     
    
     BACKGROUND 
       [0002]    This invention relates generally to gas turbine propulsion systems and more particularly to gas turbine propulsion systems that are convertible between two modes of operation. Rotorcrafts such as helicopters rely primarily on large rotating blades to produce both the lift necessary to stay aloft and the thrust necessary for propulsion. In order to produce both thrust and lift, the blades rotate in a plane generally parallel to the velocity vector of the craft, with blades advancing and retreating in the direction of flight of the craft. Due to low relative air speed or excessive angle of attack of the retreating blades, various flight instabilities and aerodynamic inefficiencies arise which limit the maximum, safe airspeed of a typical rotorcraft. In order to overcome the maximum velocity ceiling of rotorcrafts, various designs have incorporated secondary propulsion system to provide additional thrust. For example, rotorcrafts have incorporated rear propellers that rotate in a plane generally perpendicular to the velocity vector of the aircraft that produce only forward (or backward) thrust. Such secondary propulsion systems require input power that is typically siphoned off of the power supplied to the primary rotor blades. 
         [0003]    Gas turbine propulsion systems produce a large amount of rotating shaft power available to both primary and secondary propulsion systems, and thus are popular choices for dual propulsion rotorcrafts such as helicopters. Typically, a turboshaft design is used wherein a gas generator is used to drive a gas generating turbine to compress inlet ambient air and sustain combustion, and a power turbine that drives a free shaft, which is then coupled to the rotor blades through a gearbox. A supplemental output from the free shaft can also be coupled to a secondary propulsion system, such as a propeller, to produce additional thrust. Typically, the supplemental output from the free shaft is mechanically coupled with the secondary propulsion system with a mechanical clutch actuation system. To engage the secondary propulsion system, clutch-type actuation mechanisms mechanically couple the secondary propulsion system with the primary propulsion system. Thus, not only does the secondary propulsion system directly reduce the amount of power available to the primary propulsion system, the mechanical clutch coupling limits free operation of the secondary propulsion system since the secondary propulsion system must rotate at speeds dictated by the free shaft. Thus, there is a need for a convertible propulsion system that provides greater flexibility in distributing power between primary and secondary propulsion systems. 
       SUMMARY 
       [0004]    A method of powering a rotary-wing aircraft according to an exemplary aspect of the present disclosure includes selectively coupling and uncoupling a first power turbine which drives a rotor system and a second power turbine which drives a secondary propulsion system to change a power distribution between the rotor system and the secondary propulsion system. 
         [0005]    A method of powering a rotary-wing aircraft according to an exemplary aspect of the present disclosure includes moving a retractable port to control flow of combustion gases downstream of a first power turbine into a bypass passage to selectively bypass a second power turbine and change a power distribution between a rotor system driven by the first power turbine and a secondary propulsion system driven by the second power turbine. 
         [0006]    A method of powering a rotary-wing aircraft according to an exemplary aspect of the present disclosure includes changing a pressure within a pressure chamber between a first power turbine and a second power turbine to change a first pressure ratio over the first power turbine and a second pressure ratio over the second power turbine to change a power distribution between a rotor system driven by the first power turbine and a secondary propulsion system driven by the second power turbine. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0007]      FIG. 1  shows a rotorcraft having a convertible gas turbine propulsion system of the present invention. 
           [0008]      FIG. 2A  shows a schematic diagram of the convertible gas turbine propulsion system of  FIG. 1  operating in low speed or hover mode. 
           [0009]      FIG. 2B  shows a schematic diagram of the convertible gas turbine propulsion system of  FIG. 1  operating in high speed or cruise mode. 
           [0010]      FIG. 3A  shows a schematic diagram of a first embodiment of a pneumatic actuator configured for operating the gas turbine propulsion system of  FIG. 2A  in low speed or hover mode. 
           [0011]      FIG. 3B  shows a schematic diagram of the first embodiment of the pneumatic actuator of  FIG. 3A  configured for operating the gas turbine propulsion system of  FIG. 2B  in high speed or cruise mode. 
           [0012]      FIG. 4A  shows a cross-sectional view of a second embodiment of a pneumatic actuator configured for operating the gas turbine propulsion system of  FIG. 2A  in low speed or hover mode. 
           [0013]      FIG. 4B  shows a cross-sectional view of the second embodiment of the pneumatic actuator of  FIG. 4A  configured for operating the gas turbine propulsion system of  FIG. 2A  in high speed or cruise mode. 
           [0014]      FIG. 5  shows a cross-sectional view of a second embodiment of the convertible gas turbine propulsion system of the present invention including a two-stage secondary power turbine having a variable vane actuation system. 
           [0015]      FIG. 6  shows a perspective view of the aft end of the convertible gas turbine engine of  FIGS. 4A and 4B . 
       
    
    
     DETAILED DESCRIPTION 
       [0016]      FIG. 1  shows one embodiment of a rotorcraft in which convertible gas turbine propulsion system  10  of the present invention can be used.  FIG. 1  shows helicopter  12  having convertible gas turbine propulsion system  10  configured for driving main rotor  14  and pusher propeller  16 . Main rotor  14  and pusher propeller  16  allow helicopter  12  to operate in a variety of modes, such as forward and reverse vectoring, vertical take-off and landing, hovering and cruising at pilot commanded forward speeds. As such, helicopter  12  is able to fulfill various mission requirements for commercial, civil and military operations. Although convertible gas turbine propulsion system  10  is shown being used for driving main rotor  14  and propeller  16  in helicopter  12 , propulsion system  10  can be used in conjunction with a variety of rotorcraft to drive a variety of propulsion systems. For example, pusher propeller  16  could be replaced with an exhaust nozzle, and main rotor  14  could be replaced with a turbo-prop for use in a fixed wing aircraft. 
         [0017]    Convertible gas turbine propulsion system  10  includes gas generator  18 , first power turbine  20  and second power turbine  22 . Gas generator  18  comprises a gas turbine engine for generating high pressure, high energy gases for turning first power turbine  20  and second power turbine  22 . First power turbine  20  provides rotating shaft energy to main rotor  14  through horizontal shaft  24 , gearbox  26  and vertical shaft  28  such that main rotor  14  rotates in a generally horizontal plane parallel to the plane in which helicopter  12  travels. Due to rotation in the generally horizontal plane, main rotor  14  always produces some amount of upward lift. Forward thrust is produced by, among other things, adjusting the pitch and tilt of the rotor blades. Second power turbine  22  provides rotating shaft energy to pusher propeller  16  through horizontal shaft  30  such that pusher propeller  16  rotates in a generally vertical plane perpendicular to the plane in which helicopter  12  travels. When operating, pusher propeller  16  always provides forward thrust to helicopter  12  due to its rotation in the vertical plane. It is necessary, however, to eliminate forward thrust produced by pusher propeller  16  such that helicopter  12  is able to operate in hover mode. As such, convertible gas turbine propulsion system is provided with exhaust system  32  to permit selective operation of second power turbine  22 . In high speed and cruise modes, exhaust system  32  directs combustion gases through second power turbine  22  to obtain additional forward thrust from pusher propeller  16 . In low speed and hover modes, exhaust system  32  permits combustion gases to bypass second power turbine  22  such that helicopter  12  can operate similar to that of a conventional helicopter with the required thrust being produced by main rotor  14 . Exhaust system  32  permits second power turbine  22  to be mechanically uncoupled from first power turbine  20  thereby allowing performance characteristics of first power turbine  20  and second power turbine  22  to be individually designed and adjusted. 
         [0018]      FIGS. 2A and 2B  show schematic diagrams of convertible gas turbine propulsion system  10  of  FIG. 1  operating in low speed and high speed modes, respectively. Convertible gas turbine propulsion system  10  includes gas generator  18 , first power turbine  20 , second power turbine  22  and exhaust system  32 . First power turbine  20  is connected to main rotor  14  through shaft  24 , gear box  26  and shaft  28 . Second power turbine  22  is connected to pusher propeller  16  through shaft  30  and gearbox  33 . Gas generator  18  includes compressor section  34 , combustor  36 , gas generator turbine  38  and gas generator casing  40 . Exhaust system  32  includes forward duct  42 , inner duct  44 , retractable port  46  and outer duct  48 , which acts as a secondary exhaust duct to conduct exhaust gas out of first power turbine  20  and second power turbine  22 . Gas generator casing  40 , which acts as a primary exhaust duct, comprises an annular passageway surrounding gas generator  18  such that inlet air can be routed into compressor section  34  to produce compressed air and exhaust gas can be directed out of gas generator turbine  38 . With the addition of a fuel, a combustion process is carried out within combustor  36  to produce combustion gases. Forward duct  42  guides the combustion gases through first power turbine  20 , while inner duct  44 , retractable port  46  and outer duct  48  selectively guide the combustion gases through second power turbine  22 . 
         [0019]    Compressor section  34  comprises a series of rotating blades that drive inlet air A 1  past a series of stationary vanes to increase the pressure of inlet air A I . A fuel from injectors (not shown) is added to inlet air A I  within combustor  36  such that a combustion process can be carried out to produce high energy combustion gases A G  for turning gas generator turbine  38 , first power turbine  20  and second power turbine  22 . Gas generator turbine  38  is connected to compressor section  34  through shaft  50  and provides rotating shaft power to compressor section  34  such that compressed air is continuously supplied to combustor  36  to sustain the combustion process. Casing  40  surrounds the gas flow path of gas generator  18  and is sealed such that combustion gases A G  can be delivered to first power turbine  20  and second power turbine  22  to perform useful work. As a result of the combustion process, combustion gases A G  are delivered to forward duct  42  in a high pressure, high energy state having pressure P 1 , which is significantly higher than ambient pressure P A . 
         [0020]    Forward duct  42  is connected to the downstream end of gas generator casing  40  and concentrically surrounds first power turbine  20 . Forward duct  42  receives combustion gases A G  from gas generator turbine  38 . Combustion gases A G  exit gas generator turbine and impinge upon the blades of first power turbine  20  to produce rotational movement of first power turbine  20 . As combustion gases A G  travel from combustor  36  and gas generator turbine  38  they are continuously depressurizing and ultimately expanding to ambient pressure P A  as the turbine blades extract kinetic energy from combustion gases A G . Thus, after passing through first power turbine  20 , combustion gases A G  reach pressure P 2  in pressure chamber  52  between first power turbine  20  and second power turbine  22 , which is less than P 1 . Accordingly, a pressure differential is produced across first power turbine (P 1 −P 2 ), which defines the work capability of first power turbine  20  and, depending on the torque of main rotor  14 , dictates the speed at which first power turbine  20  rotates. First power turbine  20  rotates shaft  24 , which through gearbox  26  and shaft  28  drives rotor  14 . The pressure differential (P 1 −P 2 ) is modulated by controlling fuel flow to combustor  36 . Thus, the speed of rotor  14  is controlled by directly controlling the operating characteristics of gas generator  18  and first power turbine  20 . Gas generator  18  and first power turbine  20  function in a conventional manner and operate as a turboshaft gas turbine engine. 
         [0021]    Inner duct  44 , retractable port  46 , and outer duct  48  dictate the mode in which combustion gases A G  leave first power turbine  20  to exit convertible gas turbine propulsion system  10  as exhaust gas A E . Specifically, exhaust system  32  modulates the magnitude of pressure P 2  between first power turbine  20  and second power turbine  22  within pressure chamber  52 . Outer duct  48  is connected at the downstream end of forward duct  42  to form an outer annular exhaust path for combustion gas A G  and exhaust gas A E . Outer duct  48  extends radially outward from the downstream end of first power turbine  20  to beyond the perimeter of inner duct  44 , and axially past the downstream end of second power turbine  22 . Inner duct  44  is disposed within outer duct  48  and extends from the upstream end of second power turbine  22  past the downstream end of second power turbine  22 . Together, outer duct  48  and inner duct  44  form bypass passage  54  around second power turbine  22 . Retractable port  46  provides a means for selectively routing combustion gases A G  into inner duct  44  or permitting combustion gases A G  to flow around inner duct  44 . In the embodiment shown, retractable port  46  comprises a moveable annular ring that can be disposed between forward duct  42  and inner duct  44  to force combustion gases A G  to flow through second power turbine  22  ( FIG. 2B ), or can be disposed between outer duct  48  and inner duct  44  to permit combustion gases A G  to bypass second power turbine  22  ( FIG. 2A ). 
         [0022]      FIG. 2A  shows a schematic diagram of convertible gas turbine propulsion system  10  of  FIG. 1  operating in low speed or hover mode. Retractable port  46  is positioned such that combustion gases A G  are able to bypass second power turbine  22  and enter passage  54 . Passage  54  opens to the atmosphere such that pressure P 4  within passage  54  is approximately equal to ambient pressure P A . With retractable port  46  retracted, chamber  52  is also open to passage  54  and the atmosphere such that pressure P 2  is also approximately equal to ambient pressure P A . Pressure P 3  at the downstream end of second power turbine  22  is also exposed to ambient pressure P A  as exhaust system  32  opens to the atmosphere to expel exhaust gas A E . Thus, the pressure difference (P 2 −P 3 ) across second power turbine  22  is approximately equal to zero and negligible work is extracted from second power turbine  22 . Due to the negligible pressure differential, pusher propeller  16  is not driven by combustion gas A G  and is free to “windmill” in the atmosphere as helicopter  12  travels. Pusher propeller  16  is typically locked down with a mechanical brake for, among other reasons, safety. Accordingly, pusher propeller  16  generates no forward thrust and helicopter  16  is propelled by output of first power turbine  20 , namely, the rotation of main rotor  14 . Thus, helicopter  16  is operable in hover mode or other low speed modes similar to that of a conventional helicopter. As discussed earlier, horizontally rotating rotor blades are limited in the amount of speed they can produce. Thus, to achieve speeds beyond what is available from rotor  14 , retractable port  46  of exhaust system  32  is moved into position to direct combustion gas A G  through second power turbine  22  to drive pusher propeller  16 . 
         [0023]      FIG. 2B  shows a schematic diagram of convertible gas turbine propulsion system  10  of  FIG. 1  operating in high speed and cruise mode. Retractable port  46  is positioned between forward duct  42  and inner duct  44  such that a contiguous duct is formed around first power turbine  20 , chamber  52  and second power turbine  22 . Retractable port  46  prevents combustion gases A G  from entering bypass passage  54  and outer duct  48 , and forces combustion gases A G  to enter second power turbine  22 . Additionally, the position of retractable port  46  prevents ambient air from mixing with combustion gases A G  between first power turbine  20  and second power turbine  22  within pressure chamber  54 . As a result, combustion gases A G  are forced through second power turbine  22 . Combustion gases A G  enter the upstream end of second power  22  turbine at pressure P 2 , expand and depressurize, then exit the downstream end of second power turbine  22  at pressure P 3 . 
         [0024]    Since pressure chamber  54  between first power turbine  20  and second power turbine  22  is sealed from ambient pressure P A , pressure P 2  remains greater than ambient pressure P A , but less than pressure  P1 , such that first power turbine  20  extracts energy from combustion gases A G . The downstream end of second power turbine  22  is open to atmospheric pressure P A  such that as combustion gases A G  work their way through second power turbine  22 , they lose pressure such that pressure P 3  is approximately equal to ambient pressure P A , which is less than pressure P 2 . Thus, second power turbine  22  extracts energy from combustion gases A G  and converts it to rotation of shaft  30  to drive pusher propeller  16 . Thus, combustion gases A G , which would otherwise be expelled from first power turbine  20  and released to the atmosphere in a conventional gas turbine engine, are put to work again to drive pusher propeller  16 . Pusher propeller  16  provides additional horizontal thrust to helicopter  12  such that velocities above what can be provided by rotor  14  alone can be achieved. 
         [0025]    Second power turbine  22  extracts energy from the combustion process that would otherwise be available for driving first power turbine  20 . Thus, at high forward speeds, work available from first power turbine  20  that would have been inefficiently used by main rotor  14  is redirected to second power turbine  22 , and is used more efficiently by propeller  16  and results in better overall fuel economy. Power distribution between main rotor  14  and propeller  16  is changed by changing pressure P 2  within pressure chamber  52  to change the pressure ratios (P 1 −P 2 ) and (P 2 −P 3 ). This can be accomplished by appropriately sizing the aerodynamic flow areas of first power turbine  20  and second power turbine  22 , or by varying the flow area of second power turbine  22  through the incorporation of variable area vanes (such as is discussed later with respect to variable vanes  138  of  FIG. 5 ). In one embodiment, the power is split 40/60 between first power turbine  20  and second power turbine  22 , respectively, during high speed operation of convertible gas turbine propulsion system  10  and is achieved by appropriately sizing the aerodynamic flow areas of first power turbine  20  and second power turbine  22  ( fixed area vanes  106  are used, as shown in  FIGS. 4A and 4B ). However, specific distribution of power between first power turbine  20  and second power turbine  22  is based on design needs and other factors such as the rotational speed of shaft  24 , the gas flow rate through gas generator  18 , the inlet and outlet temperatures and pressures of first power turbine  20 , the number of turbine stages in first and second power turbines  20  and  22 , the use of variable vanes, and the desired power output of second power turbine  22 . It is contemplated that up to eighty percent of the power generated from propulsion system  10  could be generated by second power turbine  22 . To return propulsion system  10  to low speed operation, exhaust system  32  is simply returned to the configuration of  FIG. 2A , where overall fuel efficiency increases due to improved efficiency of main rotor  14  operating in the low speed and hover mode like that of a conventional helicopter. 
         [0026]    In another embodiment of convertible gas turbine propulsion system  10 , second power turbine  22  is omitted in favor of a thrust producing exhaust nozzle. In such a configuration, second power turbine  22 , shaft  30  and propellers  16  would be removed from inside inner duct  44  and propulsion system  10  completely, and inner duct  44  would be configured as an exhaust nozzle. For example, inner duct  44  could be configured as a conventional converging or converging-diverging nozzle to accelerate the flow of the exhaust gas, as is known in the art. Retractable port  46  would similarly operate to open and close bypass passage  54 . With retractable port  46  open, the pressure differential across the exhaust nozzle would not be large enough to accelerate the mass flow of exhaust gas A E  through the nozzle shaped from inner duct  44 . With retractable port  46  closed, however, the pressure differential across the exhaust nozzle would be large enough to produce additional thrust from exhaust gas A E . For a typical turboshaft configuration of first power turbine  20  and gas generator  18 , exhaust gas A E  would be subsonic such that a convergent exhaust nozzle would be used. In other embodiments, however, divergent or convergent-divergent nozzles could be used. 
         [0027]      FIGS. 3A and 3B  show schematic diagrams of a first embodiment of a pneumatic actuator configured for operating gas turbine propulsion system  10  in low speed or high speed modes, respectively. Gas turbine propulsion system  10  includes pusher propeller  16 , gas generator  18 , first power turbine  20 , second power turbine  22 , shaft  24 , gear box  26 , shaft  28 , shaft  30 , exhaust system  32 , compressor section  34 , combustor  36 , gas generator turbine  38 , gas generator casing  40 , forward duct  42 , inner duct  44 , retractable port  46 , outer duct  48 , shaft  50 , chamber  52  and passageway  54 , which function similarly to what is described with respect to  FIGS. 2A and 2B . As such, second power turbine  22  receives input power from gas generator  18  and first power turbine  20 , and supplies output power to propeller  16  when actuator  56  pushes retractable port  46  rearward to the closed position. In the embodiment shown, pneumatic actuator  56  includes plenum  60 , thruster  62 , strut  64  and spring  66 . Retractable port  46  includes forward guide  68 , aft guide  70  and bypass opening  72 . Pneumatic actuator  56  receives compressor bleed air A C  from compressor section  34  through valve  58  to actuate retractable port  46 . Additionally, plenum  60  introduces air A C  into nozzle  73  of thruster  62 , which accelerates air A C  through thruster  62  to provide axial thrust to overcome the spring force and push retractable port  46  rearward. 
         [0028]    In  FIG. 3A , valve  58  is closed such that retractable port  46  is in the forward, open position. Retractable port  46  is configured for sliding on rails or tracks between inner duct  44  and plenum  60 . The rear end of retractable port  46  is connected with spring  66 . Spring  66  is disposed between retractable port  46  and strut  64 . Strut  64  comprises a support structure for supporting outer duct  48  about inner duct  44 . A bottom portion of strut  64  extends through inner duct  44  to provide stopper  65 , which forms a backstop for spring  66 . Spring  66  is sized to push retractable port  46  in the forward direction to the forward position. In the forward position, forward guide  68  of retractable port  46  is disposed adjacent forward duct  42  above first power turbine  20 , and aft guide  70  is disposed adjacent inner duct  44  above second power turbine  22 . Bypass opening  72  is, therefore, disposed between outer duct  48  and pressure chamber  52 , which, through passageway  54 , is exposed to ambient pressure P A . Additionally, in the forward position, thruster  62  is disposed adjacent plenum  60  such that it is positioned for receiving compressor bleed air A c  from compressor section  34  through valve  58 . However, in the low speed mode shown in  FIG. 3A , valve  58  is closed such that compressor air A C  is not delivered to thruster  62 , and spring  66  biases adjustable portion to the forward position. As such, combustion gases A G  are passed through passageway  54  such that a negligible pressure differential (P 3 −P 2 ) is produced across second power turbine  22 . Thus, second power turbine  22  does not extract power from combustion gases A G , and pusher propeller  16  does not produce forward thrust. In this embodiment, the forward positioning of retractable port  46  such that pressure chamber  52  is open to ambient pressure is the default, fail safe mode of operation for exhaust system  32 . For example, in the event of failure of valve  58 , retractable port  46  is pushed by spring  66  to the forward position to open chamber  52  such that engine  10  operates to drive main rotor  14  as a conventional turboshaft, which is the most fuel efficient mode of operation. 
         [0029]    In  FIG. 3B , valve  58  is opened such that retractable port  46  is pushed to the rearward position. Valve  58  is opened to permit a portion of inlet air A l  compressed within compressor section  34  to be bled off into conduit  74 . Valve  58  is connected to compressor section  34  and plenum  60  by conduit  74 , which comprises any suitable means for conducting compressor bleed air A C . In various embodiments, suitable compressed air for pressurizing plenum  60  can be bled from any source of pressurized air within engine  10 , such as a high or low pressure compressor. In one embodiment, conduit  74  comprises stainless steel piping. Conduit  74  delivers compressor air A C  to plenum  60  such that plenum  60  becomes pressurized. The inlet of nozzle  73  is adjacent plenum  60  and the exit of nozzle  73  is open to ambient pressure P A . Nozzle  73  is angled in the forward direction as nozzle  73  extends from inlet to exit. In one embodiment, compressor air A C  is delivered to plenum  60  at sub-sonic speeds and nozzle  73  comprises a converging or converging-diverging nozzle. As, such, nozzle  73  accelerates compressor air A C  in the forward direction as bleed air A C  is expanded to ambient pressure P A . Thruster  62  reacts the forward force of compressor air A C , thus causing rearward movement of thruster  62  and retractable port  46 . The size and length of spring  66  and the position of strut  64  prevent retractable port  46  from moving too far in the aft direction and determine the force that nozzle  73  must generate. Retractable port  46  is pushed back such that opening  72  is retracted between inner duct  44  and second power turbine  22 . Forward guide  68  is retracted to separate passageway  54  from chamber  52 . As such, combustion gases A G  bypass passageway  54  and pass through chamber  52  such that an operative pressure differential (P 3 −P 2 ) is produced across second power turbine  22 . Thus, pusher propeller  16  produces forward thrust and gas turbine propulsion system  10  operates in high speed or cruise mode. In one embodiment, valve  58  is controlled by an automated control system, such as a FADEC (Full Authority Digital Engine Control), to control opening and closing of valve  58 . The automated control system regulates the discharge of pressurized air from valve  58  to adjust the amount of overlap between bypass opening  72  and bypass passage  54 . As such the pressure differential (P 3 −P 2 ) across second power turbine  22  can be varied to control thrust output of propeller  16 . To cease high speed operation and powered output of propeller  16 , valve  58  closes to stop providing compressor air A C  to plenum  60 , and a vent hole incorporated into valve  58  is opened to relieve pressure within conduit  74  between plenum  60  and valve  58 . 
         [0030]    Retractable port  46  is slid along the aforementioned rails to close off passage  54  from exhaust gas A E . In one embodiment, the rails are dovetail shaped to prevent rotation of retractable port  46  between outer duct  48  and inner duct  44 . By maintaining proper alignment, dovetail shaped rails also minimize thrust requirements of thruster  62  by reducing friction. Thruster  62  must produce enough rearward thrust to overcome the forward force of spring  66 . In other embodiments, thruster  62  includes a plurality of nozzles. The specific size, number and geometry ( converging or converging-diverging) of nozzles depends on the specific thrust required to move retractable port  46 , which depends on the size and performance requirements of gas turbine propulsion system  10 . Pneumatic actuator  56  is particularly well-suited for generating thrust levels for actuating retractable port  46  scaled for use with small-sized or low-thrust gas turbine propulsion systems. Exhaust system  32  is also configurable with other types of pneumatic actuators for use in larger scale or high-thrust gas turbine propulsion systems  10 . 
         [0031]      FIGS. 4A and 4B  show cross-sectional views of a second embodiment of a pneumatic actuator configured for operating gas turbine propulsion system  10  in low speed or high speed modes, respectively.  FIG. 4A  shows exhaust system  32  of convertible gas turbine propulsion system  10  with pneumatic actuator  75  configured for low speed or hover mode operation. Propulsion system  10  includes first power turbine  20  and second power turbine  22 , and exhaust system  32  includes forward duct  42 , inner duct  44 , retractable port  46  and outer duct  48 , which function similarly to what is described with respect to  FIGS. 2A and 2B . Exhaust system  32  also includes inner guide  76  and aft duct  78 . Pneumatic actuator  75  comprises canister  80 , cover  82 , spring  84 , push rod  86 , piston spring head  88  and compressor air manifold  90 . Actuator  75  comprises an automated piston that retracts and extends retractable port  46  into and from inner duct  44  to selectively permit flow of exhaust gas A G  from first power turbine  20  to enter outer duct  48 , thereby bypassing second power turbine  22 . 
         [0032]    Outer duct  48  is connected with forward duct  42 , which surrounds first power turbine  20 . First power turbine  20  includes vanes  92  and blades  94 , through which exhaust gas A G  flows. Stub shaft  96  connects first power turbine  20  to shaft  24 , and drives main rotor  14  ( FIGS. 2A &amp; 2B ). Inner duct  44  is positioned between outer duct  48  and second power turbine  22  to form bypass passage  54 . Outer duct  48  and inner flow guide  76  form pressure chamber  52  and direct exhaust gases A G  toward inner duct  44  such that combustion gases A G  can enter bypass passage  54  or second power turbine  22 . Inner flow guide  76  is aerodynamically designed to minimize losses and in one embodiment includes a bump to stream line the flow between first power turbine  20  and second power turbine  22 , or between power turbine  20  and bypass passage  54 . Inner duct  44  and aft duct  78  comprise annular rings that surround second power turbine  22  and guide exhaust from engine  10 , respectively. The outer surface of inner duct  44  forms the inner surface of bypass passage  54 , and the outer surface of aft duct  78  forms the outer surface of bypass passage  54 . The inner surface of inner duct  44  forms the outer surface of the flow path for second power turbine  22 , and the inner surface of aft duct  78  forms the inner surface of the flow path for second power turbine  22  Likewise, in another embodiment of the invention, the inner surface of inner duct  44  and the inner surface of aft duct  78  can be configured as an exhaust nozzle to supplement or replace second power turbine  22 . 
         [0033]    Actuator  75  is connected to the aft end of inner duct  44 , such as with threaded fasteners  98 . Actuator  75  is disposed within aft duct  78 , which is connected to the aft end of outer duct  48 , such as with threaded fasteners  100 . Aft duct  78  is supported by bearing housing  102 , which is part of bearing assembly  104 . Second power turbine  22 , which includes vanes  106  and blades  108 , drives output shaft  110 , which is supported within exhaust system  32  by bearing assembly  104 . The inner surface of inner duct  44  comprises a stationary rail to which vane  106  of second power turbine is secured. Blades  108  are attached to a rotor connected to output shaft  110 , which is connected to shaft  30  such as with a spline for driving propeller  16  ( FIG. 1 ). Aft duct  78  also includes a plurality of exit guide struts  111 , which are disposed on either side of canister  84  within aft duct  78 , to guide combustion gases A G  out of propulsion system  10 . Although  FIGS. 4A and 4B  show second power turbine including one set of turbine blades, as can be seen in  FIG. 5 , inner duct  44  includes space such that second power turbines having multiple stages can be readily incorporated into convertible gas turbine propulsion system  10 . 
         [0034]    In  FIG. 4A , retractable port  46  of exhaust system  32  is retracted into inner duct  44  by actuator  75  such that exhaust gas A G  is able to bypass second power turbine  22 . Inner duct  44  comprises a hollow compartment, or plenum  112 , which receives compressor bleed air from conduit similar to how plenum  60  receives compressor air A C  from conduit  74  in  FIGS. 3A and 3B . Compressor bleed air is used to actuate retractable port  46  within compartment  113 . Plenum  112  also receives cooling air from cooling air manifold  90  to prevent overheating of inner duct  44  and actuator  75  from impingement of exhaust gas A G . Conduit for cooling and actuation air can be routed through engine  10  to plenum  112  using any suitable pathway or number of pipes as needed, depending on design needs. 
         [0035]    Inner duct  44  comprises an annular cowling that includes compartment  113  for receiving retractable port  46 . Retractable port  46  includes head  114 , tail  116 , elongate main body  117 , and leaf seal  118 . Elongate main body  117  connects head  114  to tail  116 . Head  114  is configured for mating with nose seal  120  of outer duct  48 . Leaf spring  118  seals off compartment  113  to prevent exhaust gas A G  from entering inner duct  44 , and to provide piston surface area for overcoming the force of spring  84 . Spring  118 , which fits around elongate main body  117 , also centers and retains retractable port  46  within compartment  113 . Tail  116  is configured to receive push rod  86  from within canister  80 . Canister  80  comprises an elongate cylinder mounted to inner duct  44  axially downstream of compartment  113 . Canister  80  provides a piston cylinder for housing push rod  86  and piston spring head  88 . Push rod  86 , which comprises a rigid beam, extends through the forward end of canister  80  and the aft end of compartment  113  to engage tail  116 . Seal  122  surrounds push rod  86  to prevent air flow between canister  80  and compartment  113 . Spring  84  is disposed within canister  80  between piston spring head  88  and spacer  124 . Piston spring head  88  is affixed to the downstream end of push rod  86  such that spring  84  biases push rod  86  in the aft direction. Push rod  86  is affixed to tail  116  such that push rod  86  biases retractable port  46  in the aft direction. In the retracted position, spring  84  pushes piston spring head  88  in the aft direction to withdraw retractable port  46  from nose seal  120  and into compartment  113 . Plenum  112  includes port  126 , which permits compressor bleed air to enter compartment  113  to force retractable port  46  forward. In this embodiment, the rearward positioning of retractable port  46  such that pressure chamber  52  is open to ambient pressure is the default, fail safe mode of operation for exhaust system  32 . For example, in the event of failure of plenum  112 , retractable port  46  is pushed by spring  84  to the rearward position to open chamber  52  such that engine  10  operates to drive main rotor  14  as a conventional turboshaft, which is the most fuel efficient mode of operation. 
         [0036]      FIG. 4B  shows exhaust system  32  of convertible gas turbine engine  10  with pneumatic actuator  75  configured for high speed operation. In  FIG. 4B , retractable port  46  of exhaust system  32  is extended from inner duct  44  by actuator  75  such that exhaust gas A G  is prevented from bypassing second power turbine  22 . Compressor bleed air enters plenum  112  from the previously mentioned conduit (not shown). Compressor bleed air fills plenum  112 , which is sealed at its forward and aft ends, to force retractable port  46  forward by impinging on leaf seal  118 . Seal  122  prevents compressor bleed air from leaking from compartment  113  into canister  80  such that piston force is not degraded. Leaf spring  118  seals the forward portion of compartment  113  near the aft end of retractable port  46 . Leaf spring  118  engages the inner and outer diameter walls of compartment  113  to form a piston head. Leaf spring  118  traps compressor bleed air and provides an area upon which compressor bleed air pushes. The force with which leaf spring  118  and push rod  86  are pushed is equal to the pressure of compressor bleed air multiplied by the cross-sectional area of compartment  113 . The pressure of compressor bleed air is controlled by compressor section  34  and valve  58 . The area of compartment  113  is constrained by size requirements of actuator  75 . The cross-sectional area of compartment  113  and the pressure of the compressor bleed air are selected to produce enough forward force to overcome the rearward force of spring  84 . Thus, compressor bleed air pushes retractable port  46  out from compartment  113  and into passageway  54 . 
         [0037]    Push rod  86  and piston spring head  88  are also pulled in the forward direction within canister  80 . Piston spring head  88  includes vent  128  to prevent pressurization of air within canister  80 . Spring spacer  124  controls the stroke length of push rod  86  and can be replaced with different sized spacers for use in different embodiments of actuator  75 . Retractable port  46  is pushed far enough in the forward direction such that head  114  engages nose seal  120 . Leaf spring  118  is placed far enough back on retractable port  46  such that head  114  engages nose seal  120  to prevent further forward advancement of retractable port  46 . Retractable port  46  is rigid to form a firm seal with nose seal  120  on outer duct  48 . Head  114  is shaped to seal off bypass duct  54  when extended from inner duct  44 . In the embodiment shown, head  114  includes a generally flat surface to engage nose seal  120 , and a nose portion to engage outer duct  48 . Nose seal  120  comprises a flexible seal that biases head  114  to retain the position of retractable port  46  against outer duct  48 . As such, substantially all of exhaust gas A G  is prevented from entering bypass passage  54  with leakage held to a minimum. Thus, a pressure differential (P 3 −P 2 ) is produced across second power turbine  22  such that useful work can be extracted from exhaust gas A G  to turn output shaft  110  and propeller  16 . As with actuator  56  of  FIGS. 3A and 3B , the distance that retractable port  46  is withdrawn from outer duct  48  can be varied through the use of valve  58  and a FADEC to control the pressure differential (P 3 −P 2 ) across second power turbine  22 . 
         [0038]    As with actuator  56  of  FIGS. 3A and 3B , actuator  75  utilizes a resource readily available within gas turbine propulsion system  10  to obtain the energy necessary to displace retractable port  46 : compressed air. Compressed air is commonly bled from compressor sections within gas turbine engines for various cooling and actuating purposes. Actuators  56  and  75  are readily adapted to such systems to provide actuation power and to provide cooling of the actuators. Actuators  56  and  75  do not require supplies of oil for actuation, cooling or lubrication and thus do not require additional linkages or complexities associated with such sub-systems. Thus, actuators  56  and  75  provide easily integrated, lightweight and cost effective systems for displacing retractable port  46  and activating convertible gas turbine propulsion system  10  in a hot operating environment. Additionally, in other embodiments of the invention, the pneumatic thruster actuation of actuator  56  of  FIGS. 3A and 3B  and the pneumatic piston actuation of actuator  75  of  FIGS. 4A and 4B  can be combined to provide supplemental or redundant actuation systems. 
         [0039]      FIG. 5  shows a cross-sectional view of a second embodiment of convertible gas turbine engine  10  of the present invention including a two-stage secondary power turbine  130  having a variable vane actuation system  132 . Propulsion system  10  includes first power turbine  20 , forward duct  42 , inner duct  44 , retractable port  46 , outer duct  48 , pneumatic actuator  75 , flow guide  76 , aft duct  78 , stub shaft  96 , fasteners  98  and  100 , bearing housing  102 , bearing assembly  104 , output shaft  110 , exit guide struts  111  and nose seal  120 , which function similarly to what is described with respect to  FIGS. 4A and 4B . A portion of push rod  86  of pneumatic actuator  75  is shown, however, other components are omitted from  FIG. 5  for clarity. Power turbine  130  is connected with output shaft  110  in a similar manner as is second power turbine  22 . Power turbine  130 , however, includes first stage blades  134  and second stage blades  136 , which are inter-disposed with variable stator vanes  138  and stationary vanes  140 . First stage blades  134  and second stage blades  136  are connected to rotor disks, which are connected to each other with threaded fastener  142 . As such, first stage blades  134  and second stage blades  136  rotate to drive output shaft  110  when driven by exhaust gas A E , such as when retractable port  46  is extended into bypass passage  54 . Stationary vane  140  is suspended from the inner surface of inner duct  44  and suspended between first stage blades  134  and second stage blades  136 . Variable stator vanes  138  are disposed between inner duct  44  and inner shroud  144 , which are connected to the aft end of flow deflector  76  with threaded fastener  146 . Variable vanes  138  include inner and outer trunnions  150  and  148 , respectively, upon which variable vanes  138  rotate. Inner trunnions  150  rotate within inner shroud  144  and outer trunnions  148  rotate within the bottom surface of inner duct  44 . Each inner trunnion  150  includes a crank arm  152  that is connected to inner sync ring  154 , which coordinates the rotation of variable vanes  138  between inner shroud  144  and inner duct  44 . Crank arms  152  apply rotational torque to inner trunnions  150  when activated by variable vane actuation mechanism  132 . 
         [0040]    Outer trunnions  148  are connected to actuation mechanism  132 , which comprises linkage  155 , idler  156 , and actuator  158 . Actuation mechanism  132  extends from outside of outer duct  48  to inside inner duct  44  through exit guide strut  111 . Actuator  158 , which in the embodiment shown comprises a hydraulic or pneumatic cylinder, is mounted to propulsion system  10  outside of outer duct  48  such that actuator  158  remains stationary with respect to outer duct  48  and is positioned in a cool environment conducive to satisfactory operation of actuation mechanism  132  and other conventional actuation systems. Piston rod  160  extends from actuator  158  and connects with idler  156 . Idler  156  extends through outer duct  48  and into exit guide strut  111 , where idler  156  is pinned at pin  162 . The outer end of idler is connected with piston rod  160  and the inner end of piston rod  160  is connected to linkage  155 . Piston rod  160  moves the outer end of idler  156  such that idler  156  rotates about pin  162 . Linkage  155  extends axially from idler  156  to outer trunnions  148 . A trunnion of a master variable vane includes a master crank arm such that rotational torque applied to the master variable vane is transmitted to variable vanes  138  through torque ring  154 . For example, piston rod  160  is pulled within actuator  158  through hydraulic or pneumatic pressure activation, which causes idler  156  to pull linkage  155  rearward from inner duct  44 . Linkage  155  pulls the master crank arm, which rotates the master variable vane. The master variable vane rotates one of crank arms  152 , which then rotates the other crank arms  152  through torque ring  154  to rotate the full array of variable vanes  138 . As such, the amount of energy extracted from second power turbine  130  can be varied. Specifically, the pressure differential across second power turbine  130  can be varied with variable vanes  138  to regulate the power input to propeller  16  and hence the amount of forward thrust produced. 
         [0041]      FIG. 6  shows a perspective view of the aft end of convertible gas turbine propulsion system  10  of  FIGS. 4A and 4B . Propulsion system  10  includes gas generator casing  40 , which shrouds gas generator  18 ; second engine case  42 , which shrouds first power turbine  20 ; outer duct  48 , which shrouds second power turbine  22 ; and aft duct  78 , which shrouds inner duct  44 . Inner duct  44  comprises an annular structure that is supported between outer duct  48  and bearing housing  102  by exit guide struts  111  to form bypass passage  54 . Struts  111  include annular passages that receive canisters  80  of actuation mechanism  75  and, as such, form downstream extensions of inner duct  44  (as is seen in  FIGS. 4A and 4B ). Output shaft  110  is disposed within bearing housing  102  and is connected with second power turbine  22 , the blades of which are disposed forward of exit guide struts  111  between inner duct  44  and bearing housing  102 . Propeller shaft  30  is connected with output shaft  110  to drive propellers  16  ( FIGS. 1-2B  and  4 ). Inner duct  44  comprises a hollow structure that houses retractable port  46 , which is extendable from inner duct  44  to guide flow of exhaust gas A E  through second power turbine  22 . Inner duct  44  also provides a platform onto which to mount actuators  75  (not shown), including canisters  80 . Canisters  80  include springs that push retractable port  46  forward to close off bypass passage  54 . The number of actuators  75  and canisters  80  is selected to generate enough rearward spring force to retract retractable port  46 . Likewise, the number of actuators  75  and canisters  80  is selected to generate enough forward pneumatic force to extend retractable port  46  from inner duct  44 . Actuators  75  and canisters  80  are distributed evenly around inner duct  44  such that retractable port  46  easily extends from and contracts within inner duct  44  without binding. Canisters  80  include springs  84  ( FIGS. 4A &amp; 4B ) that pull retractable port  46  to the retracted, fail safe position, such that engine  10  operates as a conventional turboshaft and second power turbine  22  within inner duct  44 , or the exhaust nozzle formed by inner duct  44  ceases operation. In the embodiment shown, six actuators having six canisters  80  are used to push retractable port  46  from inner duct  44 . In one embodiment, outer duct  48 , aft duct  78  and canisters  80  extend the length of engine  10  approximately 19 inches (.about.48.3 cm). Variable vane actuation system  132  of  FIG. 5  can also be incorporated into propulsion system  10  of  FIG. 6 . For example, actuation system  132  could be integrated into one of the six shown exit guide struts  111 , such that in  FIG. 6 , an idler  156  would extend radially from one of exit guide struts  111  to connect with an actuator  158  disposed axially along outer duct  48 . 
         [0042]    The various embodiments of convertible gas turbine propulsion system  10  of the present invention provide a means for splitting power between primary and secondary propulsion systems. Each propulsion system can be individually modulated to produce a desired amount of propulsive thrust such that, for example, vertical and horizontal thrust levels can be controlled to increase engine fuel efficiency and performance at high forward speeds. The convertible gas turbine propulsion system includes actuators that utilize available engine resources, such as pneumatic power, to distribute power between the first and second propulsion systems such that additional bulky and heavy actuation systems are unnecessary. 
         [0043]    Although the present invention has been described with reference to preferred embodiments, workers skilled in the art will recognize that changes may be made in form and detail without departing from the spirit and scope of the invention.