Abstract:
An aircraft ( 10 ) having turboprop propulsion has a plurality of gas turbine engines ( 11 ), each with a two stage reduction gearbox ( 30,41 ) with the engine output shaft ( 17,17 A) inline with the propeller drive output shaft ( 19,51 ) to drive a propeller ( 12 ) in front of the engine in a tractor mode of propulsion. The input gear reduction stage ( 20 ) and output gear reduction stage ( 21 ) share a plurality layshafts ( 22,38 ) mounted in fixed circularly-spaced relation to each other about the axis of the output shafts in the mechanical housing in the aircraft. Each layshaft of said plurality of layshafts has a layshaft first end toward the rear, input end of the gearbox, and a layshaft second end toward the output, front end of the gearbox. The layshaft gears are arranged to avoid any net thrust loading of the layshafts. Spur ( 26 ) and double helical ( 27 ) are used in FIG.  4,  while spur gears  37  and  48  are used in FIG.  6  with suitable helix settings to neutralize end thrust on layshaft  38.  Cantilever layshaft input gear ( 37 ) mounting in rear bearings ( 39 ) enables it to share some layshaft output gear  48  load sharing with the front bearing ( 40 ). A torquemeter ( 61 ) is conveniently situated at the gearbox input end. The engine output gear ( 25,36 ) and gearbox output gear ( 28,49 ), and propeller ( 12 ) all have colinear rotational axes.

Description:
The present application is a continuation of PCT/US99/21228 filed Sep. 15, 1999, currently pending. The PCT application was based on a provisional patent application Ser. No. 60/100,933, filed Sep. 18, 1998, and the benefit of Ser. No. 60/100,933 is claimed in the PCT application. The content of PCT/US99/21228 and U.S. provisional application No. 60/100,933 are incorporated herein. 
    
    
     BACKGROUND OF THE INVENTION 
     The present invention relates generally to the design and construction of a gearbox for a gas turbine engine. More particularly, the present invention has one embodiment wherein the gearbox defines an inline two-stage reduction gearbox for a gas turbine turboprop engine. 
     Gas turbine turboprop engine designers generally couple a gear reduction gearbox with the engine in order to reduce the output shaft speed, and increase the torque delivered to an output device, such as a propeller. It is well known that gas turbine engines are high speed rotary equipment having components including an output shaft revolving at speeds from about 5,000 to 50,000 revolutions per minute. Sometimes, in order to harness the power from the output shaft of the gas turbine engine a gear reduction gearbox is coupled to the engine to decrease shaft rotation speed and increase output torque. Gear reduction gearboxes include gear sets therein for reducing the shaft speed during the transmission of power from the gas turbine engine to the propeller. 
     The application of gas turbine turboprop engines as a propulsion means for an aircraft often creates design parameter conflicts, such as the need for a durable long life gear train and the necessity to minimize the volume and weight of the respective engine. Prior designers of gas turbine engine gear reduction gearboxes have generally used multi-stage gearboxes to effectuate significant shaft speed reduction. Even with the variety of prior gas turbine gear reduction gearboxes there remains a need for an improved gear reduction gearbox. The present invention satisfies this need in a novel and unobvious way. 
     SUMMARY OF THE INVENTION 
     One form of the present invention contemplates an apparatus, comprising: an aircraft; at least one gas turbine turboprop engine coupled to the aircraft, the at least one gas turbine engine having an inlet end and an exhaust end, and a first member for transmitting power from the engine; a two stage reduction gearbox positioned proximate the inlet end and coupled to the engine, the gearbox having an input gear reduction stage coupled with and driven by the first member and a double helical output gear reduction stage coupled with a second member for transmitting power; and at least one propeller coupled to and driven by the second member. 
     Another form of the present invention contemplates an apparatus, comprising: an aircraft; at least one gas turbine turboprop engine coupled to the aircraft, the at least one gas turbine engine having an air inlet end and an exhaust end, and a first power transmission member for transmitting power from the engine; a gearbox positioned proximate the inlet end and coupled to the engine, the gearbox having two stages defined by an input gear reduction stage and an output gear reduction stage, the input gear reduction stage is coupled with and driveable by the first power transmission member and the output gear reduction stage is coupled with and drives a second power transmission member for transmitting power, the input gear reduction stage is defined by one of a spur or double helical gearing and the output gear reduction stage is defined by double helical gearing; and at least one propeller is coupled to and driven by the second power transmission member. 
     One aspect of the present invention contemplates a combination, comprising: an aircraft; at least one gas turbine engine coupled to the aircraft, the at least one gas turbine engine having an engine first end and an engine second end, and an engine output member for transmitting power from the engine; and a reduction gearbox positioned proximate the engine first end and coupled to the engine, the gearbox comprising: a mechanical housing; two gear reduction stages defined by an input gear reduction stage and an output gear reduction stage; a plurality of layshafts coupled with and having at least a portion thereof disposed within the mechanical housing, each of said plurality of layshafts having a layshaft first end and a layshaft second end; the input gear reduction stage includes a plurality of input gears, each of the plurality of layshafts having one of the input gears coupled to the layshaft first end and driven by the engine output member, the output gear reduction stage including a plurality of output gears, each of the plurality of layshafts having one of the output gears coupled to the layshaft second end, the plurality of output gears engaging and driving a gearbox output member that is coupled to a propeller; and wherein the plurality of input gears and the plurality of output gears eliminate thrust loading from the plurality of layshafts. 
     Another aspect of the present invention contemplates an apparatus, comprising: an aircraft; a gas turbine turboprop engine coupled to the aircraft, the gas turbine engine having an inlet end, an exhaust end, and a first power transmission member for transmitting power from the engine; a reduction gearbox positioned proximate the inlet end and coupled to the engine, the gearbox including: a mechanical housing; the gearbox having two stages of reduction defined by an input gear reduction stage and an output gear reduction stage, the input gear reduction stage defined by a first gear coupled to the first power transmission member and a plurality of second gears in meshing engagement with the first gear; the output gear reduction stage defined by a third gear and a plurality of fourth gears in meshing engagement therewith; a plurality of layshafts that are rotatably coupled with the mechanical housing, each of the plurality of layshafts having one of the plurality of second gears coupled at one end and one of the plurality of fourth gears coupled at the other end; the third gear coupled to and drives a second power transmission shaft having a propeller coupled thereto; and the input gear reduction stage is defined by one of a spur or double helical gearing and the output gear reduction stage is defined by a double helical gearing. 
     A preferred form of the invention includes torquemeter components inline with the engine output member and a power input gear of the first gear reduction stage. 
     The preferred form also mounts the input gears of the first stage layshafts in overhanging or cantilever manner from the input end of the layshafts, and the output gears of the second gear reduction stage in location between layshaft support bearings. The layshaft support bearings are located closer together and the output end bearing is relatively large in diameter, for better bearing load distribution and for longer bearing life. 
     One object of the present invention is to provide an improved propeller gear reduction gearbox. 
     These and other objects will become more apparent from the following description. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     FIG. 1 is an illustrative view of an aircraft having a plurality of gas turbine turboprop engines coupled thereto. 
     FIG. 2 is a schematic view of one of the gas turbine turboprop engine comprising a portion of FIG.  1 . 
     FIG. 3 is an illustrative view of one embodiment of a propeller gearbox comprising a portion of the FIG. 2 gas turbine turboprop engine. 
     FIG. 4 is a sectional view of one embodiment of the propeller gearbox comprising a portion of the FIG. 2 gas turbine turboprop engine. 
     FIG. 5 is a sectional view of an alternative embodiment of a propeller gearbox comprising a portion of the FIG. 2 gas turbine turboprop engine. 
     FIG. 6 is a sectional view of a flier alternative embodiment of a propeller gearbox comprising a portion of the FIG. 2 gas turbine turboprop engine. 
    
    
     DESCRIPTION OF THE PREFERRED EMBODIMENT 
     For the purposes of promoting an understanding of the principles of the invention, reference will now be made to the embodiment illustrated in the drawings and specific language will be used to describe the same. It will nevertheless be understood that no limitation of the scope of the invention is thereby intended, such alterations and flier modifications in the illustrated device, and such further applications of the principles of the invention as illustrated therein being contemplated as would normally occur to one skilled in the art to which the invention relates. 
     With reference to FIG. 1, there is illustrated an aircraft  10  having a plurality of gas turbine turboprop engines  11  coupled thereto for providing means for propelling the aircraft. In one embodiment the aircraft is a cargo aircraft, such as but not limited to the FLA turboprop being developed by Airbus. However, it should be clear that the present invention is applicable to turboprop aircraft in general and not a specific aircraft unless specifically stated. While FIG. 1, shows an aircraft having four engines it is not intended to limit the present disclosure to an aircraft having only four engines unless specifically stated. Aircraft having other quantities of engines from  1 - 6  are contemplated herein. Each of the plurality of gas turbine turboprop engines  11  includes a propeller  12  in front of the engine for tractor type propulsion. In the preferred embodiment, the propeller defines a single rotation propeller. However, it is clearly understood that the present invention is not limited to one particular aircraft design. 
     With reference to FIG. 2, there is illustrated a block diagram of one embodiment of the gas turbine turboprop engine  11 . Gas turbine engine  11  has forward inlet end  11   a  and an exhaust end  11   b.  In a preferred embodiment, the gas turbine turboprop engine  11  includes a compressor  13 , combustor  14 , turbine  15  and power turbine  16 . One example of a gas turbine engine turboprop engine is the model DART available from Rolls-Royce. A person of ordinary skill in the art will appreciate that there are a multitude of ways to link the components together. Additional compressors and turbines could be added with intercoolers connecting between the compressors and reheat combustion chambers between the turbines. Power turbine  16  has a power transmission member  17  coupled thereto for delivering power to the propeller gearbox  18 . The propeller gearbox  18  defines a single input, such as power transmission  17  member, and a single output, such as power transmission member  19 , gearbox for driving the propeller  12 . Input power transmission member  17  and output power transmission  19  are aligned with an axis X. In one embodiment, the output power transmission member  19  rotates in a clockwise direction as indicated by arrow Y as viewed from the exhaust end  11   b  of the engine  11 . 
     With reference to FIGS. 3 and 4, there is illustrated a form of the inline gear reduction system for the gas turbine turboprop engine  11 . The gearbox  18  has two stages of reduction, which include a first stage of reduction  20  and a second stage of reduction  21 . The first stage of reduction has a driven output gear  25  that is coupled to the power transmission member  17  coming from the gas turbine engine  11 . Driven output gear  25  meshes with and drives a plurality of input gears  26 . In the illustrated form of the present invention, four equally spaced gears  26  that engage the output gear  25  define the plurality of input gears  26 . In this embodiment, the gear components for the first stage reduction  20  are preferably selected from spur or double helical gear sets, and preferably the gearing for the first stage reduction  20  of this embodiment is defined by a spur gear set. The term double helical gearing as utilized herein includes but is not limited to herringbone, a double helical gear having no separation between two helical portions, and a gear having a separation between the two helical portions. Referring to FIG. 4, there is illustrated a double helical gear having a separation zone  100  between the two helical portions  101  and  102 . 
     The second stage of reduction  21  may also be referred to as the output stage and is defined by a single gearbox output gear  28  which meshes with a plurality of output gears  27 . Preferably, four equally spaced gears define the plurality of output gears  27 . The output gears  27  are utilized to drive the gearbox output gear  28  which is coupled to an output power transmission member  19  that drives the propeller  12 . In the preferred embodiment the second stage/output stage  21  defines a double helical gearing, arrangement. A plurality of layshafts  22  is positioned within a mechanical housing and they extend substantially parallel with the axis X. Each of the plurality of layshafts  22  has a first end  23  and an opposite second end  24 . The first end  23  having an input gear  26  coupled thereto and the second end  24  having an output gear  27  coupled thereto. It is understood that other quantities of output gears  27 , input gears  26 , and layshafts  22  are contemplated herein as needed by system parameters. 
     With further reference to FIG. 4, there is illustrated a sectional view of the gearbox  18  coupled to the gas turbine turboprop engine  11 . The inline dual stage gear reduction system is located proximate the inlet end  11   a  of the gas turbine engine  11 . A plurality of mounting members  29  couple the mechanical housing  30  and supporting structure of the gearbox  18  to the gas turbine engine  11 . It is understood herein that the mechanical housing  30  may be formed integral with the gas turbine engine  11  housing, or may be a separate component with the mechanical housing containing the gearbox  18  therein. The layshafts  22  are supported at each of their first end  23  and their second end  24  by a bearing  31 . In this embodiment, the bearings are rolling element radial load bearing and, because a double helical second stage  21  and a spur or double helical first stage  20  is utilized, the thrust loading on the plurality of layshafts  22  is eliminated. The first stage input gears  26  and second stage output gears  27  are located between the bearings  31  which support the layshafts for rotation in the mechanical housing/structure  30 . In this embodiment, the gearbox output gear  28  is defined as an external gear. The external gear  28  is coupled to the output power transmission member  19  that drives the propeller  12 . The design and construction of the gearbox  18  allows the access for prop pitch control mechanisms through the passageway  35  extending through the gas turbine engine  11  and which is inline with the: output power transmission member  19 . 
     In the illustrated embodiment, the output gear  28  and output power transmission member  19  are coupled together as by splines, for example. Radial load bearing  33  and thrust bearing  32  support the output member  19  of the assembly to rotate relative to the mechanical housing. In the preferred embodiment, the bearings  32  and  33  are rolling element bearings, bearing  32  being a ball bearing, and bearing  33  being a roller bearing. 
     With reference to FIG. 5, there is illustrated an alternate embodiment of an inline two stage reduction gearbox  180  that is substantially similar to the gearbox  18  and like feature numbers will be used to define like elements. The distinction that will be discussed with relation to FIG. 5 is that the second/output stage  210  of the gearbox  180  utilizes a dual helical gearing wherein the output gear  34  is an internal gear. More specifically, the output gear  34  defines a ring gear that is coupled to the output power transmission member  19 . Owing to the utilization of an internal gear  34 , the mechanical structure and bearings supporting the layshaft  220  have been rearranged. But the layshaft input gears  26 , and the layshaft output gears  34  are between the bearings  31  as in the FIG. 4 embodiment. It is understood herein that other mechanical designs can be utilized to integrate an internal gear/ring gear  34  into an inline two-stage gear reduction gearbox utilizing a double helical gearing for the output stage. 
     With reference to FIG. 6, the gas turbine engine output shaft  17 A is splined at  17 B to gear box input shaft  35  to which input gear  36  is splined at  36 A. Therefore, power output from the gas turbine is delivered to gear  36  which is in mesh with four circularly spaced gears  37 , this combination comprising the first stage reduction. The gears  37  are on layshafts  38 . Four such layshafts, circularly spaced about axis X are believed preferable, but as in the other embodiments, other numbers may be acceptable. The layshafts are mounted in bearings  39  and  40 , whose outer races are mounted in the gear box housing  41  which is attached to the engine housing  11  as by bracket  43  at the top and  44  at the bottom. Of course, other attachment points are also used at appropriate circularly spaced locations about the axis X. The outer races of the bearings  39  are received in apertures in the wall or bulkhead  46  fixed in the gear box housing in a manner very similar to the mounting of the bearings  33  in the embodiments of FIGS. 4 and 5. But in this embodiment of FIG. 6, in contrast to the embodiments of FIGS. 4 and 5, the bearings at both ends of the layshafts are large, and closer together. This is facilitated by the conical wall gear web providing cantilever mounting of the gear  37  on the layshaft input end, overhanging the input end bearing  39 . Each of the layshafts  38  has a helical gear  48  thereon engaged with the gear box output gear  49  which is splined to the output power transmission member  51  at  52 . The inner end of the member  51  is supported by roller bearings  53  whose outer race is mounted in a central aperture in the web  46 . The output end of the member  51  is mounted in the combination of roller bearing  54  and ball bearing  56  whose outer races are mounted in the gear box housing  41 . 
     The gears in FIG. 6 are all helical. While double helical can be used in both reduction stages, as can be done in the first two embodiments, single helical is preferable. Thrust loading of the layshafts is avoided by setting the helix angles on the gears  39  and  40  so that the end thrust developed at the input gears  39  is balanced by the end trust developed in the output gears  40 . 
     Referring further to FIG. 6, a torquemeter assembly  61  is provided in the engine housing. It comprises two tubular features. The first is a cylindrical wall  62 A extending axially from the web of the input gear  36 . The second is the tube  62 B centered on axis X and supported by ball bearing  63 . The rear end of tube  62 B is received on the front end of engine shaft at  17 F and fixed to it by a spanner nut  62 C. Spanner nuts are used at various other locations in these gear boxes to hold gears on shafts, or shafts together at splines, and to hold bearing races on tubes, in ways known in the art. Wall  62 A and tube  62 B have facing ends at gap  64 . The ends are notched in registry with each other as represented generally by square notches at  66 . Sensors  67  and  68  located adjacent the notches are used to detect and quantify any displacement of the notches in the wall  62 A relative to the facing notches in tube  62 B. This information can be transmitted to a computer to indicate the amount of twist in the shaft  17 B and, thereby, the torque being transmitted, in a manner known in the art. Such information is used in connection with engine control. While it is known to use torquemeters of some types for engine control, it is not believed not to have been done. inline between a gas turbine engine output shaft and its reduction gearbox input gear shaft colinear with the engine output shaft. 
     The cantilever mounting of the layshaft input gears enables the bearings  39  and  40  to be larger in diameter without enlarging the gearbox housing. Also, the very large loads usually occurring at the output end of the layshafts can be better distributed, with a greater part of the loads being taken by the input end bearing  39  than in the previous embodiments. Therefore this embodiment is expected to provide longer bearing life under normal conditions and under oil-off conditions too, than is likely in the embodiments of FIGS. 4 and 5. 
     In all three embodiments, there is access to propeller pitch mechanism from the back of the propeller shaft For example, in FIG. 6, the rear end of pitch control components  71  is at  72 , radially inboard of the bearing  53 . It is accessible through the space  74  behind the wall  46 . In the FIGS. 4 and 5 embodiment, where the pitch control components are not shown but would be located as in the FIG. 6 illustration, the rear end would be accessible through an opening in the wall of portion  30 S of the housing at a space or spaces between the layshafts as at  30 A, for example. 
     Embodiments of the gearbox of the present invention have been designed for use with gas turbine turboprop engines having horsepower between 5,000 and 20,000 horsepower. A more preferred form of the present invention is designed for utilization with a gas turbine engine developing 11,260 horsepower. Further, the dual stage gear reduction gearbox is designed for section ratios in the range of 5:1-20:1. A more preferred form of the present invention uses an overall reduction ratio of 8.63:1. The output speed of the rotating propeller is preferably within a range of 700 to 1,500 revolutions per minute and more preferably has a maximum speed output of about 850 revolutions per minute. However, other horsepower, gear reduction ratios and output speeds are contemplated for the propeller gearbox of the present invention. 
     While the invention has been illustrated and described in detail in the drawings and foregoing description, the same is to be considered as illustrative and not restrictive in character, it being understood that only the preferred embodiment has been shown and described and that all changes and modifications that come within the spirit of the invention are desired to be protected.