Convertible aircraft engine

A convertible aircraft engine, capable of turboshaft and/or turbofan modes of operation, is provided in a configuration that combines certain components to permit variation in bypass flow to match engine airflows with the mode of engine operation. The provided components permit the engine to decrease the power requirements of the engine's forward fan when the engine is operated in the turboshaft mode. To permit a decrease in fan power requirements, the fan is split into an outer portion and an inner portion separated by a rotating shroud. Airflow into the fan's outer portion is controlled with a part span inlet guide vane and an outlet guide vane. The guide vanes can be used to lessen the load on the outer portion of the fan while the inner portion continues to accelerate and compress a normal airflow into an engine compressor for supplying the engine's core. The power normally used to drive the fan outer portion is used instead to power a transfer shaft in the turboshaft mode.

BACKGROUND OF THE INVENTION 
This invention relates to aircraft gas turbine engines that are capable of 
conversion from turbofan to turboshaft modes of operation and vice versa. 
Considerable attention has been directed at various efforts to develop an 
aircraft engine that is capable of both turbofan and turboshaft operation. 
Commonly referred to as convertible aircraft engines, powerplants that 
have been conceived with this dual capability would be desirable for 
powering an aircraft in a vertical direction as a helicopter and 
alternatively for powering an aircraft in a forward direction as a typical 
turbofan-powered jet. The obvious advantage of dual mode operation is that 
the aircraft could take off and land like a helicopter but could also fly 
forward at relatively high speeds like a jet-powered airplane. 
One example of a convertible engine concept is disclosed by W. J. Stein et 
al in U.S. Pat. No. 3,351,304. As evidenced by this disclosure, W. J. 
Stein et al conceived an aircraft powered by turbofan engines that are 
connected to a helicopter-type, vertical-lift rotor by a primary gear 
train with appropriate reduction gearing. The engine additionally has a 
secondary gear train, which, when actuated, motors the rotor at lower 
r.p.m. to reduce resistance during turbofan-powered forward-thrust 
operation with a manually operated clutch to selectively actuate either 
the primary or secondary gearing. The gearing and shifting arrangement 
permits the speed of the fan to be modulated through the gear ratios. The 
advantage is that in one phase of flight operation the engine is turning 
the fan at low r.p.m., thereby permitting a major part of the engine power 
to be delivered as a shaft horsepower to the rotor system, while at the 
same time restricting the thrust produced by the fan to a low level. In 
the alternate phase of operation, the gearing permits the fan speed to be 
increased while at the same time limiting the amounts of shaft horsepower 
diverted to the rotor system. 
As might be expected, this type of system tends to be inefficient because 
the engine thermodynamics do not match the actual engine cycle. Also, the 
requirements of the gearing and clutching system are extreme and would 
require large and heavy equipment that is undesirable in aircraft 
applications. 
Since the time period of the Stein et al invention, numerous technical 
advances have been made in the area of multi-cycle aircraft engines. Some 
of these advances permit variable bypass operation of a turbofan engine. 
This new technology has fostered innovations relating variable bypass 
system applications to convertible-type engines. 
SUMMARY OF THE INVENTION 
Therefore, it is an object of the present invention to provide a 
convertible engine that is capable of powering an aircraft forward like an 
airplane or in a lifting mode like a helicopter in an efficient manner 
without the use of massive gearing and clutching systems. 
It is another object of the present invention to provide a convertible 
engine that varies bypass flow for the purpose of providing a lesser 
bypass airflow during turboshaft operation and a greater bypass airflow 
during turbofan operation to match the actual engine cycle with 
operational requirements of the aircraft. 
Briefly stated, these and other objects are attained in a convertible 
turboshaft/turbofan engine arrangement that combines certain components 
that permit variation in bypass flow in an engine that has a transfer 
shaft for providing shaft horsepower to an external location, such as a 
helicopter rotor. This combination of components allows the engine to 
effectively decrease the power requirements of the engine's forward fan 
when the power requirements of the transfer shaft are increased. 
A decrease in fan power requirements is accomplished with a split fan 
having an outer portion and an inner portion separated by a rotating 
shroud. Airflow into the fan's outer portion is controlled with a part 
span inlet guide valve (IGV). This part span IGV can be used to lessen the 
load on the outer portion of the fan while the inner portion continues to 
accelerate and compress a normal airflow into an engine compressor for 
supplying the engine's core. The power normally used to drive the fan 
outer portion is used instead to power the transfer shaft. Notably, even 
when the guide vanes are completely closed, there will always be some 
bypass flow. Regardless of IGV position, the bypass flow is mixed with the 
core engine flow in an engine tailpipe with a chute-type mixer. The mixing 
of flows improves specific fuel consumption during turbofan operation and 
cools the exhaust to lessen infrared signature out the engine's exhaust 
nozzle during turboshaft operation. Compared to a separated exhaust 
system, the mixed exhaust system in a closed IGV mode reduces the back 
pressure on the low pressure turbine, thereby increasing the shaft power 
available and reducing the residual thrust of primary exhaust.

DETAILED DESCRIPTION 
Referring now to FIG. 1, a typical gas turbofan engine 10 is shown for the 
purpose of briefly describing its internal workings. Air enters the engine 
10 at an inlet 12 that directs the air into a fan 14 for accelerating the 
inlet airflow. The fan 14 drives the air downstream (from left to right in 
FIG. 1) into both a bypass duct 16 and a compressor 18 for a core engine 
20. The air entering the bypass duct 16 is used to provide a majority of 
the forward thrust generated by the engine 10 and is called bypass air 
because it bypasses the engine's core 20. An outlet guide vane 21 is 
provided for controlling fan airflow into the bypass duct at varying fan 
speeds. 
The remaining portion of the inlet air that does not enter the bypass duct 
is directed into the compressor 18 where it is compressed to support a 
core engine combustion process that takes place inside the combustor 22. 
To create this combustion, fuel is injected into the highly compressed air 
in the combustor 22 and ignited to provide high-energy combustion gases 
that drive a high pressure turbine 24. The high pressure turbine 24 
converts this extracted energy into shaft horsepower for driving the 
compressor 18. 
A low pressure turbine 26 is disposed downstream of the high pressure 
turbine 24 in a position to receive the exiting flow of hot gases leaving 
the high pressure turbine 24. The low pressure turbine is so named because 
the combustion gases have dropped in pressure somewhat after part of their 
energy has been dissipated in the high pressure turbine. Additional energy 
is extracted by the low pressure turbine from the combustion gases, again 
for conversion into shaft horsepower, but this time to drive the fan 14. 
From the low pressure turbine 26, the core engine combustion gases flow 
into an engine exhaust 28. An alternate approach would be to provide a 
long fan duct and mixing section where the bypass flow can be mixed with 
the core engine exhaust flow to increase engine efficiency. 
The engine 10 in FIG. 1 is a good example of a typical turbofan engine used 
to propel aircraft in a forward direction at speeds that approach Mach 1. 
As stated earlier, in this type of engine, the majority of the forward 
propulsive thrust is provided by the fan 14. 
A cross section of a turboshaft type of engine used to power a helicopter 
aircraft would, in many ways, look quite similar to the cross section 
shown in FIG. 1. The major differences in a turboshaft engine would be 
that the fan 14 and bypass duct 16 would be eliminated, and the low 
pressure turbine 24, with the use of shafts and appropriate gearing, would 
power a helicopter rotor to lift the aircraft instead of a fan to drive 
the aircraft forward. 
Referring now to FIG. 2, a convertible engine 30 is shown that has been 
conceived for the purpose of providing a gas turbine engine that is 
capable of exclusively turbofan or exclusively turboshaft operation or 
both simultaneously. This powerplant is intended to supply the right 
amount of shaft horsepower to the helicopter rotor for liftoff and landing 
and, additionally, be convertible to a turbofan mode of operation for 
forward flight when the helicopter is unloaded. 
The convertible engine 30 is provided with a fan 32 ahead of a core engine 
or gas generator 34 in a configuration similar to that of existing 
turbofan engines. A low pressure turbine 36 supplies power to the fan 32 
through a low pressure turbine shaft 38 mounted concentrically to the gas 
generator 34. 
A bevel gear set 40 is provided on this same shaft 38 for the purpose of 
transferring power through a power transfer shaft 42, which is housed in a 
suitable strut 43 through a fan inlet 44. The power transfer shaft 42 
would be mechanically connected to a main gear set and helicopter rotor 
(not shown). When the aircraft is flying as a helicopter, a feature of 
this invention is that the shaft 38 can transfer its power through the 
bevel gear set 40 and transfer shaft 42 to a helicopter rotor instead of 
to the fan 32. 
Inlet guide vanes 46 are provided just downstream of the fan inlet 44 but 
ahead of the fan 32. The basic characteristics of guide vanes are known to 
those skilled in the art, and it is known that vanes can be rotated to 
different positions to control the volume of inlet airflow into the fan 
32. When the aircraft is operating as a helicopter in a turboshaft (lift) 
mode of operation, the inlet guide vanes 46 can be rotated to a "closed" 
position to reduce the inlet airflow and thereby reduce the load on the 
fan 32 with the fan still rotating. This permits a majority of the 
mechanical power from the shaft 38 to be transferred to the helicopter 
rotor instead of the fan 32. 
Outlet guide vanes 48 which are fixed in a conventional turbofan can be 
made variable as a means of further reducing the power absorbed by the fan 
32. 
In order to provide for additional control of the engine's airflow and 
thereby help increase the engine's efficiency, the fan 32 is split into an 
outer portion 50 and inner portion or fan hub 52. The inner and outer 
portions of the fan are separated by a rotating shroud 54. The inner 
portion 52 supercharges the engine core 34 with a compressed airflow that 
is separated by the shroud 54 from being affected by the airflow past the 
outer portion 50. The reason for this separation is that this allows the 
core engine supercharging and inlet flow conditions to be relatively 
independent of the fan inlet guide vane position. Since the vanes 46 are 
part span inlet guide vanes, meaning the vanes physically extend into the 
engine's bypass stream region only, the vane position will tend to affect 
the fan outer portion airflow only. The airflow that flows radially 
inwardly of the vanes 46 is then relatively unaffected by vane position. 
It has been found to be quite important to use part span inlet guide vanes 
to maintain reasonable airflow conditions entering the core engine 34 over 
the entire range of inlet guide vane angles. The shroud 54 tends to 
further promote this airflow separation, which generally increases the 
engine performance at inlet guide vane closures. 
There can be some special airflow considerations with a split fan 
arrangement especially when matching the fan inner portion airflow with 
the core engine air requirements over various power settings since core 
engine flow will vary even though the fan speed is held constant. 
Apparatus for matching airflows is shown in detail in FIG. 3. A bleed valve 
56 is provided between an engine bypass duct 58 and an engine core inlet 
60. By incorporating this bleed valve 56, an outlet is provided to permit 
redirection of excess air supercharged by the fan inner portion 52 when it 
is not required by the engine core 34. It is expected that bleeding of fan 
airflow would be required only at part power engine settings when fan 
speed is held at maximum levels. 
Referring again to FIG. 2, it is noteworthy that an inlet 64 to the fan 
inner portion 52 and a centerbody housing 66 for the bevel gear set 40 can 
both be arranged to provide a line-of-sight separation to prevent foreign 
objects from entering the engine core. This line-of-sight separation is 
not a necessary aspect of the engine 30, but can be desirable in most 
practical applications. 
At a downstream end of the convertible engine 30, the bypass air from duct 
58 is mixed with the core engine primary exhaust flow in an engine 
tailpipe 68 with a chute-type mixer 70. The mixed exhaust then passes out 
through a fixed jet nozzle 72. In a turbofan forward flight mode of 
operation, the mixed exhaust provides a specific fuel consumption benefit. 
In a turboshaft lift mode of operation, the air that passes through the 
outer portion 50 of the fan 42 and becomes bypass flow will tend to dilute 
the primary exhaust flow. The lower temperature of the resulting exhaust 
out the nozzle 72 reduces the infrared signature which would especially 
benefit military use of the convertible engine. 
This mixed exhaust arrangement offers another advantage. In the turboshaft 
mode of operation, back pressure on the low pressure turbine 36 is reduced 
when the bypass flow is reduced by closing the inlet guide vanes 46. This 
increases the power output of the convertible engine 30 and reduces the 
residual thrust of the engine. Residual thrust is generally undesirable in 
the turboshaft mode of operation. It has been suggested that, without the 
mixed exhaust, a variable primary nozzle might be required to reduce this 
residual thrust to acceptable levels. 
A modification that can be made to the embodiment shown in FIG. 2 is an 
addition of an inner part span, variable inlet guide vane 71 positioned 
forward of the fan inner portion 52. This additional inlet guide vane 
would control airflow intended for the engine compressor 34 and will 
assist in matching the fan inner portion airflow to the core engine 
requirements. This optional modification is shown in FIG. 3 but not in 
FIG. 2. 
Referring now to FIG. 4, an alternate embodiment of the present invention 
is shown in cross section to aid in comparison with the embodiment shown 
in FIG. 2. The additional features shown in FIG. 4 include a booster stage 
74 driven by the low pressure shaft 38. The booster stage 74 is needed if 
the fan hub cannot produce enough pressure ratio for the purpose of 
compressing the flow from the fan hub 52 to the level required by 
additional aircraft systems that require compressed air, such as a vehicle 
rotor circulation control system. 
In explanation, some of the new aircraft concepts that have recently been 
conceived use aircraft wings as both a rotary blade for lifting the 
aircraft and as a fixed wing when the aircraft flies like a jet. 
Compressed air is required to provide circulation control of the wing. 
This air could be provided by a separate compressor but there may be a net 
saving if the air is supplied by the main engine. 
The fan hub 52 and booster stage 74 are sized to supply air to both the 
engine core 34 and the vehicle wing or rotor (not shown). An additional 
inner variable inlet guide vane 76 ahead of the fan hub 52 is provided to 
assist in matching the airflow characteristics of the fan hub 52 and 
booster 74 with the airflow requirements of the core 34 and rotor 
circulation control system. The compressed air for the rotor circulation 
control system is collected in a scroll 78 and passed through a suitable 
auxiliary duct 80 to the rotor or wherever the compressed air is required. 
A variable bleed port 82 might also be provided for matching airflow from 
the booster stage with the requirements of the rotor circulation control 
system. 
The advantages of the embodiment of the convertible engine 73 shown in FIG. 
4 is that it can supply all three major power requirements of a 
convertible engine powered aircraft with a single powerplant. That is to 
say the engine 73 can provide: first, shaft power for the vehicle rotor to 
lift the aircraft; second, a fan airflow for thrust to power the aircraft 
as a jet; and third, compressed air for circulation control of the vehicle 
rotor. 
While two embodiments of the present invention are described in detail, it 
can be readily appreciated that various changes could be made in the 
embodiments shown without departing from the scope of the invention. 
Therefore having described the preferred embodiment of the invention, 
though not exhaustive of all possible equivalents, what is desired to be 
secured by Letters Patent of the United States is claimed below.