Note: Descriptions are shown in the official language in which they were submitted.
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HIGH MACH NUMBER UNDUCTED FAN ENGIN~
This invention relates generally to aircraft gas
turbine engines and, more particularly, to turboprop
engines capable of achieving a relatively high altitude
and mach number.
BACKGROUND OF TH~ INVENTION
. .
Gas ~urbine engines for use on aircraft fall
generally into three categories. These include turbojet,
turbofan, and turboprop engines. All three engines have a
core which includes a compressor, combustor, and turbine
in serial flow relationship. The compressor compresses
air entering the core to a relatively high pressure. This
high pressure air is then mixed with fuel in the combustor
and burned to form a high energy gas stream. This gas
stream passes through the turbine where work is extracted
to drive the compressor.
These three gas turbine engines differ in the
manner in which they produce thrust. A turbojet uses the
reactive force of the gas stream itself to provide the
required thrust. A turbojet engine is capable of
operating over a wide range of flight speeds including
supersonic (Mach Number ~ 1). However, it is also
characterized by relatively low thrust at take-off and
generally poor fuel efficiency.
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In contrast to the turbojet, the turbofan and
turboprop engines generate thrust primarily through a
propulsor radially disposed with respect to the core
engine. In such engines, the high energy gas stream
leaving the core turbine is expanded through a second
turbine, known as a power turbine or low pressure turbine,
which drives the fan or propulsor. Although some thrust
is produced by the residual gas stream exiting the core
nozzle, most of the thrust produced is generated by the
propulsor. The ratio of the mass of air which passes
through the propulsor (the air bypassing the core engine)
to the mass of air passing through the core engine is
known as the bypass ratio (~). The bypass ratio (~) is a
rough measure of the ratio of thrust of the propulsor to
thrust of the core exhaust.
Although the turboprop and turbofan engines
operate under similar thermodynamic cycles, the two
engines represent significantly different approaches to
the design of gas turbine engines for aircraft. The
conventional turboprop engine includes a relatively large
diameter propeller with several blades for moving a
relatively large volume of air and imparting to it a
relatively small pressure rise. In contrast, the turbofan
engine includes a significantly smaller diameter fan
~5 section encased within a cowling. The fan has a
relatively large number of blades and imparts a relatively
higher pressure rise to the volume of air passing
therethrough, than the turboprop. For example, the
pressure rise or pressure ratio of a typical propeller is
on the order of 1.1 whereas the pressure ratio of a
typical fan is about 1.7.
The differences between the three engines
described above may be measured by the performance of each
engine over a range of flight conditions, such as speed
and altitude. Important measures of the performance of
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the engine are engine efficiency and thrust. Engine
efficiency includes a constituent representing how
efficiently the heat energy of the fuel is converted into
kinetic energy and a constituent measuring how efficiently
the kinetic energy is converted to propulsive work. In
other words, engine efficiency is made up of the thermal
or internal efficiency of the engine and by propulsive or
e~ternal efficiency of the engine.
Simply stated, propulsive efficiency is the ratio
of the wor~ done by the engine over the useful energy
imparted to the engine airflow. Algebraically, this may
be expressed by the formula:
Ne = 2Vo/(VO + Vj), (1)
where Ne = propulsive efficiency~
VO = aircraft speed,
and Vj = exhaust velocity.
As can be seen, propulsive efficiency approaches 100% as
Vj approaches VO That is, propulsive efficiency
becomes high when the velocity of the exhaust gases
approaches the velocity of the aircraft.
The thrust generated by an aircraft engine is
proportional to the mass of air moved by the engine
multiplied by the difference between the exit velocity and
aircraft speed. This may be represented by:
Fn = Wa(Vj - vo)/gc~ (2)
` where Fn = net thrust,
Wa = mass flow,
and gc = a constant.
It is clear that large thrusts are obtainable by having a
large mass flow or a large difference between the exhaust
gas velocity and aircraft speed. Referring now to both
equations (1) and (2), it becomes clear that for a given
thrust it is more efficient to give a large mass of air a
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small increase in velocity relative to the aircraft
speed. Such is a characteristic of the turbofan and
turboprop engine which distinguishes them from turbojet
engines.
A difference between a conventional turboprop and
turbofan engine is that the turboprop exhibits a rapid
fall off in engine efficiency at higher aircraft speeds.
This is due to the supersonic flow relative to each
propeller blade which increases the drag as the tip speed
approaches the speed of sound. In contrast, the turbofan
is capable of achie~ing higher overall efficiencies at
high flight mach numbers because a diffuser section of the
turbofan cowling reduces the speed of the incoming air
below that of the aircraft speed. However, a limitation
of the turbofan engine is that increased bypass ratios for
increased propulsive efficiency require larger cowlings
which result in excess weight and drag which seriously
degrades aircraft fuel burn efficiency.
A recent improvement over the turbofan and
turboprop engines is the so-called unducted fan engine,
such as disclosed in Cdn. Serial Number 438,676 filed
Oct. 7, 1983 - Johnson.- The unducted fan engine includes
features of a conventional turboprop such as no cowling
and variable pitch blades which are thin and swept to
yield good efficiency at high aircraft speed as well as
features of a turbofan such as increased number of blades
per row and a lower tip diameter than a turboprop. In
terms of bypass ratio, this places the unducted fan engine
somewhere between a conventional turboprop and turbofan
engine. For example, bypass ratios for the unducted fan
on the order of 35:1 but in the range of 10:1 to 60:1 may
be typical.
A complicating factor in the design of high
bypass ratio engines, and particularly the conventional
turboprop engines, is the phenomenon ~nown as lapse rate.
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Lapse rate refers to the decrease in net thrust which
occurs as the engine increases in Mach number and
altitude. Referring again to equation (2), it can be seen
that at take-off when VO is low that the net thrust will
be the product of mass flow Wa and Vj. As the
aircraft climbs in altitude, several things occur. First,
VO increases with the speed of the aircraft with a
smaller increase in Vj. Thus, the difference Vj -
VO decreases. Second, with increased air speed the ram
effect of air being pushed into the engine increases the
density and mass flow of the air. However, with increased
altitude the decrease in air density more than offsets the
ram effect. Thus, mass flow Wa decreases as the cruise
condition is reached.
High bypass ratio engines have a relatively large
lapse rate. This means that a conventional high bypass
ratio engine sized for take-off will be unable to achieve
the mach number and altitude of an equivalently sized
lower bypass ratio engine with less lapse. Although it is
~O possible within a certain range to design a high bypass
ratio engine which meets a given flight speed and
altitude, such an engine will be oversized for take-off.
In the past, turboprop engines accepted the large
lapse rate and settled for the lower speeds and altitudes
~5 which could efficiently be attained. The combination of
high propulsive e~ficiency, high bypass ratio, and a lapse
rate which allows an engine to be sized for take-off
conditions as well as meet thrust requirements for mach
numbers in excess of 0.6 and altitudes in excess of 30,000
feet is now possible in connection with the unducted fan
concept.
OBJ~CTS OF THE INVENTION
It is an object of the present invention to
provide a new and improved aircraft gas turbine engine.
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It is another object of the present invention to
provide a new and improved unducted fan engine capable of
achieving mach numbers in excess of 0.6 and altitudes in
excess of 30,000 feet.
It is yet another object of the present invention
to provide a relatively high bypass ratio engine with a
reduced lapse rate.
SUMMARY OF THE INVENTION
The present invention is an aircraft gas turbine
engine capable of cruise operation at altitudes in excess
of 30,000 feet and at mach numbers in excess of 0.6. The
engine comprises a core engine including a compressor
which is effective for producing a maximum pressure ratio
in excess of 40:1. The engine further comprises a
propulsor having an air bypass ratio between 10:1 and
60:1. The engine further comprises means for varying the
pressure ratio so that the increase in pressure ratio from
take-off to cruise is in excess of 20%.
BRIEF DESCRIPTION OF THE DRAWINGS
FIGURE 1 is a graph showing the lapse rate of
engines with different bypass ratios.
FIGURE 2 is a graph showing the effects of
turbine inlet temperature on net thrust.
FIGURE 3 is a schematic view of a gas turbine
engine according to the present invention.
DETAILED DESCRIPTION OF THE INVENTION
Figure 1 shows the lapse rate of engines with
different bypass ratios (~). Lapse rate is represented as
a plot of corrected thrust versus mach number. Corrected
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thrust is represented by the Fn/Po, where Fn = net
thrust and PO = ambient pressure. As can be seen, for
high bypass ratios, the lapse rate or rate of decrease of
corrected thrust for increasing mach numbers increases.
For B = 0 ~a pure turbojet) the lapse rate is relatively
small compared to B = 5 (a typical turbofan) and B = 100
(a typical turboprop). The present invention relates to
an unducted fan engine with B between 10:1 and 60:1 with a
preferred range of between 20:1 and 40:1 and a preferred
value within that range of approximately 35:1.
According to the present invention, the lapse
rate curve of a typical turbofan represents the ideal case
to match properly with efficient high speed aircraft
design thrust requirements; with Fto representing the
thrust required for take-off and FCr representing the
thrust required for cruise at M = 0.8. Point C shows
FCr for M = 0.8. It can be seen that a typical turbofan
with a bypass ratio of about 5 provides a lapse rate curve
such that Fto and FCr are in correct proportion to
each other. An unducted fan engine designed with a core
engine for providing Fto equivalent to that of the
turbofan would follow a lapse rate curve shown by dashed
line A in Figure 1. Properly sized for take-off, this
engine would be unable to meet the thrust requirements
FCr at M = 0.8 (shown by point B). In order to reduce
the lapse rate curve A so that point B corresponds with
point C, in accordance with this invention, it is proposed
to modify the core engine.
Figure 2 shows the effect of changes in the
corrected turbine inlet temperature on corrected net
thrust. Figure 3 is a schematic view of a gas turbine
engine according to one form of the present invention with
specific engine stations called out. For reference, core
engine 10 includes a compressor 12, combustor 1~, and
turbine 16 in serial flow relationship. Located aft of
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turbine 16 is a power turbine 18 which is effective for
turning unducted fan 20. A more detailed view of such an
engine is shown in Cana~ian application Serial
Number 438,676.
Station 2 is located just forward of compressor
12, station 3 is located at compressor discharge aft of
compressor 12, and station 41 is located aft of combustor
14 and forward of-turbine 16. For a fixed al~itude and
constant mach number, the temperature at station 2 or T2
is constant. T41 is a function of fuel flow to
combustor 14. As shown by Figure 2, an increase in T
for a fixed T2 will result in an increase in corrected
net thrust Fn/Po.
Ideally, in order to increase corrected net
thrust, T41/T2 will be increased. However, increases
in this ratio increase the speed of turbine 16 which
increases the speed (Nc) of compressor 12. Increases in
Nc result in an increase in the overall pressure ratio
(P3/P2) of compressor 12, within the limits of the
compressor design. Increasing Nc beyond the design
point will result in a choke condition with no further
increase in pressure ratio being realized. Pressure ratio
increases result in temperature increases (T3) at
compressor discharge. In the past, compressors have been
designed for maximum pressure ratios of perhaps 34:1.
According to the present invention, it is
proposed to provide a compressor with maximum pressure
ratio in excess of 40:1. Preferably, the compressor will
be operable over a range of between 6:1 to at least 45:1.
Take-of thrust will be achieved with a pressure ratio of
approximately 30:1 and high altitude thrust will be
obtained with an increase in overall pressure ratio to
about 40:1 or higher. In a preferred embodiment, the
increase in pressure ratio from take-off to cruise is in
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excess of 20%. Means for varying this pressure ratio are
further provided in the form of a fuel flow control, i.e.
varying T41/T2
In operation, the present engine will be capable
5 of achieving cruise operation at altitudes in excess of
30,000 feet and mach numbers in excess of 0.6. Its
propulsive efficiency (due to increased bypass ratio)
should exceed any prior art engine capable of attaining
these flight conditions. Furthermore, the engine will be
minimally sized for both take-off and cruise condition
and, therefore, be capable of attaining relatively low
specific fuel consumption, and low weight.
It will be clear to those skilled in the art that
numerous modifications, variations, and full and partial
equivalents can now be undertaken without departing from
the invention as limited only by the spirit and scope of
the appended claims.
