Note: Descriptions are shown in the official language in which they were submitted.
13DV-8754
COUNT~R ROTATION POWER rURBINE
FIF.LD OF THE INVENTION
This invention relates to gas turbine engines and,
more particularly, to a new and improved gas turbine engine
including means for efficiently transferring the energy of
combustion gases into a net engine thrust.
BACKGROUND OF THE INVENTION
_ _ _ _
While not limited thereto the present invention is
particularly applicable to gas turbine engines such as used for
the propulsion of aircraft.
Several types of gas turbine engines are currently
available for powering aircraft. The turbcfan and the
turboprop are two examples of such engines. The turbofan
engine includes a core engine, i.e., gas generator, for
generating combustion gases which are expanded through a power
turbine to driYe a fan, whereas the turboprop engine includes a
gas generator and power turbine which drives a propeller.
Conventional turboprop engines differ from turbofan
engines in several fundamental respects. For example,
turboprop engines typically have a much greater blade diameter
than turbofan engines. This allows the blades to move a
relatively large mass of air for producing thrust.
Furthermore, for a given energy input to the blades, a
relatively small velocity increase will be imparted to the air
~ 13DV-8754
passing therethrough. Small velocity increases translate to
high engine propulsive efficiencies. Simply stated, propulsive
efficiency is a measure of how much available energy is
converted to propulsive force. Large velocity increases to air
passing through propulsor blades result in "wasted" kine-tic
energy and lower propulsive efficiency.
Turbofan engines move a somewhat smaller mass of air
than do turboprops for the same energy input and impart a
larger velocity component to the air in order to achieve the
required thrust. This results in a lower propulsive
efficiency. rurbofan engines also include a nacelle radially
surrounding the fans~ This creates an additional drag on the
engine which degrades overall engine efficiency. However, the
nacelle defines an inlet which diffuses the airstream entering
the fan thereby slowing its speed. In this manner~ air enters
the fan with a relatively low axial velocity which is generally
independent of flight speed. Such low axial velocities
decrease blade drag losses thereby making higher cruise speeds
attainable.
Intermediate-sized transport aircraft, for example,
100 to 180 passenger transports, typically utilize turbofan
engines for propulsion. Turbofans provide the relatively high
thrust required for powering these aircraft at relatively high
alti~udes and at cruise speeds of about Mach 0.6 to about Mach
0.8. For aircraft designed for lower cruise speeds,
conventional turboprops are typically used inasmuch as they can
provide superior performance and efficiency. For example,
significant reductions in fuel burn, i.e. 3 the amount of fuel
consumed per passenger mile, are possible from the use of the
aerodynamically more efficient turboprop over the turbofan.
Accordingly, it would be desirable to combine the
advantages of the turbofan and the turboprop for obtaining a
compound engine having improved overall engine efficiency at
aircraft cruise speeds typical of turbofan powered aircraft.
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The overall efficiency of an aircraft gas turbine
engine is the product of thermal efficiency, transfer
efficiency9 and propulsive efficiency. Thermal efficiency is
related to -the core engine and is a measure of how effectively
the energy in the fuel is converted to available energy in the
core engine exhaust gases. Transfer efficiency is related to
the structural engine components excluding the core engine and
is a measure of how effectively core engine exhaust gas energy
is converted into kinetic energy imparted to the air stream.
Engine components which impact transfer efficiency include the
propulsor blades, gearbox, power turbine, and engine nacelle.
Accordingly, it is desirable to obtain a compound engine having
relatively high transfer and propulsive efficiencies at
relatively high subsonic Mach numbers.
A simple scaled up version of a conventional turboprop
engine suitable for powering an intermediate-sized transport
aircraft at the cruise speeds and altitudes typical of tur~ofan
powered aircraft would require a single propeller of about 16
feet in diameter. It would also require the capability of
generating about 15,000 shaft horsepower, which is several
times the power output of conventional turboprop engines.
A conventional turboprop engine built to these
requirements would further require the development of a ~
relatively large and undesirably heavy reduction gearbox for
transmitting the required power and torque at relatively low
speed to the propeller. Such gearboxes tend to introduce
losses which reduce the engine transfer efficiency. The
rotational speed of the large diameter propeller is a limiting
factor for keeping the helical velocity of the propeller tip,
i.e.~ aircraft velocity plus tangential velocity of the
propeller tip, below supersonic speeds. This is desirable
inasmuch as a propeller tip operating at supersonic speeds
generates a significant amount of undesirable noise and results
in a loss of aerodynamic efficiency.
13DV~~754
Gas turbine engines effective Eor driving propellers
or fans without the use of a reduction gearbox are kn~wn in the
prior art. They typically include relatively low speed,
counterrotating turbine rotors having relatively few blade row
stages driving a pair of counterrotating fans or propellers.
These engines comprise various embodiments that utilize the
fans or propellers for merely augmenting the thrust genera~ed
from the exhaust jet.
Such augmentation may be effective for some purposes.
~lowever, thrust augmentation requires that significant thrust
is being produced by the exhaust gases exiting the power
turbine and core nozzle. This reduces overall engine
efficiency by degrading propulsive ef~iciency.
For propelling a modern, intermediate-sized aircraft
that requires relatively large power output, a practical and
relatively fuel efficient new generation engine having
significant performance increases over conventional turbofan
and turboprop engines and these counterrotating turbine rotor
engines is required.
Accordingly, one object of the present invention is to
provide a new and improved gas turbine engine.
Another object of the present invention is to provide
a new and improved gas turbine engine for powering an aircraft
at cruise speeds in excess of Mach 0.6 and less than 1.0 with
improved overall engine efficiency.
Another object of the present invention is to provide
a new and improved gas turbine engine including a power turbine
having counterrotating rotors.
Another object of the present invention is to provide
a new and improved gas turbine engine including a power turbine
having a plurality of counterrotating turbine blade row stages
wherein substantially all output power is obtained from
expanding combustion gases through the stages and substantially
little power remains in the exhaust gases leaving the engine.
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Another object ot the present invention is to provide
a new and improved gas turbine engine wherein output power is
obtainable ~ithout the use of a reduction gearbox.
~ nother object of the present invention is to
pro~ide a new and ilrlproved gas turbine engine effective for
powering counterrotating airfoil members such as prop~llers.
M~ARY OF THE INVF.NTION
The prese~ invention comprises a new and improved gas
turbine engine comprising a gas generator effective for
generating combustion gases and means for efficiently
transferring the energy of the gases into a net engine thrust.
The means include a counterrotating power turbine with first
and second counterrotating propellers. The power turbine
includes a first rotor having a plurality of first turbine
blade rows extending radially outwardly therefrom and a second
rotor having a plurality of second turbine blade rows extending
radially inwardly therefrom. The first and second rotors are
arranged so as to define inner and outer flowpath surfaces,
respectively, for the combustion gases flowing through the
power turbine. The power turbine is effective for receiving
the combustion gases and extracting substantially all the
output power therefrom for driving the first and second rotors
in counterrotating directions.
The first and second counterrotating propellers each
have a plurality of blades attached to first and second
rotatable nacelle rings, respectively. The first and second
propellers are directly coupled to and driven by the first and
second rotors, respectively, and are disposed radially
outwardly of the power turbineO Each of the blades has a
relatively high hub radius to tip radius ratio and relatively
low thickness to chord ratio.
; ~a~ 13D~J-875
--6--
According to another form of the present invention,
the means include an annular casing disposed radially ou-twardly
of the gas generator and forming an outer contour. The contour
has forward, intermediate, and aEt portions. The forward
portion defines an inlet optimally designed for the gas
generator. The aft portior. defines an aerodynamically smooth
tr~nsition to the second rotatable nacelle rin~. The
intermediate portion defines the maximum radius of the casing
which exceeds the hub radius of each of the first and second
propellers.
Acco-rding to another form, the present invention is a
gas turbine engine comprising a gas generator effective for
generating combustion gases and means for efficiently
transferring the energy of the gases into a net engine thrust.
The means include a power turbine, first and second
counterrotating propellers, and an annular nacelle. The power
turbine includes a first rotor having a plurality of first
turbine blades extending radially outwardly therefrom and a
second rotor having a plurality of second turbine blades
extending radially inwardly therefrom. The first and second
rotors are arranged so as to define inner and outer flowpath
-surfaces, respectively, for the combustion gases flowing
through the power turbine. The power turbine is effective for
receiving the co~bustion gases and extracting substantially al1
output power therefrom for driving the first and second rotors
in counterrotating directions. The first and second-
counterrotating propellers each has a plurality of blades
attached to first and second rotatable nacelle rings at first
and second radii, respectively. The first and second
propellers are directly coupled to and driven by the first and
second rotors, respectively, and disposed radially outwardly of
the power turbine. ~ach of the blades has a relatively high
hub radius to tip radius ratio and relatively low thickness to
chord ratio. The annular nacelle is disposed radially
outwardly of the gas generator and forms an outer contour, the
~ }~ 13DV-8754
contour having forward, in-termediate, and aft portions. The
forward portion defines an inlet op-timally designed for the gas
generator. The aft portion forms an aerodynamically smooth
transition to the second rotatable nacelle ring. The
intermediate portion defines the maximum radius of the nacelle
which exceeds each of the first and second radii.
BRIEF DESGRIPTION OF THE VRAWINGS
The invention, together with further cbjects and
advantages thereof, is more particularly described in the
following detailed description taken in conjunction with the
accompanying drawings in which:
Figure 1 is a schematic representation of a gas
turbine engine according to one embodiment of the present
invention including a power turbine having counterrotating
rotors effective for driving counterrotating aft mounted
propellers.
Figure 2 illustrates an aircraft including two gas
turbine engines such as in Figure 1 mounted to an aft end
thereof.
Figure 3 is a view illustrating an alternative
arrangement for mounting a gas turbine engine such as
illustrated in Figure 1 to a wing of an aircraft.
Figure 4 is a view of a gas turbine engine according
to another form of the present invention~
Figure 5 is a more detailed view of the gas generator
of the engine shown in Figure 4.
Figure 6 is a more detailed view of the power turbine
of the engine shown in Figure 4.
Figure 7 is an enlarged view taken along the line 7-7
in Figure 4.
~ 13DV-~754
_TAILED DLSCRIPTION
Illus~r~ted in Figure 1 is a gas turbine engine 10, or
unducted fan engine, according to one embodimen~ of the present
invention. The engine 10 includes a longi~udinal centerline
axis 12 and an annular casing 14 disp~sed coaxially ab~ut the
axis 12. The engine 10 also includes a conventional gas
generator 16, which, for example, can comprise a booster
c~mpress~r 18, a compressor 20, a combustor 2Zs a high pressure
turbine (HPT) 24, and an intermediate pressure turbine (IPT) 26
all arranged coaxially about the longitudinal axis 12 of the
engine 10 in serial, axial flow relationship. A first annular
drive shaft 28 fixedly interconnects the compressor 20 and the
HPT 24. A second annular drive shaft 30 fixedly interconnects
the booster compressor 18 and the IPT 26.
In operation, the gas generator 16 is effective for
providing pressurized aîr frorn the booster 18 and the
compressor 20 to the combustor 22 where it is mixed with fuel
and suitably ignited for generating combustion gases. The
combustion gases drive ~he HPT 24 and the IPT 26 which in turn
drive the compressor 20 and the booster 18, respectively. The
combustion gases are discharged from the gas generator 16
through the IPT 26 at a mean discharge radius Rl from the
longitudinal axis 12.
Attached to an aftmost end of the casing 14 and ~ft of
the gas generator 16 is an annular support member 30. The
support member 30 extends radially inwardly and in an aft
direction from the aft end of the casing 14. The support
member 30 includes a plurality of circumferentially spaced
strut members 3Z extending radially inwardly from the aft end
of the casing 14 and an annular hub member 34 fixedly attached
to radially inller ends of the strut members 32 and extending in
an aft direction. The strut members 32 are effective for
supporting the hub member 34 and channeling combustion gases
from the gas generator 16 to a power turbine 36 constructed in
accordance with one embodimer,t of the present invention.
~ L~ 13D~-8754
g
The energy of the combustion gases discharged from the
gas generator will be eEficiently transferred into a net engine
thrust by means described ~lore fully hereafter. Such means
include the power turbine 36, or simply low pressure turbine
(LPT) 36, which is rotatably mounted to the hub member 34.
The LPT 36 includes a first annular drum rotor 38
rotatably mounted by suitable bearings 40 to the hub member 34
at forward and aft ends 42 and 44 thereof. The first rotor 38
includes a plurality of first turbine blade rows 46 extending
radially outwardly therefrom and spaced axially thereon.
The LPT 36 also includes a second annular drum rotor
48 disposed radially outwardly of the first rotor 38 and the
first blade rows 46. The second rotor 48 includes a plurality
of second turbine blade rows 50 extending radially inwardly
therefrom and spaced axially thereon. The second rotor 48 is
rotatably mounted to the hub member 34 by suitable bearings 52
disposed at radially inner ends of a forwardmost blade row 50a
of the second blade rows 50 and at radially inner ends of an
aftmost blade row 50b which is rotatably disposed on the first
rotor 38 mounted to the hub member 34.
As shown in Figure 1, an annular flowpath for
combustion gases flowing through blade rows 46 and 50 is
bounded by first drum rotor 38 and second drum rotor 48. In
addition to bounding the flowpath, first and second drum rotors
38 and 48 define inner and outer flowpath surfaces 38a and 48a,
respectively. In this manner, LPT 36 is lighter than typical
prior art turbines which include relatively large disks.
Each of the first and second turbine blade rows 46 and
50 comprises a plurality of circumferentially spaced turbine
blades, with the first blade rows 46 alternately spaced or
interspersed with respective ones of the second blade rows 50.
Combustion gases flowing through the blade rows 46 and 50 flow
along a mean flowpath radius R2 which, by definition~
represents a blade radius at~which resultant work loads of the
LPT 36 are assumed to be concentratedO ~or example, radius
R2 can be deined as the mean pitch line radius of all the
blade rows of the LPT 36.
A~J~ . o~r~
13DV-8754
-10-
Combustion gases being discilarged from the gas
generator 16 at the mean flowpa-th radius Rl are channeled
through the strut members 32 to the I,PT 36. The LPT 36 is
effective for expanding the combustion gases through the first
and second turbine blade rows 46 and 50 along the mean flowpath
radius R2 for extracting substantially all output powe~ from
the gases for driving the first and second rotors 38 and ~8 in
counterrotating directions at rotational speeds relatively
lower than ~hose of the first drive shaft 28.
The gas generator 16 and the LPT 36 as above arranged
and described results in a new and improved gas turbine engine
having counterrotating rotors effective for providing output
shaft power at relatively low rotational speedsO Significant
features of the present invention include the complimentary
arrangement of the engine elements. More specifically, the HPT
24 is disposed aft of the combustor 22 for first receiving the
relatively high pressure combustion gases being discharged
therefrom. The HPT 24 is most efficient when it and the first
drive shaft 28 are designed to rotate at about 10,000 to 15,000
RPM in a 15,000 shaft horsepower engine. This rotational speed
efficiently utilizes the high pressure combustion gases from
the combustor 22.
The combustion gases after passing through the HPT 24
are at a reduced, intermediate pressure. The intermediate
pressure gases then flow through the IPT 26 which further
reduces the pressure of the gases to a relatively low pressure
while most efficiently extracting power for rotating the second
drive shaft 30 and the booster compressor 18 at speeds
relatively lower than those of the HPT 24.
Finally, the low pressure combustion gases are
channeled to the LPT 36 where they are further expanded and
substantially all of the remaining energy thereof is extracted
for rotating the first and second rotors 38 and 48 for
providing output shaft power. Little energy remains for the
generally inefficient thrust produced by the relatively high
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velocity gases in the exhaust jet discharged from the LPT 36.
Furthermore, inasmuch as the LPT 36 is the last element in the
engine 10, it is subject -to the lowest temperature combustion
gases and therefore, thermally induced stresses are reduced.
For more efficiently extracting energy from the
combustion gases in the LPT 36 it is preferable that the mean
flowpath radius R2 thereof be greater than the mean discharge
radius Rl of the gas generator 16. In the embodiment
illustrated in Figure 19 the mean flowpath radius R2 is about
double the magnitude of the mean discharge radius Rl. This
arrangement is effective for placing the turbine blade rows 46
and 50 at an increased radius from the longitudinal axis 12 for
increasing the rela~ive tangential velocities thereof for
reducing blade loading thereby efficiently extracting power
from the gases flowing thereover.
In the exemplary embodiment shown in Figure 1, the LPT
36 is effective for driving counterrotating, oppositely pitched
forward propellers 5~ and aft propellers 56. More
specifically, extending from an aftmost end of the first rotor
38 is an aft blade row 46a which extends radially outwardly to
about the radial position of the second rotor 48. Attached to
radially outer ends of the aft blade row 46a is an annular
shroud member 58 including an aft rotatable nacelle ring 128
adapted for the smooth flow of air thereover. The aft
propellers 56 are suitably attached to the shroud member 58.
Similarly, the forward propellers 54 are suitably attached to
an annular shroud member with forward rotatable nacelle ring
126 which is attached to a forward end of the second rotor 48.
Suitable pitch varying means 60 are provided for independently
controlling the pitch of the forward and aft propellers 54 and
56. Each annular nacelle ring which surrounds the power
turbine and the plurality of propeller blades mounted on the
ring form a propeller system.
13DV-8754
A most significant feature of the present invention is
a gas turbine engine 10 including an LPT 36 effective for
providing relatively high output power and torque at relatively
low rotational speeds without the use oE a reduction gearbox.
A reduction gearbox, and related accessories, would add a
significant amount of weight and complexity to an engine
capable of generating the relatively large thrust required for
powering a transport aircraft such as the 150 passenger
transport. Moreover, any losses attributable to the gearbox
reduce the transfer efficiency.
Speed reduction is required where a gas turbine engine
is used for driving airfoil members such as propellers or
fans. A conventional low pressure turbine (not shown) includes
a single rotor typically rotating at about 10,000 to 15,000
RPM. These rotational speeds must be reduced to relatively low
speeds of about 1,000 to about 2,000 RPM for driving airfoil
members. Propellers and fans are designed for moving a
relatively large amount of air at relatively low axial speeds
for generating thrust, and operate more efficiently at the
relatively low rotational speeds. Additionally, the low
rotational speeds are required for limiting the helical tip
speed of the propellers to below supersonic speeds.
According to the present invention, by allowing the
second rotor 48 in Figure 1 of the LPT 36 to rotate in a
direction opposite the first rotor 38, two output shafts, first
rotor 38 and second rotor 48, are provided which rotate at
about one quarter the speed of a single rotor, conventional LPT
of an equivalent output power, thereby providing speed
reduction.
Furthermore, additional speed reduction is obtainable
by increasing the number of the first and second ~urbine blade
rows 46 and 50, i.e., the number of stages. By increasing the
number of blade rows, the amount of energy extracted per stage
is reduced. This allows for a reduction of the speed of the
rotor and the aerodynamic loading of the blades on each row.
13DV-8754
-13-
Thus, in order to ob-tain the desired reduced speeds and
efficiently extract (by reduced blade loading) substantially
all remaining power from the combustion gases, an increased
number of stages would be required.
~ lowever, a fewer number of stages could be used for
accomplishing these objec~ives by having increased values of
the ratio R2/Rl or providing the combustion gases to the LPT
36 at a larger mean flowpath radius R2. Too many stages are
undesirable because of the increased complexity, size and
weight therefrom, and an LPT 36 having fewer stages and a
relatively high R2/Rl ratio is undesirable because of the
increased frontal area and weight attributable thereto. As
above-described and in accordance with the present invention,
it has been determined that an R2/Rl ratio of about 2~0 is
preferable.
Furthermore, in the embodiment illustrated in Figure 1
for driving the counterrotating propellers 54 and 56, the LPT
36 having about 14 stages is preferred for obtaining output
shaft speeds of the first and second rotors 38 and 48 of about
1200 RPM. This speed is much less than the rotational speeds
of the first and second drive shafts 28 and 30. Moreover, and
in accordance with the present invention, LPT 36 has a total
number of rows of blades effective to maintain the tip speeds
of the propeller blades below sonic velocity.
The reduction in speed of the rotors 38 and 48 of the
LPT 36 results in a second order reduction of centrifugally
generated stresses. For example, a one quarter reduction in
speed results in a seven-sixteenths reduction in centrifugal
stress. This is significant in that the LPT 36 requires less
material for accommodating centrifugal stress ~hich results in
a lighter LPT 36. For example, use of drum rotors 38 and 48
rather than disks significantly reduces weight. The overall
effect of using a counterrotating LPT 36 is a significant
reduction in engine weight as compared to an engine including a
conventional LPT and reduction gearbox.
13DV-~754
-14-
~ leans for improving transfer e~ficiency may also
include a seal 53 which is disposed between casing 14 and
second drum rotor 48. By this arrangement 3 the leakage or flow
of combustion gases between stationary casing 14 and rotor 48
will be reduced. This arrangelnent provides a single seal in
the relatively high pressure region of the flowpath approximate
to strut members 32 and forward of -the LPT 36. No other
relatively high di~meter leakage areas exist until just aft of
the aftmost blade row 50b. At such aft location, the pressure
of the combustion gases is greatly reduced, and; thus, any
leakage in this region will be small relative to leakage
locations further upstream.
Means for improving transfer efficiency further
include counterrotating propellers 54 and 56, aft mounted to
the engine 10 radially outwardly of both the first rotor 38 and
the second rotor 48. These propellers have a hub radius R3 and
a tip radius R4 from the longitudinal axis 12. What is meant
by "hub radius" is the distance measured from engine centerline
12 to the outer surface of the rotatable nacelle ring from
which each propeller blade extends. In a like manner, "tip
radius" is the distance measured from engine centerline 12 to
the r~dially outer end of each propeller blade. Mounting the
propellers 54 and 56 radially outwardly of the second rotor 48
increases the hub to tip ratio R3/R4 of the propellers to a
relatively high value when compared to conventional gear driven
propellers which typically have a small hub radius and thus
relatively low hub to tip ratio, This arrangement provides an
improvement in aerodynamic performance. For example, hub
radius to tip radius ratio is greater than about 0.4 and
between about 0.5 to Q.4 in a preferred embodiment.
Furthermore, the propellers do not obstruct the flow of
combustion gases discharged from the LPT 36, which would
otherwise reduce engine performance and require cooling schemes
for preventing thermal damagè to the propellers 54 and 56.
13DV-~754
~ ther features of the blades of propellers 54 and 56
are best show~ in Figures 4 and 7. Each blade is swept back
toward the tip. Such sweep reduces the relative Mach number of
the tip which reduces losses at cruise Mach numbers in excess
of 0.6. Each blade is Eurther provided with a twist ~rom root
to tip to pro~ide proper chord orientation for increased blade
speed with increase in radius. ~ach blade has relatively low
thickness (T) to chord (C) ratio, as shown by the blade section
in Pigure 7. For example, T/C is less than .14 at the blade
hub and is about .02 at the tip.
The use of two propellers over a single propeller
allows for propellers of lesser diameter. For example, at
aircraft cruise speeds of about Mach 0.7 to about Mach 0.8, two
propellers with diameter of about 12 feet and rotational speed
of about 1200 RPM will generate an equivalent amount of thrust
to a single propeller of about 16 feet at a rotational speed of
about 900 RYM. The reduced diameter results in reduced
propeller tip speeds and noise therefrom.
In the embodiment of engine 10 having a power turbine
with about 14 stages, it is also preferred that Rl/R49 R2/R4,
and R3/R4, equal about 0.18, 0.3S, and 0.45, respectively.
However, the number of stages of the LPT 36 can range between
about 10 and about 18 stages, and Rl/R4, R2/R4, and R3/R4 can
range between about 0.2 to 0.16, 0.4 to 0.3, and 0.5 to 0.4,
all respectively. These relationships are preferred for
obtaining an engine 10 suitable for most efficiently driving
the counterrotating propellers 54 and 56 at rotational speeds
of about 1200 RPM.
The embodiment of the engine 10 illustrated in Figure
1 results in additional advantages. For example, by mounting
the propellers 54 and 56 to the aft end of the engine 10, an
annular inlet region 62 of the engine 10 is relatively free of
flow disturbing obstructions. Accordingly, the inlet region 62
and an annular nacelle 64 surrounding the engine 10 can be
suitably designed for obtaining increased aerodynamic
performance of air entering the engine 10 as well as flowing
thereover.
13DV-8754
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Annular nacelle 64 contributes to the transfer
eficiency of engine 10. Nacelle 64 forms an outer contour
which includes forward, aft, and intermediate portions 120,
122, and 124, respectively. The outer contour is the only
surface defining the flowpath of air to propellers 54 and 56.
Forward portion 120 defines an inlet for inlet region 62
optimally designed for gas generator 16 without concern for
flow disturbing obstructions. Aft portion 124 for~ns an
aerodynamically smooth transition to forward rotatable nacelle
ring 126. Intermediate portion lZ2 defines the maximum radius
R5 of casing which is greater than the hub radius R3 of
propeller 54 (R3 also being the radius of forward rotatable
nacelle ring 126). With R5 greater than R3, flow over nacelle
64 will diffuse as it passes intermediate portion 1~2 thereby
reducing thc velocity of air near the hub of propeller 54.
This reduces losses and improves the efficiency of the
propeller.
Illustrated in Figure 2 is an aircraft 66 including
two engines 10 driving counterrotating propellers, such as the
one illustrated in Figure 1, mounted to an aftmost end of the
aircraft 66. Aft mounted counterrotating propeller engines 10
according to the present invention are effective for providing
an aircraft 66 having improved performance and fuel ~urn.
Furthermore, the engines 10 have reduced weight when compared
with a conventional turboprop engine sized for identical thrust
output. Reduced propeller noise is realizable which allows for
a reduction in the amount of noise attenuation modifications to
the aircraft, and thus additionally reduces total aircraft
weight.
Illustrated in Figure 3 is an alternative arrangement
for mounting counterrotating propeller engines 10, such as the
one illustrated in Figure 1, to a wing 68 of an aircraft (not
shown). In this embodiment, the hub member 34 of the engine 10
is extended in an aft direction and suitably mounted to the
wing 68. A stationary, annular exhaust duct 70 is suitably
13DV-8754
-17-
secured to the hub member 34 for suitably channeling the
exhaust gases of the engine 10, for example, under the wing
68. The embodiment of the engine 10 illustrated in Figure 3
clearly illustra-tes a significant advantage of the support
member 30 of the engine 10. More specifically, the support
member 30 is not only effective for mounting the LPT 36 in the
engine 10 but is also effective for mounting the entire engine
10 to a wing 68 of an ~ircraft.
Illustrated in Figures 4-7 is a more detailed cross
sectional view of an actual gas turbine engine 10 according to
a preferred embodiment of the present invention. Engine 10
comprises a gas generator 16 for generating combustion gases.
Detail of gas generator 16 is shown in Figure 5 with like
numbers for similar components carried forward from Figure 1.
Engine 10 further comprises means for efficiently
transferring the energy of the combustion gases into a net
engine thrust which includes LPT 36, forward and aft C/R
propellers 54 and 56, respectively, and annular casing 64.
Power turbine or LPT 36 is shown in greater detail in
Figure 6 with like numbers for similar components carried
forward from Figure 1. Although basically the same as LPT 36
shown in Figure 1, the Figure 6 LPT 36 includes several
different features. These include a plurality of inlet guide
vanes 49a located axially forward of first and second blade
rows 46 and 50. Similarly, outlet guide vanes 49b are located
axially aft of blade rows 46 and 50. Inlet guide vanes 49a are
effective for imparting a circumferential swirl to the
combustion gases whereas outlet guide vanes 49b are effective
for removing substantially all circumferential swirl from gases
passing therethrough. In this manner, more work may
efficiently be extracted from the forward and aftmost blade
rows of LPT 36 thereby improving its efficiency.
' The blades of aft and forward counterrotating
propellers 56 and 54 are attached to first and second rotatable
nacelle rings 128 and 1269 at first and second radii R6 and R7,
13DV-8754
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respectively. Radii R6 and R7 correspond to the hub radii of
propellers 56 and 54, respectivelyO Aft propeller 56 is
direc~ly coupled to and driven by first rotor 38 and forward
propeller 54 is directly co~lpled to and driven by second rotor
48. Annular nacelle rings 126 and 128 form the only surfaces
controlling the airflow in the region of the propeller blades.
Counterrotating propellers 54 and 56 are disposed
radially outwardly of LPT 36. In a preferred embodiment, each
of forward propeller 54 and aft propeller 56 is axially
positioned between forward and aft ends Gf LPT 36. In this
manner, improved dynamic stability of the engine is achieved.
Located forward of LPT 36 are a plurality of strut
members 32 which extend radially inwardly through the flowpath
and are fixedly attached at their radially inner ends to
annular hub member 34. In this manner, strut members 32 are
effective for both supporting hub member 34 and channeling
combustion gases from the gas generator to LPT 36.
First annular drum rotor 38 includes radially inwardly
extending carrier members 130, 132, and 134. Each of carrier
members 130, 132, and 134 is generally conically-shaped with
the radially inner ends o members 130 and 132 being connected
by a generally cylindrical carrier member 136. Rotor 38 is
rotatably mounted to hub member 34 by roller bearing 138 and
thrust bearing 139. Roller bearing 138 is located generally in
the forward portion of LPT 36 at the conjunction of carrier
members 130 and 135. Thrust bearing 139 is located in the
generally aft portion of LPT 36 and at the radially inner end
of carrier member 134. Hub member 34 is provided with a
generally cylindrical forward hub member portion 34a and a
generally cylindrical aft hub member portion 34b extending
radially from hub member 34 proximate bearings 138 and 139,
respectively. In this manner, hub member 34 provides improved
support for rotor 38.
2~
- 19 - 13~V-8754
Second rotor 48 includes generally conical carrier
members 140 and 142. ~otor 48 is supported to carrier
member 136 of rotor 38 by differential thrust bearing 144
and differential roller bearing 146. Differential thrust
bearing 14~ is located at the radially inner end of
carrier member 140 and differential roller bearing 146 is
located at the radially inner end of carrier member 142.
In operation, rotor 3~ will rotate about annular
hub member 34 in a first direction. At the same time,
rotor 48 will rotate in a seGond direction opposite to
that of the first direction. By the use of differentia~
bearings 144 and 146, rotor 48 is maintained in spaced
axial and radial relationship with rotor 38 while
simultaneously being counterrotatable therewith.
Figure 6 further discloses a pitch change
mechanism 150.
While there have been described herein what are
considered to be preferred embodiments of the present
invention, other embodiments will occur to those skilled
in the art from the teachings herein.
For example, the gas generator 16 of Figure l
without a booster compressor 18 and IPT 26 can also be
used for generating combustion gases~ Furthermore,
inasmuch as the counterrotating LPT 36 is effective for
providing relatively large output power and torque at low
speeds, gas turbine engines incorporating such LPTs can be
used for powering ships, generators, and large pumps, for
example, which can be designed for having counterrotating
input shafts suitably attached to the first and second
rotors 38 and 48 of the LPT 36.
Furthermore, although the invention has been
described as applied to a 15,000 shaft horsepower
engine, it can also be sized for other engine
classes. For example, in a smaller, 1500 shaft
horsepower engine, powering shorter propellers 54
ji,~
,.
~,~4~
13DV-8754
-20-
and 56, the HPT 24 would be designed to operate at about 30,000
RPM. The first rotor 38 and the second rotor 48 of the LPT 36
of Figure 1 would be correspondingly designed to operate at
about a 10 to l speed reduction, i.e., at about 3,000 RPM. The
propellers 54 and 56, although operating at about 3,000 RPM,
have reduced tip radii ~4 and therefore the helical tip
speeds can be main~ained below supersonic speeds.
