U.S. patent application number 11/682077 was filed with the patent office on 2007-08-30 for nested core gas turbine engine.
Invention is credited to Sudarshan Paul Dev.
Application Number | 20070201974 11/682077 |
Document ID | / |
Family ID | 22866973 |
Filed Date | 2007-08-30 |
United States Patent
Application |
20070201974 |
Kind Code |
A1 |
Dev; Sudarshan Paul |
August 30, 2007 |
NESTED CORE GAS TURBINE ENGINE
Abstract
A fan for creating lift or thrust having a fan hub and fan
blades depending from the fan hub. The fan blades have slots formed
therein with openings facing substantially aft, relative to a
rotation direction of the fan blades, wherein air blowing from the
fan hub into the fan blades and out of the openings of the slots
contribute to the aerodynamic performance of the fan blades to
enhance the aerodynamic performance of the fan blades.
Inventors: |
Dev; Sudarshan Paul;
(Seymour, CT) |
Correspondence
Address: |
PERMAN & GREEN
425 POST ROAD
FAIRFIELD
CT
06824
US
|
Family ID: |
22866973 |
Appl. No.: |
11/682077 |
Filed: |
March 5, 2007 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
11201441 |
Aug 10, 2005 |
7219490 |
|
|
11682077 |
Mar 5, 2007 |
|
|
|
10635956 |
Aug 7, 2003 |
6988357 |
|
|
11201441 |
Aug 10, 2005 |
|
|
|
09947002 |
Sep 5, 2001 |
6647707 |
|
|
10635956 |
Aug 7, 2003 |
|
|
|
60230891 |
Sep 5, 2000 |
|
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Current U.S.
Class: |
415/77 |
Current CPC
Class: |
F05D 2220/328 20130101;
F02C 3/16 20130101; F02C 3/064 20130101; F05D 2220/80 20130101;
Y02T 50/672 20130101; F02C 3/145 20130101; F02C 3/067 20130101;
F05D 2220/323 20130101; F02C 3/045 20130101; F02K 3/068 20130101;
F05D 2220/325 20130101; F05D 2220/327 20130101; Y02T 50/60
20130101; F02C 3/10 20130101; F02C 7/10 20130101; F05D 2220/90
20130101; F05D 2220/324 20130101; F05D 2220/326 20130101; F02K 3/00
20130101; F23R 3/38 20130101 |
Class at
Publication: |
415/077 |
International
Class: |
F01D 1/04 20060101
F01D001/04 |
Claims
1. A fan for creating lift or thrust comprising a fan hub and fan
blades depending from the fan hub, the fan blades having slots
formed therein with openings facing substantially aft, relative to
a rotation direction of the fan blades, wherein air blowing from
the fan hub into the fan blades and out of the openings of the
slots contribute to the aerodynamic performance of the fan blades
to enhance the aerodynamic performance of the fan blades.
2. The fan according to claim 1, wherein the air blowing out of the
openings of the slots is at least in part exhaust from an
engine.
3. The fan according to claim 2, wherein the air blowing out of the
openings of the slots is at least in part exhaust from a gas
turbine engine.
4. The fan according to claim 3, wherein the openings of the slots
are located on an upper surface of the fan blades.
Description
RELATED APPLICATION(S)
[0001] This application is a continuation from co-pending
application Ser. No. 11/201,441, filed Aug. 10, 2005 which is a
continuation from application Ser. No. 10/635,956 filed Aug. 7,
2003, now issued Pat. No. 6,988,357, which is a continuation from
application Ser. No. 09/947,002, filed Sep. 5, 2001, now issued
Pat. No. 6,647,707, which claims the benefit of U.S. Provisional
Application No. 60/230,891, filed Sep. 5, 2000, which is
incorporated by reference herein in its entirety.
BACKGROUND
[0002] 1. Field
[0003] The disclosed embodiments relate to gas turbine engines and
fans.
[0004] 2. Previous Developments
[0005] At large power levels (thousands of horsepower), turbine
engines are the most compact and lightest power systems available,
and have completely taken over the market for large aircraft.
However, scaled-down versions of these conventional gas turbine
engines offer relatively poor power/weight ratio and high specific
fuel consumption. FIG. 2A illustrates that for small engines
power/weight ratios versus rated power is low. FIG. 2B illustrates
that for small engines specific fuel consumption versus rated power
is high. There are a variety of reasons why small gas turbine
engines do not perform as well as the larger engines. These
include:
[0006] REYNOLDS NUMBER EFFECTS: Due to smaller characteristic
dimensions, small compressor and turbine blades suffer from larger
friction coefficients and greater aerodynamic losses.
[0007] THICKNESS/CHORD RATIOS: Due to physical difficulties in
manufacturing very thin blades with adequate strength, airfoils
used in small engines typically have larger thickness/chord ratios
and relatively blunt leading edges. This causes larger aerodynamic
losses due to profile drag and wave drag.
[0008] LARGE RELATIVE TIP CLEARANCES: Due to differences in
relative centrifugal and thermal growths of rotors and shrouds, and
the effects of scaling, small compressors and turbines suffer from
larger relative tip clearances (ratios of absolute tip clearance to
blade span). This in turn causes large tip leakage losses and lower
component performance and efficiency.
[0009] LOWER CYCLE PRESSURE RATIOS: Small turbine engines with high
cycle pressures need extremely small blade heights. These are
difficult to manufacture, have large tip clearance losses, and
suffer from boundary layers occupying a large fraction of passage
heights. Consequently, small engines are limited to low pressure
ratios, resulting in lower specific power (per unit mass flow), and
low cycle efficiency.
[0010] LOWER PEAK CYCLE TEMPERATURES: Large gas turbine engines can
have intricate cooling passages in their large nozzle vanes and
turbine blades. These convection, film and transpiration cooling
schemes allow gas temperatures significantly higher than the
structural capability of conventional turbine materials (metals),
for high specific power and cycle efficiency. Similar cooling
schemes are too complex and expensive for the small blade sizes of
small gas turbines, causing them to be limited to lower
temperatures and lower performance levels.
[0011] The exemplary embodiments addresses these problems, and
mitigates at least some of them, for improved power density and
efficiency as will be described in greater detail below.
Accordingly, amongst the objects of the exemplary embodiments is to
provide a lightweight/high-power density engine, having the ability
to use military-standard high-energy fuels, such as JP8 or JP5. The
engine has low observables including noise, smoke and infra-red
signatures, and adequate life, to enable its use for reusable air
vehicles, and affordable cost, to enable its use for
expendable/attritable air vehicles. The engine can also be scaled
up, as well as down, offering higher power/weight compared to
current gas turbine engines of conventional design.
SUMMARY OF THE EXEMPLARY EMBODIMENTS
[0012] In accordance with one exemplary embodiment a fan for
creating lift or thrust is provided. The fan has a fan hub and fan
blades depending from the fan hub. The fan blades have slots formed
therein with openings facing substantially aft, relative to a
rotation direction of the fan blades, wherein air blowing from the
fan hub into the fan blades and out of the openings of the slots
contribute to the aerodynamic performance of the fan blades to
enhance the aerodynamic performance of the fan blades.
BRIEF DESCRIPTION OF THE DRAWINGS
[0013] The foregoing aspects and other features of the exemplary
embodiments are explained in the following description, taken in
connection with the accompanying drawings, wherein:
[0014] FIGS. 1-1A respectively are a schematic cross-sectional view
and a schematic perspective cut-away view of a gas turbine engine
incorporating features in accordance with a first embodiment;
[0015] FIG. 1B is a perspective view of the gas turbine engine in
FIG. 1;
[0016] FIG. 1C is a perspective view of the front section of an
outer casing of the turbine engine in FIG. 1;
[0017] FIG. 1D is a perspective view of a front rotor of the
turbine engine in FIG. 1;
[0018] FIG. 1E is a perspective view of a stator section of the
turbine engine in FIG. 1;
[0019] FIG. 1F is a perspective view of a rear rotor of the turbine
engine in FIG. 1;
[0020] FIG. 1G is a perspective cut-away view of a rear end portion
of the turbine engine in FIG. 1;
[0021] FIGS. 2A-2B are graphs respectively illustrating
power/weight ratios versus rated power, and specific fuel
consumption (SFC) versus rated power for small engines of the prior
art;
[0022] FIG. 3 is a cross-sectional view of a gas turbine engine in
accordance with a second embodiment;
[0023] FIG. 4 is a graph showing variation of ignition delay time
at a number of air temperatures with respect to pressure in
accordance with the prior art;
[0024] FIGS. 5 and 6 are respectively schematic cross-sectional
views of a conventional engine with centrifugal compressors and
wrap-around burners, and a conventional engine with axial
compressors and in-line burners;
[0025] FIGS. 7-10 respectively are schematic cross-sectional views
of a turbo-jet engine, turbo-fan engine, high-bypass ration
turbo-fan engine, and ultra-high bypass ratio turbo-fan engine in
accordance with other embodiments;
[0026] FIGS. 11-12 respectively are schematic cross-sectional views
of the propulsion systems of high speed air vehicles in accordance
with still other embodiments;
[0027] FIGS. 13 and 14-14A respectively are schematic top plan,
elevation, and bottom plan views of an unmanned aerial vehicle
(UAV) in accordance with yet another embodiment;
[0028] FIGS. 14B-14C respectively are schematic side elevation and
rear elevation views of the UAV in FIG. 13 in a first mode of
operation (e.g. cruise mode), and FIGS. 15A-15B respectively are
schematic side elevation and rear elevation views of the UAV in
FIG. 13 in a second mode of operation (e.g. hover mode);
[0029] FIGS. 16-17 are graphs respectively illustrating the
relationship of thrust to engine diameter and engine frontal area
for field engines of the prior art and gas turbine (nested core)
engines according to the exemplary embodiments;
[0030] FIG. 18-19 are graphs respectively illustrating SFC at rated
thrust versus operating pressure ration (OPR), and thrust versus
OPR for field engines of the prior art and gas turbine engines of
the exemplary embodiments;
[0031] FIGS. 20-21 are graphs respectively illustrating SFC at
rated thrust versus rated normal thrust, and length/diameter ratio
versus engine diameter for field engines of the prior art and gas
turbine engines of the exemplary embodiments;
[0032] FIGS. 22-23 are graphs respectively illustrating thrust
versus engine volume and bulk density (engine weight/cylindrical
volume) versus engine diameter for field engines of the prior art
and gas turbine engines of the exemplary embodiments;
[0033] FIGS. 24-25 are graphs respectively illustrating thrust
versus weight, and thrust/weight versus thrust for field engines of
the prior art and gas turbine engines of the exemplary
embodiments;
[0034] FIG. 26 is a schematic cross-sectional view of a gas turbine
engine in accordance with another embodiment, particularly useful
for a larger (scaled-up) engine;
[0035] FIG. 27 is a schematic cross-sectional view of a gas turbine
engine in accordance with yet another embodiment, also particularly
useful for a larger (scaled-up) engine;
[0036] FIG. 28 is a schematic cross-sectional view of a gas turbine
engine in accordance with still another embodiment, also
particularly useful for a larger (scaled-up) engine;
[0037] FIGS. 29-29A are a schematic cross-sectional views of a gas
turbine engine in accordance with still other embodiments;
[0038] FIGS. 30A-30D are respectively schematic front elevation,
plan, rear elevation and side elevation views of a high speed air
vehicle embodiment according to the exemplary embodiments;
[0039] FIGS. 31A-31D are respectively schematic front elevation,
plan, rear elevation and side elevation views of the high speed air
vehicle in FIG. 30A; and
[0040] FIGS. 32A-32D are respectively schematic front elevation,
plan, rear elevation and side elevation views of another high speed
air vehicle embodiment according to the exemplary embodiments.
DETAILED DESCRIPTION OF THE EXEMPLARY EMBODIMENTS
[0041] Referring to FIG. 1-1A, there are shown respectively a
cross-sectional view and cutaway perspective view of a gas turbine
engine 10 incorporating features of the exemplary embodiments.
Although the exemplary embodiments will be described with reference
to the single embodiment shown in the drawings, it should be
understood that the exemplary embodiments can be embodied in many
alternate forms of embodiments. In addition, any suitable size,
shape or type of elements or materials may be used.
[0042] Referring still to FIGS. 1-1A, the gas turbine engine 10 is
a nested core gas turbine engine projected to produce about 20 lbf
of thrust from a 2.75'' diameter, 2.75'' long package. Some
alternate embodiments comprising features of the exemplary
embodiments are shown in the engines in FIGS. 7-10, including a
turbojet, a low-bypass turbofan, a high-bypass turbofan, and an
ultra-high bypass (UHB) lift rotor which will be described in
greater detail below.
[0043] The engine 10 shown in FIGS. 1-1A, generally has a
compressor section 12, a turbine section 16, and a combustion
chamber section or area 52. The compressor section 12 may include a
first compressor stage 13, and a second compressor stage 14. The
turbine section 16 may include first and second turbine stages 17,
18. As seen in FIGS. 1-1A, the nested core engine 10 preferably
comprises an outer casing 24, a front rotor 32, a rear rotor 42,
and a precombustor 50. The front rotor 32 and rear rotor 42 are
rotatably mounted in series inside the outer casing 24. The
precombustor 50 is located in the casing 24 to the rear of the rear
rotor 42. The front rotor 32 defines at least in part the first
stage 13 of the compressor section 12 as well as the first stage 17
of the turbine section 16. The rear rotor 42 defines at least in
part the second stage 14 of the compressor section 12 as well as
the second stage 18 of the turbine section 16. The front and rear
rotor 32, 42, when mounted in the casing 24, also form the
combustion chamber section 52 of the engine 10.
[0044] As can be seen from FIGS. 1-1A, the turbine section 16, is
substantially nested within or surrounded by the compressor section
12 of the engine 10. In addition, the combustion chamber section 52
of the engine is nested within or substantially surrounded by the
turbine section 16 of the engine 10. This concentric or coaxial and
nested configuration reduces the size of the engine relative to
conventional engines and can, therefore, allow for a reduction in
weight.
[0045] For the embodiment shown in FIGS. 1-1A, the engine 10 may be
connected to a fuel source JP which supplies fuel such as JP-5 or
JP-8 to an inlet 22 of the compressor section 12. For alternate
embodiments, the fuel supply may be connected to any other suitable
location in the engine. FIGS. 26-29 show some alternate locations
for fuel injection; additional other locations may be used. The
fuel mixed with inlet air is drawn through two stages 13, 14 of the
compressor section 12 into the precombustor 50. The fuel and the
compressed air from the compressor section is ignited in the
precombustor 50. The fuel air mixture exits the precombustor 50
into the combustion chamber section 52 of the engine 10 where
combustion is completed. The heated air in the combustion chamber
section 52 then exits from the combustion chamber section into the
surrounding turbine section 16. The heated air passes through both
stages 17, 18 of the turbine section and then exhausts out of the
casing. In the embodiment shown in FIGS. 1-1A, the engine 10 may be
provided with a starter or igniter such as for example a cartridge
igniter 56 used to initiate gas and air flows through the engine as
well as initiate combustion in the combustion chamber section 52
and the precombustor 50 of the engine. In the embodiment, the
cartridge casing burns away after the solid fuel in the cartridge
burns out. In alternate embodiments, the engine may use any
suitable starting system. For example, air and/or hot gases may be
introduced into the engine at the location where the cartridge
starter is shown to be located.
[0046] In greater detail now, and with reference also to FIG. 1B-1C
, and 1E, the outer casing 24 is preferably made up of several
sections to facilitate fabrication, though in alternate embodiments
the casing may be a one piece member formed by spin casting, for
example. The casing 24 may include a front, or inlet section 24I, a
first stator section 24C, a second stator section 24D, and a rear
or exhaust section 24R. As seen in best in FIG. 1B, the exterior of
the casing 24 is substantially cylindrical, though the casing
exterior may have any other desired shape. The inlet section 24I of
the casing 24, a perspective view of which is shown in FIG. 1-C,
defines the inlet 22 of the gas turbine engine 10. The inlet 22 is
sized to provide suitable mass flow to the engine throughout its
operating range without choking. Inlet area is dependent on engine
thrust. FIGS. 16-17 are graphs which relate thrust to engine
diameter, and engine frontal area. As can be realized from FIGS.
16-17, the frontal area (i.e. inlet size) for a given engine thrust
is less for the nested core engine 10 than for conventional small
engines and comparable for large engines. This is especially the
case for engines with an engine diameter under 1000 mm (i.e. under
40 inch diameter). The lip 70 of the inlet section 24I has a
suitable shape to maintain substantially undisrupted flow across
the entire inlet opening 22 throughout the engine operating range.
In the embodiment, the inlet section 24I has a center member or
nose cone 26 which is supported from the outer portion of the inlet
section by struts or vanes 20. Struts 20, which may be made of
metal or ceramic material, are equally distributed around the inlet
22, and are provided in sufficient number to provide effective
inlet protection to foreign object damage (FOD) without disrupting
air flow into the inlet throughout the operating range of the
engine. The FOD struts 20 may be keyed at opposite ends of the
inlet section and nose cone 26 or may be attached by any other
suitable means to the inlet section of the outer casing.
[0047] As seen in FIGS. 1 and 1A, the nose cone 26 is substantially
a one piece member which houses a rolling element bearing 30. The
nose cone 26 has a fuel port or fuel entry 28 formed at the front
of the member, to which a fuel supply line JP may be connected. For
example, the fuel port may be provided with a suitable fitting such
as a union, which may for example be threaded into the port to
which the fuel line may then be secured. The fitting may be mounted
within a bearing (not shown) allowing the nose cone to rotate
relative to the fuel line. In the embodiment, the nose cone 26 has
a bearing recess 31 formed into a rear face of the nose cone which
holds the roller bearing 30. The bearing recess 31 communicates
with the fuel port through conical transition section 29. The
bearing rotatably holds a stub shaft member of the front rotor 32
as will be described below. The bearing 30 may have ceramic rolling
elements, although other materials may be used instead.
[0048] The engine may be used at the aft end of small air vehicles
such as mini-cruise missiles (as shown in FIGS. 11-12) and Unmanned
Aerial Vehicles (UAVs), which may have fuel tanks 202, 202'
integral with the fuselage 203, 203'. Accordingly, the engine 10
preferably has a connection (such as a union) to the fuel entry 28
in the nose cone or center member 26 for coupling the fuel pipe in
the shortest manner from the fuel tank 202, 202'.
[0049] From the fuel entry 28, fuel enters the region of the
rolling element thrust bearing 30 for the front compressor 32,
where the fuel cools the bearing and also provides a measure of
lubrication to the bearing. The bearing 30 is lightly loaded
because the front rotor 32 is substantially supported at the outer
shroud directly from the outer casing as will be described
below.
[0050] The fuel then enters radial slots on the front face of the
front rotor 32, and is slung radially outward, with pressure
generated by the centrifugal action of the rotor as it spins.
[0051] Still referring to FIGS. 1-1A, the inner surface of the
inlet section is provided with a groove 23 which houses a suitable
foil bearing 100 which supports an integral shroud 35 of the first
rotor. The shroud 35 is nested in the inner portion of groove 23 as
shown in FIG. 1. Also as seen in FIG. 1E, the first stator section
24C of the outer casing 24 preferably includes an outer portion
124O and an inner ring portion 124C which define compressor pathway
40. Stator vanes 38 are disposed in pathway 40 between the inner
and outer portions of the first stator section 24C. The second
stator section 24D of the outer casing also has an inner groove
similar to groove 23 formed on the inside pathway surface of the
casing. The groove 27 also houses a foil bearing 100 which supports
an integral shroud 45 of the rear rotor 44. As seen in FIG. 1, the
shroud 45 is nested within an inner portion of groove 27. In the
embodiment, the second stator section may also include an inner
ring portion 124D which defines compressor pathway 41 (See FIG.
1A). Stator vanes 39 are disposed in the pathway 41 between the
inner portion and outer portion of the second stator section of the
casing.
[0052] Referring now also to FIG. 1G, the exhaust section 24R of
the outer casing 24 defines a cross-over passage 102 which holds
the precombustor 50 therein. The rear or bottom end 54 of the
exhaust section 24R is substantially closed by plate member, or an
integral rear member 124B except for a hole 104 for cartridge
igniter 56, and exhaust passages or outlets 54E formed into the
rear member 124B. As can be seen in FIG. 1G, the igniter hole is
essentially centered and may be threaded in order to allow the
cartridge igniter 56 to be threadably mounted into the igniter hole
104. The exhaust passages 54E are suitably sized to allow movement
of the exhaust gases from the engine throughout the operating range
of the engine, and are distributed circumferentially around the
rear member 124B of the outer casing. The precombustor 50 may be
mounted to the outer member 124B of the casing 24 as will be
described in greater detail below. As shown in FIGS. 1-1A, the
inner surface 106 of the exhaust section 24R is suitably shaped to
provide the cross-over passage 102 with an outer transition for
distributing flow of compressor gas exiting the compressor pathway
41 substantially evenly across the height of the cross-over passage
102.
[0053] Referring now to FIGS. 1-1A and 1D, the front rotor 32
generally comprises center section 32C, compressor blades 34,
turbine blades 36, inner turbine ring 112, and outer filament-wound
shroud 35. The center section preferably has a general bell-shape
which defines an annular plenum 110 at the rear face of the center
section. The rear face of the center section has a center diverging
cone 108 as shown in FIGS. 1-1A. The front of the center section 32
has a stub shaft 32S. The shaft 32S is sized to be a rotatably
mounted to the bearing 30 in the nose cone 26 as shown in FIGS.
1-1A. The front of the center section also includes fuel slots or
passages 33. The fuel slots 33 extend substantially radially
outwards from the stub shaft. In alternate embodiments the fuel
passages may be through bores in the center section. Each fuel slot
has an entry port substantially aligned with bearing 30 when the
rotor is mounted to the nose cone. The exit end of the fuel slot 33
is at the root of the compressor blades 34 on the front rotor 32.
When the front rotor is mounted to the casing, the front face of
the rotor center section 32C makes a seal with the rear face of the
nose cone 26 (see FIG. 1) so that fuel flowing through the slots 33
flows within the slots without substantially spilling over the
radial edges of the slots. The center section may be made of any
suitable heat resistant and insulating material such as
carbon/carbon composite or ceramic. The lower edge of the center
section 32C may include a filament wound portion or band that
supports the compressor and/or turbine blades by hoop tension. The
reinforcing filaments, 114, may be made of any suitable
reinforcement material, such as carbon or alumina/silica fibers.
The filaments also enable application of a hoop prestress to the
filament wound portion of the center section 32C. The compressor
blades 34 are mounted to the outside of the center section 32C. The
compressor blades 34 on the front rotor 32 form a first compressor
stage 13 of the nested core engine 10.
[0054] The engine 10 preferably uses a mixed-flow first compressor
stage 13 with a pressure ratio in the range of 2.5:1 to 3:1, using
a rotor tip speed in the range of about 465 to 515 m/sec (1525 to
1690 ft/sec). The isentropic work coefficient is about 0.40,
considered reasonable for the mixed-flow compressor.
[0055] Efficiency of the compressor stage 13 is enhanced by having
reduced axial flow velocity, and long blade chords. Conventional
axial-flow gas turbine engines are generally designed for fighter
aircraft, and then scaled down if needed to the small sizes.
However, this is not optimal for the small engines, because fighter
engines are designed to maximize air flow swallowing capacity. This
causes them to have high axial flow velocity, and the blades then
need to have short chords for short axial length. When these blades
are scaled down, the blade chords are too short for efficiency
(Reynolds Number effects, blade thickness/chord ratios for
manufacturability, blade relative roughness ratios, etc.).
[0056] The engine 10 preferably has reduced axial velocity, and
thus compressor blades 34 may have long blade chords, to retain
aerodynamic efficiency. A side effect of this change is to reduce
the flow diffusion angle between the blades 34, resulting in
reduced flow separation, greater diffusion efficiency, higher
pressure rise per stage as well as greater stage efficiency in
comparison to conventional engines. Also, the long-chord blades 34
can be thicker and stronger, and hence impart a measure of
FOD-resistance against the smaller insects and debris that gets
through the anti-FOD vanes at the front of the outer casing. The
compressor blades preferably have titanium leading edge
covers/coatings (not shown) for extra FOD-resistance.
[0057] The filament-wound outer shroud 35 is preferably a one-piece
member that encloses the compressor section of rotor 32, and
provides great hoop strength against the radial centrifugal loads
on the blades 34. The reinforcement filaments or fibers are similar
to reinforcement fibers 114 in the outer edge of the center section
32C, and may be made of any suitable material. The filaments in the
outer shroud 35 may be pretensioned generating a pre-load in the
outer shroud thereby placing the compressor blades 34 in axial
compression. This reduces the tensile and bending loads on the
compressor blades 34 during engine operation, easing the structural
requirements of the blades. Also, because the blades 34 are
shrouded, the long-chord blades avoid the large tip-vortex losses
associated with conventional small compressors.
[0058] The outer shroud 35 is sized to be conformal to groove 23 in
the inlet section 24I (See FIG. 1B) of the outer casing. The shroud
has an outer seating surface which is configured to rotatably seat
against the foil bearing 100 located in the groove 23 of the inlet
section 24I (see FIG. 1B). Minor leakage of air may occur between
the rotating shroud 35 and the stationary engine casing 24. Seals
(not shown) may be provided between the outer shroud of the front
rotor 32 and the outer casing to control air leakage as desired. In
comparison to conventional engines, it is much easier to seal the
continuous surface of the shroud 35 than it is to ensure very tight
running clearances, despite engine transients, around the tips of
individual blades in the compressor sections of conventional gas
turbines. Air leakage between shroud 35 and casing 24 may be used
for providing air for self-pressurizing and cooling the foil air
bearings 100, discussed below, that support the engine rotor 32.
The foil bearings 100 support the front rotor 32 against both
radial and forward axial loads. Accordingly, the roller bearing on
the stub shaft of the front rotor 32 is lightly loaded, providing
support for aftward axial loads that are normally of a smaller
magnitude and/or transient in nature, and allows the front rotor 32
to operate substantially as a shaftless rotor.
[0059] Turbine blades 36 are located on the inside of the center
section 32C. In particular the blades 36 are disposed as shown in
FIG. 1, inside the annular plenum 110 at the rear face of the
center section. The inner turbine ring 112 is located generally in
the annular plenum 110 of the center section 32 as shown in FIG. 1.
The turbine blade tips are connected to the inner turbine ring 112.
The front edge of the inner turbine ring 112 is suitably rounded to
form, in combination with the diverging center cone 108 of the
center section, an annular turbine nozzle as shown in FIG. 1 for
the turbine section of the front rotor 32. The front rotor thus
includes an integral compressor section and an integral turbine
section. During engine operation the outer compressor section is
the cold section of the rotor through which cold compressor air
flows, and the inner turbine section is the hot section of the
front rotor through which hot combustion gases flow.
[0060] Thus, the front compressor rotor 32 has an integrated
rotating-nozzle turbine which is the first turbine stage 17 of the
engine 10, as shown in the FIGS. 1-1A. This integration eliminates
the back-face disk pumping losses for both the compressor and the
turbine, and also eliminates the critical speed problems of
high-speed shafts. Heat transfer through the common interface is
reduced by use of ceramic construction, at least in the center
section 32C of the front rotor 32, that has lower conductivity than
typical metals. The unique rotatable nozzle first turbine stage 17
preferably has no stationary nozzles. It takes in hot combustion
gases at low relative velocity, turns and expands the gases, and
ejects the gases at high relative velocity for tangential thrust
(torque) to drive the first compressor stage 13. The turbine 17 is
expected to have competitive efficiency at somewhat lower work
coefficient than conventional turbines, and offers a very
significant benefit of immunity to the thermal pattern factors of
conventional combustors. For the same material temperature
limitation, this can allow a significant increase in peak cycle
temperature in turbine 17. The turbine nozzle 17 can also be
un-cooled, due to ceramic construction and absence of Pattern
Factor effects, for significant savings in cost compared to
conventional engines with cooled turbines.
[0061] Referring now to FIGS. 1-1A, and 1F, the rear rotor 42 (a
perspective view of which is shown in FIG. 1F) of engine 10
generally comprises middle ring 80, compressor blades 44, turbine
blades 46, inter turbine ring 82, and outer shroud 45. As seen in
FIGS. 1-1A, in the embodiment, the rear rotor 42 also has
turbulator blades 48 (the turbulator blades are not shown in FIG.
1F for clarity). The compressor blades 44 are attached
substantially radially outward from the middle ring 80. The outer
shroud 45 extends around the tips of the compressor blades 44. The
compressor blade tips are anchored to the shroud 45. The center
ring is preferable a one-piece member made of ceramic or any other
heat resistant/insulative material. The outer surface of the middle
ring 80, and the inner side of the shroud 45 are suitably shaped to
form an appropriate compressor passage for air emerging from
pathway 40 and the first stator section 24C of the outer casing
(see FIGS. 1-1A). The second compressor stage 14 of the engine is
provided by the compressor section (i.e. the middle ring 80,
compression blades 44, and outer shroud 45) of the rear rotor
42.
[0062] Similar to the first stage 13, the second stage compressor
14 is also a mixed-flow device, with reduced flow angles and
long-chord blades 44. Independent spools for two compressors (i.e.
front rotor 32 for the first stage, and rear rotor 42 for the
second stage) offer better off-design matching and greater
resistance to blade stalls. The compressor blades 44 may have
titanium leading edge covers/coatings for extra FOD-resistance
similar to blades 34 of the first compressor stage.
[0063] The second compressor stage 18 may preferably employ a
pressure ratio in the range of 2.0:1 to 2.5:1, using a rotor tip
speed in the range of about 470 to 560 m/sec (1540 to 1840 ft/sec).
The isentropic work coefficient may be about 0.40, considered
reasonable for a mixed-flow compressor.
[0064] Shroud 45 is generally similar to shroud 35 of the front
rotor 32. The shroud 45 may be a one piece filament-wound member
with reinforcing filaments or fibers substantially similar to
reinforcing fibers 114 in shroud 35. The shroud may be pre-stressed
to maintain the compressor blades in compression. The shroud is
sized to conform to the receiving groove 27 in section 24D of the
outer casing as shown in FIG. 1. Seals (not shown) may be provided
to seal air gaps between the shroud 45 and groove walls. The outer
surface of the shroud is configured to seat against foil bearings
100 in groove 27. Foil bearings 100 are only one of two foil
bearings used to support the rear rotor 42 axially and radially
(bearing 100' supports the inner turbine ring as will be described
below. Air leakage between the shroud 45 and groove 27 during the
engine operation supplies air to foil bearing 100. As seen in FIGS.
1-1A, the front and rear edges of the middle ring 80 form a general
lapped or rabbet interface with the adjoining faces of the inner
sections 124C, 124D immediately in front and to the rear of the
rear rotor 42. This minimizes leakage between the inner turbine
section 16 and surrounding outer compression section 12, and
minimizes the adverse aerodynamic consequences of such leakage.
[0065] The configuration of the nested core engine 10 results in
reduced potential for internal air leaks. A gas turbine engine with
a conventional layout having two mixed-flow compressors and two
axial flow turbines, would have 4 locations for blade tip leaks,
and 7 locations for hub leaks. The engine 10, on the other hand,
has only 2 places for shroud leaks (which are easier to control
than blade tip leaks), and has 5 places for inter-shroud leaks. The
net leakage of air is therefore expected to be less than in a
conventional engine.
[0066] Most of the leak sites in the engine 10 have relatively low
pressure differential across them, and also use the air leaks to an
advantage in providing air for the foil bearings as previously
described. There are two sites in the engine 10 that have high
pressure differential across them: at the exit hub of the
compressor section of the rear rotor 42, and between the hub of the
turbine section of the rear rotor and the combustor 52. Brush seals
115 are used at these locations.
[0067] As seen in FIG. 1, in the embodiment, the exit or rear edge
of the middle ring 80 has a seal groove or channel 84 formed
therein. The seal groove 84 interfaces with brush seals 115 mounted
on the inner section 124d of the outer casing 24. In alternate
embodiments, the brush seals may be mounted on the rotor and seated
against seals surfaces on the casing. Brush seals 115 may for
example comprise fine (0.003'' .PHI.) cobalt or ceramic fiber
brushes running against hardened surfaces of groove 84. The brushes
offer greatly reduced leaks compared to the conventional knife-edge
or labyrinth seals. The orientation of the brush seals 115 in FIG.
1 is for example purposes only, and the brush seals may have any
other suitable orientation rubbing against any other suitable
sealing surface on the rear rotor.
[0068] The compressor and turbine sections 12, 16 of the engine 10
have either a presence of fuel vapors and/or partially burnt gases
that offer some further lubrication to both brush seals 115, and to
foil bearings 100, 100', as noted below. For example, it has been
found that in the presence of hot Nickel or hot Silicon Nitride,
very fine soot particles are formed that can act as a
lubricant.
[0069] The engine 10 can function without lubrication, using only
air for the foil bearings. However, the engine 10 takes advantage
of the potential formation of lubricious soot from fuel vapors and
partially burnt hydrocarbons to further enhance the life of the
foil air bearings and the brush seals.
[0070] The turbine blades 46 of the rear rotor 42 are captured
between the middle ring 80 and the inner turbine ring 82. The inner
surface of the middle ring 80 and outer surface of the inner
turbine ring 82 define the turbine section of the rear rotor. Thus,
the rear rotor 42 includes an integral compressor section on the
outside, and a turbine section on the inside, the compressor
section substantially surrounding the turbine section. The turbine
section of the rear rotor 42 is the second turbine stage 18 of the
engine 10. The inner surface of the middle ring 80 is shaped to
blend smoothly with the inner side of section 124C of the stator
inner ring (see FIG. 1). The outer surface of the inner turbine
ring 82 is shaped to form a suitable axial-flow turbine. The inner
turbine ring 82 has a front extension portion 84 which projects
inside and overlaps with the trailing portion of inner turbine ring
112 of the front rotor 32 (see FIG. 1). The overlap aids in
reducing leaks between the engine turbine sections, and combustion
chamber as noted previously. The trailing edge 86 of the inner
turbine ring 82 of the rear rotor 42 has a channel formed therein
which defines a circumferential seating surface. This seating
surface in the inner turbine ring 82 is rotatably seated against
foil bearing 100' (similar to foil bearings 100 described
previously) mounted on an annular support flange of the
pre-combustor 50. Accordingly, the rear rotor 42 is supported
axially and radially both at the shroud 45 by foil bearing 100, and
at the inner turbine rotor 82 by foil bearing 100' (see FIG. 1).
The engine 10 may also include a brush seal (not shown) which
interfaces with the channel and the trailing edge 86 of the rear
rotor. The inner surface of the inner turbine ring of the rear
rotor is shaped to define a portion of the combustion chamber 52 of
the engine. As shown in FIG. 1, turbulator blades 48 are
substantially short, straight blades which extend inwards from the
inner surface of the inner turbine ring 82. The turbulator blades
48 extend sufficiently inwards to generate effective stirring of
gases in the combustion chamber section 52 when the rear rotor 42
is spinning.
[0071] As noted before, FIG. 1 also shows the structural
arrangement for the rotating components of the engine. In a
conventional gas turbine engine, the blades of compressors and
turbines are attached to disks, either directly or through
fir-tree-root geometries. Operation of such conventional engines,
with the rotation of the compressors and turbines, creates tensile
stresses in the blades and the disks, with the internal structure
of the disk resisting the tensile stresses. Such conventional
engines can thus be said to have an endo-skeletal structure. In
contrast, the engine embodiment shown in FIG. 1, can be said to
have an exo-skeletal structural arrangement. In this embodiment,
the first rotor has only a partial structural disk 32, near the
base of the blades, and the second rotor 42 has only a narrow ring
82, instead of a disk, near the base of the blades. In the
embodiment shown in FIG. 1, structural strength for retaining the
compressor blades 34 and 44, and turbine blades 36 and 46,
respectively of the first and second rotors, is provided by
fiber-reinforced structural rings 35, 32R and 45, for the first and
second rotors, respectively. As a clarification, the structural
ring 35 helps retain compressor blades 34, the structural ring 32R
helps retain turbine blades 36, and the structural ring 45 helps
retain both the compressor blades and the turbine blades. The use
of these structural rings places the blades in compression, which
enables the use of blade materials that have low tensile strength
but high compressive strength, such as ceramics that are capable of
operation at high temperatures. The structural rings themselves are
expected to develop high tensile hoop stresses from the blade
centrifugal loads, and thus the use of fiber reinforcements, 114,
which typically have high tensile strengths, offers an optimal
structural arrangement. The fibers may be of a chopped/whisker
variety for low cost, or the rings may have filament wound fibers
for high strength.
[0072] Referring now to FIGS. 1-1A, and 1G, the precombustor
generally comprises a top foundation or support plate 90, and tubes
92. The tubes 92 are mounted at the front edge to the support plate
90. The support plate 90 may be a generally annular and
substantially flat plate. In alternate embodiments, the top plate
of the precombustor may have any other suitable shape. The outer
edge of the support plate 90 rests against the inner side of the
inner section 124D of the outer casing as shown in FIG. 1. At the
inner edge of the annulus, a collar or circumferential flange 94
projects forward from the support plate 90. As shown in FIG. 1, the
collar 94 is sized and shaped to be conformally received inside the
channel 86 in the trailing edge of the turbine ring 82 of the rear
rotor 42. As seen best in FIG. 1, the collar 94 has a groove for
seating the foil bearing 100', as stated before, and may be
provided with a brush seal (not shown) to seal the second stage
turbine rotor 18. The support plate 90 is perforated with apertures
96 for tubes 92. The ends of tubes 92 may be press fit, bonded or
otherwise fitted into the apertures 96 to mount the tubes to the
support plate. Tubes 92 extend substantially the height of
cross-over passage 102 in the exhaust section 24R of the casing 24.
The rear end of the tubes 92 is received into corresponding exhaust
holes 54E in the casing. Thus, the top support plate 90 forms the
upper surface of the precombustor 50 in the exhaust section 24R of
the engine 10. Tubes 92 have a generally cylindrical shape, though
in alternate embodiments the tubes may have any other suitable
shape. The tubes may be of generally large size and few in number
for low-pressure drag across the precombuster 50. The outer and
inner surfaces of the tubes 92 are coated by a platinum and
palladium catalyst.
[0073] Tubes 92 may be made of ceramic, such as silicon carbide,
alumina, mullite, zirconium (high temperature resistance) or
cordierite (excellent thermal shock resistance), to withstand the
relatively high temperature with minimum weight penalty, and to
provide naturally high surface area for the catalyst coatings. It
is synergic that the tubes 92 have high pressure air on the
outside, and lower pressure gases on the inside, imposing
compressive stresses in the tubes, synergistically using the high
compressive strength of structural ceramics. The tubes 92 have a
suitable wall thickness allowing adequate heat transfer through the
wall so that the tubes 92 also act as a recuperator.
[0074] As shown in FIGS. 1-1A, the combustion chamber section 52 of
the engine 10 is generally formed by the rear face of the front
rotor 32, the inner surface of the turbine ring 112 of the front
rotor 32, the inner surface of the turbine ring 82 of the rear
rotor 42, and the inner region inside the innermost tubes 92 of the
pre-combustor 50. The combustion chamber section 52 is at the core
of the engine within the engine hub (i.e. inner rings 112, 84 of
the front and rear rotors 32, 42) and is surrounded by the turbine
section 16 of the engine 10. The turbulator blades 48 (on the rear
rotor 42) project into the combustion chamber section 52 as shown
in FIG. 1. The turbulator blades 48 induce air and gases to flow
into all regions of the combustion chamber by their centrifuging
action when the rear rotor 42 is spinning. Tip vortices shed by the
blades 48 stir the air and gases in the combustion chamber, to
ensure uniform post-combustion temperatures and the absence of any
Pattern Factor and Radial Temperature Profiles, which create
thermal problems for the combustors and turbines in conventional
gas turbine engines.
[0075] FIGS. 1 and 1A show a starter cartridge, 56, that is a means
to start the engine. Other means may be used to start the engine.
The cartridge may be a small rocket-type generator of hot gases. As
shown in FIGS. 1 and 1A, the starter cartridge 56 is threaded into
the engine. Alternative means of fixing the cartridge into the
engine may be used. The starter cartridge is shown to be located
along the axis of the engine, but locations offset from the axis
may be used. The location and orientation of the starter cartridge
is designed to feed the hot gases from the cartridge directly into
the combustion chamber of the engine, such that the exhaust gases
will pre-warm the combustion region for efficient combustion of
fuel and air, and the exhaust gases will also drive the turbines
downstream of the combustor, to thereby drive the compressors and
pump air (and fuel) into the engine combustion region. The starter
cartridge may also have a combustible casing, to minimize blockage
of the flow of air and fuel through the combustor after the
cartridge has burnt out.
[0076] As shown in FIG. 1, in the embodiment, the engine core has a
cartridge starter 56 embedded at the bottom 124B of the engine
casing 24. When ignited, the rapidly burning cordite charge
releases hot, high-velocity gases that flow through the combustion
chamber section 52 and impinge against the turbine blades 36, 46 of
the front and rear rotors 32, 42. This causes the turbines 17, 18,
and the attached compressors 13, 14 to start spinning, inducing air
flow through the engine 10. The hot cartridge gases also begin
heating the catalyst-coated tubes 92 in the pre-combustor 50
downstream of turbine 18. Simultaneously, fuel begins spraying (in
the direction indicated by arrow F1 in FIG. 1) into the compressor
13 and reaction begins on the catalyzed surfaces of the rapidly
heating tubes 92 of the pre-combustor 50. This rapidly starts the
engine 10 as the cartridge 56 burns out. As noted before, the
cartridge 56 has a combustible casing for the cordite or equivalent
charge. Hence the air and gas flow path inside the combustion
chamber section 52 becomes unobstructed as the engine 10 spools up
and the cartridge 56 burns out. In alternate embodiments, the
engine may have a small, high-speed, permanent magnet
motor/generator in the nose cone. This motor/generator can be
cooled by fuel flow, for benefits of high power density and
lightweight. Direct mounting of electric motor/generator and fuel
pump on the engine shaft can eliminate the auxiliary gearbox, for
considerable savings in cost and improved reliability for the small
engine.
[0077] In the embodiment shown in FIG. 1A, the starter cartridge,
56, has a unique shape of the starter cartridge base, 56B, such
that the shape of the base forms a toroidal recirculation region,
56T (seen best in FIG. 1G), for the air and fuel, to help stabilize
combustion, after the cartridge with the combustible casing has
burnt out. In alternate embodiments, the base of the cartridge may
have any other shape, including a low-cost flat shape.
[0078] As the engine 10 spools up, suction generated at the inlet
22 draws air (in the direction indicated by arrow A in FIG. 1) into
the inlet 22, and then into the first compressor stage 13. FIG. 1
shows the process of introduction of the fuel, F1, into the engine,
wherein the fuel first cools and lubricates the rolling element
bearing 30, and is then centrifuged radially along a fuel feed
surface (such as slots 33, though any surface may be used) outward
by the compressor rotor to be injected as a then sheet-like spray
into the air flow path. Fuel flows (in the direction indicated by
arrow FE) under pressure and/or gravity feed from the fuel source
into the nose cone 26 and then through bearing 30 into the fuel
slots 33 at the front face of the front rotor. From the fuel slots
33, fuel is sprayed (as indicated by arrows F1) into the air stream
A in the compressor 13. As noted before, the centrifugal action of
the spinning rotor 32 generates outward pressure on the fuel in
slots 33 to spray fuel substantially across the compressor inlet.
As the fuel enters the flow path into the compressor section 12 of
the engine, the combination of the pressure difference and the
relative velocity between the fuel and the air causes the fuel to
be atomized in the compressor into small droplets and/or mist
particles. The effect of liquid fuel introduction at the inlet face
is equivalent to flying a turbine engine through light rain, and
causes little performance degradation of the compressor 12 of the
engine 10. A great advantage accrues from this manner of fuel
introduction for the small engine: countering the effects of the
square-cube law on the engine combustor, by converting the entire
engine into a combustor, as noted below. The first compressor stage
13 compresses the fuel air mixture as described previously. From
the first compressor stage 13, the mixed flow moves through pathway
40 (in the direction indicated by arrow AC1) through the
intermediate stator 38, into the second compressor stage 14. The
fuel air mixture is further compressed in the second compressor
stage 14 and then flows as indicated by arrow AC2 through the rear
stator 39 into the transition 106 at the rear of the engine casing
24 (see FIG. 1). Air from the compressor is transferred (in
direction indicated by arrow PC) across the exhaust gas flow path
by flowing the pressurized air over hollow tubes 92 of the
pre-combustor 50 that have exhaust gases flowing axially inside the
tubes.
[0079] The fuel for the engine injected into the inlet face of the
first compressor 13, and atomized by the shearing action between
the centrifuged fuel droplets and the high relative speed of the
air flow A, is substantially evaporated by the time the fuel
reaches the catalytic pre-combustor 50. The catalytic surfaces of
the pre-combustor initiate combustion, and raise the temperature of
the fuel-air mixture from the compressor delivery temperature
(.apprxeq.600 F/600 K) to about 1000 F/800 K. This temperature is
high enough to initiate rapid combustion in the main combustion
chamber 52, but low enough to be within acceptable temperatures
(.apprxeq.1200 K) for the catalyst. The catalyst surfaces in the
pre-combustor are maintained above its operating temperature by the
exhaust gases flowing through the tubes 92.
[0080] The burning mixture of air and partially oxidized fuel flows
from the catalytic pre-combustor 50 to the fully-stirred lean-burn
main combustion chamber 52 in the central part of the engine
10.
[0081] Conventional combustors, fed by warm air from compressors
and cold fuel injected into the warm airstream, cannot sustain
stable combustion under all operating conditions if they were to
operate under homogenous conditions of premixed fuel and air, due
to flammability limits and due to the problems of auto-ignition,
flashback and acoustic resonance. Therefore, conventional
combustors operate on the rich-lean system, or the newer
rich-quench-lean systems for low NOx.
[0082] The engine 10 aims to operate to use the pre-mixed,
pre-vaporized system, but uses the catalytic pre-combustor to
initiate combustion. The lower compressor pressures and exit
temperatures, combined with substantially complete homogeneity of
the fuel-air mixtures, helps ensure the absence of flash-back. The
resulting partially burnt hot gases have much wider flame stability
limits, and thus can operate using the lean-premix system. Such
catalytic lean-premix combustors have been demonstrated in
laboratories. Catalytic combustors also avoid acoustic
resonance.
[0083] There is some concern that, for conventional, large,
high-pressure, gas turbine engines, fuel cannot be introduced too
early in the combustor. This is because the very short combustion
delay period at the high air delivery temperatures can cause
flashback in the pre-combustion region. FIG. 4 shows the variation
of delay period with air temperature.
[0084] For the engine 10, it is synergistic that fuel is introduced
early in the compressor 12, but flashback is prevented because the
low temperatures cause the delay period to be longer (>1 sec)
than the residence time (.apprxeq.1 msec) in the compressor 12.
However, once catalytic pre-combustion has warmed up the gases, the
delay period becomes short enough (.apprxeq. msec) to allow
completion of combustion within the residence time (.apprxeq.
several msec) for the combustion chamber 52.
[0085] A major advantage of the lean pre-mixed combustion chamber
is the general complete absence of soot or smoke. In conventional
engines, soot is caused by pyrolysis of large fuel droplets before
they can vaporize. In the engine 10, fuel is vaporized, by the
whipping action of compressor blades and the large relative air
velocities, well before the temperatures get hot enough for
pyrolysis. The absence of exhaust smoke offers a stealth advantage
for military systems. It also implies an absence of carbon balls,
that can cause hot-section erosion in conventional engines. In
addition, lean pre-mixed combustors have extremely low NOx, which
would be useful for commercial applications of the engine, such as
for APUs.
[0086] The airflow swallowing capacity of engines decreases with
square of engine linear dimensions, while the combustor volume
diminishes with the cube of engine scale. Small engines also have
lower cycle pressure ratios, needing larger combustor volume.
However, the specific heat release rate (e.g. BTU/hr per ft.sup.3
per atm. pressure or kW/m.sup.3 per atm.) is rather limited by the
combustion chemistry of fuel-air mixtures. The result is that, in
conventional engines, combustors occupy an increasingly larger
fraction of the total engine volume as engines are scaled down.
[0087] For the engine 10, almost the entire engine can be used as
the combustor, with the fuel mixing region in the compressor 12,
the catalytic pre-combustor 50 in the cross-over region, the main
combustion chamber 52 within the central part of the engine (that
is wasted in conventional engines), combustion continuing in the
contra-rotating turbine 16 that is immune to pattern factors (if
the engine was to have one), and the combustion reaching
completion, for final suppression of unburned hydrocarbons, within
the tubes 92 of the crossover region. The engine 10, thus,
overcomes the square-cube laws of scaling down.
[0088] From the main combustion chamber section 52, the hot
combustion gases are directed (as indicated by arrows C in FIG. 1)
into the rotating nozzle of the first turbine stage 17. For optimum
efficiency in creating torque, the exit gas (indicated by arrow
AT1) relative jet velocity from the nozzle 17 should be about twice
the rotor blade velocity for the first turbine stage. Thus, the
gases issuing from first turbine stage still have a large
tangential velocity component. This is used to create torque for
the second turbine stage 18, without the use of an intervening
nozzle. The second turbine stage 18 may have some additional
expansion of gases for torque generation (to power the rear rotor),
though this will be minimized in order to minimize loss of kinetic
energy in the exhaust gases. In combination, the two turbine stages
17, 18 essentially act as one conventional turbine nozzle and rotor
set, except that the nozzle is allowed to rotate and hence generate
torque to drive its own compressor. This reduces the number of
bladed stages for the turbine section. Any loss in turbine
efficiency because the first turbine acts as a free-to-spin nozzle
may be largely compensated for by the absence of blade tip
clearance losses (due to integral shroud 35 nested in the casing
surface).
[0089] Exhaust gases flow (as indicated by arrow E) through tubes
92 in the crossover region of the pre-combustor 50, as shown in the
FIG. 1. The tubes 92 may also be coated on the inside with a
Platinum/Palladium oxidation catalyst, to fully suppress any
unburned hydrocarbons, and any smoke if the engine ever makes smoke
(with lean pre-mixed combustion, the engine 10 is expected to not
produce any smoke).
[0090] Referring now to FIG. 3, there is shown a nested core gas
turbine engine 10A in accordance with another embodiment. Except as
otherwise noted below, engine 10A in FIG. 3, is substantially
similar to engine 10 described above and shown in FIGS. 1-1G, with
similar features having similar reference numbers. Engine 10 also
includes an outer casing 24A, a front rotor 32A, a rear rotor 42A,
and pre-combustor 50A. The front and rear rotors 32A, 42A have
integral compressor and turbine sections. In this case however, the
front rotor 32A, does not have an inner turbine ring similar to
ring 112 of the front rotor 32 in engine 10. Instead, the engine
casing 24 A has an intervening stator nozzle 25A between the first
turbine stage 17A and the second turbine stage 18A on the rear
rotor. The stator nozzle 25A comprises an inner ring 125A and vanes
23A which are mounted to the casing as shown in FIG. 3.
[0091] FIGS. 5 and 6 are schematic representations of the two
common conventional engines: centrifugal compressors with
wrap-around burners, and axial compressors with in-line burners.
The two engines and the engine 10 shown in FIG. 1 have the same air
flow swallowing capacity, and the about the same power. However,
the engine 10, 10A has about half the frontal area of a centrifugal
flow engine, and about half the length of an axial flow engine. The
weight of the engine 10, 10A may be one-third that of the
competition, because the engine 10, 10A does not have the heavy,
solid shafts of the conventional engines. In addition, the engine
10, 10A avoids the shaft critical speed problems of small gas
turbine engines. By using low-expansion, high-modulus ceramic
materials for the structure, it gains control over the tip
clearances in the rotor systems, and hence has significant
performance advantages.
[0092] The engine 10, 10A may have many derivatives, all using a
common core, but being coupled to different tail sections. This
will enable the core engine 10, 10A to be used for a variety of
applications, with a variety of thrust systems optimum for each air
vehicle.
[0093] An example of a turbojet 400 incorporating features of the
exemplary embodiments is shown in FIG. 7. This embodiment has a
thrusting nozzle 480 attached to the aft end of the core engine
similar to engine 10. In this embodiment, the engine 400 is
projected to produce about 20 lbf (9 kgf) of thrust, at a Specific
Fuel Consumption .apprxeq.1.5 lbm/hr/lbf (1.5 kg/hr/kgf). This
engine 400 will be suitable for propulsion of high speed air
vehicles 200, for example precision-targeted Mini-Cruise Missiles
launched from 70 mm (2.75 in.) rocket tubes (see FIG. 11). The
starter and launch-thrust booster cartridges 482, 484, with
combustible casings 486 and 488, are built into the engine.
[0094] As noted above, FIG. 7 shows an embodiment of a turbojet
engine 400 that has a core gas turbine engine similar to engine 10
described previously, and hence similar features are similarly
numbered. In this embodiment, a rocket type starter cartridge, 484,
is used as a boost cartridge for the turbojet engine, which may be
used to launch or accelerate the turbojet engine and an air vehicle
(similar to the air vehicle 200 shown in FIG. 11) powered by the
turbojet engine. In the embodiment shown in FIG. 7, ignition
systems, 491 and 492, for the core engine starter cartridge and the
turbojet engine boost cartridge respectively, are connected to a
common ignition initiation system, 495. Further, in the embodiment
shown in FIG. 7, the starter cartridge 482 used to start the core
engine is similar to the boost cartridge 484. In alternate
embodiments, the starter cartridge and the boost cartridge may be
different.
[0095] An example of a low bypass turbo-fan engine 500
incorporating features of the exemplary embodiments is shown in
FIG. 8. The thrusting nozzle can be replaced by a an integrated
turbo-fan 588, in which a turbine 590 in the exhaust gas flow path
drives a fan 592 directly radially outside the turbine. To minimize
shock losses at the high tip speeds, the fan blades are to have
sharp leading edges combined with blade sweep and lean. There is a
slight gain in thrust (to 21 lbf), and a slight reduction in SFC
[to .apprxeq.1.45 lbm/hr/lbf (0.9 kg/hr/kgf)]. The main advantage
of this version is the gain in propulsive efficiency of the system
due to Wake Ingestion Propulsion of the airframe. The noise is
about the same as the turbojet, because the lower noise from lower
exhaust gas velocity is countered by noise from the transonic fan
blades. This version will also be more expensive than the turbojet,
and is better suitable to long-range missiles or non-expendable
targets. The starter and launch-thrust booster cartridges are built
into the engine.
[0096] For long-range/high-endurance air vehicles, and for V/STOL,
hover-capable air vehicles, higher Bypass Ratio is more desirable,
to reduce fuel consumption and to reduce noise. This is achieved by
additional derivatives of the engine as discussed below.
[0097] An example of a high-bypass twin contra-rotor fan engine 600
incorporating features of the exemplary embodiments is shown in
FIG. 9. The aft section of this engine has a two-stage
contra-rotating turbine 694 driving integrated shroud-fans 696 at
relatively low speeds. This allows the fan diameter to be larger
without incurring shock losses on the blade tips. The result can be
higher thrust (.apprxeq.25 lbf), lower fuel consumption
.apprxeq.1.2 lbm/hr/lbf (1.2 kg/hr/kgf) for longer range/endurance,
and lower noise. This embodiment is likely to be well suited to
very long-range cruise missiles 200' (see FIG. 12) and completely
reusable reconnaissance air vehicles. FIG. 9 shows such an aft fan
with a shroud 698 for low tip losses and enhanced propulsive
efficiency. The shroud 698 is supported by a grill 699 attached to
the mid-section 624 of the core engine that also provides F.O.D.
resistance by blocking ingestion of birds and bugs.
[0098] An example of an ultra-high bypass lift-fan engine 700
incorporating features of the exemplary embodiments is shown in
FIG. 10. Pressurized gases can be taken from a core engine and
ducted to hollow tubes 795 within the blades 796 of a rotor, where
the gas is allowed to escape through tip jets or aft-facing slots
798 in the rotor blades. Such a rotor is ideal to generate lift for
a hover-capable air vehicle because of the absence of
counter-torque eliminates the need for a tail-rotor. The exemplary
embodiment of the ultra-high-bypass engine 700 shown in FIG. 10 has
the core gas turbine engine similar to engine 10 described
previously. Hence, similar features are similarly numbered. In this
embodiment, the fan is driven by hot gases from the core engine,
wherein the hot gases flow from the core engine via a rotating
plenum 794, though passages 795 of the hollow fan blades 796, and
are effluxed from the passages via aft-facing tip jets 797 or
aft-facing slots 798 on the upper surface of the fan blades 796.
Gases issuing from the tip jets 797 will provide torque to help
drive the fan in a rotational direction. Gases issuing from the
slots 798 will provide torque to help drive the fan in a rotational
direction as well as enhance the lifting effectiveness of the fan
blades 796.
[0099] Other benefits include greatly increased thrust (.apprxeq.35
lbf), lower fuel consumption .apprxeq.0.9 lbm/hr/lbf (0.9
kg/hr/kgf) for longer range/endurance, and lower noise. The main
disadvantage is losses incurred in the ducts to the blade jets.
However, this is balanced by elimination of losses associated with
a power turbine, and the losses in the multi-stage gearboxes used
in conventional helicopters. One example of a UAV embodiment 800
using engine 700 is shown in FIGS. 13, 14A-14C and 15A-15B. The
configuration of the UAV shown in FIGS. 13, 14A-14C and 15A-15B is
merely an example of a suitable UAV configuration using the
ultra-high bypass lift-fan engine 700. The ultra-high bypass
lift-fan engine may be used with any other suitable UAV
configuration. As seen in FIGS. 13-14A, the engine 700 is
substantially centrally mounted in a vertical duct of the engine.
The lower opening of the duct has doors or louvers which are opened
during hover (see FIGS. 15A-15B), and closed or partially closed
when the UAV 800 is moving horizontally (see FIGS. 14B-14C). The
air passage has a horizontal exhaust which can be opened or closed
with a flap at the exhaust opening as shown in FIGS. 14B-14C.
[0100] FIG. 11 shows an example of an application of the Nested
Core Engine used for a small, high-speed air vehicle, 200, such as
for example a missile or an Unmanned Aerial vehicle. As seen in
FIG. 11, in this embodiment the vehicle 200 has a body 203, with an
internal fuel tank 202, wherein the fuel tank is connected directly
to the fuel inlet of the Nested Core Engine, as also shown in FIG.
7 and described before.
[0101] FIG. 12 shows an example of another application of the
Nested Core Engine with a power turbine in the engine exhaust
driving an aft fan, used for a small, high-speed air vehicle 200',
such as for example a missile or an Unmanned Aerial vehicle. As
seen in FIG. 12, in this embodiment the vehicle 200' has a body
203', with an internal fuel tank 202', wherein the fuel tank is
connected directly to the fuel inlet of the Nested Core Engine, as
also shown in FIGS. 7 and 9, and described before.
[0102] The nested core gas turbine engine 10 described above and
shown in FIGS. 1-1A, and 3, is a new gas turbine engine that offers
about half the frontal area compared to the typical small turbine
engines with centrifugal compressors, and offers about half the
length compared to typical gas turbine engines with axial flow
compressors. Simultaneously, an engine comprising features of the
exemplary embodiments can offer one-third the weight, as well as
dramatic improvements in component and overall efficiencies that
compensate for the adverse effects of scaling down engines.
[0103] In the embodiments described above, the turbine section of
the engine is substantially nested within the compressor section of
the engine, and again, the combustor section of the engine is
nested within the turbine section of the engine. This gives the
engine an appearance of having been telescoped into itself,
offering very significant reductions in engine overall volume and
weight. Additional advantages of the nested core gas turbine engine
10, 10' in comparison to conventional gas turbine engines are
illustrated in the graphs shown in FIGS. 18-25.
[0104] A nested core gas turbine engine comprising features
described above can be the basis for new applications, such as
mini-cruise missiles, targets & drones, and cruising &
hovering Unmanned Aerial Vehicles, using turbojet, low-bypass
turbofan, high-bypass turbofan and slot/tip-driven lift rotor
versions of the engine.
[0105] It should be understood that the foregoing description is
only illustrative of the invention. Various alternatives and
modifications can be devised by those skilled in the art without
departing from the invention. Accordingly, the exemplary
embodiments are intended to embrace all such unforeseeable
alternatives, modifications and variances which fall within the
scope of the appended claims.
[0106] Some examples of such alternative embodiments are shown in
FIGS. 26-29, and 29A. These embodiments show the nested core
engines 900-900D with combustor sections 952, 952A, 952B, 952C, and
952D nested substantially radially within turbine sections
(917-919, 917A-919A, 917B-918B, 917C-918C, and 917D-918D), and the
turbine sections nested substantially radially within compressor
sections (913-916, 913-915A, 913B-915B, 913C-915C, and 913D-915D).
These embodiments have a greater number of compressor and turbine
stages compared to the embodiment for the engine 10, 10' discussed
earlier and shown in FIGS. 1-1A and 3.
[0107] Fuel for the embodiments shown in FIGS. 26-29 is introduced
into the engines at alternative locations as shown (indicated by
arrows F1, F2, F3, and F4 in FIGS. 26-29 respectively), and may be
staged, as a function of engine operating condition, for optimum
operation of the engines. FIG. 29A shows additional alternative
locations for the fuel introduction at different locations (in the
directions indicated by arrows FF) upstream of the pre-burner
950D.
[0108] Foil bearings for the engines shown in FIGS. 26-29, 29A are
located outside the compressor shrouds, and have a generally
conical configuration to support axially forward as well as axially
aftward rotor thrust loads.
[0109] FIGS. 30A-30D, 31A-31D and 32A-32D show examples of
high-speed aircraft embodiments 1000-1000' that use alternative
embodiments of the nested core engines in a lift-fan configuration,
deriving benefit from the short axial length of the nested core
engines. Alternative aircraft embodiments can be made using the
nested core engines in similar aircraft configurations.
* * * * *