U.S. patent application number 16/656529 was filed with the patent office on 2020-09-24 for micro-turbine gas generator and propulsive system.
The applicant listed for this patent is JETOPTERA, INC.. Invention is credited to Andrei Evulet.
Application Number | 20200300166 16/656529 |
Document ID | / |
Family ID | 1000004873623 |
Filed Date | 2020-09-24 |
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United States Patent
Application |
20200300166 |
Kind Code |
A1 |
Evulet; Andrei |
September 24, 2020 |
MICRO-TURBINE GAS GENERATOR AND PROPULSIVE SYSTEM
Abstract
A propulsion system includes a first compressor in fluid
communication with a fluid source. A first conduit is coupled to
the first compressor, and a heat exchanger is in fluid
communication with the first compressor via the first conduit. A
second conduit is positioned proximal to the heat exchanger. A
combustor is in fluid communication with the heat exchanger via the
second conduit and is configured to generate a high-temperature gas
stream. A third conduit is coupled to the combustor, and a first
thrust augmentation device is in fluid communication with the
combustor via the third conduit. The heat exchanger is positioned
within the gas stream generated by the combustor.
Inventors: |
Evulet; Andrei; (Edmonds,
WA) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
JETOPTERA, INC. |
Edmodns |
WA |
US |
|
|
Family ID: |
1000004873623 |
Appl. No.: |
16/656529 |
Filed: |
October 17, 2019 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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15368428 |
Dec 2, 2016 |
|
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16656529 |
|
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|
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62263407 |
Dec 4, 2015 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F05D 2220/32 20130101;
F23R 3/52 20130101; F02K 1/002 20130101; F23R 3/007 20130101; F05D
2260/211 20130101; F02K 3/10 20130101; F05D 2300/6033 20130101;
F02K 1/36 20130101; F05D 2260/4023 20130101; Y02T 50/60 20130101;
F23R 3/32 20130101; F02C 7/143 20130101; F02C 7/10 20130101; F02C
3/04 20130101; F02C 6/12 20130101 |
International
Class: |
F02C 7/143 20060101
F02C007/143; F02C 3/04 20060101 F02C003/04; F02C 7/10 20060101
F02C007/10; F02C 6/12 20060101 F02C006/12; F23R 3/32 20060101
F23R003/32; F02K 1/00 20060101 F02K001/00; F02K 1/36 20060101
F02K001/36; F02K 3/10 20060101 F02K003/10; F23R 3/00 20060101
F23R003/00; F23R 3/52 20060101 F23R003/52 |
Claims
1. A propulsion system, comprising: a first compressor in fluid
communication with a fluid source; a first conduit coupled to the
first compressor; a heat exchanger in fluid communication with the
first compressor via the first conduit; a second conduit positioned
proximal to the heat exchanger; a combustor in fluid communication
with the heat exchanger via the second conduit and configured to
generate a high-temperature gas stream; a third conduit coupled to
the combustor; and a first thrust augmentation device in fluid
communication with the combustor via the third conduit, the heat
exchanger being positioned within the gas stream generated by the
combustor.
2. The propulsion system of claim 1, wherein the heat exchanger is
disposed within the third conduit.
3. The propulsion system of claim 1, further comprising a turbine
coupled to the first compressor and positioned between the
combustor and the heat exchanger.
4. The propulsion system of claim 3, wherein the turbine comprises
ceramic matrix composites.
5. The propulsion system of claim 3, further comprising a second
compressor coupled to the turbine.
6. The propulsion system of claim 5, wherein the turbine is fixedly
coupled to the first compressor and is coupled to the second
compressor via a clutch.
7. The propulsion system of claim 1, further comprising a swiveling
connector coupling the third conduit to the first thrust
augmentation device.
8. The propulsion system of claim 1, further comprising a second
thrust augmentation device in fluid communication with the
combustor.
9. A combustor, comprising: a first toroidal casing circumscribing
an axis and having an inlet configured to receive fluid, the first
casing defining a first internal chamber in fluid communication
with the inlet; a second toroidal casing disposed within the first
internal chamber and circumscribing the axis, the second casing
having an outer wall defining a second internal chamber, the outer
wall having a plurality of orifices formed therethrough, the
orifices providing fluid communication between the first and second
chambers; a plurality of fuel injectors positioned to inject fuel
into the second chamber through the orifices; and outlet structure
defining at least one channel in fluid communication with the
second chamber, the at least one channel being oriented parallel to
the axis.
10. The combustor of claim 9, wherein the orifices are oriented at
an oblique angle with respect to the outer wall.
11. The combustor of claim 9, further comprising an ignition source
positioned within the second chamber.
12. The combustor of claim 9, wherein the outlet structure
comprises an inner wall converging toward the outer wall and is
configured to urge high-temperature fluid flowing about the axis
within the second chamber through the at least one channel.
13. The combustor of claim 9, further comprising a plurality of
funnel elements disposed within the orifices and extending into the
second chamber, the funnel elements tapering from the first chamber
to the second chamber.
14. The combustor of claim 9, wherein the injectors extending into
the second chamber.
15. The combustor of claim 9, wherein the second casing comprises
ceramic matrix composites.
16. The combustor of claim 9, further comprising an air source in
fluid communication with the inlet, the air source being heated by
fluid emitted by the second chamber through the at least one
channel.
Description
PRIORITY CLAIM
[0001] This application is a continuation of U.S. application Ser.
No. 15/368,428 filed on Dec. 2, 2016; which claims priority from
U.S. Provisional Patent Application No. 62/263,407 filed on Dec. 4,
2015, the above-referenced applications are hereby incorporated by
reference as if fully set forth herein.
COPYRIGHT NOTICE
[0002] This disclosure is protected under United States and/or
International Copyright Laws. .COPYRGT.2016-2019 Jetoptera, Inc.
All Rights Reserved. A portion of the disclosure of this patent
document contains material which is subject to copyright
protection. The copyright owner has no objection to the facsimile
reproduction by anyone of the patent document or the patent
disclosure, as it appears in the Patent and/or Trademark Office
patent file or records, but otherwise reserves all copyrights
whatsoever.
BACKGROUND
[0003] Micro-turbines have become increasingly popular for aviation
propulsion. FIG. 1, an illustration of the current dominant design
of micro-turbines as disclosed by Thomas Kamps, "Model Jet
Engines," Third Ed., ISBN 978-1-900371-93-3, shows a centrifugal
compressor with a case (21) and rotor (22); a diffuser (23);
bearings and lubrication (25); a shaft (26) connecting the
compressor wheel with the axial turbine (32); an annular, reverse
flow combustor containing an outer liner (28) and an inner liner
(29) contained inside a casing (35); a nozzle (31) to accelerate
the combustion gases and direct them to the turbine (32); and an
exhaust nozzle (34) to accelerate the exhaust and generate the
thrust via a jet. FIG. 1 illustrates the design of virtually all
hobby and small jet engines in the range under 100 lbf, albeit the
same design is used also for up to 1000 lbf of thrust.
[0004] While engineers are implementing sophisticated and high-cost
technologies to enable large jet engines to maximize their
efficiency, present day micro-jet engines continue to lack such
technology. For example, micro-jet engines are life limiting
because there is a lack of turbine cooling air flow. Moreover, the
size of the micro-jet engines requires them to spin at very high
speeds--typically well over 100,000 Rotations Per Minute (RPM), but
some approaching and exceeding 150,000 RPM. As a result of
less-sophisticated technology, micro-jet engines cannot achieve
demanding thermodynamic cycles that involve high firing
temperatures and pressure ratios. The fuel consumption of micro-jet
engines typically exceeds 1.5 lb fuel per hour and lbf of thrust,
as compared to large jet engines of the high bypass type which can
have fuel consumption as low as 0.5 lbs of fuel per hr and lbf of
thrust.
BRIEF DESCRIPTION OF THE DRAWINGS
[0005] FIG. 1 illustrates in cross-section a conventional
micro-turbine used in aviation propulsion.
[0006] FIG. 2 illustrates in rear cross-section an embodiment of
the present invention.
[0007] FIG. 3 illustrates in elevated side cross-section the
embodiment illustrated in FIG. 2.
[0008] FIG. 4 illustrates in side cross-section the embodiment
illustrated in FIG. 2.
[0009] FIG. 5 illustrates schematically one embodiment of a
propulsion device oriented in cruise position.
[0010] FIG. 6 illustrates an embodiment of the present invention
with a helicoidal heat exchanger, two stages of centrifugal
compressors, a toroidal combustor and a Coanda ejector.
[0011] FIG. 7 illustrates a Coanda nozzle propulsion system with
augmenter.
[0012] FIG. 8 illustrates the thermodynamic cycle of a conventional
mini-turbojet.
[0013] FIG. 9 illustrates a conventional turbojet.
[0014] FIG. 10 illustrates the thermodynamic cycle when modified
with a regenerative heat exchanger and ejector.
[0015] FIG. 11 illustrates the heat exchanger streams.
[0016] FIG. 12 illustrates the ratio of augmentation obtained
through experimental data using the augmentation devices disclosed
in this application as compared to the pressure ratio of the
exhaust gas supplied to the plenum and ambient pressure.
[0017] FIG. 13 illustrates in elevated side cross-section the
embodiment illustrated in FIG. 2.
[0018] FIG. 14 illustrates in elevated side cross-section the
embodiment illustrated in FIG. 2.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
[0019] This application is intended to describe one or more
embodiments of the present invention. It is to be understood that
the use of absolute terms, such as "must," "will," and the like, as
well as specific quantities, is to be construed as being applicable
to one or more of such embodiments, but not necessarily to all such
embodiments. As such, embodiments of the invention may omit, or
include a modification of, one or more features or functionalities
described in the context of such absolute terms. In addition, the
headings in this application are for reference purposes only and
shall not in any way affect the meaning or interpretation of the
present invention.
[0020] One or more embodiments provide a propulsion system that
includes some or all of the following features:
[0021] A gas generator defining at least one compressor, a
combustion chamber, a turbine and a thrust augmentation device;
[0022] A first compressor defining an intake opening, at least one
bleed port provided with a valve in connection with a fluid
pressurized network of conduits, and one outlet to a secondary
compressor or a combustion chamber or both via at least a volute
and a compressor discharge conduit;
[0023] A turbine connected to at least the first compressor in a
fixed manner and to a second compressor in a fixed manner or via a
clutch;
[0024] A toroidal combustor that receives the air received from the
compressor discharge conduit after it is preheated in a heat
exchanger and gradually introduces the preheated air tangentially
to the main axis of the gas generator into a sleeve formed by a
scrolled casing and a mostly toroidal shaped liner via
prevaporizing mixers and generating an overall circumferential
movement of the air and combustion products inside a liner;
[0025] A plurality of pre-vaporization mixers of fuel and air
distributed around the circumference of the combustor scroll and
receiving the fuel from fuel conduits and mixing the fuel and air
in Venturi passages inside the mixers;
[0026] A plurality of rectangular slots in communication with the
fuel injectors and the pre-vaporization mixers that accelerate the
mixtures of fuel and air and inject them in a tangential manner
towards the diametrical centerline of the toroidal liner;
[0027] A converging channel in communication with the toroidal
liner and guiding the hot gases from a mainly
circumferential/angular reacting flow direction to mostly axial
direction and collinear to the main axial direction of the said gas
generator;
[0028] The turbine receives a stream of hot gases from the
combustor in the mainly axial direction and expands the gases while
extracting power to drive the compressors;
[0029] A heat exchanger that receives the hot gases from the
turbine and preheats the compressor discharge air, delivering it to
the combustor, and guides the cooled hot gases towards a swiveling
joint;
[0030] A swivel connector that transmits the pressurized hot gases
supplied by the combustor outlet towards a thrust augmenting device
as motive gas;
[0031] A thrust augmenting device containing a mixing section, a
throat section and a diffusor, and receiving the pressurized gases
to use as a motive gas to generate thrust by fluidically entraining
ambient air, mixing it with the motive air and ejecting it at high
velocities via the diffuser; and
[0032] A series of thrust augmenting devices each containing a
mixing section, a throat section and a diffusor, whereby they
receive the compressed air from the compressor via the bleed valve
and the fluid network and use the pressurized air as motive gas to
generate thrust by fluidically entraining ambient air, mixing it
with the motive air and ejecting it at high velocities via the
diffusors;
[0033] The first compressor is coupled to the shaft, second
compressor and turbine via the clutch and its output air is
directed to the thrust augmenting devices at take-off, hovering and
landing mission points and decoupled at cruise conditions;
[0034] A single compressor supplies with air both the combustor via
the heat exchanger and the thrust augmenting devices via a fluid
network and the compressor bleed valve to generate thrust in more
than one location of the propulsion system;
[0035] The liner and the turbine consist of Ceramic Matrix
Composites; and
[0036] The fluid network is in communication with the bleed valve
and can modulate the flow to multiple thrust augmentation devices
to assist the attitude control of the aircraft powered by the
propulsion system.
[0037] A method of flying an aircraft or hovercraft may
include:
[0038] Accelerating the gas generator to maximum power with open
compressor bleed valves supplying several thrust augmenting devices
and balancing the attitude of the aircraft by closing and opening
control valves distributing said compressed air to the thrust
augmenting devices and for vertical hovering, take-off and
landing;
[0039] Supplying the remaining gas generator core flow exiting said
heat exchanger to a thrust augmenting device as motive fluid to
balance the attitude of the aircraft and generate the thrust used
to vertically take off, hover or vertically land;
[0040] Balancing the gas generator speed and power with the bleed
proportion of the compressor and fluid network actuation of the
control valves to control the attitude of an aircraft, hovercraft
or any other flying device using said gas generator;
[0041] Accelerating or decelerating the gas generator to produce
more or less flow to the thrust augmentors supplied with compressed
air from the compressor bleed and hot exhaust gas from the
turbine;
[0042] Opening or closing the compressor bleed valve to supply or
block a portion of the compressed air to the thrust augmenting
devices in communication with the fluid network;
[0043] Opening or closing control valves that distribute the
compressed air to thrust augmenting devices to control roll, yaw
and pitch; and
[0044] The ejectors contain one or more fuel injection nozzles for
augmentation of thrust during short periods of time.
[0045] Embodiments of the present invention disclosed in this
application relate to a micro-turbine (also known as a micro jet
engine) that specifically operates as a gas generator. Rather than
seeking to maximize thrust by accelerating a mass of gas to the
highest velocity possible like a typical jet engine, the preferred
embodiment of the present invention produces several streams of
pressurized, hot gases into ejectors and creates force used in all
phases of flight. In one embodiment of the present invention, a new
regenerative cycle and components thereof are disclosed in this
application, such as (i) a novel compressor and/or several stages
of compressors that may or may not be coupled mechanically or via a
clutch to the main shaft; (ii) a novel combustion system, utilizing
heated air to minimize the fuel requirements to meet the
specifications; and (iii) novel materials to maximize the
performance, such as use of Ceramic Matrix Composites (CMC) in the
components.
[0046] In another embodiment of the present invention, the turbine
stage is designed to extract power without expanding the flow to
close to atmospheric pressure, as would be the case in typical
turbojets. Instead, the flow expands to pressures higher than those
pressures typical of turbojet nozzle entry. The pressure at the
exit of the turbine stage is hence higher than for typical
turbojets--and purposefully so--to use in a Coanda type ejector as
motive fluid. In other embodiments of the present invention, the
disclosed technology allows for eliminating certain components
altogether, such as the throttling nozzle to the turbine stage or a
propulsive nozzle for accelerating the hot gases.
[0047] The novel gas generator is designed from the principles of
centrifugal (compressor) and axial (compressor and turbine)
turbomachinery. The thermodynamic cycle is of the regenerative
type, with the compressor discharge air being routed to the
combustor via a heat exchanger placed in the exhaust area from the
turbine stage and before the exit flange of the gas generator.
[0048] In an embodiment of the present invention, several stages of
compression may be applied to the air entering the system using,
for example, a clutch to engage a first compressor at take-off and
landing stages or when hovering the aircraft. The air compressed by
the first compressor may be routed to ejectors and/or may be used
for other purposes, including being directed into the intake of the
secondary nozzle or used for cooling, augmentation of thrust, cabin
pressurization, or other uses. As with typical turbocharger
compressors, the first compressor may have at peak operation a
pressure ratio preferably 2.5 or more. A valve may be present on
the compressor discharge volute to direct the compressed air to
either the secondary compressor or outside the gas generator, as
need may be.
[0049] The second compressor may use its own air intake or may
ingest a portion or all of the air from first stage compressor.
This second stage compressor, similar to the first one, may employ
a pressure ratio of at least 3, but preferably 5 or more. Hence, at
take-off, landing or hovering, the overall pressure ratio may
exceed 7.5:1.
[0050] In one embodiment, the compressor is connected mechanically
to an axial turbine, and they spin at the same rate on the same
shaft.
[0051] The compressed air outlet from the second compressor is
routed to the back of the engine via insulated conduits at
appropriate velocities, as will be described in greater detail with
reference to FIG. 6, and towards a heat exchanger placed at the
exit of the gas generator inside the exhaust duct. This results in
increased efficiency because the heat exchanger picks up heat from
the hot gas exhaust exiting the gas generator at more than two
atmospheres pressure and transferring that heat to the air supplied
to the combustor (in other words, preheating the air going into the
combustor). The heat exchanger itself may be compact, utilizing
spirals and manifolds to increase the surface area and the
residence time of the compressed air so that a significant heat
pickup can occur. Temperatures greater than 1000 F can be obtained
before combustion. The heating of the air is advantageous because
it reduces the fuel consumption of the cycle, and hence the fuel
burn, by at least 30%.
[0052] Moreover, because the turbine is not cooled, current CMC
materials allow for temperatures of about 2000 F (1750 F for
metals) to enter the turbine (Turbine Entry Temperature or TET).
Thus, in an example where there is a 1 lb/s flow, with a Pressure
Ratio (PR) of 4 and a metal nozzle (1750 F maximum TET), the
combustor would need to add approximately 59% of the fuel to reach
the TET regenerated from 1000 F to reach 1750 F if the present
invention were utilized as compared to a situation where there is
no heat recovery. In addition, if better materials are used, such
as a turbine manufactured with CMC materials that can tolerate 2000
F TET, the fuel efficiency can further increase another 35%. Table
1 shows the comparison of the fuel to air ratios needed to reach
the same firing temperature TET in the both cases. Table 2
exemplifies a case where the heat exchanger may be fitted on a 100
lbf thrust system.
TABLE-US-00001 TABLE 1 to the left is the metallic version of a gas
generator that fires the turbine at 1750 F, to the right is the CMC
version firing to 2000 F. More than 1/3 of the fuel can be saved in
this manner. Non-Regenerated Regenerated Non-Regenerated
Regenerated T inlet Combustor [F] 430 816 T inlet Combustor [ 430
816 TET [F] 1750 1750 TET [F] 2000 2000 Pressure [psia] 60 60
Pressure [psia] 60 60 Fuel-to-Air Ratio 0.0198 0.0143 Fuel-to-Air
Ratio 0.0241 0.0186 Savings 27.7% Savings 23.0%
TABLE-US-00002 TABLE 2 to the left is the metallic version of a gas
generator that fires the turbine at 1750 F, to the right is the CMC
version firing to 2000 F, as modelled with a heat exchanger that
fits on a 100 lbf engine. Non-Regenerated Regenerated
Non-Regenerated Regenerated T inlet Combustor [F] 430 1000 T inlet
Combustor [ 430 1000 TET [F] 1750 1750 TET [F] 2000 2000 Pressure
[psia] 60 60 Pressure [psia] 60 60 Fuel-to-Air Ratio 0.0198 0.0116
Fuel-to-Air Ratio 0.0241 0.0158 Savings 41.3% Savings 34.3%
[0053] For a 1 lb/s air flow gas generator operating at 60 psia
(low pressure) and firing at 1750 F, the savings per hour of flight
would be significant. When used with existing gas turbines, the
fuel consumption can drop from approximately 1.5 lb fuel per hour
per lbf of thrust to about 0.87 lb fuel per hour per lbf of thrust.
Particularly in the case of a CMC turbine, similar fuel savings can
result in levels under 1 lb of fuel per hour per lbf of thrust.
These efficiencies would allow a vehicle to fly longer in range and
duration, or faster, or both, for the same payload. Older
generation jet fuel powered turbofans such as low-bypass turbofans
exhibit similar levels of fuel burn of 0.8 lb fuel per hour per lbf
of thrust, lower than typical turbojets, but higher than high
by-pass turbofans. This means that the present invention allows the
smaller gas turbines, typically turbojets, to perform similarly to
low-by-pass turbofans with respect to regeneration of the exhaust
gases and introduction of better materials and tolerances.
[0054] In another embodiment, one compressor may be employed,
provided that there is at least one bleed port that supplies the
air used for vertical take-off, hovering, vertical landing and
other maneuvers required by the mission. The bleed port can also
provide the air used by the combustor and turbine. The bleed may
also be extended during cruise flight for various reasons. Bleeds
of up to 15% are common in large aircraft engines and a
specifically designed compressor may benefit from operability with
bleeds, albeit lowering its performance. With bleeds employed only
in limited portions of the flight, and the bleed valve closed for
most of the mission, the compressor's performance may be acceptable
for unmanned vehicles as well as light airplanes.
[0055] The preheated air, typically at pressures exceeding 50 psia
and over 1000 F is then directed to the combustor. The combustor in
the present invention is of tangential type, with the entry of the
hot, preheated air via a volute, and it is designed with
significantly increased combustion residence time. The current
designs in the prior art only allow for very short residence times
to occur, which results in flames coming out of the turbine very
often due to the combination of the short residence time and with
the small space in which the combustor is crammed. The amount of
time that the exothermal reaction of combustion is limited to and
the efficiency of the combustion process is far less than the
typical 99.5% or more on the current large turbofan combustors, for
example. With fuel not completely converted in products, and
therefore, with products of incomplete combustion exiting the gas
turbine, the efficiencies are remarkably low. This is another
contributing factor to the low efficiencies encountered in current
micro-turbines.
[0056] On the contrary, the present invention disclosed in this
application allows for a very generous time for combustion via
increased residence time due to the large volume of the combustion
chamber itself, as well as the tangential and swirling around the
combustor pattern formed by its design. While a residence time of
10 milliseconds is typical in such prior art small engines, it is
not enough to ensure complete the combustion process. A residence
time of over 20 milliseconds is possible via a geometry which
induces a combustion process in a toroidal fashion recirculating
around the axis of the combustor. With this approach, several
advantages are achieved.
[0057] The fuel injection is done in a pre-vaporized manner through
tubes that also employ co-flow of air so that it moves fast and
rich enough locally to delay the auto-ignition. About 20% of the
total air from the compressor (after bleed) is passed through the
premixing/pre-vaporized tubes. At the end of each tube, a rich
mixture of fuel and air emerges into the main combustion torus in a
co-flow arrangement without the aid of a swirling flow. These fuel
supply tubes provide long mixing lengths in order to vaporize the
fuel and inject it optionally advantageously as a gas mixed with
little air (a fuel-rich/air mixture).
[0058] The emerging fuel and air mixture then joins the general
toroidal recirculation pattern formed in the combustor and assists
with the projection of the flowfield in a circumferential
direction. Secondary air enters tangentially and at staggered
locations (in between the fueled injectors) to maintain the
recirculation and to provide the hot air required for combustion
and re-ignition of the fresh mixture of fuel and air, in a similar
fashion as explained in papers by Kalb et al. and hereby
incorporated by reference (Bruckner-Kalb, J., Krosser, M., Hirsch,
C., Sattelmayer, T., Emission characteristics of a premixed
Cyclic-Periodical-Mixing Combustor operated With hydrogen-natural
gas fuel mixtures, Journal of Engineering for Gas Turbines and
Power, Vol. 132, No. 2, pages 021505, 2010). The orifices are
preferably non-circular in nature, mostly slots emerging at an
angle to the main axis of the combustor, and may be staggered both
axially and circumferentially, as well as staged, including shut
off when supplied via two fuel line manifolds, one for low and both
for high power operation. Since the overall fuel-to-air ratio of a
recuperated, all metallic non-cooled turbine is in the range of
0.010-0.015 (see Table 1, in which case it is 0.0116), the fuel and
air mixture in the pre-vaporized tubes can be about 0.05-0.075
(near stoichiometric to fuel-rich, as the stoichiometric FAR for
jet fuel is typically about 0.068). The auto-ignition at the low
pressures of small engines (such as the present invention) is in
the range of 15 milliseconds or more, assuming a pressure of about
60 psia and a preheat temperature of 1000 F. (See Vasu, S. S.,
Davidson, D. F., and Hanson, R. K., "Jet fuel ignition delay times:
Shock tube experiments over wide conditions and surrogate model
predictions," Combust. Flame 152, 125-143 (2008)). A residence time
of less than 7 milliseconds is achievable in these premixing
elements, which can be tubes or non-circular shapes, which are
preferable, as shown in FIGS. 2-4. The mixing of the fuel and air
in these premixing tubes does not involve any swirling movement,
being purely unidirectional, and emerges in a tangential direction
to the circumference of the main reaction zone of the combustion
chamber at a velocity not less than 80 feet per second at nominal
speed.
[0059] FIG. 2 shows a combustor 311 of an embodiment as seen from
an aft position looking forward. Combustor 311 includes a first
toroidal casing 302 circumscribing an engine shaft 901 and having
an inlet 301 configured to receive fluid. The first casing 302
defines a first internal chamber 304 in fluid communication with
the inlet 301.
[0060] Combustor 311 further includes a second toroidal casing 303,
which may be made of ceramic matrix composites, disposed within the
first internal chamber 304 and also circumscribing the shaft 901.
The second casing 303 has an outer wall 306 defining a second
internal chamber 315. The outer wall 306 has a plurality of
orifices (discussed in greater detail with reference to FIG. 3)
formed therethrough that provide fluid communication between the
first and second chambers 304, 315. In an embodiment, the orifices
are oriented at an oblique angle with respect to the outer wall
306. A plurality of fuel injectors 310 are positioned to inject
fuel into the second chamber 315 through the orifices. In an
embodiment, the injectors extend into the second chamber 315.
Additionally, an ignition source 335 may be positioned within the
second chamber 315.
[0061] Combustor 311 further includes outlet structure defining at
least one channel 350 (FIG. 4) in fluid communication with the
second chamber 315. The channel 350 is oriented parallel to the
shaft 901. The outlet structure comprises an inner wall 307 (FIG.
4) converging toward the outer wall 306, which is configured to
urge high-temperature fluid flowing about the axis of the shaft 901
within the second chamber 315 through the channel 350. As will be
described in greater detail hereinafter, inlet 301 receives heated
air from a heat exchanger that is heated by fluid emitted by the
second chamber 315 through the channel 350.
[0062] Air, indicated by arrows 311, flowing in the first chamber
304 cools the second casing 303, and a portion of air 311,
indicated by arrows 312, is introduced in a circumferential manner
into the second chamber 315 via several fuel and air mixers 305
distributed around the outer wall 306. First casing 302 essentially
serves as a pressure vessel and is mechanically attached in the
front to a compressor casing and in the back to a turbine case. The
entry and flow of the air into second chamber 315 is
circumferential to the main axis of the engine, which coincides
with the engine shaft 901.
[0063] The reacting mixture 312 of air and fuel flow scrubs the
second casing 303 in a circumferential manner, with fresh supplies
of air and fuel mixtures introduced from the first chamber 304 and
injectors 310 at various circumferential locations. The injectors
310 may also preheat the fuel by immersing them in the first
chamber 304 and closer to the second casing 303. The fuel picks up
the heat and vaporizes before being delivered to the fuel and air
mixers 305. After combusting in the second chamber 315 for several
tens of milliseconds, the accelerated and completely burned gas
exits the second casing 303 into a turbine, with or without the use
of a nozzle, at a pre-determined angle of incidence.
[0064] FIG. 3 depicts the detailed mechanism of the introduction of
the fresh, preheated air into, and structure of, the mixers 305.
Slots 326 form an angle to the vertical or horizontal axes
coordinate of the shaft 901 and may receive the air through a
"scooping" function by a funnel 327. The air 311 scrubbing the
outside of the second casing 303 is gradually admitted in small,
portioned quantities 331, via the funnels 327 into the
prevaporizing/mixing slots 326. Funnels 327 are disposed within the
slots 326 and may extend into the second chamber 315. Additionally,
the funnels 327 may taper from the first chamber 304 to the second
chamber 315. The slots 326 are designed such that a diffusing
section commencing at the fuel injection plane supports the rapid
mixing of fuel delivered via an injector 310. The fuel has already
been subjected to immersion in preheated air, hence is nearly fully
prevaporized, in essence behaving like a gas being injected from a
single of multiple sources at the throat section of the slot 326,
rapidly mixing with the fuel to form a fuel-rich, hot mixture with
the air.
[0065] The slot 326 is designed so that the residence time of the
fuel and air mixes inside it before being supplied to the second
chamber 315. The residence time is typically less than 5
milliseconds or less, allowing for no auto-ignition of the fuel
inside the mixer 305 to occur. Moreover, the high velocity inside
the mixer 305 prevents flashback to occur in these small passages.
As the air is being supplied to the second chamber 315 and induced
into a circumferential flow around the first chamber 304, it is
cooling the first chamber as well as picking up heat, making the
air at the end of the 360-degree complete revolution and final
admittance into the last funnel 327 hotter than at the first
funnel.
[0066] The funnels 327 may be fine-tuned and adjusted accordingly
to ensure uniformity of supply to the second chamber 315 and
temperature uniformity within the second chamber. The fuel flow
rate supplied to each of the mixers 305 around the circumference,
however, is not constant, but changes slightly to ensure uniformity
and smooth operation. The ignition mechanism of the fresh mixture
of air and fuel admitted from each mixer 305 into the second
chamber 315 is via high temperature products of the previous,
immediately adjacent mixer. As such, once ignited via a retractable
or detachable ignition source 335, such as a glowing plug, the
system becomes stable in the reaction down to extremely low flame
temperatures, as low as 2000 F or less.
[0067] The air supplied to the combustion chamber is mainly split
into the scrubbing and cooling of the second casing 303 via flow in
the first chamber 304 and combustion air 331 supplied to the mixing
and pre-vaporizing mixers 305. The total combustion air is hence
about 60-70% of the total air supplied to the combustor 311,
whereas a remaining portion is introduced in the converging section
of the second casing 303, where the reacting flow is turned axially
and into the turbine. The introduction of this dilution air cools
the gases before entry into the turbine and to help guide the gases
to a mainly axial direction, albeit with a significant residual
circumferential orientation, into the rotating stage of the
turbine.
[0068] The ignition of a fresh, rich mixture is propagated around
the main combustion torus and it provides a reduced scrubbing of
the walls if additional air, about 60% of the compressor discharge
post-bleed, is injected though the inner and outer walls of the
torus, at an angle that is mainly same circumferential direction
with the reactive flow inside the torus, forming a protective film
for the walls. The inside of the torus is then a uniform, reacting
flow, able to stabilize at very low temperatures, which is exactly
what is needed for turndown and avoidance of the lean blowout
point. Moreover, the high temperature of the preheated air from the
heat exchanger and delivered to the combustor is hot enough to keep
the flame going while also cold enough to make the material
withstand the loads. An external heat transfer factor is ensured by
maintaining a small gap between the second casing 303 and the first
casing 302 of the combustor, maintained by spacers, such that the
air cooling the walls of the second casing 303 on the cold side is
picking up heat and carrying it to the dilution holes located near
the exit of the second casing 303. If the second casing 303 is
metallic, then a film cooling may be advantageous, but if the
second casing 303 is made out of CMCs, then the external cooling
may be reduced significantly. CMCs are particularly very strong
materials and a single shaped second casing 303 will be able to
subsist in these conditions for at least 15000 cycles.
[0069] The use of the pre-vaporized fuel in the premixing tubes is
in effect mimicking the use of a gaseous fuel like propane, which
may be used to start up the engine, as an alternative option.
Ignition is assisted by sparking plugs or pilot flames until the
propane combustion process is stable and the engine is at idle and
no longer assisted by a starter motor. Portable propane bottles are
available, and a control logic is implemented to replace the
propane (or other suitable gas) after ignition and thermal
stabilization of the heat exchanger and engine. After a few minutes
of operating with the gaseous fuel, and once the cycle is
recovering the heat and preheating the combustion inlet air to
acceptable levels, the gaseous fuel is replaced gradually by liquid
fuel such as jet, Diesel, etc. The operation continues until the
gaseous fuel is completely replaced by the liquid fuel, and the
gaseous fuel source can be detached from the engine. Similarly,
upon shut-down, any residue of liquid fuel can be oxidized in the
fuel injection tubes due to the high temperatures.
[0070] The fuel turndown, in order to control the engine, can be
achieved via two differently sized manifolds with each feeding an
odd number of injectors. For example, with the gaseous fuel
injected into one manifold for ignition on propane alone and
feeding three injectors, the second manifold starts injecting
liquid fuel via opening of a solenoid, while the propane supply is
reduced to zero. The heat addition to the engine in this phase
follows a constant value while balancing the reduction of the
gaseous fuel with the increase in liquid fuel, until the gaseous
fuel is completely replaced. At that point, the engine is at idle.
The same liquid fuel circuit increases for acceleration, and a
second transition occurs with the primary liquid fuel circuit
reducing the fuel flow while the solenoid valve of the secondary
liquid fuel manifold starts supplying the fuel that makes up for
the acceleration curve. Above the idle point (low power) and
maximum power (take-off, hovering and landing), both liquid fuel
manifolds supply liquid fuel to the combustor at all times, and one
of them is used to stage down as required by the mission.
Additional operations with reducing flow to both circuits are also
possible. The recuperator provides a constant, high temperature
supply to the combustor inlet, enough to provide a stabilized
operation and not become subject to combustion operability issues,
including avoidance of lean blow out. This type of stabilization of
the flame is known in the art, and turndown can be achieved to
significant low levels due to the high thermal inertia of the
system (high recuperated combustor inlet temperature throughout
operation, highly uniform temperature reaction zone of the
combustion process inside the toroidal liner, hot walls of the
liner all contribute to a stabilized operation).
[0071] In FIG. 13, the admission of air from inlet 301 and mixture
with fuel vapors is performed in mixers 305. The fresh mixture of
fuel and air is injected at an angle into the second chamber 315,
driving a circumferential overall reacting flow around shaft 901,
which is also the main axis of the gas generator, constantly
supplied with fresh air and fuel mixtures 903. Portions of the
first chamber 304 air are introduced into the reaction zone until
the entire combustion air is admitted to the combustion process.
The residence time is increased due to the volume of the combustion
zone and the stability of the combustion process is ensured by the
ignition of fresh mixtures 903 exposed to hot gases from previous,
adjacent mixer 305. The first casing 302 is designed so that a high
velocity of the preheated air is maintained to cool the liner.
Mixers 305 also contain fairings to maintain a low recirculation
zone behind the mixers and provide lower pressure drop.
[0072] As illustrated in FIG. 14, the second casing 303 is located
inside the first casing 302 and the fuel injectors 310 are immersed
inside the sleeve formed by the second casing 303 and the first
casing 302 so that fuel is picking up heat and is nearly
prevaporized by the time it enters pre-vaporizing mixers 305. The
second casing 303 can, due to its small size, be installed inside
the first casing 302 before the final welds of the first casing 302
are performed to trap the second casing 303 inside. In addition,
the second casing 303 may be manufactured out of CMC.
[0073] FIG. 5 illustrates one embodiment of a propulsion device
including a thrust augmentation device 500 in cruise position.
Through methods know in the art, the thrust augmenting device 500
can be swiveled at least 100 degrees around the axis perpendicular
to an axis passing through the length of the gas generator. The
propulsion device consists of a compressor shroud 801, a shaft 806,
and a compressor rotor 802 discharging the air in a volute 803 and
directing the flow, via a conduit 820, towards a compact heat
exchanger 830. The heat exchanger 830, which may be of compact,
helically coiled type, receives and preheats the compressor
discharge air from inlet 825 and guides the preheated air to the
heat exchanger outlet 812. The preheated air is then further guided
to a combustor inlet 301 where the air is directed in a
circumferential flow around the main axis of the combustor. The air
is then combusted within the second casing 303 with fresh mixtures
of air and fuel being supplied in a tangential direction around the
circumference via mixers 305 with fuel supplied from manifold
810.
[0074] The system may or may not contain a first stage nozzle 811.
In one embodiment, the nozzle is eliminated and the convergent
channel guides the gases towards the turbine rotor 812, carrying
some residual circumferential component of the velocity. After
expansion to a lower pressure, the exhaust gas at the exit of the
turbine is guided as hot gas flow toward a conduit 814 and over the
compact heat exchanger 830 to preheat the compressor discharge air.
The guiding, swiveling conduit 814 directs the gases while under
pressure to the plenum 501 of the device 500, where the exhaust gas
is used as motive fluid to generate the thrust augmentation in the
direction of flight 509. Thrust bearings and their auxiliary system
are represented by 805.
[0075] FIG. 6 illustrates a propulsion system 1000 according to an
alternative embodiment. The system contains multiple stages of
compressors, in which the first compressor 1001 may supply air to
thrust augmenting ejectors at take-off and at various stages of
flight, but may be disconnected via a clutch from a main shaft
during the rest of the mission. The second compressor 1002
compresses the air and then directs it via a conduit 1005 and
flange 1006 to a heat exchanger 1050 positioned in the exhaust area
within a conduit 1009 and intermediate a turbine 1030 and at least
one thrust augmenting ejector 1100. In an embodiment, the turbine
is manufactured from ceramic matrix composites. The heat exchanger
1050 uses the exhaust heat of the fluid provided by a combustor
1020 and exiting the turbine 1030 to increase the temperature of
the air discharged by compressor 1002 and supplied by the conduit
1005 to the heat exchanger.
[0076] The heated air then leaves the heat exchanger 1050 and
conduit 1009 via a flange 1007 and a conduit 1008 to the combustor
1020. In an embodiment, the heat exchanger is of helicoidal type.
In yet another embodiment, at cruise conditions, the exhaust gas
leaves the turbine 1030 at 27 psi and 1400 F, and, after
transferring the heat to the colder compressor, discharge air
supplied by conduit 1005 drops to 25 psi and 800 F. In this
embodiment, the heat exchanger 1050 delivers the preheated air to
the combustor 1020 via flange 1007 and conduit 1008, then boosts
the temperature of the compressor discharge air flow at 60 psi, 400
F and supplied by conduit 1005 to the heat exchanger 1050 to at
least 500 F and a pressure of 58 psi. In this way, the fuel
consumption is decreased by more than 7% and possibly over 20%
depending on the type and performance of the heat exchanger (see
Table 3 below).
TABLE-US-00003 TABLE 3 Example of a 75 lbf class propulsion system
employing the arrangement of FIG. 6 with various types of heat
exchangers and a one-hour cruise condition flight. For extended
flight times, the additional weight of the heat exchanger 1050 is
hence justified, as the negative effect of additional weight is
balanced by the benefit in fuel savings and cost of operation. Heat
Heat Heat no Heat Exchanger Exchanger Exchanger Exchanger 1 2 3
Pressure [psi] 58 58 58 58 Heat Recovered 400 500 600 700 Combustor
Inlet Temperature [F.] Firing 1750 1750 1750 1750 Temperature
uncooled Turbine [F.] Fuel-to-Air 0.0202 0.0188 0.0174 0.0160 Ratio
Fuel Savings 0.0000 5.0 10.1 15.2 per hour at cruise for a 75 lbf
powerplant [lbs] Fuel Savings 0.0000 6.90% 13.87% 20.91% per hour
at cruise for a 75 lbf powerplant [%]
[0077] In one embodiment of the present invention, no nozzle is
used to guide the hot gases into the turbine. In another
embodiment, a nozzle may be utilized for minimal re-directing of
the gases. This is particularly different from most of the
conventional gas turbine systems, where the gases typically need to
be turned and accelerated significantly into the first stage
turbine due to the uncoordinated axial recirculation processes
(stirring, nested recirculations in the axial direction) that occur
in combustors with premixers using swirlers.
[0078] In yet another embodiment of the present invention, the use
of swirlers is completely eliminated from the combustor for either
stabilization of the flame or mixing.
[0079] In still another embodiment of the present invention, the
combustion chamber offers a much larger residence time, different
from the aviation practices utilized today, and more closely
resembles the gas turbine frame combustors utilized for, for
example, power generation. What typically precludes the large
volume combustors to be used in aviation applications is the need
of compactness and weight restrictions. The residence time in
aviation applications is nearly one order of magnitude shorter than
of those from power generation frames, due to the high velocities
and short lengths requirements. In an embodiment, the flow of the
reacting flow in a circumferential mode before the re-direction in
axial direction is optionally advantageous to being able to
increase the volume significantly, hence allowing for completeness
of reaction and high efficiencies.
[0080] Furthermore, the lack of a first stage nozzle vane(s)
eliminates weight while reducing the heat transfer requirements,
while still maintaining a geometry that favors the acceleration and
introduction at an angle into the rotating stage of the
turbine.
[0081] The cycle and engine disclosed in this application can be
specifically paired with Coanda augmenting ejectors. Since the
compressor bleed air and exhaust gas emerging from the turbine are
supplied as motive fluids to specialized ejectors, it may be
desired to choke the motive fluid flow at the ejector itself to
maximize performance and minimize the fuel burn. Significantly
different from other applications, the disclosed cycle nearly
chokes the passage to the turbine but stops short of doing so, and
instead, chokes the flow at the exhaust thrust augmenting
ejector.
[0082] The turbine receives the hot gases as described and extracts
the power needed to power the compressor via mechanisms known in
the art. From the turbine efflux, the gases emerge into a compact
heat exchanger, where the hot gases are cooled while preheating the
compressor discharge air en route to the combustor.
[0083] The heat exchanger may be of various shapes to maximize the
heat transfer and endure the operation cycles. In one embodiment,
the air discharge is passed through the exhaust pipe via helicoidal
elements to maximize the heat recovery and minimize the pressure
drop. In a preferred embodiment, the heat exchanger recovers heat
to boost the compressor discharge temperatures from e.g. 400 F to
600 F and preferably to 1000 F, depending on the cycle and
application. In turn, the efficiency of the cycle increases as the
fuel burn is reduced accordingly. The architecture of the
propulsion system allows the heat recovery to be implemented via a
compact heat exchanger and a unique combustion chamber, highly
integrated in the engine.
[0084] Rather than being accelerated into a nozzle, as would be the
case in a traditional turbojet or turbofan, the resulting flow from
the heat exchanger, now at lower pressures and temperatures
compared to the entry section in the heat exchanger, but still at
higher values than the ambient pressure and temperature,
respectively, is delivered to and accelerated into the primary
nozzle of an ejector, near or at choking conditions. In one
embodiment, a Coanda ejector uses the emerging gas from the turbine
to provide a thrust by entrainment of ambient incoming air at all
points of the mission as the primary nozzle of the ejector is
connected via conduits to the turbine efflux. The ejector may or
may not swivel to allow for vertical take-off, landing and
hovering, as well as level flight propulsion.
[0085] FIG. 7 illustrates a Coanda nozzle propulsion system
combustor augmenter. At the exit of the combustor liner toroidal
structure, a small opening turns into a convergent, annular channel
accelerating the flow into the axial direction towards the turbine.
Because of the high rotational nature of the circumferential flow
inside the torus, the straightening of the flow into axial
direction is not fully achieved nor desired. With the appropriate
geometry, the need for a nozzle to accelerate the flow and guide it
into the rotating axial turbine disappears. The turbine nozzle
(also known as the first stage nozzle) is eliminated. The residual
movement of the reacting flow, further facilitated by the late
injection of dilution holes, is sufficient to guide the flow to the
turbine rotor blades at the appropriate angle. This embodiment does
not choke the flow at the exit of the combustor and inlet into the
turbine; rather, the choking of the flow occurs at the exit of the
gas generator (the location of the specialized ejector of the
Coanda type) as explained herein. In this embodiment, the hot
exhaust flowing from the turbine still has a high pressure and
temperature such that it can be used as a motive fluid for the
ejector to entrain air and augment the stream's thrust by 25-75%
when compared to the baseline thrust of the original turbojet.
[0086] The turbine itself may be manufactured from a CMC or
metallic-based materials. Cooling may or may not be employed. In
one embodiment, the CMC made turbine blades can withstand 2000 F
inlet temperatures (firing temperature) without the need of cooling
and can extract the work needed to drive the compressor between
pressures of 4 to 2 bar (60 to 30 psia) at take-off and hovering
conditions. At that operating condition, the compressor bleeds
roughly 20% of the flow to forward ejectors, and the remaining 80%
flows through the compressor and is preheated to 1000 F by exhaust
gases from the compressor discharge using the heat exchanger. The
combustor fires at 2200 F inside the CMC liner and the hot stream
is diluted to 2000 F TET. A gas turbine as described herein could
produce, for example, 500 lbf at take-off using 5 lb/s air,
bleeding 25% of that and at an efficiency of 30% at take-off and in
higher 30s% at cruise conditions. The take-off condition may also
be used for further augmenting the thrust by injecting fuel
directly into the center of the Coanda nozzles, thereby generating
a very large amount of thrust at lower efficiency for a short burst
via a pseudo-ram effect, followed by the transition to cruise
condition when bleeds are closed on the compressor and the
efficiency increases to approximately 40%. In this embodiment, the
pseudo-ram effect is generated by the vacuum created in front of
the Coanda ejector, which may be used to entrain vast amounts of
air of at least 10:1 entrainment ratio. At the same time, the
Coanda ejector entrained air and its wall jets assists with the
atomization of the fuel injected in its middle and can autoignite
the fuel and air within the diffuser section, generating additional
thrust in the process. Ignition of the fuel and air mixture inside
the diffuser is achieved by a pilot flame or a sparking plug, where
the flame stabilization is away from the walls of the diffusor in
the areas of lower velocities of the Coanda diffuser (see FIG.
5).
[0087] In FIG. 7, the plenum 501 of the Coanda ejector 500
introduces pressurized fluid which can be compressed air from the
compressor bleed or exhaust gas at higher than ambient pressure as
motive air into the ejector, forming a flow pattern resulting in
thrust 508. A fuel injector 502 is placed in front of the Coanda
ejector 500, away from the inlet but into the low-pressure area
impacted by the operation of the Coanda ejector. The injector 502
sprays/injects liquid (or gaseous) fuel along the main axis of the
Coanda ejector 500 in the shape of a spray or fuel jet 506. A pilot
flame or torch or spark igniter ignites the mixture, and the flame
507 propagates downstream, being sucked in by the ejector 500,
however stabilizing itself away from the walls 503 of the diffuser,
which are protected from direct touch with the flame via the very
high velocity wall jets 504. Because the local axial velocity in
the center is equal to the turbulent flame propagation speed in the
direction of fuel injection, the flame front stabilizes, releasing
heat that generates more thrust 509. If the fluid in plenum 501 is
hot, pressurized exhaust from a gas turbine, the fuel and exhaust
gas may be autoignited at contact with the gases, by orienting the
spray accordingly. The operation is intended to be performed for
small duration, e.g. VTOL or STOL, hovering or emergency.
[0088] Conventional jet engines employed in small aerial vehicles
less than 1000 lbs in total weight are commonly turbojets or jet
with very low bypass ratio, such as those employed by Williams
International (U.S. Pat. No. 4,598,544) or manufactured by hobby
suppliers such as Jetcat or Jetbeetle. FIG. 8 depicts the results
of the modeling of a turbojet engine similar in size with a Jetcat
model as compared with an embodiment of the present invention.
[0089] In FIG. 8, the turbojet has a specific fuel consumption of
30.9 g/(s-kN) or 1.09 lb fuel/lbf-h, in order to produce a thrust
of 300 N or 67 lbf. The present invention implements a heat
regeneration unit in the jet engine to recover the heat and reduce
the consumption of the fuel by at least 25%, so the thrust specific
fuel consumption can drop accordingly to 1.09*75% or 0.8175 lb
fuel/lbf-hr. For a two-hour mission of an unmanned aerial vehicle
with a thrust requirement of 300 N on average can use roughly 110
lbs of fuel (or about 18 gallons) while the conventional turbojet
would require 146 lbs of fuel (or 25 gallons) for the same mission.
This would result in savings of 22 miles per gallon of fuel for the
present invention versus 16 miles per gallon for the conventional
turbojet.
[0090] The implementation of the heat exchanger can further be
improved by the introduction of thrust augmenting ejectors to the
cycle.
[0091] FIG. 9 illustrates a conventional turbojet. One or more
embodiments of the present invention differ from the turbojet shown
in FIG. 9 in numerous ways. First, the compressor is bled at
take-off, hovering and landing by roughly up to 25% to feed a 2:1
thrust augmenting ejector described within this application. The
bleed valves are marked handling bleed and can be closed at cruise
condition and re-opened at hovering and landing, according to the
mission.
[0092] Second, the compressor discharge air at section 3 is routed
through a heat exchanger that is located within the exhaust section
past the turbine section 6. The unit is called heat regeneration
unit and allows the exhaust gases to heat up the combustion air
prior to its delivery to the combustor.
[0093] Third, the flow from section 3 to section 31 is modified to
allow the introduction of a heat exchanger to recover some of the
exhaust heat.
[0094] Fourth, the combustor adds heat for a smaller temperature
difference compared to the original turbojet, due to the heat
exchanger. The increase in the temperature between the combustor
inlet temperature and the outlet form the combustor is typically
400 F or 475 K, resulting in 25% reduction of the fuel needed by
the cycle, and maintaining the same inlet temperature to the
turbine (see Tables 1 and 2).
[0095] Fifth, the combustor exit section carries some residual
circumferential motion of the gases, and does not necessarily
employ a traditional first stage nozzle to accelerate the gases.
Rather, a converging section can introduce the exhaust gases to the
turbine rotor stage.
[0096] Sixth, the turbine is designed such that it expands the
exhaust gas to higher than atmospheric pressure, preferably above
1.1 pressure ratio compared to ambient conditions; FIG. 10 (in
comparison with FIG. 8) shows that the expansion process ends
somewhere between 1.5:1 and 2:1 pressure ratios. Compared to the
ambient pressure, this enhances the heat exchanger performance of
the heat regenerator represented in FIG. 11 and leave room for the
use of the ejector at the exit at section 8.
[0097] Seventh, the heat exchanger, which preferably is of the
coiled or counterflow type, recovers the heat and transfers a
portion of it into the fresh combustion air stream, while dropping
the exhaust heat temperature between sections 5 and 6. FIG. 11, one
embodiment of the present invention, shows the exhaust gas
temperature of the hot side gas provided from section 1105 after
the turbine is 1400 F and 26.5 psi at element 1106, before the heat
exchanger and 944 F and 17 psi after the heat exchanger at element
1107 of FIG. 11. The heat exchanger 1111 raises the temperature of
the compressor discharge flow supplied from the compressor
discharge plenum 1108 and supplied via conduit 1105 at 426 F and 60
psi. After the heat exchanger occurs, the fresh, recuperated
compressor discharge air reaches 856 F at 53.5 psi, before being
introduced to the combustor described in the sections above. In the
combustor described in FIGS. 2-3 and text above, the temperature is
being raised to, for example, approximately 2000 F before
introducing the flow to the turbine. The power extracted by the
turbine from 52 psi and 2000 F to 27 psi and 1400 F (plenum 1107 in
FIG. 11), combined with the respective efficiencies known in the
art for turbomachinery components, balances the need for the
compressor power input. The rise in temperature of the fresh air
exiting the heat exchanger reaching 430 F would result in
considerable savings of the fuel used in this thermodynamic cycle.
The other stream exiting the heat exchanger also contributes to
reducing the losses, as it is rejected at much lower temperatures
than the original value of the turbojet cycle, e.g. 944 F (780 K)
instead of 1500 F (1090 K).
[0098] Eighth, following the same sample calculations above, the
exhaust stream from the heat exchanger exits the heat exchanger at
17 psi and 944 F and is directed next not to a simple nozzle, as
shown in the FIG. 9 at section 8, but instead to a specially
designed ejector, where the said exhaust stream is used as motive
air. Testing data indicates an augmentation ratio of >1.25 for a
pressure ratio of <1.25 for an axisymmetric ejector and an
augmentation ratio of over 2 for a flat ejector for a pressure
ratio of <1.5. Hence, the thrust augmentation expected from the
conditions described above is between 1.25 and 2.0 for the present
invention and for very small pressure ratios required.
[0099] Ninth, the thrust augmentation can produce between
67*1.25=84 lbf and 67*2=134 lbf. In comparison, the reduction of
the flow to achieve the same level of thrust needed would be 0.454
kg/s (1 lb/s)*0.80 (i.e. 20% less flow) for the same thrust needs,
and the same fuel-to-air ratio to match the cycle conditions
including the TET.
[0100] FIG. 10 illustrates the thermodynamic cycle when modified
with a regenerative heat exchanger and ejector. The heat
regeneration transfers heat from the turbine exhaust stream and
reduces the amount of heat addition required by the combustor. The
combustion heat addition evolution becomes 3'-4. The turbine
expansion process also changes to a 4'-6' evolution, providing
enough power to run the compressor. The heat exchanger determines
the evolution to point 8' on almost an isobaric process, followed
by the ejector evolution 8'-s8' in isentropic manner, in the
primary nozzle of the ejector. At the end of the evolution, when
the pressure has dropped to s8', the mixing process begins with the
secondary fluid, which is the ambient air, evolution s8'-mix. The
point called `mix` is at a slightly higher pressure than ambient,
and the final evolution is a nearly isentropic expansion of the
flow to the ambient exit static pressure.
[0101] The various embodiments of the present invention disclosed
in this application, including the regenerative cycle and the
reduction in flow combined with the introduction of the ejector
technologies can generate fuel savings of more than 25% and a
smaller rotating core. As such, an aerial vehicle may achieve more
than 25% more miles per gallon when compared to the turbojet,
without the use of large fans or other moving parts. The fuel
consumption can be dropped according to the present invention to
less than 0.7 lbs fuel per pound force and hour, allowing an aerial
vehicle powered by the system disclosed in this application to
travel 400 miles on less than 90 lbs of fuel, or roughly 27
mpg.
[0102] FIG. 11 illustrates an example of the heat exchanger
streams, which include: the exhaust from the turbine into plenum
1107 and flowing through conduit 1103 to the heat exchanger at 1400
F and 26.5 psi and exiting the heat exchanger 1111 through conduit
1104 at 944 F and 17 psi; the fresh compressed air from the
compressor at 426 F and 60 psi is supplied at plenum 1108 and flows
through conduit 1105 to the heat exchanger 1111, exiting via
conduit 1106 to the combustor at 856 F and 53.5 psi; and, as
represented by element 1101, the motive fluid nozzle (primary
nozzle) of the ejector that is permanently connected to the system.
The conduit supplying element 1101 in FIG. 11 is the element 501 in
FIG. 7.
[0103] An optionally advantageous element in one embodiment of the
present invention is the use of a compressor that employs bleed
valves. The opening of the bleed valves during operation results in
a drop in pressure and lowers of the working line, away from the
stall line. While the pressure drops, the compressor can still be
accelerated and the flow can be increased, albeit at lower
efficiencies. If the bleed flows are used for non-propulsive
reasons (cabin pressurization, overboard bleed, etc.), then the
specific fuel consumption increases because the compressed air, for
which power was consumed, is not contributing to generating thrust.
However, if the bleed flows are used for thrust augmentation at
1.5-2.5 times the thrust otherwise obtained with the same flow via
expansion through a nozzle at the end of the cycle, the bleed can
contribute significantly to the thrust and particularly to the
vertical thrust required in vertical take-off and landing
applications. The present invention allows the system to be
flexible and engage the compressor bleed powered thrust
augmentation ejectors at various stages of the flight. If the
intention is to operate the bleed valves fully open at vertical
take-off and landing in order to power the thrust augmentation
ejectors, as well as hovering, the fuel consumption may accordingly
increase. However, since for most applications, this may be a small
portion of the mission, the opening of the valves to allow these
maneuvers is acceptable, and bleed valves supplying ejectors may be
closed or minimized at cruise conditions.
[0104] The operation with open bleed valves on simple compressors
such as the small turbojets used in hobby applications, i.e. less
than 300 lbf thrust, is impacting the TET because the TET is
limited due to materials capabilities and less air is supplied to
the combustor in case of the bleeding. It is, however, possible to
operate the turbine for a limited amount of time at exceeding
temperatures compared to nominal values without significant impact
to the maintenance interval of the system. Moreover, recent
advancements in materials science and the introduction of CMCs to
the combustor liners and turbine nozzles and rotors allow the
possibility to exceed the turbine inlet temperatures several
hundreds of degrees over the similar metallic turbines without
significant impact on the life of the turbine.
[0105] The overall performance of the system depends on the
efficiency of the thrust augmenting ejectors as well. Simply
bleeding off the compressor stream has demonstrated that using a
simple nozzle instead of an ejector can achieve 2-3 times the
entitlement. In one embodiment, the system is sized for 1 lb/sec
air flow at maximum speed, 4:1 pressure ratio, 10% compressor bleed
powering "cold" ejectors in the front of the system, and 90% hot
gas supplied to an exhaust (also known as the "hot" ejector). The
cold ejectors produce 11 lbf thrust (i.e. 110 lbf/lb/s) and the
remaining 90% of the hot flow produces 100 lbf/lb/s resulting in 90
lbf of thrust. The additional thrust boost resulting from the fuel
injection in the thrust augmentation ejectors (as disclosed herein)
results in the boosted thrust increasing to 20 lbf for the cold
ejectors and 150 lbf for the hot ejector. The conventional hobby
turbojet produces only 50 lbf of thrust, so without fuel injection,
the described ejectors can augment the thrust by (11+90)=101 lbf,
more than doubling the thrust for the same amount of fuel consumed.
The result is a fuel burn savings of more than 50% compared to the
original product. The additional fuel injection in the thrust
augmentation ejectors reduces the fuel efficiency and conversely
increases the fuel consumption for a small duration of the mission,
such as take-off or hovering or landing. However, the thrust
augmentation becomes 170 lbf versus the original 50 lbf, which can
be also vertically directed for vertical take-off, hovering or
landing, or assisting for Short Take-Off and Landing (STOL.)
[0106] The device augments the thrust between 1.25-2.0 times the
otherwise simple expansion to the ambient. FIG. 12 illustrates the
ratio of augmentation obtained through experimental data using the
augmentation devices disclosed in this application as compared to
the compilation of other ejectors in the NTIS publication ADA098620
of the Vought Corporation Advanced Technology Center, published in
September of 1979.
[0107] As shown in FIG. 12, the thrust augmentation performance of
the present invention is in the range 1-2 pressure ratio ranges
between 1.5-3, outperforming most other ejectors in the prior art.
At maximum speed of the compressor and maximum power of the present
invention gas generator, it is expected that the hot exhaust gas
can produce at least 50% more thrust than the original small
turbojet simple exhaust nozzle, and the thrust augmenting device
(element 500 in FIG. 5) of the invention can be oriented for
vertical take-off and landing, and/or hovering, making the use of
the gas generator. Moreover, the compressor bleed flows can be
conveniently directed through a network of conduits to swiveling
thrust augmenting devices that can produce between 2-3 times the
thrust of the simple turbojet propelling nozzle. If desired,
further augmentation may be obtained via injection of fuel and its
ignition in the said thrust augmenting devices (as explained
herein).
[0108] While the preferred embodiment of the invention has been
illustrated and described, as noted above, many changes can be made
without departing from the spirit and scope of the invention.
Accordingly, the scope of the invention is not limited by the
disclosure of the preferred embodiment. Instead, the invention
should be determined entirely by reference to the claims that
follow.
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