GB2420615A - Thermo acoustic pressure rise pulse combustor - Google Patents
Thermo acoustic pressure rise pulse combustor Download PDFInfo
- Publication number
- GB2420615A GB2420615A GB0520278A GB0520278A GB2420615A GB 2420615 A GB2420615 A GB 2420615A GB 0520278 A GB0520278 A GB 0520278A GB 0520278 A GB0520278 A GB 0520278A GB 2420615 A GB2420615 A GB 2420615A
- Authority
- GB
- United Kingdom
- Prior art keywords
- fuel
- pressure
- combustor
- thermo acoustic
- pulse
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
- 239000000446 fuel Substances 0.000 claims abstract description 93
- 238000002485 combustion reaction Methods 0.000 claims abstract description 48
- 238000002347 injection Methods 0.000 claims abstract description 11
- 239000007924 injection Substances 0.000 claims abstract description 11
- 238000004200 deflagration Methods 0.000 claims abstract description 9
- 230000000630 rising effect Effects 0.000 claims description 8
- 238000000034 method Methods 0.000 claims description 7
- 230000001939 inductive effect Effects 0.000 claims description 4
- 230000001747 exhibiting effect Effects 0.000 claims description 2
- 238000005474 detonation Methods 0.000 abstract description 6
- 239000000126 substance Substances 0.000 abstract description 6
- 238000005086 pumping Methods 0.000 abstract description 5
- 238000010248 power generation Methods 0.000 abstract description 4
- 239000007788 liquid Substances 0.000 abstract description 2
- 239000007787 solid Substances 0.000 abstract description 2
- 239000000203 mixture Substances 0.000 description 8
- 238000006243 chemical reaction Methods 0.000 description 7
- 239000012530 fluid Substances 0.000 description 4
- 230000002123 temporal effect Effects 0.000 description 3
- 230000008878 coupling Effects 0.000 description 1
- 238000010168 coupling process Methods 0.000 description 1
- 238000005859 coupling reaction Methods 0.000 description 1
- 230000000737 periodic effect Effects 0.000 description 1
- 239000000376 reactant Substances 0.000 description 1
- 230000002459 sustained effect Effects 0.000 description 1
- 238000004056 waste incineration Methods 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23C—METHODS OR APPARATUS FOR COMBUSTION USING FLUID FUEL OR SOLID FUEL SUSPENDED IN A CARRIER GAS OR AIR
- F23C15/00—Apparatus in which combustion takes place in pulses influenced by acoustic resonance in a gas mass
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C5/00—Gas-turbine plants characterised by the working fluid being generated by intermittent combustion
- F02C5/10—Gas-turbine plants characterised by the working fluid being generated by intermittent combustion the working fluid forming a resonating or oscillating gas column, i.e. the combustion chambers having no positively actuated valves, e.g. using Helmholtz effect
- F02C5/11—Gas-turbine plants characterised by the working fluid being generated by intermittent combustion the working fluid forming a resonating or oscillating gas column, i.e. the combustion chambers having no positively actuated valves, e.g. using Helmholtz effect using valveless combustion chambers
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K7/00—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof
- F02K7/02—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof the jet being intermittent, i.e. pulse-jet
- F02K7/04—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof the jet being intermittent, i.e. pulse-jet with resonant combustion chambers
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23C—METHODS OR APPARATUS FOR COMBUSTION USING FLUID FUEL OR SOLID FUEL SUSPENDED IN A CARRIER GAS OR AIR
- F23C2205/00—Pulsating combustion
- F23C2205/10—Pulsating combustion with pulsating fuel supply
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Fluidized-Bed Combustion And Resonant Combustion (AREA)
Abstract
A thermo acoustic pressure rise pulse combustor comprises injecting fuel 8 in pulses into a combustion chamber 4 so that a cyclical pressure fluctuation is exhibited and induces deflagration of the fuel, and the fuel is injected with a phase relationship with the pressure rise in the combustion chamber 4. Fuel can be pulsed at the same frequency as the pressure fluctuation or at a phase angle delay of between 90 and 210 degrees from the minima of the pressure fluctuation. Combustion chamber 4 pressure can be measured by sensors and a feedback circuit couples the sensors with the fuel injection 8 so that fuel is pulsed in response to the pressure in the combustion chamber 4. The combustor may be a valve-less pulse combustor, a valved pulse combustor, a pulse detonation engine and they all may be used for a gas turbine engine, direct propulsion, pumping, power generation or in a bypass duct of a turbofan engine. The fuel may be solid, liquid or gaseous. Controlling the timing, duration and/or amplitude of the fuel injection permits an increase of converting chemical energy into mechanical energy. The fuel injector may be a car fuel injector.
Description
1 2420615
COMBUSTOR
This invention relates to pressure rise combustors and in particular thermo acoustic pressure rise combustors and valve-less pulse combustors.
Combustion chambers fall into two categories: constant pressure and pressure rise. In a constant pressure combustor fuel at steady state is continually combusted and the hot exhaust gas allowed to expand without constraint. While there may be some pressure loss or pressure fluctuations due to resonance within the chamber these variations are kept small. Examples of constant pressure combustors are: turbo- annular combustors for gas turbines, ram jets or dump combustors. Such combustors are not said to be thermo acoustic.
In a pressure rise combustor the pressure within the combustor varies widely and in a periodic manner. A pressure rise combustor utilises unsteady combustion to produce an exhaust gas stream which has a higher mechanical energy, or stagnation pressure, than that of the inlet stream. The produced mechanical energy can be extracted as, for example, thrust or shaft work. Pressure rise combustors may be further divided into thermo acoustic combustors, where at least the outlet of the combustion chamber is open to atmosphere and the acoustics of the combustor are such that deflagration of the fuel / air mixture acts against an induced pressure wave to further increase the pressure in the chabe, andmechanical combustors in which the fuel / air rrixture is constrained within an enclosed combustor and deflagration or detonation of the fuel / air mixture acts against a piston or other mechanical device Pressure rise combustors may be used to provide propulsion. An example of a thermoacoustic pressure rise combustor is a pulse-jet.
A pressure rise combustion system may be applied to a gas turbine and offers a potentially increased thermodynamic performance. In a conventional gas turbine combustion chamber, i.e. non pressure rise, there is a pressure loss of typically 5% of the engine overall pressure ratio and there is no conversion of thermal energy to mechanical energy.
In a pressure-rise combustor there is a stagnation pressure rise due to the conversion of chemical energy into mechanical energy. Pressure-rise combustion can be used with solid, liquid or gaseous fuels.
A pulse-jet may be valved or valve-less and utilise unsteady combustion in an acoustically resonant combustion chamber to produce a pressure rise. Fuel is steadily supplied to the combustion chamber and the timing of the unsteady combustion heat release is dictated by the aerothermo-acoustical coupling in the working fluid. The phase angle and amplitude of the unsteady heat release is governed by the internal fluid mechanics of the system i.e. temporal and spatial variations of mixing processes, strain rates, convection of reactants and ignition sources. The combustion chambers burn the fuel in a deflagration process rather by detonation. The deflagration enables the combustion process to be selfsustaining in that once initial ignition is effected the acoustics within the system generate a cyclical combustion process without requiring further energy input to re-ignite an injected air / fuel mixture. This is in contrast to detonation combustors where an air/fuel mixture is detonated through input from an external energy source such as a spark plug, the chamber evacuated of the products of the detonation, a new air/fuel mixture supplied to the chamber and detonated through input from the external energy source. Each combustion event can be said to be isolated from an earlier and subsequent combustion event and consequently significant energy must be input by the spark plugs to ensure operation of the combustor for a sustained period.
it is an object of the invention to seek to provide an improved thermo acoustic deflagration combustion device.
According to an aspect of the present invention there is provided a thermo acoustic pressure rise combustor having an air inlet, an air outlet, a combustion chamber and an associated fuel injector from which in use fuel is injected into the combustor chamber in pulses, the combustor chamber in use exhibiting a cyclical pressure fluctuation having a pressure rising portion and a pressure falling portion induced by deflagration of the fuel, the timing of the start of the pulse of fuel being in a phase relationship with the pressure rising portion.
The fuel flow may be pulsed such that the fuel addition and / or heat addition to the combustor is non-sinusoidal.
The fuel may be pulsed at the same frequency as the pressure fluctuations.
The cyclical pressure fluctuation may be sinusoidal, the pressure rising portion starting at the minima of the sinusoidal pressure fluctuation and extending to the maxima of the sinusoidal pressure fluctuation, the pressure falling portion starting at the maxima of the sinusoidal pressure fluctuation and extending to the minima of the sinusoidal pressure fluctuation, wherein the timing of the start of the pulse of fuel is phased from the minima by between 00 and 210 , and preferably by between 90 and 210 .
The timing of the end of the pulse of fuel may be phased from the timing of the start of the pulse of fuel by a phase angle of between 300 and 180 . The phase delay of the timing of the start of fuel injection may vary between cycles.
The thermo acoustic pressure rise combustor may further comprising pressure sensing means for sensing the pressure within the combustor chamber. Preferably the pressure sensor is functionally connected to the fuel injector such that fuel is pulsed in response to a sensed pressure value within the combustor casing.
According to a second aspect of the present invention there is provided a method of operating a thermo acoustic pressure rise combustor comprising the steps introducing air and fuel into the combustor and inducing the air and fuel to ignite and deflagrate thereby creating a cyclical pressure fluctuation and pulsing fuel into the injector at the same frequency as the pressure fluctuation.
According to a third aspect of the present invention there is provided a method of operating a thermo acoustic pressure rise combustor comprising the steps introducing air and fuel into the combustor, creating a cyclical pressure fluctuation by inducing the air and fuel to ignite and deflagrate, measuring the pressure in the combustor and pulsing fuel into the injector at a phase angle delay of between 0 and 210 to the minima of the pressure fluctuation.
Thermal efficiency is the quantity of mechanical energy that is extracted from combustion products divided by the quantity of heat liberated to the fluid through combustion.
This is different to combustion efficiency which is defined as the quantity of heat liberated to the fluid through combustion divided by the calorific content of the fuel.
Beneficially, the conversion of chemical energy into mechanical energy within the system can be improved and thus the thermal efficiency of the system is improved.
Preferably the fuel is pulsed at a defined phase relationship with the pressure rise. Preferably the heat addition to the combustor is nonsinusoidal. Preferably the phase relationship is allowed to vary and can be controlled.
The pressure-rise combustor may be a valve-less pulse combustor and may be used as a combustor for a gas turbine, for direct propulsion, for pumping, for power generation or in the bypass duct of a turbofan engine.
The pressure-rise combustor may be a valved pulse combustor and may be used as a combustor for a gas turbine, for direct propulsion, for pumping, for power generation or in the bypass duct of a turbofan engine.
The pressure-rise combustor may be a pulse detonation engine and may be used as a combustor for a gas turbine, for direct propulsion, for pumping, for power generation or in the bypass duct of a turbofan engine.
The invention will now be described, by way of example only, with reference to the following figures, in which: Figure 1 depicts a valveless pulse-jet combustor.
Figure 2 depicts a pressure wave response in time of the combustor of Figure 1.
Figure 3 depicts the pressure fluctuations in the combustor chamber and fuel pulse delay Figure 4 is a graph of RMS unsteady pressure fluctuation for a pulsed flow and a continuous flow The valve-less pulse-jet of Figure 1 comprises an air inlet 2, a combustion chamber 4 and a tailpipe 6. Air enters the combustion chamber from the air inlet 2 and fuel is injected from a fuel injector 8. An igniter 10 is initially used to ignite the air/fuel mixture in the combustion chamber.
Where a fuel and air mixture is ignited in the chamber in a deflagration process a pressure wave is initiated that travels within the combustion chamber 4, tailpipe 6 and air inlet 2 as depicted in figure 2. The fuel is ignited at point 10 and this induces a pressure wave that travels axially forward 12a and axially rearward 12b. At the open ends of the inlet and tail-pipe the pressure wave reflects and expansion waves 14a, 14b travel back along the combustion device to the opposite open ends where the expansion waves reflect and the resultant pressure waves 16a, 16b intersect at point 18 within the combustion chamber.
In a conventional pulse-jet fuel is either continuously supplied to the combustion chamber at a constant rate for the period of time between point 10 and point 18, or the fuel supply rate is allowed to fluctuate due to the oscillating pressure in the combustion chamber. As the inlet is of a smaller length to that of the tail pipe air is drawn into the combustion chamber and mixes with the fuel. At point 18, where the pressure waves intersect the pressure within the combustion chamber is at a maximum and induces ignition in the air / fuel mixture. This creates a further pressure wave and the cycle 10 to 18 repeats.
The combustion chamber is acoustically resonant and the aero-thermalacoustic operation of the pulse combustor causes the combustion to naturally oscillate at around 200 Hz. The unsteady pressure within the combustion chamber oscillates as a sinusoidal variation, as shown in Figure 3 which, with a continuous supply of fuel gives rise to a constant RS unsteady pressure within the combustion chamber as shown as the double line in Figure 4.
In accordance with the invention fuel is pulsed into the combustor at the same frequency as the combustion process. The pressure of the combustion chamber is measured and a feedback circuit couples the sensor with the fuel injector.
The efficiency of the combustion process is governed by how heat is added to a volume of gas. When the heat is added in a non-steady manner to a volume of gas experiencing unsteady pressure fluctuations the mechanical energy content of the volume of the gas can be increased.
The temporal phase angle between the unsteady pressure fluctuations and the unsteady heat addition is minimised to increase the conversion of chemical energy into mechanical energy. The temporal phase angle is minimised through selectively varying the timing, duration and amplitude of unsteady fuel injection into the combustion chamber.
Figure 3 further depicts the unsteady pressure fluctuations with the pulses of fuel injected at the frequency of the combustion process. A fuel injector allows injection to begin at the rising edge 20 and terminates injection at the falling edge 22 a time period T later which, in this example corresponds to the duty cycle. The rising edge of the fuel pulse is positioned at a time delay t1 which corresponds to a phase angle from the minima of the of the sinusoidal pressure fluctuation within the combustor By correctly controlling the timing, duration and / or amplitude of the fuel injection, the conversion of chemical energy into mechanical energy is increased. If the timing, duration and / amplitude of the fuel injection is incorrectly controlled the conversion of chemical energy into mechanical energy is reduced.
Figure 4 depicts a graph of RMS unsteady pressure in the combustor against phase angle delay where fuel is pulsed into the combustor at a rate equivalent to l.6g/s for a period equivalent to the duty cycle. Also plotted is the RMS unsteady pressure observed for a fuel flow into the combustor at a continuous rate equivalent to l.6g/s.
It may be observed from the graph that pulsing the fuel flow provides an increase to the RMS unsteady pressure especially where the phase angle is between 00 and 210 .
Between a phase angle t1 of 90 and 210 the RMS unsteady pressure is approximately 4kPa greater than the RMS unsteady pressure of the pulse jet having a continuous injection of fuel throughout the duty cycle. The increase equates to an increase in mechanical energy produced by the combustor of approximately 27% over that where the fuel is fed into the chamber at a constant rate The amplitude of the RMS unsteady pressure is indicative of the quantity of mechanical energy stored in the acoustic wave.
A pressure sensor measures the pressure within the combustor and the amplitude and phase of the fuel injection is varied in accordance with the measured pressure via a feedback circuit with the fuel being injected a time period t1 after the measured pressure is sensed by the pressure sensor. Alternatively, an external driver controls the fuel injector. There may be one or many fuel injectors at the same or different axial locations.
The RMS unsteady pressure may be further increased by reducing the length of time within the over which the fuel is injected.
The results in the exemplified system was obtained from a valve-less pulse jet. Inlet had a length 140mm and inside diameter 39mm, the combustion chamber a length 145mm, and inside diameter 75mm giving a combustor volume of 640cc. The tail pipe had a length 945mm and was tapered from an initial inside diameter of 22.5mm to an exit inside diameter of 60mm.
The fuel injector was one suitable for a car and periodically injected fuel at a rate of l.6g/s. The front end of the combustion chamber is defined by a tapered combustor head and four injectors were arranged at a 900 spacing.
Initial ignition is effected by a spark plug located part way along the combustion chamber.
Whilst the present invention has been described with respect to a valveless pulse jet, the invention is also applicable to pulsed combustors used in a gas turbine. The partial conversion of thermal energy to mechanical energy beneficially allows the pressure of the hot gas entering the turbine to be increased above the pressure of the gas entering the combustor i.e. there is no pressure loss in the combustor. The increased pressure of gas entering the turbine allows more work to be done on the turbine and thus improves on the efficiency of the overall engine. Improved efficiency lowers fuel costs and the cost of running the engine.
Pulse jets may be used on their own or as part of a propulsion system e.g. as a reheat system, thrust augmenter or VTOL system.
A gas turbine incorporating the invention may be used in power generators, pumping systems or marine propulsion as well as aircraft propulsion.
The pulsed combustor may also be applied in waste incineration and micro combined heat and power systems.
Claims (11)
1. A thermo acoustic pressure rise combustor having an air inlet, an air outlet, a combustion chamber and an associated fuel injector from which in use fuel is injected into the combustor chamber in pulses, the combustor chamber in use exhibiting a cyclical pressure fluctuation having a pressure rising portion and a pressure falling portion induced by deflagration of the fuel, the timing of the start of the pulse of fuel being in a phase relationship with the pressure rising portion.
2. A thermo acoustic pressure rise combustor according to claim 1, wherein the fuel flow is pulsed such that the fuel addition and / or heat addition to the combustor is non- sinusoidal.
3. A thermo acoustic pressure rise combustor according to claim 1 or claim 2, wherein the fuel is pulsed at the same frequency as the pressure fluctuations.
4. A thermo acoustic pressure rise combustor according to claim 1, claim 2 or claim 3, wherein the cyclical pressure fluctuation is sinusoidal, the pressure rising portion starting at the minima of the sinusoidal pressure fluctuation and extending to the maxima of the sinusoidal pressure fluctuation, the pressure falling portion starting at the maxima of the sinusoidal pressure fluctuation and extending to the minima of the sinusoidal pressure fluctuation, wherein the timing of the start of the pulse of fuel is phased from the minima by between 00 and 210
5. A thermo acoustic pressure rise combustor according to claim 4, wherein the timing of the start of the pulse of fuel is phased from the minima by between 900 and 210 .
6. A thermo acoustic pressure rise combustor according to claim 4 or claim 5, wherein the timing of the end of the pulse of fuel is phased from the timing of the start of the pulse of fuel by a phase angle of between 30 and 1800.
7. A thermo acoustic pressure rise combustor according to any preceding claim, wherein the phase delay of the timing of the start of fuel injection varies between cycles.
8. A thermo acoustic pressure rise combustor according to any preceding claim, further comprising pressure sensing means for sensing the pressure within the combustor chamber.
9. A thermo acoustic pressure rise combustor according to claim 8, wherein the pressure sensor is functionally connected to the fuel injector such that fuel is pulsed in response to a sensed pressure value within the combustor casing.
10. A method of operating a thermo acoustic pressure rise combustor comprising the steps introducing air and fuel into the combustor and inducing the air and fuel to ignite and deflagrate thereby creating a cyclical pressure fluctuation and pulsing fuel into the injector at the same frequency as the pressure fluctuation.
11. A method of operating a thermo acoustic pressure rise combustor comprising the steps introducing air and fuel into the combustor, creating a cyclical pressure fluctuation by inducing the air and fuel to ignite and deflagrate, measuring the pressure in the combustor and pulsing fuel into the injector at a phase angle delay of between 90 and 210 to the minima of the pressure fluctuation.
Priority Applications (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US11/262,930 US7637096B2 (en) | 2004-11-25 | 2005-11-01 | Pulse jet engine having pressure sensor means for controlling fuel delivery into a combustion chamber |
| US12/500,887 US20090286189A1 (en) | 2004-11-25 | 2009-07-10 | Combustor |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GBGB0425901.6A GB0425901D0 (en) | 2004-11-25 | 2004-11-25 | Combuster |
Publications (3)
| Publication Number | Publication Date |
|---|---|
| GB0520278D0 GB0520278D0 (en) | 2005-11-16 |
| GB2420615A true GB2420615A (en) | 2006-05-31 |
| GB2420615B GB2420615B (en) | 2007-06-13 |
Family
ID=33561333
Family Applications (2)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| GBGB0425901.6A Ceased GB0425901D0 (en) | 2004-11-25 | 2004-11-25 | Combuster |
| GB0520278A Expired - Fee Related GB2420615B (en) | 2004-11-25 | 2005-10-06 | Pressure rise combustor |
Family Applications Before (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| GBGB0425901.6A Ceased GB0425901D0 (en) | 2004-11-25 | 2004-11-25 | Combuster |
Country Status (1)
| Country | Link |
|---|---|
| GB (2) | GB0425901D0 (en) |
Cited By (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101985904A (en) * | 2010-10-28 | 2011-03-16 | 西北工业大学 | Detonation pipe for high-frequency pulse detonation engine and control method thereof |
| US10094288B2 (en) | 2012-07-24 | 2018-10-09 | Icr Turbine Engine Corporation | Ceramic-to-metal turbine volute attachment for a gas turbine engine |
| WO2019204389A1 (en) * | 2018-04-17 | 2019-10-24 | North American Wave Engine Corporation | Method and apparatus for the start-up and control of pulse combustors using selective injector operation |
| US10557438B2 (en) | 2015-12-18 | 2020-02-11 | North American Wave Engine Corporation | Systems and methods for air-breathing wave engines for thrust production |
| US10995703B2 (en) | 2015-03-19 | 2021-05-04 | North American Wave Engine Corporation | Systems and methods for improving operation of pulse combustors |
| US11286866B2 (en) | 2016-09-19 | 2022-03-29 | Finno Energy Oy | Intermittent injection system for a gas turbine combustor |
| US11578681B2 (en) | 2015-03-19 | 2023-02-14 | University Of Maryland | Systems and methods for anti-phase operation of pulse combustors |
Families Citing this family (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP2430292A1 (en) | 2009-05-12 | 2012-03-21 | Icr Turbine Engine Corporation | Gas turbine energy storage and conversion system |
| US8866334B2 (en) | 2010-03-02 | 2014-10-21 | Icr Turbine Engine Corporation | Dispatchable power from a renewable energy facility |
| US8984895B2 (en) | 2010-07-09 | 2015-03-24 | Icr Turbine Engine Corporation | Metallic ceramic spool for a gas turbine engine |
| CA2813680A1 (en) | 2010-09-03 | 2012-03-08 | Icr Turbine Engine Corporation | Gas turbine engine configurations |
| US9051873B2 (en) | 2011-05-20 | 2015-06-09 | Icr Turbine Engine Corporation | Ceramic-to-metal turbine shaft attachment |
Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5361710A (en) * | 1993-10-07 | 1994-11-08 | The United States Of America As Represented By The Secretary Of The Navy | Method and apparatus for the active control of a compact waste incinerator |
| JPH08226338A (en) * | 1995-02-22 | 1996-09-03 | Toshiba Corp | Combustion vibration control device for combustor |
-
2004
- 2004-11-25 GB GBGB0425901.6A patent/GB0425901D0/en not_active Ceased
-
2005
- 2005-10-06 GB GB0520278A patent/GB2420615B/en not_active Expired - Fee Related
Patent Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5361710A (en) * | 1993-10-07 | 1994-11-08 | The United States Of America As Represented By The Secretary Of The Navy | Method and apparatus for the active control of a compact waste incinerator |
| JPH08226338A (en) * | 1995-02-22 | 1996-09-03 | Toshiba Corp | Combustion vibration control device for combustor |
Cited By (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101985904A (en) * | 2010-10-28 | 2011-03-16 | 西北工业大学 | Detonation pipe for high-frequency pulse detonation engine and control method thereof |
| US10094288B2 (en) | 2012-07-24 | 2018-10-09 | Icr Turbine Engine Corporation | Ceramic-to-metal turbine volute attachment for a gas turbine engine |
| US10995703B2 (en) | 2015-03-19 | 2021-05-04 | North American Wave Engine Corporation | Systems and methods for improving operation of pulse combustors |
| US11578681B2 (en) | 2015-03-19 | 2023-02-14 | University Of Maryland | Systems and methods for anti-phase operation of pulse combustors |
| US10557438B2 (en) | 2015-12-18 | 2020-02-11 | North American Wave Engine Corporation | Systems and methods for air-breathing wave engines for thrust production |
| US11434851B2 (en) | 2015-12-18 | 2022-09-06 | North American Wave Engine Corporation | Systems and methods for air-breathing wave engines for thrust production |
| US11286866B2 (en) | 2016-09-19 | 2022-03-29 | Finno Energy Oy | Intermittent injection system for a gas turbine combustor |
| WO2019204389A1 (en) * | 2018-04-17 | 2019-10-24 | North American Wave Engine Corporation | Method and apparatus for the start-up and control of pulse combustors using selective injector operation |
| US11585532B2 (en) | 2018-04-17 | 2023-02-21 | North American Wave Engine Corporation | Method and apparatus for the start-up and control of pulse combustors using selective injector operation |
| US11592184B2 (en) | 2018-04-17 | 2023-02-28 | North American Wave Engine Corporation | Method and apparatus for the start-up and control of pulse combustors using selective injector operation |
Also Published As
| Publication number | Publication date |
|---|---|
| GB2420615B (en) | 2007-06-13 |
| GB0425901D0 (en) | 2004-12-29 |
| GB0520278D0 (en) | 2005-11-16 |
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Effective date: 20161006 |