Continuous Combustion Chamber TRIZ Cases

A self-cooled orifice structure uses segmented end walls to create a cooling cavity that directs swirling air flow through dedicated inlet and outlet apertures.

Pre-buckled fuel tubes in gas turbine injectors relieve compressive stress from differential thermal expansion, preventing structural failure.

An oblique intermediate portion deflects radiant heat away from the fuel-air mixture area, preventing autoignition and flash-back in gas turbine combustors.

Inspection ports in the gas turbine combustor liner enable debris removal from the annular region without compromising structural rigidity.

Mechanically attached swirler housing and nozzle eliminate brazing to reduce manufacturing time and cost.

Alignment plate with aligned holes and an extending tube guides installation rods through combustor flange features to reduce manual force during assembly.

Integrated louver and cooling holes direct airflow from an air gap to the hot surface, eliminating complex floating collars.

Oblique cooling passages in a turbine transition piece frame distribute fluid uniformly, resolving uneven heat transfer and preventing thermal cracking.

An adjustable bypass flow channel manages compressor air delivery to maintain constant combustion chamber pressure loss and material temperature.

A variable frequency Helmholtz damper system adjusts purge medium temperature to tune damping resonance.

Radial resonating tubes on the inner cap counteract high dynamic pressure oscillations from lean premixed combustion to prevent hardware failure.

Extended flow sleeve directs ambient air to cool the combustion liner in a stoichiometric exhaust gas recirculation turbine system.

A diffusion fuel nozzle injects separate fuel and oxidant flows into a turbine combustor to produce a stable diffusion flame.

Fuel nozzle assembly recirculates cooling fluid to premixing region, stabilizing flame and reducing emissions from direct discharge.

Rails with anti-rotation notches define cavities that capture leaking coolant air, reducing leakage while maintaining cooling efficiency.

Direct cooling of securing elements via through openings reduces gas turbine cooling air consumption and lowers NOx emissions without sealing gaps.

Asymmetric lobe-shaped nozzle outlets expand surface area to reduce NOx and CO emissions while extending equipment life.

An aerospace vane uses an integral longitudinal wall passage to eliminate separate fuel tubes and seals, reducing assembly complexity and material usage.

Fuel injection holes located between radial and axial swirlers atomize fuel via turbulent airflow, reducing flashback risk and emissions in gas turbine engines.

Segmented fuel pegs create a mixing zone that controls combustor temperature to reduce NOx emissions without sacrificing fuel efficiency.

Overlapping frame segments and extended vanes minimize thermoacoustic communication while reducing cooling fluid requirements.

Purging the hot gas circuit with inert gas dilutes hydrogen fractions, preventing explosive mixtures and equipment damage.

Oriented effusion holes counteract recirculating hot gas flow in the dilution zone wake area, preventing oxidation of heat shield panels.

A loading assembly transfers mechanical loads between adjacent turbine transition ducts using movable connection elements.

Interlocking tabs and slots eliminate stagnation channels in turbine engine combustors, reducing leakage paths and enhancing durability.

A ductile wear liner with a lower modulus of elasticity than the support hook distributes loads across the seal assembly interface.

Non-planar micro-core cooling channels follow airfoil curvature to resolve poor thermal management on curved engine components.

One-piece bent annular parts decouple radial movements between hot and cold components, reducing assembly complexity while managing thermal gradients.

A modified fuel nozzle stem shifts natural frequency outside the engine running range.

Single air circuit merges multiple cooling functions into one passage, reducing total cooling air usage and boosting turbomachine efficiency.

Staggered inlet and outlet orifices in a double partition divergent bowl guide swirling cooling air to reduce thermal stress on the inner partition.

Curved swirl vanes on a fuel nozzle outer surface create vortex flow to mix compressed air and fuel, preventing backfire by uniformizing flow velocity.

Mini heat shields create a purge cavity that shifts critical gaps away from leading edges, reducing overheating and leakage.

Segmented cooling inlets distribute medium across a gas turbine transition piece, preventing performance loss from rising coolant temperatures.

CMC combustor panels and monolithic air seals form a closed ring to eliminate circumferential gaps and reduce thermal expansion leakage.

Inclined surfaces on transition piece inner walls expand the flow passage area, suppressing Karman vortex streets and reducing pressure fluctuations.

A propulsion turbine nozzle defines a cooling circuit that extracts oxidizer to cool the structure.

A combustor with a lean pre-nozzle fuel injection system and premixing annulus for improved fuel-air mixing.

An integrated perforated plate transfers structural loads to reduce weight and simplify assembly in gas turbine combustion chambers.

A combustor cap assembly separates cooling and fuel flows via a baffle, reducing thermal stress and emissions.

Expansion dampers in the fuel manifold and pigtail lines absorb hydraulic instability from engine vibrations to ensure consistent fuel delivery.

Segmented combustor liner panels discharge compressed fluid via distinct hole rows to purge jet wakes, reducing hot spots and NOx emissions.

Internal diffuser directs coolant flow to a metering location using an accumulation diverter, increasing pressure through ram force.

Curved guide faces on downstream members redirect combustion gas flow away from cavities to maintain turbine efficiency.

Segmented fastening reduces tolerance stacking to prevent cracks from vibratory stresses.

An aft-curved lip on swirler vanes increases axial velocity and reduces shear between airflows, preventing flame holding and flashback in hydrogen combustion.

Controller modifies compressor discharge temperature to shift combustion dynamic frequencies away from turbine bucket natural frequencies.

Swirler generates recirculation zone to stabilize pilot flame, reducing flow losses and improving ignition reliability during reheat operations.

Segmented conical members reduce radial height while accommodating thermal expansion, preventing oil coking in mid turbine frame bearings.

Angular fuel injection orifices create recirculation zones that extend residence time, ensuring complete CO oxidation and stable combustion.