Continuous Combustion Chamber TRIZ Cases

Axially staged combustion system resolves fuel flexibility and complexity trade-offs by segmenting primary and secondary combustion zones.

Integrated transfer passages merge pilot air and transfer fuel flows, reducing NOx emissions by stabilizing the pilot flame.

A cryogenic fuel injector vaporizes liquid natural gas within a jet turbine section to enable efficient combustion.

Segmented fuel delivery places fast-reacting hydrogen centrally surrounded by ammonia, reducing nitrogen oxide emissions while maintaining ignition reliability.

Radial barriers and axial baffles in a combustor tube bundle ensure uniform fuel distribution, reducing thermal stresses and emissions.

A coaxial dihydrogen injection device uses inner and outer swirlers to generate a recirculation zone that stabilizes the flame aerodynamically.

A heat shield arrangement uses a recessed fastening bolt to enable individual component replacement.

Dilution slots and a fence direct air toward the chamber center to reduce hot spots and lower NOx emissions.

Segmented combustor modules and intermediary heat exchangers stabilize combustion while utilizing solar thermal energy.

Strategic narrowing parts suppress carbon monoxide generation during partial load operations by ensuring uniform temperature distribution.

Inclined swirl channels match injector helix angles to shear fuel layers, eliminating recirculation that causes soot deposits and nitrogen oxide emissions.

Specific inner and outer diameter ratios between shell flanges control joint gaps to reduce gas leakage while maintaining assembly ease.

Asymmetrical perforation patterns mitigate non-uniform thru-flow variations across the air inlet plane to ensure uniform fuel-air mixture distribution.

Circumferential fins on a glow plug heater rod increase surface area, resolving wet-condition reliability issues in gas turbine engines.

Segmented combustor cap assemblies reduce thermal cracking by allowing independent segment movement, mitigating fatigue in gas turbines.

An air layer between a heat shield cover and the fuel injection portion suppresses temperature increase, preventing fuel coking in gas turbine engines.

A pre-mix burner integrates an ignition element into the air supply duct to ignite pilot gas directly.

Quench aperture grommets channel cooling air through combustor walls to reduce thermal stresses on turbine engine components.

An engagement part connects the inside turn duct and nozzle guide vane to transmit axial loads.

Varied outlet shapes and inclined flow paths in a micro-mixer bundle prevent spontaneous ignition by reducing radiant heat transfer from the combustion chamber.

Hexagonal swirl mesh injects fuel directly into gas turbine combustors, reducing pressure losses and NOx emissions while stabilizing flames.

A stepped heat shield directs cooling air through a cavity to film cool downstream components and effuse upstream.

An internal air circuit in swirler vanes injects high-pressure gas to eliminate suction side flow deficits and increase flame holding margins.

Extracting spraybars from the main flowpath eliminates pressure losses during dry operation while maintaining thrust augmentation capability.

A gas turbine burner adjusts pilot-air flow to regulate combustion operating points.

Hollow heat shield filled with inert gas prevents coking in compact gas turbine pilot arrangements.

Stacked metallic and ceramic spacers compensate for thermal expansion mismatches during brazing, preventing stress-induced breakage in aeronautical assemblies.

A swirl-stabilized burner design uses an inerting front to protect nozzle openings from high temperatures.

A push-lock pin uses a ball lock mechanism and spring to secure ceramic insulation tiles in aircraft engines.

Porous transpiration cooling buffers thermal damage in oxy-fuel combustors, enabling higher operating temperatures and pressures for efficient power generation.

Segmented cooling circuits balance heat pickup on both venturi sides while reducing pressure drop to enhance combustion performance.

Stiffening plates reinforce an annular combustor housing to maintain mechanical stability while reducing component count.

A gas turbine combustor uses clocked inner and outer dilution passages to de-swirl combustion products and reduce peak temperatures.

Segmented machining of the outer body and plenum resolves manufacturing cost contradictions while enabling effective premixing and impingement cooling.

Segmented plenums adjust fuel ratios across varying loads, resolving flame stability loss at low temperatures while maintaining reduced emissions.

Flared aperture walls reduce coating stress and prevent delamination in high-temperature gas turbine environments.

An insert device mixes fuel and gas using internal conduits before injection into engine cylinders.

Segmented liner holes generate targeted quench jets to rapidly mix fuel-rich gases, reducing NOx emissions without increasing structural complexity.

Segmented water injection stabilizes flame temperature while secondary fuel burns non-combusted products to limit CO, UHC, and NOx emissions.

A segmented fuel control system dynamically switches nozzle groups to optimize combustion characteristics during gas turbine operation.

Radial flanges with axial thermal conductors cool the seal third portion, preventing melting and material deposition on downstream components.

Pre-sintered coupons replace removed sections in cast-iron turbine diaphragms, avoiding welding heat damage while restoring joint integrity.

Alternating clockwise and counter-clockwise swirl in circumferential rows accelerates fuel-oxidizer mixing, reducing hot streaks at combustor walls.

Radial turbulators on a reduced-thickness combustor liner improve heat transfer while minimizing cooling air consumption in gas turbine transition regions.

Segmented dome-side cooling passages direct counter-swirled air through the cavity to enhance circulation and turbulence.

Stacked brazed joint structure with deformable titanium spacer accommodates thermal expansion differences between metal and ceramic parts.

Tangential fuel injection maintains swirling motion in the annular combustor, eliminating diffusors and deswirler vanes to reduce engine footprint.

A Helmholtz resonator embedded in a gas turbine combustor absorbs acoustic waves via baffle plate apertures.

Angled slots in an annular air mixer create differential offsets to induce swirling flows, reducing NOx emissions.

A unitary fuel nozzle design merges concentrically-aligned channels and inlets into a single monolithic body.