Continuous Combustion Chamber TRIZ Cases

Varying radial spacing in a turbine diffuser nozzle minimizes circumferential speed differences to reduce head losses and stabilize combustion.

Interrupting thermal conduction paths with a thermal break reduces silicide formation at the interface.

Integrated fuel injector heat exchanger merges thermal exchange with structural mounting to simplify retrofitting into existing gas turbine engines.

A combustion device routes fuel ammonia to a catalyst reduction unit, utilizing the gas as a reducing agent for exhaust nitrogen oxides.

A spring-biased seal carrier moves dynamically to maintain wedge seal contact against turbine housing surfaces.

A thermal coupling assembly transfers heat from fuel conduits to a sink using a conductive element and insulative wrap.

Segmented injector rings distribute over 200 injection sites radially to resolve low site count limits and optimize exhaust emission management.

Elongated depressions in flame holders accommodate thermal expansion, reducing stress concentrations caused by high temperature gradients.

A turbofan engine integrates the Haber Process to synthesize ammonia and urea from exhaust gases.

Guide vanes redirect high-momentum cooling flow into wake regions behind bluff bodies, resolving thermal performance issues in gas turbine liners.

A gas turbine system pressurizes injection water using boosted hydrogen fuel pressure through a dedicated passage.

Differentiated cooling hole inclinations create opposing swirl flows around fuel supply holes, preventing hot spots while maintaining stable combustion.

An ammonia decomposition apparatus thermally breaks down ammonia using exhaust heat, supplying hydrogen fuel to reduce carbon dioxide emissions.

Polygonal nozzle passage geometry reduces dimensional deviations and assembly time in additive manufactured turbine engine fuel injectors.

A combustor dome bifurcates airflow into inner and outer streams to distribute combustion gases through the chamber.

Axially movable blocking member adjusts port area to resolve combustion capacity versus flame lifting trade-off across varying Btu inputs.

Segmented radial walls coupled via a connecting member define a flow cavity that controls leakage and pressure drop in heat engines.

A compliant seal spans gaps between ceramic matrix composite heat shields to maintain contact during operation.

Cooled compressor air purges residual fuel from the gas turbine distribution system, preventing hot gas recirculation and ignition risks during oil mode.

A control system adjusts heated water flow to a fuel gas heater using real-time temperature feedback for precise thermal regulation.

Segmented cooling flow reduces pressure losses in transition nozzles while maintaining component lifetime through targeted film cooling.

Upstream and downstream outlet segmentation preserves structural rigidity while maintaining comprehensive cooling coverage for gas turbine components.

An off-axis elbow reorients the combustion chamber flow to reduce axial length, resolving maintenance accessibility issues in high-pressure engine cases.

Segmented plating elements tilted relative to gas flow guide cooling air distribution, reducing thermal stress and wear on turbine components.

A jet engine combustor ramp with an inclination angle of at least 13.5 degrees promotes fuel-air mixing.

Fuel nozzle swirl vanes with mixed radial profiles stabilize flames and prevent flashback in gas turbines.

Radial and axial staging with a trapped vortex cavity reduces NOx emissions by optimizing local equivalence ratios across combustion zones.

Segmenting the backside cavity via a midrail allows strategic cooling hole placement in hot areas, reducing rail strike risk during manufacturing.

Intermediary push rod transmits uniform force to align combustion liners, preventing hula seal damage from torsional loading.

Nested cavities in a single liner damp multiple combustion frequencies, reducing installation space and cost.

Segmented turbulators attach via brazing or resistance welding, solving manufacturing bottlenecks on complex 3D geometries.

Differential axial length micro-mixer nozzles mix fuel and air to mitigate high-frequency dynamic-tone vibrations in turbomachine combustors.

An inner reactor housing separates heat and catalyst beds from the outer wall to minimize thermal loss in liquid monopropellant thrusters.

A fuel injector with a bluff body portion anchors the flame within a gas turbine combustor.

Annular trapped vortex cavity pilot stabilizes flames using rotating vortices, reducing flow losses and extending flameholder life in gas turbine engines.

Periodic air supply valves and dual discharge passages suppress cat back phenomenon while reducing energy consumption.

Inverting platefin placement to the cooler outer liner prevents thermal erosion and collapse of the cooling air passage gap.

A venturi structure energizes the wake region downstream of turbine components to restore uniform airflow distribution.

Air opening allows flow through cavity and pilot fuel nozzle to purge residual fuel, preventing coking in compact gas turbine arrangements.

Intermediate drain holes move collected liquid from the main damping volume to a secondary chamber, preventing deflagration risks in gas turbine combustors.

A double-jet film cooling structure uses a separating section to form swirl flows that push the cooling medium against the wall surface.

Segmenting the front panel into distinct hot and cold walls resolves the trade-off between material usage and mechanical stability in gas turbine combustors.

Segmenting the closed annular chamber into parallel linear modules enables stable detonation across wide thrust ranges.

Segmented pilot cone assemblies reduce thermal stresses and extend service life by allowing independent component movement.

A gas turbine control system measures fuel composition using fast infrared analyzers to dynamically distribute fuel across burner groups.

Asymmetric diffuser film cooling holes prevent manufacturing overlap while maintaining minimal clearance distances for improved turbine cooling.

Rotational pivot motion via an annular spacer reduces axial translation distance by 3.6 mm, resolving engine length and weight constraints.

Plunged inlet holes in inner and outer liners create counter-swirl doublet flows to reduce combustion temperature non-uniformity.

A spring-loaded cap assembly maintains seal contact between a tube sleeve and an aft plate, ensuring even cooling of tube tips despite thermal growth.

A unitary flow path assembly merges combustor and turbine boundaries into a single structure to reduce engine weight.