Continuous Combustion Chamber TRIZ Cases

A ceramic matrix composite heat shield mounts to a metallic combustor shell via a floating burner seal and retainer flange.

Segmented outlet ports disperse pilot fuel-air mixtures radially to stabilize main flames and reduce NOx emissions in gas turbines.

A parallel bypass valve routes fuel around the oil cooler when temperatures exceed limits, eliminating thermal recirculation loops and reducing aircraft weight.

Segmented downstream injectors limit secondary fuel volume to shorten high-temperature residence time, reducing NOx while maintaining engine efficiency.

Slidable fluid conduits mitigate vibration and bending stresses in bundled tube fuel injectors.

A continuous annular flow sleeve surrounds the combustion liner to route compressed working fluid for impingement and convective cooling.

Late injection assemblies in transition ducts shift hot gas flow to eliminate first stage nozzles.

An extractable insert with surface protrusions retains a flow sleeve via friction, eliminating on-site welding for field replacement.

Segmenting the flowpath into a detonation zone and conventional chamber stabilizes combustion across varying dynamic pressures.

An annular support frame with radial spokes reduces bending stress on cantilevered bundled tube segments by limiting vibration-induced deflection.

A model-based fuel-air ratio control system estimates combustor airflow in real time to adjust engine actuators.

Paired feed holes direct cooling air into separate exit cone and splash plate channels, preventing overheating from inadequate thermal management.

A unitary grommet with a bevel surface controls air flow through quench ports, reducing variability that causes combustion instability and emissions.

Cooling air holes in communication tubes direct film cooling along inner walls, suppressing metal temperature increases at the welding line.

Dual-hub spark ignition and centrifugal flameholding stabilize combustion, resolving power output limits in high-g rotary ramjet environments.

Radial passages in the baffle plate create uniform premixing to stabilize flame and reduce NOx emissions.

Integrates a triangular support body with slot-like recesses into gas turbine tiles to manage residual stresses during additive manufacturing.

Non-linear leakage paths reduce cooling air loss by up to 50% through segmented protrusion-recess interfaces that impede gas flow without complex structures.

Segmented box-like structures in the combustion chamber wall extend coolant residence time for efficient heat extraction while reducing total air consumption.

Effusion apertures in the shell reduce thermal stress on igniter bodies without increasing device complexity.

Varying pin fin densities guide coolant flow to resolve durability issues caused by circumferential rails.

Inner and outer air swirlers form recirculation zones that mix fuel with hot oxidizer, reducing NOx emissions while preventing thermo-acoustic instabilities.

A combustor liner jet wall increases cooling flow velocity to reduce compressed air requirements and lower NOx formation.

Asymmetric aperture sizing resolves misalignment issues in double-walled gas turbine combustion chambers, ensuring uniform coolant distribution.

Extracting feed ducts from external flows into flame-holder arms reduces head losses while maintaining thermal protection.

Fermat spiral pedestals eliminate weak lines in combustion tiles, preventing cracking under thermal loads.

Segmented injectors with vortex generators create turbulent eddies that improve combustion efficiency while reducing NOx emissions and preventing flashback.

Segmented panels with misaligned holes create a collecting space that traps particulates, resolving the trade-off between reliability and device complexity.

Lobed mixers blend combustion gases from dual annular liners to manage temperature profiles and resolve complexity issues in two-stage combustors.

Cascade holes and ribs in the ferrule generate low-energy vortex pairs that reduce flow instabilities and combustor dynamics.

Annular cooling aperture body directs airflow to film cool turbine engine quench aperture grommets and heat shields.

Radial cooling channels in a tubular wall structure resolve the conflict between mechanical strength and cooling efficiency in hot gas path components.

Staging fuel into the transition zone creates stable combustion zones that maintain temperatures below the thermal NOx formation threshold.

Streamlined fins in a combustor panel manage secondary flows from inner wall reaction forces, improving heat transfer efficiency.

Segmented quench apertures with varying intraset and interset distances create specific flow patterns to enhance fuel-air mixing in gas turbine combustors.

Staged mixing with swirled vanes prevents flashback and autoignition risks while maintaining axial flow velocity across varying hydrogen fuel compositions.

Grooved rope seals segment cooling air to control thermal stress distribution and extend CMC component life.

Jagged trailing edges on guide elements generate vortices to eliminate fuel accumulation in wake regions and improve combustion efficiency.

An external fluid circuit bypasses fuel injectors during non-base load operations, preventing air dilution and harmful emissions.

A jacketed core integrates catalyst material directly into molded component walls to define internal passages.

Removable ball bearing housing enables bushing radial movement within the fuel supply manifold assembly.

An aircraft burner uses an annular gap to direct hydrogen fuel into a protective film on the nozzle walls.

A cooling hole design featuring a metering portion and an expanding diffuser portion directs cooling air laterally across engine component surfaces.

Dividing detected fuel data into multiple databases reduces model discontinuity, enabling precise flow rate adjustments to prevent combustion vibrations.

Ejector mixer blends detonation exhaust with bypass air to cool gases, reducing nitrogen oxide emissions and fuel consumption.

Segmented CMC heat shields eliminate cooling fluid requirements by using adapters and collars to align with dome apertures, reducing turbine emissions.

A rotating detonation combustion system uses a fuel passage between outer walls to heat incoming fuel before injection into the detonation chamber.

Impingement plates direct cooling jets onto effusion plates to reduce overheating and thermal mechanical fatigue in gas turbine end caps.

A recessed dual fuel concentric nozzle design prevents emulsion penetration into the gas fuel duct during liquid fuel operation.