Continuous Combustion Chamber TRIZ Cases

A forward-facing step on a turbine band raises static pressure, limits airflow reversal, and reduces hot-gas ingestion into cavities.

Independent fuel and air paths let one injector mix liquid and gaseous fuels while reducing injector-set complexity and failure risk.

Radial rotary vanes separate fuel injection paths to mix fuels with different flame speeds and stabilize flame position in the combustor.

Bolted flange joints between wall cases are shielded from the hot flowpath to reduce thermal gradients and connection stress.

A switchable gas channel alternates between air and gaseous-fuel injection while liquid fuel remains available, helping prevent overheating and soot.

An adjustable rocket ejector raises inlet-air pressure and temperature, helping ramjet combustion begin at lower vehicle speeds.

Slot-shaped dilution openings and swirl vanes spread air, intensify mixing, and quench hot zones that drive NOx in gas-turbine combustors.

Secondary exhaust gases mix into the primary chamber to lower oxygen and flame temperatures, reducing NOx in turbine-engine combustion.

Brazed hollow members reduce selected turbine cooling passage exits where excess cooling exists, avoiding repeated filling and reopening.

Catalytic ammonia cracking supplies combustible cracked gas for initial ignition, while controlled ammonia flow helps prevent incomplete combustion.

Gyroid-like porous geometries provide frequency-targeted acoustic damping in combustor components while supporting durable, retrofit-friendly designs.

Independent fuel branches use dedicated heat exchangers to adjust fuel temperature by injection stage, helping limit hydrogen flashback risk.

Circumferential protrusions and recesses improve reactant mixing, reduce pressure losses, and sustain the detonation wave at low power.

This AFS injector sends high-pressure air and fuel toward the mixing chamber together for rapid premixing and lower emissions.

Multiple fuel nozzles, a venturi injector, and controller-based fuel management support ammonia, hydrogen, or natural gas with lower NOx.

Two fuel-oil heat exchangers and an interposed fuel pump manage fuel temperature, oil cooling, and combustion efficiency across varying fuels.

Cooling holes route air around bolt heads and shanks to limit thermal distress in combustor deflectors and extend assembly life.

An outward-diverging inner passage helps address poor hydrogen fuel-air mixing in turbine combustion through controlled flow expansion.

A head groove connects the liner gap and combustion chamber, helping route cooling air to suppress fixture-head heating during secure liner attachment.

Pulsed terahertz reflections assess coating properties and estimate remaining life without destructive testing.

Tangential feed conduits create swirling hydrogen–gas flow in a turbine injector to improve mixing and limit flame backflow.

Cooling paths through the backstop and annular wall reduce hot spots around gas turbine combustor dilution holes.

A rich hydrogen-air mixture enters the inner channel while outer air forms a lean flame, balancing stability, NOx, and flashback resistance.

Cooling holes in hollow caps direct air around turbine coupling studs, limiting hot-gas exposure and thermal distress.

Variable fluid injection and flame stabilizers shift combustion axially, reducing turbine-stage temperature variation during expansion.

A cold fuel-injection zone insulates the hot combustion zone in an afterburner, reducing external noise while sustaining thrust.

Multiple sub-element mixers vary local air-fuel ratios in staged gas-turbine combustion, helping reduce NOx and particle emissions.

Separate combustor liners isolate adjacent fuel–oxidant flow fields, while swirling flow improves mixing and reduces unburnt components.

Thermal compression tanks use heat and controlled valves to pressurize liquid hydrogen for aircraft combustor delivery without traditional pumps.

Selective nozzle groups and a shared fuel manifold tune combustor injection across power levels while improving mixing and reducing NOx emissions.

A second combustor switches between steam injection and fuel-air combustion to address CO/NOx emissions and sustain thrust augmentation.

Annular plenums, pre-filming, and vortex induction spread fuel through the combustor to limit hot and cold zones, NOx, and wear.

Asymmetric steam injection disrupts combustion-chamber acoustic modes to reduce engine noise and stabilize turbine combustion.

A radially diverging annular passage with air and fuel swirlers improves hydrogen fuel-air mixing for turbine combustion efficiency.

The backstop and annular peripheral wall route cooling air to the liner panel, limiting hot spots and temperature variation around gas turbine combustor dilution passages.

Steam mixed with hydrogen in the nozzle moderates flame speed and peak temperature, limiting flashback and NOx emissions.

Raising gas turbine fuel to at least 135°C at combustor entry improves spray behavior while controlled oil-to-fuel heat transfer protects components.

A piston check valve stabilizes differential pressure in an aircraft gas-turbine fuel supply while supporting precise flow control.

Mirrored detachable struts create recirculation zones and oblique shock waves to stabilize flames, supporting transition to oblique detonation.

When a reservoir fault is detected in flight, pumping or pressure equalization redirects fuel to a second reservoir for uninterrupted powerplant delivery.

Cold-side fastening and deflector arms protect combustor dome joints from hot gases, reducing thermal distress and air leakage.

Separate purging of liquid and gaseous fuel lines increases workload; a shared screw compressor and pipeline network stabilizes multi-turbine purging.