A new generalised model for liquid film cooling in rocket combustion chambers

An experimental investigation is conducted to bring out the effects of coolant injector configuration on film cooling effectiveness, film cooled length and film uniformity associated with gaseous and liquid coolants. A series of measurements are performed using hot air as the core gas and gaseous nitrogen and water as the film coolants in a cylindrical test section simulating a thrust chamber. Straight and compound angle injection at two different configurations of 30°–10° and 45°–10° are investigated for the gaseous coolant. Tangential injection at 30° and compound angle injection at 30°–10° are examined for the liquid coolant. The analysis is based on measurements of the film-cooling effectiveness and film uniformity downstream of the injection location at different blowing ratios. Measured results showed that compound angle configuration leads to lower far-field effectiveness and shorter film length compared to tangential injection in the case of liquid film cooling. For similar injector configurations, effectiveness along the stream wise direction showed flat characteristics initially for the liquid coolant, while it was continuously dropping for the gaseous coolant. For liquid coolant, deviations in temperature around the circumference are very low near the injection point, but increases to higher values for regions away from the coolant injection locations. The study brings out the existance of an optimum gaseous film coolant injector configuration for which the effectiveness is maximum.

Influence of coolant injector configuration on film cooling effectiveness for gaseous and liquid film coolants

: A one-dimensional analytical model of liquid film cooling in rocket engine combustion chambers is developed and compared with existing data. The vapor generated at the liquid interface greatly decreases the convective heat flux and is treated as a 'transition' process. The radiant heat is absorbed at the walls and transmitted to the liquid film by boiling, which can lead to burnout of the film. Downstream of the liquid film, the vapor provides continued thermal protection to the wall and is treated as a gaseous film cooling process. A standard correlation is expressed in differential form to allow for the nonuniform free-stream flow in the nozzle. Liquid film lengths are well predicted by the model. Downstream of the dry-out point, wall temperatures are well predicted up to the start of convergence. Downstream of the converging turn, a correction term, correlated by a centrifugal parameter, is required to account for increased mixing at the converging turn. Downstream of the throat, wall temperatures decrease very quickly because of acceleration of the boundary- layer gases, an effect not included in the model.

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Liquid Film Cooling in Rocket Engines

Review of some chemical propulsion technology advances suitable for future unmanned spacecraft applications. Discussed system varieties include liquid space-storable propulsion systems, advanced liquid monopropellant systems, liquid systems for rendezvous and landing applications, and low-thrust high-performance solid-propellant systems, as well as hybrid space-storable systems. To optimize the performance and operational characteristics of an unmanned interplanetary spacecraft for a particular mission, and to achieve high cost effectiveness of the entire system, it is shown to be essential that the type of spacecraft propulsion system to be used matches, as closely as possible the various requirements and constraints. The systems discussed are deemed to be the most promising candidates for some of the anticipated interplanetary missions.

Unmanned Planetary Spacecraft Chemical Rocket Propulsion

An evaluation of an existing analytical heat transfer model was made to develop the technology of boundary film/conduction cooled rocket thrust chambers to the space storable propellant combination oxygen difluoride/diborane. Critical design parameters were identified and their importance determined. Test reduction methods were developed to enable data obtained from short duration hot firings with a thin walled (calorimeter) chamber to be used quantitatively evaluate the heat absorbing capability of the vapor film. The modification of the existing like-doublet injector was based on the results obtained from the calorimeter firings.

R. N. Gurnitz

Papers

Boundary cooled rocket engines for space storable propellants