A subscale thrust - augmented - nozzle (TAN) rocket engine was designed, built, and hot-fire tested to demonstrate the validity of the TAN concept. As described in more detail in the immediately preceding article, in a TAN, during operation at sea level, thrust is augmented through injection and burning of secondary propellants (a fuel and an oxidizer) within the nozzle, downstream of the nozzle throat. The secondary propellants can be the same as, or different from, the primary propellants (the fuel and oxidizer burned during operation, in a near vacuum, in the absence of thrust augmentation).
In this demonstration effort, emphasis was placed on the design and analysis of the TAN injector, which is a cooled nozzle section, downstream of the throat, through which the secondary propellants are injected. The issues involved in designing the TAN injector are similar to those involved in designing conventional primary propellant injectors, with additional design challenge posed by the flow of combustion products over the TAN injector.
The subscale engine was designed to burn gaseous hydrogen and gaseous oxygen as the primary propellants and RP-1 (essentially, kerosene) and liquid oxygen (LOX) as the secondary propellants. The TAN injector was designed and built following a platelet approach, according to which flow passages are etched into thin metal sheets, which are then bonded together to form a monolithic structure. In addition to passages for RP-1 and LOX, the TAN injector contained passages that were parts of a water cooling circuit used during main (primary)- injector operation without TAN propellant flow. A water-cooled nozzle was attached downstream of the TAN injector to enable acquisition of performance data at an area ratio more representative of that of a launch rocket engine.
A thermal model of the TAN, implemented by use of commercial finite-element software, was used to perform a thermal analysis to determine whether the water cooling circuit afforded adequate cooling margin to ensure the thermal integrity of the TAN injector. Another thermal analysis was conducted to ensure that the RP-1-propellant-injection circuit maintained a wall temperature below its design coking-temperature limit. Temperature predictions were made to enable evaluation of pressure and of cycle life of the TAN injector structure based on low cycle fatigue criteria. The gas-side boundary conditions for the thermal analysis were based on a combination of test data and results of computational fluid dynamics simulations. The coolant-side boundary conditions were based on conventional correlations for the secondary propellants.
The subscale TAN rocket engine was subjected to a total of 28 tests at various combinations of operational parameters that included the main chamber pressure, propellant mixture ratios, and relative flow rates of the TAN (secondary) and main (primary) propellants. The TAN injector was found to be in excellent condition after the tests (see figure). Thermal data gathered during the tests consisted primarily of heat-load data for the main chamber, the TAN injector, and the portions of the nozzle other than the TAN injector. The calculated water flow rate for each component was based on data from a pre-hot-fire water-flow test of that component.
The bulk temperature rise of each component was calculated from readings of thermocouples located at the coolant inlet and outlet of the component. The bulk temperature rises of the water coolant for the main chamber (the chamber wherein the primary propellants are burned, upstream of the nozzle throat) were found to be fairly close to pre-test-predicted values because the main injector had been previously characterized. TAN operation was found to exert only a minimal effect on the main chamber. The heat load in the TAN injector and the non-tan-injector parts of the nozzle was found to be as much as 40 percent higher during TAN operation than during non-TAN operation. The increases in heat loads associated with TAN operation were found to depend on such parameters as the ratio of TAN flow to main-chamber flow and the main-chamber mixture ratio.
This work was done by Fred A. Ferrante and Felix F. Chen of GenCorp Aerojet for the Air Force Research Laboratory. For further information, download the free white paper at www.defensetechbriefs.com under the Mechanics/Machinery category. AFRL-0010
This Brief includes a Technical Support Package (TSP).
Thermal Design and Analysis of a Rocket-Engine TAN Injector
(reference AFRL-0010) is currently available for download from the TSP library.
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