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The Rotating Detonation Engine has been seen as the next step for rocket propulsion applications with the advent of the Rotating Detonation Rocket Engine, an engine configuration developed by the Air Force Research Laboratory. In an effort to flight-test this engine and provide a dataset to train detonation-based simulation, the Rotating Detonation Rocket Engine has been tested in a collaborative effort including the University of Central Florida. For this testing, a thrust stand was developed to obtain the key thrust and impulse data necessary for advancing the engine to flight readiness. This thrust stand utilized the small-scale of the Rotating Detonation Rocket Engine to motivate an axial-loading measurement approach and the integration of an automatic-calibration subassembly, altogether which allows for incredibly accurate thrust measurements from an engine. Results using this thrust stand for two similar engine configurations are shown to validate the operation of the thrust stand.
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Effect of Nozzle Throat Diameters on Kerosene-NOS Liquid Rocket Performance
This paper presents experimental results for the performance effects of different converging- diverging graphite nozzle throat diameters on an in-house developed kerosenenitrous oxide liquid rocket test stand. The project aims to enhance the performance and efficiency of small-scale liquid rocket engines by experimentally investigating the effects of nozzle throat diameter on thrust and specific impulse. By confirming the correlation between nozzle geometry and the experimental data, it provides valuable insight for improving propulsion systems and components used in experimental rocketry such as sounding rockets. This study will evaluate two different nozzle throat diameters under varying propellant pressures and mass flow rates. The liquid rocket test stand consists of an external aluminum casing with a combustion chamber measuring 20” in length with an outer diameter of 76 mm and an internal diameter of 1.66”. The nozzle throat diameter tested will be 58/64” and 60/64”, each with a fixed exit diameter of 1.82”. Experimental results were collected over a range of total mass flow rates using data acquisition systems and analyzed using graphs and trend lines. The results indicate that as the throat diameter increases, the thrust output and specific impulse increase, although the results are inconclusive due to leaks and a back flame during testing, possibly skewing the results. The ablative wear was analyzed based on the nozzle throat size and mass flow rate. The knowledge gained from this study can be used to prevent future accidents for small-scale liquid rocket engine test stands and verify if the trends seen will be applicable to different nozzle materials and find the optimum nozzle throat diameter.
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- Award ID(s):
- 1911370
- PAR ID:
- 10637661
- Publisher / Repository:
- American Institute of Aeronautics and Astronautics
- Date Published:
- ISBN:
- 978-1-62410-755-9
- Format(s):
- Medium: X
- Location:
- Multiple Locations
- Sponsoring Org:
- National Science Foundation
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- Free Publicly Accessible Full Text
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Accepted Manuscript1.0
- Conference Paper:
- https://doi.org/10.2514/6.2025-99101
