A novel in-flight thrust modulation technique for SPRM: proof of concept | ||
| International Conference on Aerospace Sciences and Aviation Technology | ||
| Volume 20, Issue 20, May 2023, Pages 1-16 PDF (958.56 K) | ||
| Document Type: Original Article | ||
| DOI: 10.1088/1742-6596/2616/1/012020 | ||
| Authors | ||
| A M Abdelraouf* 1; O K Mahmoud2; M A Al-Sanabawy3 | ||
| 1M.Sc. Student, Rocket Department, Military Technical College, Cairo, Egypt. | ||
| 2Dr. Eng., Military Technical College, Cairo, Egypt. | ||
| 3Assoc. Prof., Military Technical College, Cairo, Egypt. | ||
| Abstract | ||
| Thrust modulation (termination or reduction) is one of the most sophisticated topics in rocket motors. Thrust termination is employed when combustion must be terminated to ensure proper stage separation, to avoid motor explosion, to attain certain range, or for research purposes. While liquid and hybrid propellant rocket motors have the ability to extinguish thrust, thrust termination in solid propellant rocket motors (SPRMs) need the use of specialized technology. Inflight thrust termination of (SPRMs) can be achieved by sudden and sufficient increase in throat area. Increasing throat area causes high depressurization rate which consequently leads to extinguishing the rocket motor. In contrast, in-flight thrust reduction can be theoretically attained either by reducing chamber pressure to a new equilibrium level or by reducing the divergent section length. The objective of the present paper is to prove the concept of using Linear Shaped Charge (LSC) for thrust modulation; a technique novel to SPRM applications. Thrust termination concept is proved via a set of static firing tests on a standard test motor. In addition, a mathematical model is developed to predict the drop in thrust due to cutting the nozzle at any arbitrary location. The model is validated by comparing with published experiments for thrust termination and reduction. A parametric study is conducted based on model results. It is confirmed that by cutting the nozzle along its divergent section, thrust reduction less than 20% can be attained. Further reduction can be achieved by cutting the nozzle along its convergent section. | ||
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