بررسی تجربی اثر اسپین بر نرخ سوزش پیشرانه کامپوزیت دارای ذرات آلومینیوم

نوع مقاله: مقاله پژوهشی

نویسندگان

1 سازمان فضایی ایران

2 استادnدانشگاه صنعتی شریف

3 دانشگاه صنعتی شریف

چکیده

در این پژوهش به بررسی تجربی اثر شتاب بر نرخ سوزش یک پیشرانه جامد کامپوزیتی بر پایه HTPB دارای ذرات آلومینیوم به‌عنوان یکی از عوامل تعیین‌کننده فشار محفظه پرداخته شده است. بدین منظور از یک سامانه‌ی گریز از مرکز استفاده شد. با انتخاب گرین درون‌سوز، بردار شتاب در زمان سوزش همواره عمود بر سطح پیشرانه اعمال گردید. در این آزمایش‌ها فشار محفظه از 30 تا 80 بار و شتاب نیز از g2 تا g60 تغییر کرد. متغیر قابل‌اندازه‌گیری فشار محفظه احتراق بوده که جهت ارتباط آن به نرخ سوزش از کد تحلیل صفر بعدی استفاده گردید. در آزمایش‌هایی که شتاب کمتر از g5 بوده، نرخ سوزش تغییر محسوسی نداشته، اما در آزمایش‌هایی که شتاب در بازه g30 تا g60 قرار گرفته، نرخ سوزش از مقدار پایه شروع شده و در انتهای سوزش، به 5/1 برابر مقدار پایه خود می‌رسد.

کلیدواژه‌ها


عنوان مقاله [English]

Experimental Investigation on Spin Effect upon Burning Rate of Aluminized Composite Propellant

نویسندگان [English]

  • alireza mohammadi 1
  • Mohammad Farahani 2
  • Masood Goodarz 3
1 Iranian Space Agency
2 Sharif University of Technology
3 Sharif University of Technology
چکیده [English]

This research is conducted to study acceleration effect on the burning rate augmentation of an aluminized solid propellant based on HTPB as an effective factor determining combustion chamber pressure. A centrifugal experimental setup was designed to obtain a uniform acceleration field by rotating the test motor around its longitudinal axis. A cylindrical port propellant grain was used in the test motor which had an inner diameter of 30 mm, an outer diameter of 60 mm and a length of 52 mm, so acceleration vector was always perpendicular to inner burning surface of propellant. Inner radius of tube was small, so the magnitude of acceleration increased as the grain burned back. The pressure of combustion chamber was changed from 30 bar 80 bar by changing the nozzle throat and the magnitude of acceleration changed from 2g to 60g by changing the rotational speed of solid rocket motor. Pressure of combustion chamber was measured. An analytical 0D code was used to compute burning rate augmentation of propellant in acceleration field. Nozzle throat diameter is another parameter controlling the pressure field of combustion chamber of solid rocket motor. Thermomechanical erosion of nozzle throat is significant in aluminized propellant, so erosion of graphite throat insert was measured and taken into account in 0D code. Burning time of all tests was below 2 sec. due to small web of the grain. As investigated, at low acceleration level (below 5g), burning rate of propellant is not sensitive to acceleration while in the condition in which acceleration changed from 30g to 60g, burning rate augmentation ratio increased from 1 to 1.5. The transient behavior of burning rate augmentation in acceleration field was obvious in obtained results due to short burn time of SRM. At high acceleration level, inner cylindrical surface of inhibitor was coated with aluminum oxide particles.

کلیدواژه‌ها [English]

  • Acceleration
  • Burning Rate
  • Solid Rocket Motor
  • Composite Propellant