بررسی اثر ابعاد هندسی پاشنده بر عملکرد محفظه احتراق رانشگر دومولفه‌ای

نوع مقاله : مقاله پژوهشی

نویسندگان

1 دانشکده مهندسی هوافضا دانشگاه صنعتی امیرکبیر

2 دانشگاه صنعتی شریف

3 پژوهشکده سامانه های حمل و نقل فضایی

چکیده

استفاده از احتراق پیشرانه‌های خودمشتعل در رانشگرها، به ­دلیل دمای بالای محصولات احتراق، سبب افزایش ضربه ویژه می‌شود. در این مقاله، با استفاده از یک نرم‌افزار توسعه داده‌ شده، فرایند احتراق درون رانشگر دومولفه‌ای به‌صورت یک‌بعدی و با استفاده از سینتیک شیمیایی شبیه‌سازی می‌شود. در این راستا، مدل‌هایی برای پاشش، تبخیر قطرات، تشکیل فیلم مایع و محاسبات مربوط به انتقال حرارت از فیلم‌های مایع و گازی و احتراق به­ کار گرفته شده است. با استفاده از این نرم‌افزار، رفتار رانشگر آستریوم با سوخت منومتیل‌هیدرازین و اکسنده تتراکسید نیتروژن شبیه‌سازی شده است. با بهره‌گیری از مکانیزم شیمیایی گسترده 1619مرحله‌ای، نتایج شبیه‌سازی عملکرد رانشگر در دبی‌های مختلف اعتبارسنجی شده است. سپس، اثر ابعاد هندسی پاشنده بر فرایند تبخیر قطرات و نیز احتراق مورد بررسی دقیق قرار گرفته است. نتایج نشان می‌دهد که بزرگ­شدن پاشنده سبب افزایش طول تبخیر قطرات شده و ساختار شعله درون محفظه احتراق تغییر می‌کند، به ­نحوی که محصولات احتراق با دمای بالاتر وارد نازل شده و درنتیجه ضربه ویژه رانشگر افزایش می‌یابد. 

کلیدواژه‌ها

موضوعات


عنوان مقاله [English]

Investigation of injector dimension on the performance of combustion chamber of a bi-propelant thruster

نویسندگان [English]

  • Masoud EidiAttarZade 1
  • محمد فرشچی 2
  • Atiyeh Sarabadani 2
  • hamed khosrobeygi 2
  • ghazal davarnia 2
  • Alireza Ramezani 3
1 Aerospace Engineering Department Amirkabir University of Technology
2 Sharif University of Technology
3 Space Transportation Research Institute, Iranian Space Research Center, Tehran, Iran
چکیده [English]

Combustion of hypergolic propellants increases the specific impulse in the thrusters due to high temperature products. In this paper, the combustion process will be investigated through the axis of a bi-propellant thruster by an in-house code with chemical reaction mechanism. This code includes several models for injection, droplet evaporation, liquid film, combustion and heat transfer through liquid and gas films. The Astrium thruster with MMH as fuel and NTO as oxidizer has been simulated. By implementing a detail mechanism with 1619 steps, the thruster has been simulated at different total mass flow rates and results have been validated by experimental data. Then, injector dimension effects on the droplet evaporation and combustion have been investigated. Results show that by increasing the injector dimension, the droplet evaporation length increases, so the flame structure changes in the combustion chamber. Therefore, the combustion products enter the nozzle with higher temperature and as a result, the thruster specific impulse increases.

کلیدواژه‌ها [English]

  • Thruster
  • Hypergolic
  • Swirl injector
  • Monomethylhydrazine
  • Nitrogen Tetroxide
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