Fuel and Combustion

Fuel and Combustion

Numerical Simulation of the Effects of Inlet Shape on the Combustion Chamber Performance of a Hypersonic Projectile

Document Type : Original Article

Authors
mostafa mahmoodi 1
mohsen shojaei 2
afshin valimohammad 2
1 Propulsion group , Malek Ashtar University
2 Malek Ashtar University
10.22034/jfnc.2026.549461.1444
Abstract
In this study, an air intake was designed for the combustion chamber of a supersonic projectile operating at flight conditions of Mach 3.4 and an altitude of 15 km. Initially, a three-dimensional design methodology for the supersonic air intake is presented. Subsequently, the designed intake was integrated with the combustion chamber, and its performance was evaluated. To ensure the accuracy of the analyses, the simulation process was validated against the results of an existing combustion chamber. The results indicate that the air intake’s performance closely aligns with the calculated theoretical values. The maximum error observed was 6.25%, corresponding to the Mach number at the first section of the intake. The obtained total pressure recovery factor also showed a 2.43% difference compared to the calculated values. The performance of the combustion chamber, using the airflow supplied by the intake, was examined, yielding a combustion temperature of 1298 K and a combustion efficiency of 83.5%. Furthermore, the distance between the air inlet and the fuel injection point was reduced to investigate its effect on combustion. It was determined that at a distance-to-diameter ratio of 1, the combustion efficiency increased by 3.9%.

Highlights

· In the integrated analysis, the air intake efficiency exhibited a deviation compared to the calculated theoretical value. This discrepancy is attributed to the three-dimensional nature of the geometry, the precise modeling of flow viscosity, and the inherent viscous and shock losses along the flow path.

· The maximum deviation in the Mach number within the air intake occurred at the first cross-section, which did not exceed  relative to the theoretical value. Considering the negligible difference between the simulated mass flow rate entering the combustion chamber and the required mass flow rate for combustion, it can be stated that the air intake design was executed correctly.

· To evaluate the performance of the combustion chamber, based on the mass-weighted averaging method, a combustion temperature of  K was obtained inside the chamber. Using this temperature alongside the theoretical combustion temperature, the combustion chamber efficiency was determined to be .

· Based on the efficiencies calculated within the semi-free stream flow simulation, it can be concluded that a suitable air intake for the combustion chamber of a ducted ramjet was successfully designed for the specified flight conditions.

· By reducing the distance between the fuel injection location and the air intake from  times the diameter to  times the diameter, the mixing of fuel and oxidizer in the initial region of the combustion chamber improved, thereby increasing the effective residence time available for chemical reactions. This modification led to an earlier onset of combustion, reduced the loss of unburned fuel downstream, and ultimately resulted in an enhanced combustion efficiency.

Credit Author Statement

All authors contributed equally to the conceptualization of the article and writing of the original and subsequent drafts.

Data Availability Statement

Not applicable

Ethical Considerations

The authors avoided data fabrication, falsification, and plagiarism, and any form of misconduct.

Funding

This research did not receive any specific grant from funding agencies in the public, commercial, or not-for-profit sectors.

Keywords
Hypersonic Projectile
Combustion Chamber
Supersonic Air Inlet
Numerical Simulation

Subjects

در این پژوهش یک ورودی هوای مافوق‌صوت برای یک محفظه احتراق مشخص طراحی شد. محفظه احتراق مربوط به موشک داکت رم جت است. برای ورودی هوا الزاماتی از شرایط پروازی، قبیل دبی ورودی هوا و عدد ماخ ورودی به محفظه احتراق تعیین شد که به کمک آن یک ورودی هوا به طور کامل طراحی شد. در گام بعد نیاز بود تا ورودی هوای طراحی شده به‌صورت عددی و سه‌بعدی شبیه‌سازی شود. از آنجا که در این تحلیل احتراق نیز دخیل بود، یک مرجع مربوط به تحلیل محفظه احتراق موشک داکت - رم اعتبارسنجی شد که نتایج نشان دهنده دقت مناسب در تحلیل بود. سپس ورودی هوا و هندسه محفظه احتراق به صورت یکپارچه در نرم‌افزار فلوئنت شبیه‌سازی شد. به طور کل نتایج بدست آمده را می‌توان به صورت زیر ارائه داد:

1. برای اعتبارسنجی تحلیل محفظه احتراق از مرجع [25] استفاده شد. نتایج شبیه‌سازی نشان می‌دهد که مقدار دمای احتراق و بازده احتراق نسبت به مقدار ارائه شده در مرجع [25] به ترتیب دارای 3/4 و 6/8 درصد است.

2. در تحلیل یکپارچه شده مقدار بازده ورودی هوا نسبت به مقدار تئوری محاسبه شده دارای اختلاف 3.8 درصدی بود. این مقدار اختلاف به واسطه سه بعدی بودن هندسه، مدلسازی دقیق لزجت جریان و افت‌های موجود در مسیر جریان است.

3. مقدار بیشینه اختلاف عدد ماخ در ورودی هوا مربوط به مقطع اول بود که مقدار آن از 25/6 درصد نسبت به مقدار تئوری فراتر نرفته است. با توجه به اختلاف کم دبی وارد شده به محفظه احتراق در شبیه‌سازی و مقدار دبی لازم برای احتراق، می‌توان گفت که طراحی ورودی هوا به درستی انجام شده است.

4. برای ارزیابی عملکرد محفظه احتراق که به روش میانگین گیری جرمی انجام شد، مقدار دمای احتراق 2117 کوین بدست آمد. با استفاده از دمای بدست آمده و مقدار دمای تئوری احتراق (محاسبه شده در نرم‌افزار CEA) بازده محفظه احتراق برابر با 5/83 درصد محاسبه شد.

5. با توجه به بازده محاسبه شده در یک شبیه‌سازی جریان نیمه‌آزاد، میتوان این جمع بندی را انجام داد که در آن شرایط پروازی برای یک موشک داکت-رم جت یک ورودی هوای مناسب برای محفظه احتراق طراحی شده است. این ورودی هوا قادر است تا هوا را در شرایط پروازی طرح، جریان اکسید کننده را به درستی وارد محفظه احتراق کند تا در محفظه احتراق عمل احتراق به خوبی صورت گیرد.

6. با کاهش فاصله محل پاشش سوخت و ورودی هوا از 75/1 برابر قطر به 1 برابر قطر، اختلاط سوخت و اکسیدکننده در ناحیه ابتدایی محفظه احتراق بهبود یافته و زمان مؤثر در دسترس برای انجام واکنش‌های شیمیایی افزایش می‌یابد. این امر منجر به شروع زودتر احتراق و کاهش اتلاف سوخت نسوخته در پایین‌دست شده و در نهایت موجب افزایش بازده احتراق می‌شود.

در گام‌های بعدی می‌توان این تحلیل را فراتر برد. در ادامه لیستی از فعالیت‌های آتی ارائه شده است:

1. تحلیل عملکرد ورودی هوا و محفظه احتراق در شرایط خارج از طرح.

2. تحلیل پارامتریک و آنالیز حساسیت برای پارامترهای هندسی ورودی هوا و محفظه احتراق

3. بهینه‌سازی نتایج بدست آمده از عملکرد ورودی هوا و محفظه احتراق

4. قراردادن ورودی و محفظه احتراق در یک هندسه کامل بالستیک موشک و تحلیل کل سیستم به‌عنوان بالستیک داخلی و خارجی.

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