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

The main objective of this paper is to numerically simulate the effects of the air intake geometry on the combustion chamber performance of a supersonic projectile. The geometry of the combustion chamber is predefined, and its combustion conditions, including the oxidizer inlet mass flow rate supplied by the air intake, are known. Therefore, based on the design requirements of the combustion chamber and the flight conditions, it is necessary to design an air intake capable of delivering the oxidizer into the combustion chamber under specific flow conditions. Following the design phase, the final geometry is modeled, and the flow behavior from the air intake to the nozzle exit is simulated using Computational Fluid Dynamics (CFD) analysis using ANSYS Fluent software. The combustion efficiency can then be determined based on the simulation results.

In general, the novelties of this research can be summarized as follows:

·         Design of an air intake for a ducted ramjet.

·         Integration of the combustion chamber and the air intake of the ducted ramjet, and simulation under semi-free stream flow conditions.

·         Numerical simulation of the integrated geometry and the simultaneous determination of the performance parameters for both the air intake and the combustion chamber.

Furthermore, the distance between the air intake and the fuel injection location was varied. This distance, defined as a ratio of the chamber diameter, was initially set to 1.75 times the diameter in accordance with the combustion chamber requirements. After conducting the initial analyses and evaluating the aforementioned outputs, this distance ratio was reduced to 1 in order to investigate its effect on the combustion efficiency.

Keywords
Hypersonic Projectile
Combustion Chamber
Supersonic Air Inlet
Numerical Simulation
Subjects

در این پژوهش برای محفظه احتراق یک پرتابه مافوق صوت یک ورودی هوا در شرایط پروازی عدد ماخ 4/3 و ارتفاع پروازی 15 کیلومتر طراحی شده است. در ابتدا یک روش طراحی برای ورودی هوای مافوق صوت به صورت سه بعدی ارائه شده است. در ادامه ورودی هوای طراحی شده به محفظه احتراق متصل شده و عملکرد آن بررسی شده است. به منظور اطمینان از صحت تحلیل‌های انجام شده، روند تحلیل با نتایج یک محفظه احتراق موجود، اعتبارسنجی شد. در ادامه هندسه یکپارچه شده شبکه‌بندی و تحلیل شد. بر اساس نتایج عملکرد ورودی هوا به مقادیر اشاره شده در نتایج محاسبه شده، نزدیک است. بیشترین مقدار خطا به میزان 25/6 درصد، مربوط به عدد ماخ مقطع اول ورودی هوا است. مقدار ضریب بازیابی فشار بدست آمده نیز در مقایسه با مقادیر محاسبه شده، 43/2 خطا دارد. عملکرد محفظه احتراق در مواجهه با هوای دریافتی از ورودی هوا مورد بررسی قرار گرفت که مقدار دمای احتراق 1298 کلوین و بازده احتراق نیز 5/83 درصد محاسبه شد. در ادامه فاصله محل پاشش سوخت و ورودی هوا کاهش یافت تا تأثیر آن بر احتراق بررسی شود که مشخص شد در نسبت فاصله 1 برابر قطر، راندمان احتراق 9/3 درصد افزایش داشته است.

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