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Propulsion and Combustion in Aerospace Systems

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A special issue of Energies (ISSN 1996-1073). This special issue belongs to the section "J: Thermal Management".

Deadline for manuscript submissions: closed (20 July 2021) | Viewed by 16485

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Prof. Dr. Jeong Yeol Choi

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Guest Editor

Department of Aerospace Engineering, Pusan National University, Busan 46241, Republic of Korea
Interests: propulsion and combustion phenonmena in rocket, scramjet, and detonation engines; detonation; supersonic combustion; turbulent combustion; supercritical combustion; high-resolution numerical methods; high-performance computing; combustion experiments; visualization of high-speed flows
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Special Issue Information

Dear Colleagues,

The Guest Editor is inviting submissions to a Special Issue of Energies on the subject area of “Propulsion and Combustion in Aerospace Systems”. All aerospace systems need propulsion to generate thrust for acceleration or for drag compensation. Thermal energy from combustion was the main source of propulsion in aerospace systems from the beginning and will continue as the mainstream in the future, differently from ground and surface transportations. Faster and higher are everlasting themes for this area, but cleaner is becoming a new one.

This Special Issue will deal with new findings and developments on propulsion and combustion in aerospace systems.

Topics of interest for publication include but are not limited to:

Gas turbine, ramjet/scramjet, rocket, and detonation engines;
Aerothermodynamics in propulsion systems;
Combustion physics in propulsion systems, including turbulent combustion, supercritical combustion, multiphase combustion, supersonic combustion, and detonation;
Fluid dynamic, turbulence ,and combustion models for propulsion systems;
Numerical and computational techniques for internal aerothermodynamics and combustion;
Visualization and measurement techniques for propulsion and combustion.

Prof. Dr. Jeong Yeol Choi
Guest Editor

Manuscript Submission Information

Manuscripts should be submitted online at www.mdpi.com by registering and logging in to this website. Once you are registered, click here to go to the submission form. Manuscripts can be submitted until the deadline. All submissions that pass pre-check are peer-reviewed. Accepted papers will be published continuously in the journal (as soon as accepted) and will be listed together on the special issue website. Research articles, review articles as well as short communications are invited. For planned papers, a title and short abstract (about 100 words) can be sent to the Editorial Office for announcement on this website.

Submitted manuscripts should not have been published previously, nor be under consideration for publication elsewhere (except conference proceedings papers). All manuscripts are thoroughly refereed through a single-blind peer-review process. A guide for authors and other relevant information for submission of manuscripts is available on the Instructions for Authors page. Energies is an international peer-reviewed open access semimonthly journal published by MDPI.

Please visit the Instructions for Authors page before submitting a manuscript. The Article Processing Charge (APC) for publication in this open access journal is 2600 CHF (Swiss Francs). Submitted papers should be well formatted and use good English. Authors may use MDPI's English editing service prior to publication or during author revisions.

Published Papers (6 papers)

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Research

25 pages, 9580 KiB

Open AccessArticle

Analysis of Propellant Weight under Re-Entry Conditions for a Reusable Launch Vehicle Using Retropropulsion

by Yongchan Kim, Hyoung-Jin Lee and Tae-Seong Roh

Energies 2021, 14(11), 3210; https://doi.org/10.3390/en14113210 - 31 May 2021

Cited by 1 | Viewed by 2537

Abstract

In this study, a minimum amount of required propellant was calculated by analyzing the sequence with various re-entry conditions. This study aims to obtain data related to variation in trajectory and required propellant weight according to various re-entry scenarios. The drag coefficient at various altitudes, velocities, and thrust was calculated through numerical simulations to raise the reliability of the results. The calculation results were compared to the optimal values extracted from the genetic algorithm. It was observed that the duration of the entry-burn phase is dominant to the total required propellant weight. As a general tendency, high entry-burn starting altitude, high ending Mach number, and low landing-burn starting thrust make the required propellant weight low. However, if the entry-burn ending condition is set to the Mach number, it is necessary to select an appropriate re-entry condition. Additionally, from comparisons with the optimized results, it was confirmed that accurate calculation of the drag coefficient is important to succeed a soft landing of RLV. Full article

(This article belongs to the Special Issue Propulsion and Combustion in Aerospace Systems)

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14 pages, 5826 KiB

Open AccessArticle

Experimental Study on Dynamic Combustion Characteristics in Swirl-Stabilized Combustors

by Donghyun Hwang and Kyubok Ahn

Energies 2021, 14(6), 1609; https://doi.org/10.3390/en14061609 - 14 Mar 2021

Cited by 3 | Viewed by 1839

Abstract

An experimental study was performed to investigate the combustion instability characteristics of swirl-stabilized combustors. A premixed gas composed of ethylene and air was burned under various flow and geometric conditions. Experiments were conducted by changing the inlet mean velocity, equivalence ratio, swirler vane angle, and combustor length. Two dynamic pressure sensors, a hot-wire anemometer, and a photomultiplier tube were installed to detect the pressure oscillations, velocity perturbations, and heat release fluctuations in the inlet and combustion chambers, respectively. An ICCD camera was used to capture the time-averaged flame structure. The objective was to understand the relationship between combustion instability and the Rayleigh criterion/the flame structure. When combustion instability occurred, the pressure oscillations were in-phase with the heat release oscillations. Even if the Rayleigh criterion between the pressure and heat release oscillations was satisfied, stable combustion with low pressure fluctuations was possible. This was explained by analyzing the dynamic flow and combustion data. The root-mean-square value of the heat release fluctuations was observed to predict the combustion instability region better than that of the inlet velocity fluctuations. The bifurcation of the flame structure was a necessary condition for combustion instability in this combustor. The results shed new insight into combustion instability in swirl-stabilized combustors. Full article

(This article belongs to the Special Issue Propulsion and Combustion in Aerospace Systems)

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20 pages, 7475 KiB

Open AccessArticle

Experimental Investigation of Detonation Propagation Modes and Thrust Performance in a Small Rotating Detonation Engine Using C₂H₄/O₂ Propellant

by Hyung-Seok Han, Eun Sung Lee and Jeong-Yeol Choi

Energies 2021, 14(5), 1381; https://doi.org/10.3390/en14051381 - 03 Mar 2021

Cited by 11 | Viewed by 3834

Abstract

A small rotating detonation engine (RDE) model and the corresponding experimental setup were constructed for the experimental investigation of the detonation propagation characteristics and thrust performance of a circular RDE. Experiments were conducted at a range of 0.3–2.5 equivalence ratio with a total mass flow rate of less than 180.0 g/s using a C₂H₄/O₂ mixture. Irregularly unstable detonative combustion occurs immediately after the detonation initiation, which includes initiation, propagation, decaying, and the merging of detonation waves. Following this, periodically unsteady detonative combustion occurs in the circular channel, resulting in the stable operation of the RDE. During stable operation, two detonation waves are predominant, rotating along the wall at a speed lower than the Chapman–Jouguet (CJ) detonation speed. The characteristic velocity efficiency is approximately 73% on average. The low characteristic velocity efficiency is presumed to be caused by the unoptimized combustion channel and the poor mixing efficiency owing to the two-dimensional injector configuration. The effect of the RDE component design and the RDE flow field characteristics need to be further investigated for improving the performance of the RDE. Full article

(This article belongs to the Special Issue Propulsion and Combustion in Aerospace Systems)

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18 pages, 19530 KiB

Open AccessArticle

Numerical Study on the Flow and Heat Transfer Characteristics of a Second Throat Exhaust Diffuser According to Variations in Operating Pressure and Geometric Shape

by Seonghwi Jo, Sanghyeon Han, Hong Jip Kim and Kyung Jin Yim

Energies 2021, 14(3), 532; https://doi.org/10.3390/en14030532 - 20 Jan 2021

Cited by 11 | Viewed by 2341

Abstract

A numerical study was conducted to investigate the flow and heat transfer characteristics of a supersonic second throat exhaust diffuser for high-altitude simulations. The numerical results were satisfactorily validated by the experimental results. A subscale diffuser using nitrogen was utilized to investigate starting pressure and pressure variation in the diffuser wall. Based on the validated numerical method, the flow and heat transfer characteristics of the diffuser using burnt gas were evaluated by changing operating pressure and geometric shape. During normal diffuser operation without cooling, high-temperature regions of over 3000 K appeared, particularly near the wall and in the diffuser diverging section. After cooling, the flow and pressure distribution characteristics did not differ significantly from those of the adiabatic condition, but the temperature in the subsonic flow section decreased by more than 1000 K. Furthermore, the tendency of the heat flux from the diffuser internal flow to the wall was similar to that of the pressure variations, and it increased with operating pressure. It was confirmed that the heat fluxes of the supersonic and subsonic flows in the diffuser were proportional to the operating pressure to the 0.8 and −1.7 power, respectively. In addition, in the second throat region after separation, the heat flux could be scaled to the Mach number ratio before and after the largest oblique shock wave because the largest shock train affected the heat flux of the diffuser wall. Full article

(This article belongs to the Special Issue Propulsion and Combustion in Aerospace Systems)

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23 pages, 7906 KiB

Open AccessArticle

Real-Gas-Flamelet-Model-Based Numerical Simulation and Combustion Instability Analysis of a GH₂/LO_X Rocket Combustor with Multiple Injectors

by Won-Sub Hwang, Bu-Kyeng Sung, Woojoo Han, Kang Y. Huh, Bok Jik Lee, Hee Sun Han, Chae Hoon Sohn and Jeong-Yeol Choi

Energies 2021, 14(2), 419; https://doi.org/10.3390/en14020419 - 13 Jan 2021

Cited by 9 | Viewed by 2739

Abstract

A large eddy simulation (LES) and combustion instability analysis are performed using OpenFOAM for the multiple shear-coaxial injector combustor DLR-BKD (in German Deutsches Zentrum für Luft–Brennkammer D, German Aerospace Center–Combustion Chamber D), which is a laboratory-scale combustor operating in a real-gas environment. The Redlich–Kwong–Peng–Robinson equation of state and steady-laminar flamelet model are adopted in the simulation to accurately capture the real-gas combustion effects. Moreover, the stable combustion under the LP4 condition is numerically analyzed, and the characteristics of the combustion flow field are investigated. In the numerical simulation of the combustion instability, the instability is generated by artificially superimposing the 1st transverse standing wave solution on the stable combustion solution. To decompose the combustion instability mode, the dynamic mode decomposition method is applied. Several combustion instability modes are qualitatively and quantitatively identified through contour plots and graphs, and the sustenance process of the limit cycle is investigated. Full article

(This article belongs to the Special Issue Propulsion and Combustion in Aerospace Systems)

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14 pages, 6716 KiB

Open AccessArticle

Experimental Study of the Combustion Efficiency in Multi-Element Gas-Centered Swirl Coaxial Injectors

by Seongphil Woo, Jungho Lee, Yeoungmin Han and Youngbin Yoon

Energies 2020, 13(22), 6055; https://doi.org/10.3390/en13226055 - 19 Nov 2020

Cited by 8 | Viewed by 2405

Abstract

The effects of the momentum-flux ratio of propellant upon the combustion efficiency of a gas-centered-swirl-coaxial (GCSC) injector used in the combustion chamber of a full-scale 9-tonf staged-combustion-cycle engine were studied experimentally. In the combustion experiment, liquid oxygen was used as an oxidizer, and kerosene was used as fuel. The liquid oxygen and kerosene burned in the preburner drive the turbine of the turbopump under the oxidizer-rich hot-gas condition before flowing into the GCSC injector of the combustion chamber. The oxidizer-rich hot gas is mixed with liquid kerosene passed through combustion chamber’s cooling channel at the injector outlet. This mixture has a dimensionless momentum-flux ratio that depends upon the dispensing speed of the two fluids. Combustion tests were performed under varying mixture ratios and combustion pressures for different injector shapes and numbers of injectors, and the characteristic velocities and performance efficiencies of the combustion were compared. It was found that, for 61 gas-centered swirl-coaxial injectors, as the moment flux ratio increased from 9 to 23, the combustion-characteristic velocity increased linearly and the performance efficiency increased from 0.904 to 0.938. In addition, excellent combustion efficiency was observed when the combustion chamber had a large number of injectors at the same momentum-flux ratio. Full article

(This article belongs to the Special Issue Propulsion and Combustion in Aerospace Systems)

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