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Aerospace
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Fluid-Dynamics and Heat Transfer in Aerospace Propulsion Systems
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Fluid-Dynamics and Heat Transfer in Aerospace Propulsion Systems

A special issue of Aerospace (ISSN 2226-4310).

Deadline for manuscript submissions: closed (31 May 2023) | Viewed by 17538

Special Issue Editors

Dr. Francesco Battista

E-Mail Website
Guest Editor
Centro Italiano Ricerche Aerospaziali (CIRA), Via Maiorise, 81043 Capua, Italy
Interests: space propulsion; plasma physics; magnetic field optimization for plasma propulsion; cryogenic propellants
Special Issues, Collections and Topics in MDPI journals
Dr. Marco Pizzarelli

E-Mail Website
Guest Editor
Agenzia Spaziale Italiana, Via del Politecnico snc, 00133 Roma, Italy
Interests: rocket propulsion; computational fluid dynamics; heat transfer

Special Issue Information

Dear Colleagues, 

Prediction of the flow dynamics and heat transfer is central to the design process of aerospace propulsion systems. The motivation for this Special Issue is to present a series of research articles covering various experimental, numerical and theoretical aspects in the study of heat transfer and fluid dynamics (among other relevant factors) for aerospace propulsion applications. The central role of fluid dynamics and heat transfer in the design process of aerospace propulsion system is recognized among researchers due to the strong impact they have on the performance and reliability of any propulsion system. This Special Issue will fill the gap especially regarding the link between these two aspects toward finding common guidelines for the advanced design architectures of propulsion systems. Authors are encouraged to submit contributions linked to those areas, describing recent achievements applied to aerospace propulsion that are also supported by relevant experiments.

Dr. Francesco Battista
Dr. Marco Pizzarelli
Guest Editors

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. Aerospace is an international peer-reviewed open access monthly 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 2400 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.

Keywords

  • propulsion systems
  • fluid dynamics
  • heat transfer

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Published Papers (6 papers)

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Research

20 pages, 7904 KiB  
Article
Experimental Firing Test Campaign and Nozzle Heat Transfer Reconstruction in a 200 N Hybrid Rocket Engine with Different Paraffin-Based Fuel Grain Lengths
by Daniele Cardillo, Francesco Battista, Giuseppe Gallo, Stefano Mungiguerra and Raffaele Savino
Aerospace 2023, 10(6), 546; https://doi.org/10.3390/aerospace10060546 - 7 Jun 2023
Cited by 5 | Viewed by 2060
Abstract
Firing test campaigns were carried out on a 200 N thrust-class hybrid rocket engine, using gaseous oxygen as an oxidizer and a paraffin-wax-based fuel. Different fuel grain lengths were adopted to extend the fuel characterization under different operating conditions, and to evaluate rocket [...] Read more.
Firing test campaigns were carried out on a 200 N thrust-class hybrid rocket engine, using gaseous oxygen as an oxidizer and a paraffin-wax-based fuel. Different fuel grain lengths were adopted to extend the fuel characterization under different operating conditions, and to evaluate rocket performances and internal ballistics in the different configurations. In addition to data collected under a 220 mm propellant grain length, two further test campaigns were carried out considering 130 mm and 70 mm grain lengths. Two different injector types were adopted in the 130 mm configuration; in particular, a showerhead injection system was used with the aim to contain high-amplitude pressure oscillations observed during some firing tests in this engine configuration. Parameters such as the chamber pressure and temperature inside the graphite nozzle, space-averaged fuel regression rate and nozzle throat diameter were measured. The results allowed for the investigation of different issues related to hybrid rockets (e.g., fuel regression rate, engine performance, nozzle ablation under different conditions). The focus was mainly directed to the nozzle heat transfer, through the reconstruction of the convective heat transfer coefficient for different tests in the 70 mm grain length engine configuration. The reconstruction took advantage of the experimental data provided by the nozzle embedded thermocouple. Then, the experimental convective heat transfer coefficient was used to validate the results from some empirical correlations. The results showed significant differences between the experimental convective heat transfer coefficients when considering tests with different oxidizer mass flow rates. Furthermore, the predictions from the empirical correlations proved to be more reliable only in cases characterized by oxidizer-rich conditions. Full article
(This article belongs to the Special Issue Fluid-Dynamics and Heat Transfer in Aerospace Propulsion Systems)
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16 pages, 4558 KiB  
Article
Velocity Increment on Incidence Angle near the Leading Edge of the Compressor Cascade
by Xiaobin Xu, Baojie Liu, Xianjun Yu and Guangfeng An
Aerospace 2023, 10(5), 461; https://doi.org/10.3390/aerospace10050461 - 16 May 2023
Viewed by 1464
Abstract
The geometry of a compressor leading edge has an important effect on the aerodynamic performance at an off-designed incidence angle. The current geometric design methods of the leading edge are usually developed based on the flow characteristics at the designed incidence angle. However, [...] Read more.
The geometry of a compressor leading edge has an important effect on the aerodynamic performance at an off-designed incidence angle. The current geometric design methods of the leading edge are usually developed based on the flow characteristics at the designed incidence angle. However, few research focuses on the quantitative rules of the leading edge flow characteristics at the off-designed incidence angle in a compressor cascade. This situation restricts the further optimization and development of the leading edge geometry design method. In this paper, starting from the research of a potential cascade theory, the singularity point, where the surface velocity approaches infinity in the leading edge region, is eliminated by applying the characteristic that the ratio of the velocity increasement on the incidence angle in the plate cascade and the isolated plate flow is finite. Secondly, the equivalent pitch lengths based on 1/cos(β) and VI caused by a diffuser deceleration in the cascade passage were employed to correct the effect of the stagger angle. Finally, by introducing the isolated flow around the thick airfoil and considering the influence of the camber line geometry, a model of the variation of the surface velocity near the leading edge under the off-designed incidence angle, named the velocity increment on incidence angle, is derived from any compressor cascade. Hence, the relation between the off-designed incidence angle and the designed incidence angle of the surface velocity in a cascade blade is established, and it depends only on the geometrical parameters. Through a verification using numerical calculations and experimental measurement, the explicit formula for the velocity increment on incidence angle proposed in this paper has high precision near the leading edge. Full article
(This article belongs to the Special Issue Fluid-Dynamics and Heat Transfer in Aerospace Propulsion Systems)
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15 pages, 3977 KiB  
Article
A Simplified Thermal Analysis Model for Regeneratively Cooled Rocket Engine Thrust Chambers and Its Calibration with Experimental Data
by Matteo Fagherazzi, Marco Santi, Francesco Barato and Marco Pizzarelli
Aerospace 2023, 10(5), 403; https://doi.org/10.3390/aerospace10050403 - 26 Apr 2023
Cited by 3 | Viewed by 5179
Abstract
An essential part of the design of a liquid rocket engine is the thermal analysis of the thrust chamber, which is a component whose operative life is limited by the maximum allowable wall temperature and heat flux. A simplified steady-state thermal analysis model [...] Read more.
An essential part of the design of a liquid rocket engine is the thermal analysis of the thrust chamber, which is a component whose operative life is limited by the maximum allowable wall temperature and heat flux. A simplified steady-state thermal analysis model for regeneratively cooled rocket engine thrust chambers is presented. The model is based on semi-empirical correlations for the hot-gas and coolant convective heat transfer and on an original multi-zone approach for the wall conduction. The hot-gas heat transfer is calibrated with experimental data taken from an additively manufactured water-cooled nozzle that is connected to a combustion chamber either fed with decomposed hydrogen peroxide or decomposed hydrogen peroxide and automotive diesel. The thrust chamber (i.e., combustion chamber and nozzle) is designed to produce about 450 N of thrust when operating with a chamber pressure of 11 bar. For this application, the calibrated model predicts the total wall heat transfer rate very accurately and the temperature distribution within the wall structure with an uncertainty of a few tens of kelvins. This level of accuracy can be considered more than adequate for the design, and generally for engineering-type thermal analysis, of similar thrust chambers. Full article
(This article belongs to the Special Issue Fluid-Dynamics and Heat Transfer in Aerospace Propulsion Systems)
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34 pages, 18673 KiB  
Article
Swirl Flow and Heat Transfer in a Rotor-Stator Cavity with Consideration of the Inlet Seal Thermal Deformation Effect
by Yu Shi, Shuiting Ding, Peng Liu, Tian Qiu, Chuankai Liu, Changbo Qiu and Dahai Ye
Aerospace 2023, 10(2), 134; https://doi.org/10.3390/aerospace10020134 - 31 Jan 2023
Cited by 5 | Viewed by 2525
Abstract
In the typical structure of a turboshaft aero-engine, the mass flow of the cooling air in the rotor-stator cavity is controlled by the inlet seal labyrinth. This study focused on the swirl flow and heat transfer characteristics in a rotor-stator cavity with considerations [...] Read more.
In the typical structure of a turboshaft aero-engine, the mass flow of the cooling air in the rotor-stator cavity is controlled by the inlet seal labyrinth. This study focused on the swirl flow and heat transfer characteristics in a rotor-stator cavity with considerations of the inlet seal thermal deformation effect. A numerical framework was established by integrating conjugate heat transfer (CHT) analysis and structural finite element method (FEM) analysis to clarify the two-way aero-thermo-elasto coupling interaction among elastic deformation, leakage flow, and heat transfer. Simulation results showed that the actual hot-running clearance was non-uniform along the axial direction due to the temperature gradient and inconsistent structural stiffness. Compared with the cold-built clearance (CC), the minimum tip clearance of the actual non-uniform hot-running clearance (ANHC) was reduced by 37–40%, which caused an increase of swirl ratio at the labyrinth outlet by 5.3–6.9%, a reduction of the Nusselt number by up to 69%. The nominal uniform hot-running clearance (NUHC) was defined as the average labyrinth tip clearance. The Nusselt number of the rotating disk under the ANHC was up to 81% smaller than that under the NUHC. Finally, a clearance compensation method was proposed to increase the coolant flow and decrease the metal temperature. Full article
(This article belongs to the Special Issue Fluid-Dynamics and Heat Transfer in Aerospace Propulsion Systems)
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19 pages, 73637 KiB  
Article
Study on the Ablation Mechanism of the First Pulse Insulation Layer in a Double-Pulse Solid Rocket Motor
by Kaining Zhang, Chunguang Wang and Weiping Tian
Aerospace 2022, 9(10), 590; https://doi.org/10.3390/aerospace9100590 - 11 Oct 2022
Cited by 3 | Viewed by 2242
Abstract
In this work, numerical simulation and experimental research were carried out on ablation mechanism of the first pulse (I pulse) insulation layer in a double-pulse solid rocket motor (SRM). Firstly, based on the internal thermal environment of the typical double pulse SRM, the [...] Read more.
In this work, numerical simulation and experimental research were carried out on ablation mechanism of the first pulse (I pulse) insulation layer in a double-pulse solid rocket motor (SRM). Firstly, based on the internal thermal environment of the typical double pulse SRM, the internal flow field in combustion chamber of the double-pulse SRM with soft type pulse separation device (PSD) under the second pulse (Ⅱ pulse) working condition was numerically simulated. The results showed that the main reason for the difference of ablation in I pulse insulation layer was the difference of gas phase velocity. Secondly, based on the simulation analysis results, the experimental system for ablation of insulation layer was developed, and the ablation performance experiments under two gas phase velocities were carried out. It was found that a brittle carbonized layer had been formed on the surface of the insulation layer after the completion of I pulse work. In addition, at the beginning of Ⅱ pulse work, the suddenly generated gas flow made a denudation effect on the carbonized layer, which consumed a part of the carbonized layer. After the carbonized layer was peeled off, the gas flow continued to ablate the matrix of the insulation layer. Finally, the simulation analysis of the ablation process of the insulation layer under two gas phase velocities was carried out. The results showed that the velocity of the fuel gas is the main factor affecting the ablation rate of the insulation layer, which was consistent with the experimental results. It is proven that the model can be used to estimate the ablation amount of insulation of solid rocket motor. The conclusion can provide a significant reference for the internal heat protection design of the double-pulse SRM. Full article
(This article belongs to the Special Issue Fluid-Dynamics and Heat Transfer in Aerospace Propulsion Systems)
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20 pages, 9744 KiB  
Article
Study on Flow and Heat Transfer Performance of a Rectangular Channel Filled with X-Shaped Truss Array under Operating Conditions of Gas Turbine Blades
by Lei Xi, Jianmin Gao, Liang Xu, Zhen Zhao, Tao Yang and Yunlong Li
Aerospace 2022, 9(10), 533; https://doi.org/10.3390/aerospace9100533 - 21 Sep 2022
Cited by 1 | Viewed by 1760
Abstract
In this investigation, the heat transfer and flow capabilities of an X-shaped truss array cooling channel under various operating conditions of gas turbine blades were thoroughly studied. The influence laws of the inlet Reynolds number, inlet turbulence intensity, wall heat flux and cooling [...] Read more.
In this investigation, the heat transfer and flow capabilities of an X-shaped truss array cooling channel under various operating conditions of gas turbine blades were thoroughly studied. The influence laws of the inlet Reynolds number, inlet turbulence intensity, wall heat flux and cooling medium (air, steam) on the heat transfer and flow performance of the X-shaped truss array channel were analyzed and summarized. The empirical correlations of friction coefficients and average Nusselt numbers with maximum deviations less than ± 14% were fitted. The results show that the inlet Reynolds number has the most significant effect on the flow and heat transfer performance of the X-shaped truss array channel. When the inlet Reynolds number increases from 20,000 to 200,000, the average Nusselt number of the X-shaped truss array channel is increased by 3.92 times, the friction coefficient is decreased by 12.88%, and the comprehensive thermal coefficient is decreased by 31.19%. Compared with the medium turbulence intensity of Tu = 5%, the average Nusselt number, friction coefficient and comprehensive thermal coefficient of the X-shaped truss array channel at Tu = 20% are increased by 3.70%, 2.51% and 2.79%, respectively. With the increase in the wall heat flux, the friction coefficient of the X-shaped truss array channel roughly shows a trend of first decreasing and then increasing, while the average Nusselt number and the comprehensive thermal coefficient show a trend of first rapidly increasing and then slightly decreasing or remaining unchanged. Compared with air cooling, the average Nusselt numbers of the X-shaped truss array channel of steam cooling are increased by 6.30% to 9.54%, and the corresponding friction coefficients and comprehensive thermal coefficients are decreased by 0.11% to 0.55% and 2.63% to 5.59%, respectively. Full article
(This article belongs to the Special Issue Fluid-Dynamics and Heat Transfer in Aerospace Propulsion Systems)
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