Aerospace

Research

25 pages, 7141 KiB

Open AccessArticle

Effects of Local Mixing Ratios and Mass Flow Rates on Combustion Performance of the Fuel-Rich LOX (Liquid Oxygen)/kerosene Gas Generator in the ATR (Air Turbo Rocket) Engine

by Yuankun Zhang, Qingjun Zhao, Bin Hu, Qiang Shi, Wei Zhao and Xiaorong Xiang

Aerospace 2023, 10(6), 545; https://doi.org/10.3390/aerospace10060545 - 7 Jun 2023

Cited by 3 | Viewed by 2147

Abstract

This paper presents a numerical simulation analysis of the flow and combustion characteristics of a fuel-rich LOX (liquid oxygen)/kerosene gas generator in an ATR (air turbo rocket) engine, examining the effects of local parameters on the combustion flow field and performance. The analysis [...] Read more.

This paper presents a numerical simulation analysis of the flow and combustion characteristics of a fuel-rich LOX (liquid oxygen)/kerosene gas generator in an ATR (air turbo rocket) engine, examining the effects of local parameters on the combustion flow field and performance. The analysis considers variations in unit injector mixing ratios and unit mass flow rates. The results indicate that as the mixing ratio in the inner-ring injectors increases (while the mixing ratio in the middle-ring injectors decreases), the oxygen concentration area near the axis zone and the 50% radius zone of the gas generator expands. Conversely, the kerosene concentration area near the axis zone decreases while gradually increasing near the 50% radius zone. In the flow direction section, there is an inverse relationship between the variation trend of local temperature and the oxygen concentration in the local area. As the oxygen concentration increases, the temperature decreases. The temperature distribution across the cross-section of the gas generator exhibits a circular pattern. When the mixing ratio (or mass flow rates) of the unit injector are perfectly balanced, the temperature distribution becomes highly uniform. A larger disparity in flow rate between the inner ring injector and the middle ring injector leads to a lower combustion efficiency. This effect differs from the effect of the mixing ratio difference between the two injector rings. Increasing the mixing ratio in the inner-ring injectors (or decreasing the mixing ratio in the middle-ring injectors) initially leads to a rise in combustion efficiency, followed by a subsequent decline. The maximum combustion efficiency of 89.10% is achieved when the mixing ratio is set to Km_-1 = 0.7 and Km_-2 = 2.79, respectively. Increasing the flow rate in the inner-ring injectors (or decreasing the flow rate in the middle-ring injectors) initially leads to an improvement in combustion efficiency, followed by a subsequent reduction. The maximum combustion efficiency of 86.13% is achieved when the mass flow rate is set to m_-1 = m_-2 = 0.1 kg/s. Full article

(This article belongs to the Special Issue Aerospace Combustion Engineering)

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22 pages, 7344 KiB

Open AccessArticle

Prediction of Tail-Off Pressure Peak Anomaly on Small-Scale Rocket Motors

by Stefano Mini, Fabrizio Ponti, Alessandro Brusa, Roberto Bertacin and Barbara Betti

Aerospace 2023, 10(2), 169; https://doi.org/10.3390/aerospace10020169 - 13 Feb 2023

Cited by 2 | Viewed by 2322

Abstract

Numerical studies intended to predict solid rocket motors anomalies are the major contributors when developing strategies to both limit expensive fire tests and to investigate and understand the physical phenomena from which anomalies can arise. This paper aims to present a mathematical–physical method [...] Read more.

Numerical studies intended to predict solid rocket motors anomalies are the major contributors when developing strategies to both limit expensive fire tests and to investigate and understand the physical phenomena from which anomalies can arise. This paper aims to present a mathematical–physical method to evaluate the pressure peak, namely Friedman Curl, occurring at the tail-off phase of small-scale rocket motors. Such phenomenon is linked to the grain solid particles arrangement (i.e., packing effect); indeed, those particles show a tendency to accumulate at a certain distance from the metallic case, implying a local burn rate increment and a combustion chamber pressure rise close to the tail-off phase. Comparisons between experimental and simulated combustion chamber pressure profiles are outlined to prove the effectiveness of the mathematical–physical approach. Simulations were carried out with an internal ballistic simulation tool developed by the authors of this work. Full article

(This article belongs to the Special Issue Aerospace Combustion Engineering)

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14 pages, 850 KiB

Open AccessArticle

Prediction of the Propulsive Performance of an Atmosphere-Breathing Electric Propulsion System on Cathode-Less Plasma Thruster

by Nabil Souhair, Mirko Magarotto, Raoul Andriulli and Fabrizio Ponti

Aerospace 2023, 10(2), 100; https://doi.org/10.3390/aerospace10020100 - 19 Jan 2023

Cited by 4 | Viewed by 4229

Abstract

Atmosphere-breathing electric propulsion (ABEP) is a type of electric propulsion system that uses the atmosphere as a propellant source instead of a stored reservoir. This technology is still in its early stages, but holds the promise of providing a clean, efficient, and sustainable [...] Read more.

Atmosphere-breathing electric propulsion (ABEP) is a type of electric propulsion system that uses the atmosphere as a propellant source instead of a stored reservoir. This technology is still in its early stages, but holds the promise of providing a clean, efficient, and sustainable propulsion system for spacecraft, enabling very low Earth orbit (VLEO) mission scenarios. To optimise the ABEP technology, accurately simulating air-based plasma chemistry plays a crucial role. In this paper, an air-based global model (GM) is presented that includes a detailed chemistry model for the various reactions that are involved in ABEP applications. The model’s goal is to forecast the performance of a cathode-less RF plasma thruster under various pressure levels and species concentrations that are typical of VLEO missions. The GM was exploited to map the performance of a fictitious ABEP based on a cathode-less RF thruster in order to assess its feasibility in VLEO. The numerical model is promising as a tool for the design of ABEP systems and for the preliminary optimization of mission scenarios. Full article

(This article belongs to the Special Issue Aerospace Combustion Engineering)

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16 pages, 4248 KiB

Open AccessArticle

Effect of Penetrative Combustion on Regression Rate of 3D Printed Hybrid Rocket Fuel

by Xiaodong Yu, Hongsheng Yu, Wei Zhang, Luigi T. DeLuca and Ruiqi Shen

Aerospace 2022, 9(11), 696; https://doi.org/10.3390/aerospace9110696 - 7 Nov 2022

Cited by 8 | Viewed by 2749

Abstract

3D printing manufacturing is used to manufacture hybrid rocket fuel grains featuring a special grid-like structure in order to control combustion performance. An innovative penetrative combustion mechanism, capable of affecting regression rate, was noticed during the combustion of low-packing density grains. The 3D [...] Read more.

3D printing manufacturing is used to manufacture hybrid rocket fuel grains featuring a special grid-like structure in order to control combustion performance. An innovative penetrative combustion mechanism, capable of affecting regression rate, was noticed during the combustion of low-packing density grains. The 3D printing manufacture was implemented using acrylonitrile-butadiene-styrene (ABS) material to clarify this mechanism and the corresponding combustion performance. Grid-like structure fuels with different packing densities were prepared to assess the effects of penetrative combustion on fuel combustion performance. The thermal decomposition of ABS was analyzed by infra-red spectroscopic analysis (FTIR) and thermogravimetric analysis-differential thermal scanning (TG-DSC). The internal structure of the ABS grains was observed by high-resolution 3D micro-computed tomography (μCT). All fuel grains were burned in a hybrid 2D radial burner, allowing visualization of the combustion process and evaluation of the ballistic parameters. The experimental results suggest that the combustion process of the ABS porous grains includes two regimes, both featuring an increased regression rate. In the normal layer-by-layer burning regime, at

G_{o x} = 45 kg / (m^{2} \cdot s)

, the regression rates of 100% and 90% ABS increased by 29.6% and 38.1%, respectively, compared with solid ABS which was manufactured by a computerized numerical control (CNC) lathe. In the fracture-led volumetric burning regime, data acquisition is more difficult, but the regression rate is again observed to increase as the packing density decreases. Full article

(This article belongs to the Special Issue Aerospace Combustion Engineering)

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13 pages, 4952 KiB

Open AccessArticle

Condensed Combustion Products Characteristics of HTPB/AP/Al Propellants under Solid Rocket Motor Conditions

by Xue-Li Liu, Song-Qi Hu, Lin-Lin Liu and Yan Zhang

Aerospace 2022, 9(11), 677; https://doi.org/10.3390/aerospace9110677 - 2 Nov 2022

Cited by 5 | Viewed by 3663

Abstract

Condensed combustion products (CCPs) generated during the combustion of aluminized propellants can reflect invaluable information about the combustion mechanisms of propellants. CCPs of hydroxyl-terminated polybutadiene/ammonium perchlorate/aluminum (HTPB/AP/Al) propellants were collected using an experimental apparatus capable of controlling pressure fluctuations within 0.3 MPa, and [...] Read more.

Condensed combustion products (CCPs) generated during the combustion of aluminized propellants can reflect invaluable information about the combustion mechanisms of propellants. CCPs of hydroxyl-terminated polybutadiene/ammonium perchlorate/aluminum (HTPB/AP/Al) propellants were collected using an experimental apparatus capable of controlling pressure fluctuations within 0.3 MPa, and their microscopic morphologies, particle size distributions, and chemical compositions were characterized using a scanning electron microscope (SEM), laser particle size analyzer, energy disperse spectroscopy (EDS), X-ray diffraction (XRD) and complexometric titration. The results showed that the size of CCPs presented a bimodal distribution, with modes at ~5 µm and ~100 µm; particles less than 2 µm were spherical, with smooth surfaces. The main components of CCPs were C, AlN, AlCl₃, Al₂O₃, Fe₂O₃ and Al, with Al₂O₃ being the most abundant. The combustion efficiency of aluminum increased by 3.27% when the size of virgin aluminum particles decreased from 23 µm to 13 µm, but the content of catocene (a burning-rate catalyst) and fine AP (1 µm) had little effect on combustion efficiency. Higher combustion efficiencies and smaller agglomeration sizes can be achieved at higher pressures, due to the positive correlation between pressure and the driving forces for aluminum particles exciting the burning surface. Full article

(This article belongs to the Special Issue Aerospace Combustion Engineering)

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18 pages, 8500 KiB

Open AccessArticle

Effect of Fuel-Injection Distance and Cavity Rear-Wall Height on the Flameholding Characteristics in a Mach 2.52 Supersonic Flow

by Zhonghao He, Hongbo Wang, Fan Li, Yifu Tian, Minggang Wan and Jiajian Zhu

Aerospace 2022, 9(10), 566; https://doi.org/10.3390/aerospace9100566 - 29 Sep 2022

Cited by 6 | Viewed by 1831

Abstract

The ethylene-fueled flameholding characteristics of a cavity-based scramjet combustor are experimentally and numerically investigated. The test facility used the air heater, which heats air from room temperature to total temperature 1477 K. A nozzle is installed behind the heater outlet to increase the [...] Read more.

The ethylene-fueled flameholding characteristics of a cavity-based scramjet combustor are experimentally and numerically investigated. The test facility used the air heater, which heats air from room temperature to total temperature 1477 K. A nozzle is installed behind the heater outlet to increase the air speed to Mach 2.52. Two cavity geometries with different rear-wall heights of 8 mm and 10 mm and two injection distances upstream of the cavities of 10 mm and 40 mm are compared to show the effect of these parameters. The CH* spontaneous emission images obtained by dual-camera synchronous shooting and the wall-pressure distribution obtained by a pressure-scan system are used to capture the flame dynamics. The global equivalence ratio range for different combination schemes is controlled from 0.14 to 0.27 in this paper. The results show that the conventional cavity (the rear-wall height is 10 mm) and the shorter injection distance can effectively decrease the lean blowoff limit of the combustor, while the rear-wall-expansion cavity (the rear-wall height is 8 mm) and the longer injection distance can effectively increase the rich blowoff limit. Compared with the injection distance, the rear-wall height of the cavity has little effect on the oscillation distribution of the shear layer-stabilized flame. However, the fuel-injection distance and cavity rear-wall height both have great influence on the spatial distribution of the flame. Full article

(This article belongs to the Special Issue Aerospace Combustion Engineering)

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20 pages, 4097 KiB

Open AccessArticle

Characterisation and Design of Direct Numerical Simulations of Turbulent Statistically Planar Flames

by Andrej Sternin, Daniel Martinez, Daniel Sternin, Oskar Haidn and Martin Tajmar

Aerospace 2022, 9(10), 530; https://doi.org/10.3390/aerospace9100530 - 21 Sep 2022

Cited by 3 | Viewed by 1582

Abstract

This work aims to provide support for the design of reliable DNSs for statistically planar flames. Improved simulation design strategies are developed. Therefore, design criteria for the simulative domain are discussed. The gained mathematical relations for all of the relevant physical quantities were [...] Read more.

This work aims to provide support for the design of reliable DNSs for statistically planar flames. Improved simulation design strategies are developed. Therefore, design criteria for the simulative domain are discussed. The gained mathematical relations for all of the relevant physical quantities were channelled into a deterministic calculation strategy for mesh features. To choose design parameter values within the mathematical formulations, guidelines were formulated. For less controllable variables, namely the viscosity and Prandtl number, a measurement technique was developed. A new determination strategy to determine characteristic points within the flame front was conducted. In order to present and compare cases with different Prandtl numbers, normalisation of the x-axis of the regime diagram was suggested. Full article

(This article belongs to the Special Issue Aerospace Combustion Engineering)

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19 pages, 12024 KiB

Open AccessArticle

Numerical Investigation on Detonation Initiation and Propagation with a Symmetric-Jet in Supersonic Combustible Gas

by Jian Dai and Linyuan Peng

Aerospace 2022, 9(9), 501; https://doi.org/10.3390/aerospace9090501 - 8 Sep 2022

Cited by 1 | Viewed by 1881

Abstract

In this study, supersonic gaseous detonation initiation and propagation by single- and symmetric-jets are compared, and the effects of symmetric-jets of different intensities on the detonation are further investigated to obtain a more comprehensive understanding of the initiation mechanism of hot jet in [...] Read more.

In this study, supersonic gaseous detonation initiation and propagation by single- and symmetric-jets are compared, and the effects of symmetric-jets of different intensities on the detonation are further investigated to obtain a more comprehensive understanding of the initiation mechanism of hot jet in supersonic mixtures. The two-dimensional reactive Navier–Stokes equations, together with a one-step Arrhenius chemistry model, are adopted to analyze the flow field structure. The results show that the bow shocks induced by symmetric-jets interacting with each other will achieve local detonation combustion. Influenced by the unstable shear layer behind the triple point, a large-scale vortex shedding is formed in the flow field, thus promoting the consumption of the unburned region. By comparing with the single-jet, it is found that the dual-jet initiation method can shorten the distance to complete initiation, but has little effect on the detonation overdrive degree. In addition, a study of the impact of jet size parameters on the symmetric-jet initiation further revealed that there is a critical value, above which the ignition decreases rapidly which is a significant advantage over single-jet. However, below this threshold, detonation initiation will rely on the energy generated by the collision of Mach stems formed at the walls, resulting in a slower ignition rate compared to a single-jet. Therefore, the use of the appropriate jet strength when using a symmetric-jet will result in a more desirable ignition velocity and a shorter distance to achieve detonation. Full article

(This article belongs to the Special Issue Aerospace Combustion Engineering)

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21 pages, 4321 KiB

Open AccessArticle

Thrust Control of Lab-Scale Hybrid Rocket Motor with Wax-Aluminum Fuel and Air as Oxidizer

by Anandu Bhadran, Joel George Manathara and P. A. Ramakrishna

Aerospace 2022, 9(9), 474; https://doi.org/10.3390/aerospace9090474 - 26 Aug 2022

Cited by 15 | Viewed by 5115

Abstract

This article explores the throttling aspect of the hybrid rocket motor through experiments using a lab-scale motor. The lab-scale motor utilizes a wax-Al based fuel and compressed air as the oxidizer. The oxidizer flow rate was modulated using a PID controller to study [...] Read more.

This article explores the throttling aspect of the hybrid rocket motor through experiments using a lab-scale motor. The lab-scale motor utilizes a wax-Al based fuel and compressed air as the oxidizer. The oxidizer flow rate was modulated using a PID controller to study the closed-loop thrust control performance of the motor. Numerical simulations and cold flow tests were carried out to identify the suitable gains for the PID control algorithm. Pressure feedback was used in the control algorithm to obtain the closed-loop thrust control. The resultant closed-loop system followed the reference pressure accurately during the step input response test of the system. The maximum error in the observed chamber pressure was

1.86

% for a reference pressure of

4.69

bar, which corresponds to a reference thrust of

117.6

N. The response of the system for a ramp input, with linear thrust variation from

78.4

N to

127.4

N, showed that the measured thrust followed the desired ramp profile with a root-mean-square error of

1.99

N. A ramp-down test with the same thrust range produced a root-mean-square error of

6.2

N. Full article

(This article belongs to the Special Issue Aerospace Combustion Engineering)

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