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1
Rabeta, Bismil, Mohammad A.F Ulhaq, Aswan Tajuddin, and Agus Sugiharto. "Simulasi Graphical User Interface Analisis Termodinamika Mesin Turboprop Menggunakan Perangkat Lunak Matlab R2020a." Jurnal Teknologi Kedirgantaraan 6, no. 2 (2021): 31–50. http://dx.doi.org/10.35894/jtk.v6i2.44.
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A turboprop engine is a hybrid engine that delivers thrust or jet thrust and also drives the propeller. This is basically similar to a turbojet except the turbine works through the main shaft which is connected to the reduction gear to rotate the propeller in front of the engine. This research was conducted to determine the development of engine performance in thermodynamic analysis so as to know the value of each parameter on a engine that has been developing for 20 to 50 years with different engine manufacturing. So that in this study a comparison of the thermodynamic analysis of the TPE-331, PT6A-42 and H85-200 engines was carried out. In the TPE331-10, PT6A-42, and H85-200 turboprop engines the value of fuel to air ratio and shaft work increases with increasing altitude while compressor work, fuel flow rate, shaft power, propeller thrust, jet thrust, total thrust, equivalent engine power and ESFC decrease with increasing altitude. Furthermore, the turbine's working value is relatively stable as the altitude increases. After that, the value of compressor work and turbine work on the PT6A-42 engine was greater than that of the TPE331-10, and H85-200 engines. However, the value of the fuel to air ratio, fuel flow rate, shaft power, jet thrust, equivalent engine power and ESFC on the H85-200 engine was greater than the TPE331-10 and PT6A engines. Furthermore, at sea level, the value of the axle, propeller thrust, and total thrust on the H85-200 engine is greater than that of the TPE331-10 and PT6A-42 engines but at an altitude of 25,000 ft, the PT6A-4 engine has a greater value than that of the TPE331-10 and PT6A-42 engines. TPE331-10, and H85-200 engines.
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Dolgopolov, S. I. "Determination of the effect of internal and external factors on the thrust spread of a cluster propulsion system." Technical mechanics 2022, no. 2 (2022): 47–58. http://dx.doi.org/10.15407/itm2022.02.047.
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The thrust spread of a stand-alone rocket engine caused by external (the pressure and temperature of the propellant components at the engine inlet) and internal (spread in the geometry and operating conditions of the engine units and assemblies) factors is known from experimental tests or can be computed by a known procedure. As a rule, liquid-propellant propulsion systems (LPPSs) of launch vehicle lower stages include a cluster of several engines, whose thrust spread cannot often be determined from firing tests due to limited capabilities of bench equipment. The aim of this work is to develop an approach to determining the thrust spread of an LPPS comprising a cluster of two and more engines. For a multiengine propulsion system, this methodological approach also includes the development of a mathematical model of engine interaction in an LPPS and calculations of an LPPS startup at different combinations of spread in the external and internal factors in cases where the parameter spreads of all engines are both identical and different. For an LPPS with two engines and a common oxidizer feed pipeline, the paper gives an example of calculating the effect of external and internal factors on the thrust spread of each engine and the LPPS as a whole during an LPPS startup. . It is shown that the calculated spread of the 90 percent thrust (combustion chamber pressure) time lies in the range – 0.0917 s to +0.0792 s (engine 1) and –0.0941 s to +0.0618 s (engine 2). The calculated variations of the combustion chamber pressure (engine thrust) from its nominal value lie in the range –6.2 percent to +7.0 percent (engine 1) and -6.8 percent to +6.3 percent (engine 2). The calculated spreads of the 90 percent thrust time and the thrust for the LPPS as a whole are far smaller (about by 40 percent) and lie in the range – 0.0733 s to +0.0457 s for the time and – 4.8 percent to +4.8 percent for the thrust (about the nominal thrust). Using Pearson’s chi-squared test, an estimate is obtained for the goodness of fit of the anticipated theoretical distributions of the 90 percent thrust time spread and the steady thrust spread to the obtained statistical ones both for the two engines and for the LPPS as a whole.
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Setiyawati, Defi. "Analisis Perbandingan Performa Saat Takeoff Pada Engine CFM56-7b Dengan Konfigurasi Thrust Rating 26300 Lbs Dan 27300 Lbs." Jurnal Teknologi Kedirgantaraan 7, no. 1 (2022): 7–14. http://dx.doi.org/10.35894/jtk.v7i1.51.
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The CFM 56-7B engine is manufactured by CFM International which is used on the B737-600/700/800/900 aircraft. This engine has several variations of the thrust rating with varying performance. Engine performance parameters include Thrust, Specific Fuel Consumption (SFC), Core Speed (N2), and Exhaust Gas Temperature (EGT). Performance testing can be done using the Engine Test Cell. However, the engine test cell is a calibrated tool, which allows deviation of the test results. Then the performance calculation is done using the formula in the Engine Shop Manual - Test 003 - Engine Acceptance Test to find out the engine performance during the takeoff phase at the highest thrust rating of 26300 lbs and 27300 lbs and compare the performance of the two engines. Comparison of the calculation results states that an engine with a thrust rating of 26300 lbs is superior to Exhaust Gas Temperature, while an engine with a thrust rating of 27300 lbs has advantages at Thrust, Specific Fuel Consumption and Core Speed (N2).
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Langston, Lee S. "For Jet Engine Wing Mounting." Mechanical Engineering 140, no. 09 (2018): S52—S53. http://dx.doi.org/10.1115/1.2018-sep4.
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The mounting of a jet engine under the wing of an airliner can be a daunting task for turbofan engineers. Thrust forces generated by gas path momentum flow changes in a jet engine are transmitted by pressure (and friction) forces on stators and struts attached to the engine case. Case engine mounts then transmit the thrust forces (as high as 100,000 pounds thrust on the largest engines) to the wing pylons to pull the plane forward. The mounts must also support the engine weight (as high as 20,000 pounds) and carry nacelle flight loads. Engine bypass ratios are increasing (12:1 on the new geared fan engines), with fan sizes ever growing (178 inch diameter fan on the new GE9X). Mounting these new engines under a wing can present new challenges. During the early days of its introduction in the late 1960’s, Boeing’s iconic 747 jumbo jet had engine mount problems. These are examined, together with their solution.
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S. Abdulhussain, Uzaldin, Taj Elssir Hassan, and Maisara Mohy Eldin Gasim. "Theoretical Performance Comparison between Inline, Offset and Twin Crankshaft Internal Combustion Engine Models." FES Journal of Engineering Sciences 2, no. 1 (2006): 26. http://dx.doi.org/10.52981/fjes.v2i1.91.
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Twin crankshaft is a new engine arrangement introduced to overcome cylinder’s liner wear problems encountered in the conventional inline crankshaft engine due to the effect of the side thrust force. The offset crankshaft arrangement was also introduced to solve the same problem. In this work a computer programs was built to obtain the theoretical performance comparison between the three engines arrangements (inline, twin and offset crankshaft engines), and compared the theatrical performance with the experimental results, which done to the engine’s models. The study results show that the twin crankshaft engine model exhibited no thrust force, and that the thrust force in the offset crankshaft model is smaller than that in the inline crankshaft engine model. These agree with experimental results obtained from the same engine model.
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Chen, Ruiyang. "Advanced development and future of jet engines." Theoretical and Natural Science 12, no. 1 (2023): 162–66. http://dx.doi.org/10.54254/2753-8818/12/20230457.
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An aerodynamic reaction engine, a jet engine, is an internal combustion engine that produces thrust through jet thrust. A jet engine is an important power unit used in aviation and spaceflight, which generates thrust by pushing air in reverse. In recent decades, various developments and improvements have significantly enhanced the efficiency and performance of jet engines. These engines are divided into turbojets, turbojets, turboprops, turboshafts, and ramjets. This article will begin with an introduction to the different types of jet engines and then detail the design, benefits, and issues of the CFM International CFM56/LEAP engine. Finally, there is some description of the future of jet engines, including materials, greater efficiency, and applications in space exploration. The development of the jet engine dates back to the 1930s. British inventor Frank Whittle and German inventor Hans von Ohain developed the first jet engines independently. Jet engines were widely used during World War II and have been continuously improved and developed. The modern jet engine has become the main power unit of commercial and military aviation, and its structure and performance have also been greatly improved.
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Sneha, Rachel Francis, and J. S. Mija. "Design of Fault Tolerant Control Laws for Jet Engines." Applied Mechanics and Materials 367 (August 2013): 96–100. http://dx.doi.org/10.4028/www.scientific.net/amm.367.96.
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Full Authority Digital Electronic Control (FADEC) system is common in all modern jet engines. In jet engine, the simplest control system is to produce desired engine thrust by changing the fuel flow. Since in flight engine thrust measurement is practically not possible, engine low pressure shaft rotational speed (N1/NL), engine pressure ratio (EPR), or exhaust jet temperature (EJT) has been effectively used as an indicator of the engine thrust. In this paper, the model used is twin spool turbofan engine. If any fault occurs in N1 sensor, the entire engine operation will be affected. Since there exists a unique aerodynamic relationship among the spool speeds, if any fault occurs in N1 sensor, engine thrust can be controlled with a certain amount of degradation using high pressure spool speed (N2/NH).Both soft and hard failures are detected using kalman filter, range, rate and comparison techniques. The effectiveness of the proposed approach is demonstrated by means of simulations.
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Jačionis, Tomas, Vytautas Urbanavičius, Andrius Katkevičius, et al. "UAV Detection Using Thrust Engine Electromagnetic Spectra." Drones 6, no. 10 (2022): 306. http://dx.doi.org/10.3390/drones6100306.
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Artificial intelligence used in unmanned aerial vehicle (UAV) flight control systems tends to leave UAV control systems without any radio communication emissions, whose signatures in an electromagnetic spectrum (ES) are widely used to detect UAVs. There will be problems in the near future in detecting any dangerous threats associated with UAV swarms, kamikaze unmanned aerial vehicles (UAVs), or any other UAVs with electrically powered thrust engines because of the UAV’s flight capabilities in full radio silence mode. This article presents a different approach to the detection of electrically powered multi-rotor UAVs. The main idea is to register the electromagnetic spectrum of the electric thrust engines of the UAV, which varies because of the changing flight conditions. An experiment on a UAV’s electric thrust engine-produced electromagnetic spectrum is carried out, presenting the results of the flight-dependent characteristics, which were observed in the electromagnetic spectrum. The electromagnetic signature of the UAV’s electric thrust engines is analyzed, discussed, and compared with the most similar behaving electric engine, which was used on the ground as a domestic electric appliance. A precision tunable magnetic antenna is designed, manufactured, and tested in this article. The physical experiments have shown that the ES of the electric thrust engines of multi-rotor UAVs can be detected and recorded for recognition. The unique signatures of the ES of the multi rotor UAV electric engine are recorded and presented as a result of the carried-out experiments. A precision tunable magnetic antenna is evaluated for the reception of the UAV’s signature. Moreover, results were obtained during the performed experiments and discussions about the development of the future techniques for the identification of the ES fingerprints of the UAV’s electric thrust engine are carried out.
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Ravi, Teja Cheepuri, Y. S. Prudhviraj Oruganti, and Rama Priya Raavi Sai. "Study of Small Gas Turbine Engines of Thrust 1-5KN." International Journal of Innovative Science and Research Technology 8, no. 1 (2023): 2114–19. https://doi.org/10.5281/zenodo.7634395.
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Small gas turbine engines present a significantlylower performance when compared to large ones. The major reasons for that are the lower component efficiencies and size effects. The present work accomplishes an investigation of data available in the open literature for small turbojet and turbofan engines, covering a range from 1 to 5 kN of thrust, to find the most important parameters for these engine configurations. This study aims to evaluate the parameters used in engines in this thrust range, for the possible development of an engine for a specific application. This overview relies on a correlation between some of the major parameters of the engine, namely thrust, specific fuel consumption, pressure ratio, turbine inlet temperature, weight, length, diameter, and mass flow.
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ANDREI, Irina-Carmen, Mihai Victor PRICOP, Mihai Leonida NICULESCU, et al. "COMPARATIVE ANALYSIS FOR PERFORMANCE PREDICTION IN CASE OF IAR 99 AIRCRAFT PROPULSION SYSTEMS." SCIENTIFIC RESEARCH AND EDUCATION IN THE AIR FORCE 25 (July 31, 2024): 152–67. http://dx.doi.org/10.19062/2247-3173.2024.25.18.
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This paper presents a comparative analysis regarding Performance Prediction at Design and Off-Design Regimes, applied for two distinct constructions of Jet Engines as Propulsion Systems solutions for the IAR 99 Aircraft. Case Study #1 is represented by Turbojet Engine Rolls Royce VIPER MK 632-41 and Case Study #2 is represented by a Mixed Flows Turbofan Engine. The thorough study is based on appropriate Applicable Theory, which is detailed in the bibliographic references and can be accessed. Applicable Theory includes a part dedicated to Engine Parameter Identification, which is necessary to calculate Brayton Diagram and the performances of the jet engines, expressed as Thrust and Specific Fuel Consumption (TSFC). Applicable Theory includes detailed Mathematical Modeling, Developments and Numerical Simulations for Turbojet Engine and Mixed Flows Turbofan Engine. Performance Prediction results from aero-thermo-gas dynamics analysis of the studied engines. The accuracy of the numerical results depends on the assumptions used for mathematical modeling, which in this case have been considered: real gas, adiabatic flow, the variation with static temperature of the specific heat, losses due to pressure drop and friction, the conditions for full expansion in the exhaust nozzle are met such that the engine can generate maximum of Thrust. Results from Numerical Simulations express Performance Prediction for the Turbojet Engine and the Mixed Flows Turbofan Engine, for the Design and Off-Design Regimes. Comparative diagrams illustrating the variation of Thrust and Specific Fuel Consumption (TSFC) with Mach number and altitude, for the studied jet engines, conclude the analysis. As final remark, the Mixed Flows Turbofan Engine represents a better option than the Turbojet Engine, from the standpoint of greater Thrust and lower TSFC
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Pismennyi, V. L. "Hyper Afterburner Jet Engines." Proceedings of Higher Educational Institutions. Маchine Building, no. 01 (718) (January 2020): 51–62. http://dx.doi.org/10.18698/0536-1044-2020-1-51-62.
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This paper introduces a thrust augmentation method for super- and hypersonic jet engines by means of applying water at the engine intake. This method expands the use of jet engines with subsonic combustion, allowing velocities up to Mach 8 and altitude up to 45 km. At velocities higher than 3–4 Mach, stagnation temperature of the air is getting higher than the critical temperature of water, which makes the existence of water at the gas turbine engine intake impossible. Water vapour as a working medium of a jet engine creates the so-called inner thermodynamic circle. This phenomenon defines the physics of the thrust augmentation method proposed. The author discusses three variants of hyper afterburner application: hyper afterburner turbojet, hyper afterburner ramjet, and hyper afterburner turbo ejecting engine. The presented basic specifications of the hyper afterburner engines qualitatively differ from those of their prototypes (engines without the hyper afterburner thrust augmentation function). The proposed thrust augmentation method of jet engines is of a special interest for the aerospace field, particularly, for creating air launch systems. It is shown that the application of hyper afterburner in turbo ejecting engines can increase velocity and altitude of the launch aircraft up to Mach 7 and 40 km respectively, thus opening new avenues in space exploration.
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Koff, B. L. "F100-PW-229 Higher Thrust in Same Frame Size." Journal of Engineering for Gas Turbines and Power 111, no. 2 (1989): 187–92. http://dx.doi.org/10.1115/1.3240235.
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The F100-PW-229 fighter aircraft engine is a higher thrust derivative of the F100-PW-220 and in the same frame size. The engine was developed from the F100 Engine Model Derivative (EMD) Program and parallel IR&D efforts. The increased thrust was achieved by increasing the flow and pressure ratio of the two-spool compression system accompanied by an increase in turbine temperature. The increased length compression system was offset by an innovative design intermediate case and a reduced length combustor to maintain overall engine axial length. The –229 engine has a thrust-to-weight ratio of 8.0 with a 20–30 percent performance increase over the –220 model across the flight map. Significant improvements in maintainability have been incorporated while retaining the proven durability and operability features of the –220 engine. The Government-industry partnership is working well, continually providing increased performance engines for our first-line fighters.
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Ryzhkov, V. V., I. I. Morozov, and E. A. Lapshin. "Computer-aided design of low-thust rocket engines using the domain-specific knowledge database and CAE / CAD systems." VESTNIK of Samara University. Aerospace and Mechanical Engineering 18, no. 4 (2020): 106–16. http://dx.doi.org/10.18287/2541-7533-2019-18-4-106-116.
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The paper presents approaches to computer-aided design of low-thrust thrust rocket engines using an extensive knowledge base that allows making basic technical decisions that determine the conceptual design of the engine, based on the developed algorithm of this process. The procedure of creating an electronic 3D-model of a low-thrust rocket engine fueled by gaseous oxygen-hydrogen in the environment of the graphical complex UNIGRAPHICS is described. 3D electronic models of the main elements of a rocket engine with a thrust of P = 25 N were obtained, with subsequent virtual assembly of all components, including the components comprised in the knowledge base, providing the development, among other things, of design documentation, creation of a production environment based on an electronic engine model, preparation for the product manufacturing and the manufacturing proper.
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Kolden, J. J. "A Method of Sizing Multi-Cycle Engines for Hypersonic Aircraft." Journal of Engineering for Gas Turbines and Power 112, no. 2 (1990): 217–22. http://dx.doi.org/10.1115/1.2906165.
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A method of sizing multi-cycle engines for integration with hypersonic vehicles has been developed. The new procedure independently sizes the inlet, each engine cycle, and the nozzle during the vehicle sizing loop to optimize propulsion/aircraft integration. Using uninstalled engine performance for each cycle of a multi-cycle engine along with inlet and nozzle performance and an estimate of aircraft drag, an iterative procedure is utilized to size each component simultaneously. A propulsion system is defined that meets the aircraft thrust requirements at all mission points. The inlet is sized to provide airflow such that the maximum Mach cruise and/or combat thrust conditions are met. Each cycle is sized independently to meet all thrust requirements while minimizing either inlet drag or engine size. Nozzle sizing must trade off thrust, drag and nozzle weight. This methodology has been incorporated into a computer code entitled “Multi-Cycle Engine Sizing Program,” MCESP.
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Oruc, Ridvan, and Tolga Baklacioglu. "Propulsive modelling for JT9D-3, JT15D-4C and TF-30 turbofan engines using particle swarm optimization." Aircraft Engineering and Aerospace Technology 92, no. 6 (2020): 939–46. http://dx.doi.org/10.1108/aeat-02-2020-0031.
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Purpose The purpose of this paper is to create high-accuracy thrust modelling for cruise flight using particle swarm optimization (PSO) algorithm. Design/methodology/approach In this study, using PSO, new thrust models with high accuracy for the cruise flight stages of Pratt & Whitney JT9D-3, JT15D-4C and TF-30 engines were created. For this aim, real Mach number, flight altitude and thrust values taken from the engine manufacturers were used. In the model, thrust is given as a function of altitude and Mach number. The sensitivity of the results given by the PSO thrust model has been examined using several different error types. Finally, the effect of some PSO parameters on the created models is examined. Findings It was observed that the model created predicted real thrust values with high precision. Practical implications The PSO thrust model can be used in the trajectory estimates of today’s aircraft with the use of accurate scaling factors. In addition, using the developed PSO thrust model together with a correct aerodynamic model provides more effective management of air traffic flow in air traffic management applications. Combining the PSO model with fuel flow-rate models will significantly increase engine efficiency and performance; thus, making a major contribution to reducing engine emissions. Originality/value The originality of this study is that it is the first thrust modelling made with PSO in the literature for turbofan engines. The use of real data in the study and the creation of models for several different turbofan engines are important for the reliability of thrust models.
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Rolt, Andrew, Vishal Sethi, Florian Jacob, et al. "Scale effects on conventional and intercooled turbofan engine performance." Aeronautical Journal 121, no. 1242 (2017): 1162–85. http://dx.doi.org/10.1017/aer.2017.38.
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ABSTRACTNew commercial aero engines for 2050 are expected to have lower specific thrusts for reduced noise and improved propulsive efficiency, but meeting the ACARE Flightpath 2050 fuel-burn and emissions targets will also need radical design changes to improve core thermal efficiency. Intercooling, recuperation, inter-turbine combustion and added topping and bottoming cycles all have the potential to improve thermal efficiency. However, these new technologies tend to increase core specific power and reduce core mass flow, giving smaller and less efficient core components. Turbine cooling also gets more difficult as engine cores get smaller. The core-size-dependent performance penalties will become increasingly significant with the development of more aerodynamically efficient and lighter-weight aircraft having lower thrust requirements. In this study the effects of engine thrust and core size on performance are investigated for conventional and intercooled aeroengine cycles. Large intercooled engines could have 3%–4% SFC improvement relative to conventional cycle engines, while smaller engines may only realize half of this benefit. The study provides a foundation for investigations of more complex cycles in the EU Horizon 2020 ULTIMATE programme.
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Afridi, Saadia, Tariq Amin Khan, Syed Irtiza Ali Shah, Taimur Ali Shams, Khawar Mohiuddin, and David John Kukulka. "Techniques of Fluidic Thrust Vectoring in Jet Engine Nozzles: A Review." Energies 16, no. 15 (2023): 5721. http://dx.doi.org/10.3390/en16155721.
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Thrust vectoring innovations are demonstrated ideas that improve the projection of aerospace power with enhanced maneuverability, control effectiveness, survivability, performance, and stealth. Thrust vector control systems following a variety of concepts have been considered for modern aircraft and missiles to enhance their military performance. Short Take-off and Landing (STOL) and control effectiveness at lower aircraft speeds can be achieved by employing Fluidic Thrust Vectoring Control (FTVC). This paper summarizes a range of ideas for FTVC that have been designed and tested both computationally and experimentally to determine the thrust vectoring performance of supersonic propulsion system nozzles. The conventional method of thrust vectoring involves mechanical means to deflect the direction of flow of the exhaust gases, whereas the most recent method involves fluidic-based thrust vectoring techniques. Fluid-based thrust vectoring has the advantages of simplicity and low weight over mechanical-based thrust vectoring, which has complex geometry and adds extra weight to the aircraft. The fluidic vectoring control nozzles are divided into seven categories: shock vector, bypass shock vector, counterflow, co-flow, throat skewing, dual throat, and bypass dual throat nozzle control. This paper provides a summary of each fluidic thrust vectoring technique with its characteristics, design, classification, and different operational criteria developed to date and compares the intrinsic characteristics of each technique. Based on the present literature, it is concluded that among all the fluidic control techniques, the bypass dual-throat nozzle control can achieve better thrust vectoring performance with large vector angles and low thrust loss.
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Yin, Yongqi, Zongchun Hu, Xiaoguang Zhu, et al. "Research on throat type variable thrust solid rocket motor." Journal of Physics: Conference Series 2820, no. 1 (2024): 012019. http://dx.doi.org/10.1088/1742-6596/2820/1/012019.
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Abstract Solid Rocket Motors (SRM) is a power propulsion device using solid propellant, which has obvious advantages in structure, reliability, and ease of use compared with other types of power propulsion devices. With the continuous development of science and technology, the real environment puts forward higher requirements for the maneuverability of rocket engines, and engines with fast start ability and power adjustment functions are becoming more and more important. Therefore, variable thrust solid rocket engines occupy a very important place in the development of future rocket engines. In this paper, the working principle of the throat plug engine is analyzed, and the functional relationship between the thrust and the displacement of the throat plug is derived. The throat bolt solid rocket motor changes the minimum cross-sectional area of the nozzle through the reciprocating motion of the throat bolt on the central axis and then changes the thrust. By establishing a Cartesian coordinate system to describe the shape of the throat plug and the nozzle in the shape of a functional curve, the problem of the minimum cross-sectional area of the nozzle of any throat plug configuration is transformed into the problem of the shortest distance between the two curves, and the established functional relationship between the thrust force and the displacement of the throat plug provides a basis for the design of the throat plug configuration.
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Jimboh, K., H. Aono, T. Chikata, Y. Hagiwara, K. Nakasu, and T. Hoshino. "Thrust Load Measurement on Aero-Engine Bearing." Journal of Engineering for Gas Turbines and Power 107, no. 1 (1985): 181–86. http://dx.doi.org/10.1115/1.3239680.
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Aero-engine bearings operate in an extraordinarily high speed range (high DN number) and severe conditions. It is especially necessary to measure and adjust the bearing thrust load in the engine development phase, but it is very difficult to measure the thrust load accurately, because bearings and bearing housings are subjected to elevated temperature and oil environment. Open space permitted for installation of thrust measurement transducers is small and limited around the bearing housing. We tried to measure the thrust load by applying “Unit Cells,” which are installed between bearing and bearing housing. “Unit Cells” which have been specially designed to measure the bearing thrust load are very small and temperature-compensated load cells. We have been successful in measuring the actual thrust load using the above “Unit Cells,” both in the steady-state and transient condition. Repeatability and hysteresis of the data have been satisfactory. We have established the effect of seal clearance on the thrust load by the measurement. We also have obtained the dynamic characteristics of the thrust load versus rotor speed in low bypass fan engines. Procedure and obtained data are presented in detail.
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Guo, Feng, Ming Liu, Guozhong He, Junhui Zhou, Jianfeng Zhu, and Yancheng You. "Analysis and Suppression of Thrust Trap for Turbo-Ramjet Mode Transition with the Integrated Optimal Control Method." Aerospace 10, no. 8 (2023): 667. http://dx.doi.org/10.3390/aerospace10080667.
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An aircraft/engine integrated optimal control method is proposed for turbine-based combined cycle (TBCC) engines based on the Gauss pseudospectral method. The optimal flight trajectory and TBCC control law are obtained for a TBCC-powered aircraft, and the “thrust trap” that occurs during turbo-ramjet mode transition is further analyzed and suppressed. Results show that the aircraft goes through the mode transition phase using a “climb-dive” trajectory, which is a strategy of applying gravity-assist and temporarily reducing the drag. Furthermore, the TBCC engine adjusts at the quickest rate to minimize thrust loss. With the coupling of the trajectory and TBCC control law, the minimum thrust during the mode transition is only 23% of the thrust before the mode transition, suggesting the “thrust trap” phenomenon. By decreasing the mode transition time from 60 s to 15 s, the minimum thrust can only increase to 30%, and the “thrust trap” phenomenon cannot be effectively suppressed. When the operating speed range of the turbine engine increases from Ma2.5 to 2.9, the minimum thrust will reach 80%, and the “thrust trap” tends to level off.
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Magdin, Aleksandr, Alexey Pripadchev, Alekcandr Gorbunov, and Dmitry Soldatov. "THRUST AUGMENTATION OF THE TURBOJET ENGINE BY UPGRADING THE NOZZLE CLUSTER." Transport engineering 2022, no. 9 (2022): 22–29. http://dx.doi.org/10.30987/2782-5957-2022-9-22-29.
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The study objective is the possibility of thrust augmentation of a turbojet engine due to upgrading the nozzle cluster by installing an external duct of the annular combustor on the engine case, acting as a nozzle. The problem to which the article is devoted is to get a new product with higher operating parameters at minimal cost by upgrading a mass-produced engine. The novelty of the work is in the introduction of an annular combustor into existing engines that are in serial production, thereby reducing the cost of producing a completely new product with greater thrust. During the research, a number of theoretical advantages of a new product based on a mass-produced engine are obtained, including increased thrust with minimal change in its weight, as well as the economic benefits of producing such units with minimal production changeover. At the stage of theoretical research of this topic, a retrofitted engine with an annular combustor has a number of advantages over the standard engine taken as a basis.
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Wang, Yeguang, Honglin Liu, and Kai Liu. "Improved Thrust Performance Optimization Method for UAVs Based on the Adaptive Margin Control Approach." Mathematics 11, no. 5 (2023): 1176. http://dx.doi.org/10.3390/math11051176.
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This study proposes a strategy for improving the thrust performance of fixed-wing UAV turbine engines from the perspective of aircraft/engine integration. In the UAV engine control process, the inlet distortion caused by the angle of attack change is taken into account, the inlet distortion index is calculated in real time by predicting the angle of attack, and the influence of the inlet distortion on the engine model is analyzed mechanically. Then, the pressure ratio command is adjusted according to the new compressor surge margin requirement caused by the inlet distortion to finally improve the engine thrust performance. To verify the effectiveness of the algorithm, an adaptive disturbance rejection controller is designed for the flight control of a fixed-wing UAV to complete the simulation of horizontal acceleration. The simulation results show that, with this strategy, the UAV turbofan engine can improve the turbofan engine thrust performance by more than 8% under the safety conditions.
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Mulyani, Sri, and Rudi Setiawan. "Analisis Performa Take Off Engine CFM56-7B Pada Thrust Tipe 26300 Lb." Prosiding Seminar Nasional Sains Teknologi dan Inovasi Indonesia (SENASTINDO) 3 (December 21, 2021): 401–10. http://dx.doi.org/10.54706/senastindo.v3.2021.134.
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Penurunan performa engine dalam beberapa kondisi memungkinkan engine tidakdapat lagi di-maintain untuk mencapai performa yang disyaratkan. Menurut CFM masalah ini dapat diatasi dengan menurunkan versi (Thrust Rate) engine, dengan begitu engine tersebut tetap dapat beroperasi dengan standar performa yang lebih rendah dari versi awal.Dalam Tugas akhir ini penulis membahas perbandingan performa take-off engine CFM56-7B dengan konfigurasi thrust rating 26300 lbs. Perhitungan performa dilakukan dengan pengolahan data Test Cell Result menggunakan formula pada Engine Shop Manual – 003 untuk mengetahui seberapa besar perbedaan yang terjadi pada kedua konfigurasi thrust rating tersebut. Pada perhitungan performa take-off engine CFM56-7B dengan konfigurasi thrust rating 26300 lb berdasarkan formula dari Engine Shop Manual-003 Engine Acceptance Test adalah Konfigurasi thrust rating 26300 lb menghasilkan thrust sebesar 26920 lb. Nilai SFC yang didapatkan untuk thrust rating sebesar 0.3885 dengan SFC margin -5.3% dan 0.382. EGT yang dihasilkan engine sebesar 905.9 ºC dan EGT margin 17.1 ºC.
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Burova, A. Yu. "Methods and Algorithms of Turbojet Engines Thrust Parameters Control Unerroric." Journal of Physics: Conference Series 2096, no. 1 (2021): 012060. http://dx.doi.org/10.1088/1742-6596/2096/1/012060.
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Abstract This report shows a real way for solving the problem of turbojet engine thrust parameters control in flight. The purpose of the research is the formalization of the digital methods and its algorithms for turbojet engines thrust parameters control unerroric. The methods of deductive digital signal processing and combined method of system analysis and approximate synthesis are used for such research. It is described the digital algorithms of unerroric methods wich are based on rotor speed control for turbojet engines of twin-engine airliner by its power plant control system. There are given the calculation formulas for those algorithms.
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Catana, Razvan Marius, Grigore Cican, and Gabriel-Petre Badea. "Thermodynamic Analysis and Performance Evaluation of Microjet Engines in Gas Turbine Education." Applied Sciences 14, no. 15 (2024): 6754. http://dx.doi.org/10.3390/app14156754.
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This paper presents a detailed study on the main parameters and performance evaluation of microjet engines, at take-off regime and at various engine working regimes, based on thermodynamic analysis of a particular engine data library, from different engine manufacturers such as JetCat and AMT Netherlands. The studied engines have the same spool design but different thrust classes ranging from 97 to 1569 N. The particular data library includes engine specifications from catalogs or data sheets as well as our own experimental data from the JetCat P80 microjet engine, obtained using the ET 796 Jet Turbine Module, a complete testing facility for gas turbine education purposes. Various ratios and differences between certain engine main parameters and performances are studied in order to calculate values through which the analyses can be performed. Even if the engines have different thrust classes, the study examines if there are close values of the ratios and differences of parameters, that can be defined as reference parameters through which the engine performance can be compared and evaluated.
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Nguyen, Le Thanh, Quoc Quan Nguyen, and Ha Hiep Nguyen. "Research on operating characteristics of single-spool turbojet enginewith the afterburner AL-21F." Ministry of Science and Technology, Vietnam 65, no. 3 (2023): 39–45. http://dx.doi.org/10.31276/vjst.65(3).39-45.
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The single-spool, single-thread, turbojet engine with afterburner AL-21F has been used on fighter aircraft for a long time, but there are few studies on this engine in Vietnam. This article uses the theoretical approach, calibrates according to the technical description and actual operation, combined with specialised software such as GasTurb and GSP to calculate engine thermodynamic parameters in some actual flight conditions. The obtained results showed that, when the flight speed increases, the specific fuel consumption increases linearly, while the initial engine thrust decreases and reaches the minimum value at M=0.3÷0.5, then continue to increase and reach the maximum at the speed limit. When increasing the height, the speed and regulating spool speed of the engine remain the same, the engine thrust decreases as well as the specific fuel consumption increases. When increasing the rotor rotation speed, the engine thrust increases linearly but the specific fuel consumption decreases gradually and remains almost the same at engine rpm above 85%. The results allowed for a better understanding of the operating process of the AL-21F engine in particular and turbojet engines in general or used to build simulation models operating close to reality.
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Rosell, Daniel, and Tomas Grönstedt. "Design Considerations of Low Bypass Ratio Mixed Flow Turbofan Engines with Large Power Extraction." Fluids 7, no. 1 (2022): 21. http://dx.doi.org/10.3390/fluids7010021.
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The possibility of extracting large amounts of electrical power from turbofan engines is becoming increasingly desirable from an aircraft perspective. The power consumption of a future fighter aircraft is expected to be much higher than today’s fighter aircraft. Previous work in this area has concentrated on the study of power extraction for high bypass ratio engines. This motivates a thorough investigation of the potential and limitations with regards to performance of a low bypass ratio mixed flow turbofan engine. A low bypass ratio mixed flow turbofan engine was modeled, and key parts of a fighter mission were simulated. The investigation shows how power extraction from the high-pressure turbine affects performance of a military engine in different parts of a mission within the flight envelope. An important conclusion from the analysis is that large amounts of power can be extracted from the turbofan engine at high power settings without causing too much penalty on thrust and specific fuel consumption, if specific operating conditions are fulfilled. If the engine is operating (i) at, or near its maximum overall pressure ratio but (ii) further away from its maximum turbine inlet temperature limit, the detrimental effect of power extraction on engine thrust and thrust specific fuel consumption will be limited. On the other hand, if the engine is already operating at its maximum turbine inlet temperature, power extraction from the high-pressure shaft will result in a considerable thrust reduction. The results presented will support the analysis and interpretation of fighter mission optimization and cycle design for future fighter engines aimed for large power extraction. The results are also important with regards to aircraft design, or more specifically, in deciding on the best energy source for power consumers of the aircraft.
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Улитенко, Юрий Александрович. "АНАЛИЗ ХАРАКТЕРИСТИК ТУРБОРЕАКТИВНОГО ДВУХКОНТУРНОГО ДВИГАТЕЛЯ С ФОРСАЖНОЙ КАМЕРОЙ СГОРАНИЯ С ВПРЫСКОМ ВОДЫ ЗА ВХОДНЫМ УСТРОЙСТВОМ". Aerospace technic and technology, № 1 (7 березня 2019): 29–38. http://dx.doi.org/10.32620/aktt.2019.1.03.
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Development of perspective high-speed aircraft inseparably depends on the level of aircraft propulsion engineering as engine performances to determine aircraft capabilities as a whole. The basic requirements to engines of high-speed aircraft are increase speed and flight height. The new generation of turbojet bypass engine with afterburner each their specific thrust and a specific impulse increases, also the application of high technologies raises leads to substantial growth of the engine cost too. At the same time, existing engines design has big reserves for modernization. The system of water injection to the input at the turbojet bypass engine with afterburner is one of the accessible ways for design improvement. Those advanced engines theoretically will allow to satisfy requirements from designers of high-speed aircraft concerning to thrust and other key parameters, at the same time to secure continuity of already existing types of power-plants. The possibility of range extension of turbojet bypass engine with classical scheme afterburner operation till Mach number 3 is considered in this article. The analysis of existing developments is carried out. Impact of water injection to the input at turbojet bypass engine with afterburner on its performance is investigated. Results of calculations for the influence of water injection to reaction mass parameters on the engine duct and its thrust characteristics are proved. Received results will allow to increase thermodynamic efficiency and to expand range extension of turbojet bypass engine with afterburner provided to use materials that applied in aviation manufacture, as well as to reduce terms of development competitive engines for high-speed aircraft at the expense of purposeful search of their rational thermodynamic and is constructive-geometrical architecture.
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Langston, Lee S. "Some Details of Jet Engine Thrust." Mechanical Engineering 138, no. 09 (2016): 76–77. http://dx.doi.org/10.1115/1.2016-sep-6.
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This article throws light on details of jet engine thrust. The momentum flux of the engine exiting flow is greater than that which entered, brought about by the addition of the energy input from combusted fuel, and giving rise to engine thrust. Thrust arises from pressure and frictional forces on these surfaces, e.g., blades, vanes, endwalls, ducts, etc. This interior force view of thrust is easy to visualize but quite another thing to actually measure. In doing research on secondary flow in gas turbine passages, researchers have measured both steady-state momentum changes and surface forces, in the much simpler case of a turbine blade cascade. The thrust values for each component in the Rolls-Royce single spool engine have been shown in this paper. It has been noted that from the compressor, gas path flow enters the engine case diffuser, where a pressure gain produces another component of forward thrust of 2,186 lbt. Newton’s second law of motion allows us to examine engine component behavior that exhibits both forward and rearward propelling forces, which results in the net thrust our airline passengers have purchased.
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Wadia, A. R., and F. D. James. "F110-GE-132: Enhanced Power Through Low-Risk Derivative Technology." Journal of Turbomachinery 123, no. 3 (2000): 544–51. http://dx.doi.org/10.1115/1.1378301.
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The F110-GE-132, originally referred to as the F110-GE-129 EFE (Enhanced Fighter Engine), presently undergoing qualification testing, is being offered at two different thrust/inspection levels with a maximum augmented thrust of 34,000 pounds. The EFE has been developed using low-risk derivative engine technology. It features a new increased airflow, high efficiency, three-stage long chord blisk fan, and an advanced radial augmentor that reduces complexity, improves maintainability, and provides increased parts life. The paper first provides a historical background of the F110 engines to relate the heritage of the F110-GE-132. The F110 engine model development roadmap is shown to illustrate the incremental low-risk approach used to provide thrust growth with improved product reliability. A detailed description of the unique power management features of the EFE engine to meet individual customer thrust and life requirements is outlined. The long chord blisk fan design, development, and test results are presented, followed by a description of the radial augmentor and the exhaust nozzle. The EFE engine has successfully completed sea level static and altitude development testing and fan aero mechanical qualification at the AEDC in Tullahoma, Tennessee.
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Chao, Zhang, Shi Aobo, Cheng Guangqi, and Sun Jian. "A hybrid dynamics modeling method for micro-turbine engine thrust vectoring." Journal of Physics: Conference Series 2746, no. 1 (2024): 012025. http://dx.doi.org/10.1088/1742-6596/2746/1/012025.
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Abstract Thrust vector control technology is widely used in modern advanced aircrafts, space shuttles and unmanned aerial vehicles, etc. An accurate thrust vector mathematical model is of great significance to enhance the thrust vector control effect. In this paper, the P200 micro-turbine engine with an axisymmetric thrust vector nozzle placed at the rear is taken as an example, A hybrid thrust vectoring dynamics modeling method consisting of engine speed model, engine thrust model, actuator dynamics model, airflow angle model and thrust loss coefficients is proposed. The engine speed model is of Linear Parameter Varying (LPV) model with each segment is characterized by a First Oder Plus Dead Time (FOPDT) model. The engine thrust model is characterized as a cubic polynomial function of engine speed. The actuator’s dynamic model is a FOPDT model. The nozzle airflow deflection angle is proportional to the airflow deflection angle. the thrust loss coefficient is approximated by a cosine relationship with the airflow deflection angle. The experimental results show that the established model can well characterize the dynamics of the thrust vector system.
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Stevenson, J. D., and H. I. H. Saravanamuttoo. "Simulating Indirect Thrust Measurement Methods for High-Bypass Turbofans." Journal of Engineering for Gas Turbines and Power 117, no. 1 (1995): 38–46. http://dx.doi.org/10.1115/1.2812779.
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As yet, there is no known reliable method for directly measuring the thrust of a turbofan in flight. Manufacturers of civil turbofans use various indirect thrust measurements to indicate performance of an engine to the flight deck. Included among these are: Engine Pressure Ratio (EPR), Integrated Engine Pressure Ratio (IEPR), Fan mechanical speed (N1), and various Turbine Gas Temperatures such as ITT or EGT. Of key concern is whether these thrust indicators give an accurate account of the actual engine thrust. The accuracy of these methods, which is crucial at take-off, may be compromised by various types of common engine deterioration, to the point where a thrust indicator may give a false indication of the health and thrust of the engine. A study was done to determine the effect of advanced engine cycles on typical values of these parameters. A preliminary investigation of the effects of common kinds of turbofan deterioration was conducted to see how these faults can affect both actual engine performance and the indirect thrust indicators.
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Pietraszek, Mariusz, and Tomasz Klemba. "Investigation of the Oxidizer Charge Impact Onto Energetic Characteristics Demonstrated by the Stream of Gunpowder Gas Ejected by Solid-Propellant Missile Engines." Research Works of Air Force Institute of Technology 33, no. 1 (2013): 203–19. http://dx.doi.org/10.2478/afit-2013-0012.
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Abstract The study discloses outputs from examinations devoted to charges of oxidizers and their impact on behaviour of missile engine operation with particular attention to measures that are undertaken to counteract stall of a helicopter engine. Homogenous solid propellant for missile engines exhibits a negative oxygen balance. When a series of rocket missiles driven by means of such fuel is launched from the helicopter board there is a hazard of the engine stall effect that may lead to killing its engines. Admixture of potassium (II) sulphate (VI) that is added into the combustion chamber of a missile engine as an inhibitor of the combustion reaction favourably alters characteristics of the engine output power and, at the same time, is irrelevant to the available thrust of the engine. Application of the oxidizer as an insert into the engine substantially improves flight safety when rocket missiles are launched from an aircraft and makes it possible to avoid significant changes in the engine design. The paper outlines the results from investigations when a charge of K2SO4 was introduced into the Mk66 missile engine as a compressed bar and the working parameters of the engine were measured. The investigations on a vertical workbench comprised measurements of the engine thrust and temperature in the stream of exhaust gases. The measurement results were compared against figures sourced from the original data sheets of engines
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Denning, R. M., and N. A. Mitchell. "Trends in Military Aircraft Propulsion." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 203, no. 1 (1989): 11–23. http://dx.doi.org/10.1243/pime_proc_1989_203_049_01.
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The major factors determining the choice of engine cycle for a combat aircraft are the requirements of the design mission and those of aircraft speed and agility. The requirement for jet-borne flight in short take-off vertical landing (STOVL) aircraft imposes further demands on cycle and configuration. The changing nature of combat aircraft requirements is the reason for changes in engine design. Specific thrust is shown to be the major parameter defining engine suitability for a particular role. An examination of mixed turbofan characteristics shows that specific thrust is also the key to understanding the relationships between engine characteristics. The future development of combat engines is discussed, in particular the implications of stoichiometric limits on cycle temperatures and the benefits of variable cycle engines are examined. Recent work on advanced STOVL (ASTOVL) aircraft is reviewed and aircraft/engine concepts designed to meet the requirements of the role are assessed. Experience shows that the technology for these advanced engines must be fully demonstrated before production to minimize the risks and costs of the development programme.
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Dunn, M. G., R. M. Adams, and V. S. Oxford. "Response of Large Turbofan and Turbojet Engines to a Short-Duration Overpressure." Journal of Engineering for Gas Turbines and Power 111, no. 4 (1989): 740–47. http://dx.doi.org/10.1115/1.3240321.
Full textAbstract:
The influences of thrust setting and overpressure level on engine operating characteristics have been obtained for two different high-thrust engines. The thrust setting was varied from engine-off to take-off rated thrust (TRT) and the overpressure was varied from 6.9 kPa (1.0 psi) to 19.4 kPa (2.8 psi). The specific engines under consideration were the Pratt/Whitney TF33 low bypass ratio turbofan and the Pratt/Whitney J57 turbojet. The experimental results suggest that overpressure has little influence on either the HP compressor speed or the exhaust gas total temperature. However, the magnitude of the overpressure has a large influence on turbine exhaust total pressure and on the inlet casing and the diffuser casing radial displacements. The J57 turbine casing was significantly influenced by the overpressure, whereas the TF33 turbine casing was relatively insensitive. The J57 inlet casing radial displacement was noticeably greater than the corresponding turbofan displacement. Even though the component radial displacements for the TF33 exceeded the steady-state red-line limit by more than 300 percent, the engine did not sustain any permanent damage. The J57 did, however, experience an internal rub at an overpressure of about 14.5 kPa (2.1 psi).
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Pizzo, Joe. "Airplane dynamics: Engine thrust, engine braking, and wing lift." Physics Teacher 26, no. 2 (1988): 122–23. http://dx.doi.org/10.1119/1.2342452.
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Pylypenko, O. V., S. I. Dolgopolov, N. V. Khoriak, and N. D. Nikolayev. "Procedure for determining the effect of internal and external factors on the startup thrust spread of a liquid-propellant rocket engine." Technical mechanics 2021, no. 4 (2021): 7–17. http://dx.doi.org/10.15407/itm2021.04.007.
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Despite of the package of measures to adjust a liquid-propellant rocket engine (LPRE) to a specified operating regime, minimum acceptable spreads in the geometrical parameters and operating conditions of its units and assemblies steel remain. These internal factors together with external ones (the pressure and temperature of the propellant components at the engine inlet) govern the engine thrust spread. To provide an acceptable engine thrust spread according to the engine requirements specification, it is important to know the spread value as early as at the stage of off-engine tryout of the engine units and assemblies. The aim of this work is to develop a procedure for calculating the effect of external and internal factors on the LPRE startup thrust spread. This paper presents a procedure for determining the effect of internal and external factors on the LPRE startup thrust spread. The procedure includes the development of a mathematical model of engine startup that accounts for the maximum number of internal factors, the choice of internal factors that produce the maximum effect on the LPRE startup thrust spread, the choice of a method for specifying the external and internal factor spread, engine startup calculations at different combinations of external and internal factor spread values, engine thrust spread determination, determining the statistical and the theoretical distributions of the 90 percent thrust time spread and the steady thrust spread, and assessing their goodness of fit using Pearson’s chi-squared test. The paper gives an example of calculating the effect of the external and internal factor spread on the LPRE startup thrust spread for a staged-combustion oxidizer-rich sustainer LPRE. Using the results of previous calculations, 12 internal factors that produce the maximum effect on the engine startup thrust spread are identified. It is shown that the calculated spread of the 90 percent thrust (combustion chamber pressure) time lies in the range – 0.08220s to +0.07300s about its nominal value, and the calculated steady engine thrust (combustion chamber pressure) spread lies in the range –6.4 percent to +6.6 percent of the nominal thrust. Using Pearson’s chi-squared test, an estimate is obtained for the goodness of fit of the anticipated theoretical distributions of the 90 percent thrust time spread and the steady thrust spread to the obtained statistical ones.
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Awang Saifudin, Awang Raisudin, and Nurul Musfirah Mazlan. "Computational Exploration of a Two-Spool High Bypass Turbofan Engine's Component Deterioration Effects on Engine Performance." Applied Mechanics and Materials 629 (October 2014): 104–8. http://dx.doi.org/10.4028/www.scientific.net/amm.629.104.
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Aircraft engines are exposed to degradation due to several factors such as environmental air pollution, fuel content and ageing or degradation of engine’s components, which are experienced within specified time. While the turbofan in operation, its components deteriorate and consequently affect its performance. This study is aimed to computationally investigate the effect of components degradation on engine performance. A high bypass turbofan engine operated at cruise is selected for this evaluation and the simulation was performed using the Gas Turbine Simulation Program (GSP). The affected components considered are turbines and compressors with deterioration rate ranging from 0% to 5%. The effect of selected deterioration rate on engine thrust and thrust specific fuel consumption (TSFC) is studied. Results obtained show an agreement with literature where reduction in engine thrust and TSFC are observed. Turbine’s fouling has been found to be more severe than erosion in terms of power and efficiency losses. However, in terms of the overall performance, the erosion effect is more severe than fouling.
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Терещенко, Юрий Юрьевич, Иван Алексеевич Ластивка, Павел Владимирович Гуменюк та Су Хунсян. "ЕФЕКТИВНА ТЯГА ТА ЗОВНІШНІЙ ОПІР АВІАЦІЙНОЇ СИЛОВОЇ УСТАНОВКИ". Aerospace technic and technology, № 5 (29 серпня 2020): 61–67. http://dx.doi.org/10.32620/aktt.2020.5.08.
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Increasing the efficiency and effectiveness of a gas turbine engine can be achieved through a comprehensive review of all tasks that determine the parameters and characteristics of an aircraft power plant and aircraft. An important place in this complex is occupied by the problem of obtaining the most efficient traction and power plant based on the integration of the parameters and characteristics of the nacelle and gas turbine engine, consisting of a universal gas generator module and a turbofan module. Reducing the negative impact of the engine nacelle module on effective traction and effective specific fuel consumption is an urgent problem that can be solved based on the results of studies of the integration parameters and characteristics of the engine nacelle of the gas generator module and the gas turbine engine with the turbine-fan extension module, namely, with the implementation of structurally layout diagram of a gas turbine engine with a modular design with a rear arrangement of a turbofan attachment. For modern power plants with bypass gas turbine engines with a large bypass ratio, the external resistance is 2-3 % of the engine thrust during cruising operation. The results of experimental studies have shown that the external resistance of power plants with bypass gas turbine engines of modern supersonic aircraft is 4-6 % of the engine thrust during cruising operation. The paper considers the issues of aerodynamic integration of a gas turbine engine and a nacelle of an aircraft power plant. Aerothermogasdynamic integration of a gas turbine engine and an aircraft provides for the coordination of the parameters of the working process and the characteristics of the gas turbine engine and the parameters and characteristics of the nacelle of the aircraft in order to obtain optimal parameters and characteristics of the aircraft in the design flight conditions. The dependences of the relative effective thrust on the flight velocity are obtained. The obtained dependencies show the influence of the external resistance of the engine nacelle on the effective thrust of the bypass engine at subsonic flight velocities. The calculations were performed to lengthen the nacelle in the range from 4 to 8.
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Landy, R. J., W. A. Yonke, and J. F. Stewart. "Development of HIDEC Adaptive Engine Control Systems." Journal of Engineering for Gas Turbines and Power 109, no. 2 (1987): 146–51. http://dx.doi.org/10.1115/1.3240017.
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The NASA Ames/Dryden Flight Research Facility is sponsoring a flight research program designated Highly Integrated Digital Electronic Control (HIDEC), whose purpose is to develop integrated flight-propulsion control modes and evaluate their benefits in flight on NASA F-15 test aircraft. The Adaptive Engine Control System (ADECS I) is one phase of the HIDEC program. ADECS I involves uptrimming the P&W Engine Model Derivative (EMD) PW1128 engines to operate at higher engine pressure ratios (EPR) and produce more thrust. In a follow-on phase, called ADECS II, a constant thrust mode will be developed which will significantly reduce turbine operating temperatures and improve thrust specific fuel consumption. A performance seeking control mode is scheduled to be developed. This mode features an onboard model of the engine that will be updated to reflect actual engine performance, accounting for deterioration and manufacturing differences. The onboard engine model, together with inlet and nozzle models, are used to determine optimum control settings for the engine, inlet, and nozzle that will maximize thrust at power settings of intermediate and above and minimize fuel flow at cruise. The HIDEC program phases are described in this paper with particular emphasis on the ADECS I system and its expected performance benefits. The ADECS II and performance seeking control concepts and the plans for implementing these modes in a flight demonstration test aircraft are also described. The potential payoffs for these HIDEC modes as well as other integrated control modes are also discussed.
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Vasiliev, Igor, Boris Kiforenko, and Yaroslav Tkachenko. "COMPARATIVE ANALYSIS OF THE EFFICIENCY OF CONSTANT POWER THROTTLED ROCKET ENGINES FOR INTERORBITAL FLIGHTS TO GEOSTATIONAR." Journal of Automation and Information sciences 6 (November 1, 2021): 66–77. http://dx.doi.org/10.34229/1028-0979-2021-6-7.
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Carrying out low-thrust transfers of spacecrafts in the near-earth space from intermediate elliptic to the geostationary orbit using electric rocket engines seems to be one of the most important tasks of modern cosmonautics. Electric rocket engines, whose specific impulse of the reactive jet is an order of magnitude more than in chemical RD, are preferable for interorbit flights with a maximum payload in the case when a significant increase in the duration of the maneuver is permissible. Ability to throttling the rocket engine thrust is traditionally considered as one of the ways to reduce both the engine mass and the required fuel assumptions for performing the specified maneuver. Using the concept of an ideal-rocket engine provides the upper estimates of the payload mass of interborbital flights for the given power level. Accounting for the properties of real engines leads to the need of considering the mathematical models with more strict limits on control functions. A study of the efficiency of three modes of thrust control of an electric propulsion rocket engine was carried out when performing practically interesting spacecraft flights from highly elliptical intermediate near-earth orbits to geostationary orbits. A mathematical model of constant power relay rocket engine has been built. The formulation of the variational problem of the Maer type is given about the execution of a given dynamic maneuver for the throttled and unregulated electric rocket engines of constant power. Using the Pontryagin maximum principle, an analysis of the optimal control functions was carried out, for which the final relations were written out, which allowed to write down the system of differential equations of the optimal movement of the spacecraft, equipped with relay electric rocket engine. The obtained numerical and quality results of the study of the effectiveness of various modes of thrust control of an electric propulsion engine to increase the payload of a given orbital maneuver confirmed the correctness of mathematical models of throttled and relay engines and, in general, the efficiency of using solutions of the averaged equations of optimal motion of a spacecraft for numerical solution of the corresponding boundary value problems in an exact formulation.
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Ramadhani, Dimitry Rizal, and Erifive Pranatal. "Analysis of Design Changes from Main Engine Diesel Engine to Electric Engine on Fishing Vessels." Journal of Applied Sciences, Management and Engineering Technology 5, no. 1 (2024): 7–17. http://dx.doi.org/10.31284/j.jasmet.2024.v5i1.5965.
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Fishing vessels in Indonesia conventionally use diesel engines as their main propulsion engine. However, the use of diesel engines has quite a bad impact on the surrounding environment. So the latest innovations using alternative energy are needed. One of them is the use of electric engines as a replacement for diesel engines in the ship's main propulsion engine. The use of this electric machine will have a friendly impact on the environment. Apart from that, you don't need to spend a lot of money to maintain this electric machine. We can see this by comparing the ship's thrust requirements with maximum speed.
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Asraff, A. K., S. Sheela, Krishnajith Jayamani, S. Sarath Chandran Nair, and R. Muthukumar. "Material Characterisation and Constitutive Modelling of a Copper Alloy and Stainless Steel at Cryogenic and Elevated Temperatures." Materials Science Forum 830-831 (September 2015): 242–45. http://dx.doi.org/10.4028/www.scientific.net/msf.830-831.242.
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High performance rockets are developed using cryogenic technology. High thrust cryogenic rocket engines operating at elevated temperatures and pressures are the backbone of such rockets. The thrust chamber of such engines, which produce the thrust for the propulsion of the rocket, can be considered as structural elements. Often double walled construction is employed for these chambers for better cooling and enhanced performance. The thrust chamber investigated here has its hot inner wall fabricated out of a high conductivity high ductility copper alloy and outer wall made of a ductile stainless steel. The engine is indigenously designed and developed by ISRO and is undergoing hot tests. Inner wall is subjected to high thermal and pressure loads during operation of engine due to which it will be in the plastic regime. Evaluation of tensile properties of the copper alloy and stainless steel up to fracture, at cryogenic, ambient and elevated temperatures in parent metal and welded forms is of paramount importance for its constitutive modelling and thermo structural analysis of the thrust chamber.
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Knyshenko, Yu V., and V. M. Durachenko. "Mathematical model of the operation of a different-scale two-component low-thrust jet engine system." Technical mechanics 2022, no. 3 (2022): 47–62. http://dx.doi.org/10.15407/itm2022.03.047.
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The aim of this work is to modify a comprehensive mathematical model of a system of two-component low-thrust jet engines using the numerical method of characteristics in the propellant pipeline system with account for different sound speeds in the oxidizer and the fuel employing a unified method of pipeline discretization. This paper presents a unified approach to a numerical implementation of the method of characteristics for both fuel components and for regular computational cross-sections (internal for structural sections with constant geometrical and elastic parameters) and terminal cross-sections at the pipeline system inlets, the section joints, and the engine inlets for each propellant components. The approach accounts for the hydraulic resistances of the propellant injectors and electric propellant valves and the actual pressures in the engine combustion chambers. The performance of the mathematical model is illustrated by the example of the predesigning of a system of different-scale low-thrust engines to control the motion of a spacecraft relative to its center of mass in pitch, yaw, and roll and transfer the spacecraft to a new orbit (higher of lower) for maneuvering and docking with another spacecraft. The computed results show the possibility of determining the key hydraulic and gas-dynamic parameters of the system in transient conditions: the pressure and propellant component flow rate distribution at the inlet of any of the engines, the combustion chamber pressure and thrust characteristics of each engine, and the mutual effect of the engines on their thrust characteristics by the example of varying the flow areas of the propellant manifolds in the steady (continuous) and unsteady pulsed operation of all engines or some of them. The proposed mathematical model may be used in the computational justification of design parameters and operating conditions in the preparation of a draft proposal or in the predesign determination of an engine system configuration. Detailed information on the hydraulic and gas-dynamic performance parameters of an engine system is an important complement to the results of a ground tryout of both single engines and an engine system in conditions that simulate the flight environment.
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Timoshenko, V. I., L. K. Patryliak, Yu V. Knyshenko, V. M. Durachenko, and A. S. Dolinkevych. "Use of a “green” propellant in low-thrust control jet engine systems." Technical mechanics 2021, no. 4 (2021): 29–43. http://dx.doi.org/10.15407/itm2021.04.029.
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The aim of this work is to analyze the state of the art in the development and use of pollution-free (“green”) propellants in low-thrust jet engines used as actuators of spacecraft stabilization and flight control systems and to adapt computational methods to the determination of “green”-propellant engine thrust characteristics. The monopropellant that is now widely used in the above-mentioned engines is hydrazine, whose decomposition produces a jet thrust due to the gaseous reaction products flowing out of a supersonic nozzle. Because of the high toxicity of hydrazine and the complex technology of hydrazine filling, it is important to search for its less toxic substitutes that would compare well with it in energy and mass characteristics. A promising line of this substitution is the use of ion liquids classed with “green” ones. The main components of these propellants are a water solution of an ion liquid and a fuel component. The exothermic thermocatalytic decomposition of a “green” propellant is combined with the combustion of its fuel component and increases the combustion chamber pressure due to the formation of gaseous products, which produces an engine thrust. It is well known that a “green” propellant itself and the products of its decomposition and combustion are far less toxic that hydrazine and the products of its decomposition, The paper presents data on foreign developments of “green” propellants of different types, which are under test in ground (bench) conditions and on a number of spacecraft. The key parameter that governs the efficiency of the jet propulsion system thrust characteristics is the performance of the decomposition and combustion products, which depends on their temperature and chemical composition. The use of equilibrium high-temperature process calculation methods for this purpose is too idealized and calls for experimental verification. Besides, a substantial contribution to the end effect is made by the design features of propellant feed and flow through a fine-dispersed catalyst layer aimed at maximizing the monopropellant-catalyst contact area. As a result, in addition to the computational determination of the thrust characteristics of a propulsion system under design, its experimental tryout is mandatory. The literature gives information on the performance data of “green”-propellant propulsion systems for single engines. However, in spacecraft control engine systems their number may amount to 8–16; in addition, they operate in different regimes and may differ in thrust/throttling characteristics, which leads to unstable propellant feed to operating engines. To predict these processes, the paper suggests a mathematical model developed at the Institute of Technical Mechanics of the National Academy of Sciences of Ukraine and the State Space Agency of Ukraine and adapted to “green”-propellant engine systems. The model serves to calculate the operation of low-thrust jet engine systems and describes the propellant flow in propellant feed lines, propellant valves, and combustion chambers. To implement the model, use was made of the results of experimental studies on a prototype “green”-propellant engine developed at Yuzhnoye State Design Office. The analysis of the experimental results made it possible to refine the performance parameters of the monopropellant employed and obtain computational data that may be used in analyzing the operation of a single engine or an engine system on this propellant type in ground and flight conditions
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Miller, Cassandra J., Prakash Prashanth, Florian Allroggen, et al. "An environmental cost basis for regulating aviation NOx emissions." Environmental Research Communications 4, no. 5 (2022): 055002. http://dx.doi.org/10.1088/2515-7620/ac6938.
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Abstract Combustion in aircraft engines results in the formation of nitrogen oxides (NOx) and carbon dioxide (CO2), among other species. NOx impacts air quality and is an indirect contributor to radiative forcing, while CO2 is a long-lived greenhouse gas. The International Civil Aviation Organization sets limits on NOx emissions from commercial aircraft, where for engines with a rated thrust greater than 89 kN the allowable NOx production per unit rated thrust is defined as a function of engine overall pressure ratio (OPR). This definition links the engine thermodynamic cycle, and implicitly fuel burn and CO2 emissions, to allowable NOx levels. These regulations have historically been evaluated and implemented with a focus on reducing adverse air quality impacts around airports, but the thermodynamic efficiency tradeoff with CO2 requires additional analysis to quantify net environmental impacts. This paper introduces a social cost basis for evaluating aviation NOx emissions regulations and quantifies the implied CO2 and NOx attributable air quality damage, climate damage, and fuel costs associated with the emissions standard. We show that higher overall pressure ratio engines operating at the current NOx regulatory limit are allowed more environmental damage per unit rated thrust than lower overall pressure ratio engines, resulting in variable social costs per unit thrust (i.e. fuel and environmental costs combined) across the engine design space. This is a consequence of the definition of the regulation today, where higher pressure ratio engines are allowed higher NOx emissions. Alternative regulation definitions are evaluated which consider the engine cycle and combustor together. Achieving constant social costs requires a regulatory limit where the increase in allowed NOx emissions tapers off at higher pressure ratios, corresponding to the diminishing marginal efficiency improvements due to increasing OPR in that region.
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Xiang, Hao. "A novel quantum genetic algorithm with the application in LS-SVR." Journal of Physics: Conference Series 2637, no. 1 (2023): 012022. http://dx.doi.org/10.1088/1742-6596/2637/1/012022.
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Abstract The characteristics of LS-SVR are analyzed. LS-SVR is fitted for modeling small samples and high dimensional data, but the performance of LS-SVR is related to the specific data distribution, the kind of kernel function, its related kernel parameter, and the penalty coefficient. In this paper, the radial basis function is applied as the kernel function of LS-SVR, and the real double-chain coding target gradient quantum genetic algorithm (DCQGA) is applied to optimize the kernel parameter and penalty item coefficient of LS-SVR, then the regression prediction model DCQGALSSVR is proposed. It is of great significance to build an accurate and reliable fault prediction model for the health monitoring and fault diagnosis of liquid rocket engines. The thrust of a liquid rocket engine is an important factor in its health monitoring. By predicting the thrust change value and comparing the predicted value with the engine thrust threshold, it can be predicted whether the engine will fail at a certain time. In this paper, the proposed DCQGALSSVR model is used to model the thrust of a liquid rocket engine. The simulation results show that the average relative error is 0.37% using LSSVR for modeling on 12 test samples, and is 0.3186% using DCQGALSSVR on the same samples. It can be seen that DCQGALSSVR is effective for the health monitoring of liquid rocket engines, so it has a certain promotion significance.
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Wang, Yiwei, and Xianghua Huang. "Performance Seeking Control of Propfan Engines Based on Modified Cuckoo Search." International Journal of Turbo & Jet-Engines 37, no. 4 (2020): 363–70. http://dx.doi.org/10.1515/tjj-2017-0034.
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AbstractPerformance seeking control benefits propfan engine by generating optimal performance in different flight status. It is based on engine model, control model and optimization model. The control scheme of propfan engine is different from those of turbofan engines, for the thrust of propfan engines is mainly produced by propfans. After analysing the control structure of propfan engines, a control scheme for the propfan engine is proposed. The control scheme works well in flight envelope and the simulation results show that the overshoot of power shaft rotation speed is less than 2 % and the settling time is less than 0.9s. Based on this, a control scheme in performance seeking mode is proposed. A Modified Cuckoo Search method, which modifies the search step size and abandonment rate, is applied in the control scheme in maximum thrust mode and minimum fuel flow mode. The control scheme in performance seeking mode can reduce 2 % fuel flow, compared with the control scheme in torque-compensation mode. Performance of the scheme is better than standard Cuckoo Search and Genetic Algorithm.
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Zolotko, Olena, Oleksandr Zolotko, Oleksandr Aksonov, Vitalii Stoliarchuk та Oleksandr Cherniavskyi. "Аналіз характеристик ежекторного режиму імпульсно-детонаційної двигунної установки комбінованого циклу прискорення". Aerospace Technic and Technology, № 6 (21 листопада 2024): 52–59. https://doi.org/10.32620/aktt.2024.6.05.
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The subjects of this article are processes in the chamber of the pulse detonation propulsion system of the combined acceleration cycle (PSCA). The goal is to identify the most significant factors and study their influence on the defining characteristics of the ejector mode of operation of the pulse detonation PSCA. The task: to consider promising concepts of multi-mode engines and to investigate the influence of the mode parameters and design factors on the coefficient of increase of the thrust-specific impulse and the coefficient of increase of the thrust of the ejector. Methods of problem solving: computational analytical and experimental approaches. The following results were obtained. An analysis of the concept of multi-mode engines in transatmospheric aircraft was performed. The most famous of them include: two-mode direct-jet engine (ramjet-scramjet engine); rocket-based combined cycle (RBCC) engine; hybrid air-breathing rocket engine (Synergetic Air-Breathing Rocket Engine, SABRE). The propulsion system of the combined cycle of acceleration combines the main advantages of air-jet and rocket engines and can provide direct access to outer space (the concept of single stage to orbit, SSTO). A new stage in the development of PSCA is associated with the use of detonation mode combustion. During detonation, the pressure in the combustion chamber of the engine increases, which allows the pump to be excluded from the propulsion system. The following acceleration cycles are integrated into the pulse-detonation propulsion system: rocket-detonation; rocket-detonation with ejector thrust enhancement; jet-detonation with a direct or oblique detonation wave. In a promising engine for transatmospheric aircraft (Sodramjet), a stable and fixed oblique detonation wave is created under hypersonic flight conditions. Conclusions. It was established that with a certain combination of the studied factors, the coefficient of increase of the thrust-specific impulse reaches its maximum value. The dependence of the ejector thrust enhancement coefficient ΦP is determined by the flight conditions. The mode of self-similarity (independence) of ΦP from the flight height occurs at a speed of M = 5 and altitudes greater than 8 km. It is proposed to use a turbopump supply system without a gas generator to supply the rocket fuel components to the detonation combustion chamber. The influence of the initial pressure in the detonation chamber on the coefficient of increase of the thrust of the ejector ΦP and the magnitude of the specific impulse was investigated by computational fluid dynamics using the TVD scheme. The environmental advantages of using the detonation cycle of combustion in rocket and space technology are analyzed.
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Rajput, Pankaj, and Iraj Kalkhoran. "Optimization of Blockerless Engine Thrust Reverser." Journal of Propulsion and Power 33, no. 1 (2017): 213–26. http://dx.doi.org/10.2514/1.b36152.
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