Specific Thrust Research Articles

Novel Nuclear Power and Fossil Fuel Hybrid Propulsion Systems for Long-Endurance Unmanned Aerial Vehicles: Configuration Comparison and Performance Optimization

Aircraft nuclear propulsion system can significantly enhance the endurance, but its large size and weight become disadvantages against the conventional fossil fuel propulsion system. In this paper, a novel nuclear power and fossil fuel hybrid propulsion system, which combines the lithium-cooled fast reactor and the kerosene-fueled turbine engine, is proposed to meet the power requirements of different flight segments. Performance assessment and comparison among different hybrid configurations is performed. Results indicate that the configuration combining the nuclear power module with a high bypass ratio turbofan engine is the most promising for aircraft. The primary optimization shows that there is a trade-off between the weight of the nuclear power module and the specific thrust, and the optimal weight of nuclear power module is 3942 kg, with the specific thrust of 195.8 N·kg-1·s-1. The aircraft equipped with the hybrid propulsion system operates at the maximum altitude of approximately 18.2 km and reaches a maximum Mach number of 0.71, allowing for sustained flight for several months in cruise segment. The turbofan engine, nuclear reactor, ancillary systems, shield, and heat exchanger account for 11%, 23%, 11%, 17% and 1% of the fuel mass of the original unmodified aircraft.

Energy and exergy analyses on the high-pressure bleeding-air variable cycle designed for a wide-speed-range airbreathing propulsion system

The performance of the traditional turbojet decreases sharply with the increase of the Mach number in the supersonic flight, especially for the turbine-based combined cycle (TBCC) engine in the transition region of the turbojet and scramjet engine. This paper proposes a new high-pressure bleeding-air variable cycle engine (HB-VCE) concept that aims to ease the sharp contradiction between thrust output and fuel consumption of TBCC in the transition region. Through the energy and exergy analyses, the HB-VCE shows greater potential in energy utilization under high Mach flight conditions. The performance of HB-VCE significantly improves as the bleed ratio increases, with a more pronounced effect at higher Mach numbers. At Mach 1.5, the HB-VCE increases the specific thrust (ST) by 2.84 % and reduces the specific fuel consumption (SFC) by 2.73 %. At Mach 2.4, it increases the ST by 6.60 % and reduces the SFC by 6.52 %. Additionally, the controlling law of the critical bleeding ratio constrained by the turbine outlet temperature is proposed. Finally, the effect of the turbine inlet temperature, overall pressure ratio, and Mach number on the critical bleeding ratio is analyzed. This paper provides a promising HB-VCE concept with greater energy utilization for the TBCC design in the transition region.

Numerical investigation of micropropulsion systems for CubeSats: Gas species and geometrical effects on nozzle performance

In the present investigation, a numerical investigation is conducted to investigate cold-gas micropropulsion devices for CubeSat orbital control. Different micronozzle geometry configurations and gas species are analyzed to assess their influence on the flowfield structure, surface properties, and microthruster performance parameters. Due to the small length scales, the Direct Simulation Monte Carlo (DSMC) method is used to simulate rarefied argon and nitrogen in rectangular and curved micronozzles. The results indicate that nitrogen gas leads to a 23% increase in specific impulse at a relatively insignificant thrust cost. On the other hand, the use of a curved geometry increases the specific impulse and thrust generated by 23% and 35%, respectively. The results indicate that using lighter gases for cold gas microthrusters increases the micropropulsion efficiency. In addition, curved geometries significantly improve the overall performance of such devices, in contrast to rectangular geometries.

Teaching Aero-Engine Performance: From Analytics to Hands-On Exercises Using Gas Turbine Performance Software

Abstract In this paper, we present an approach for teaching turbojet engine performance that combines mathematical analysis with the use of performance software to provide students with a comprehensive and in-depth understanding of engine performance and how it is affected by design parameter choices. In the first part, we present the derivation of a universally valid analytical model for turbojet engine cycle design, which is the basis of the propulsion courses at RWTH Aachen University. The analytical model allows for the calculation of specific thrust as well as thermal and propulsive efficiencies as a function of pressure ratios and burner exit temperature. These are all expressed as differentiable functions. Their mathematical extremes form a basis for examining the intricate relationships in turbojet engine design. In the second part, we move from theory to practice. We utilize the GasTurb software, computing turbojet engine cycles for a wide range of cycle design parameters. The analytically derived trends are validated and errors introduced by simplifications are assessed. Differences between the analytical approach and software-based performance simulations of real engines, like temperature-dependent gas properties and secondary air systems, are emphasized and discussed, thus providing a bridge to real-world applications. In summary, this paper describes a two-step approach to teaching turbojet engine design using analytics and performance software. Each approach alone can be used to enrich and extend existing performance courses. Combining the approaches has significant benefits, improving understanding and inspiring students to pursue further research in the field of propulsion.

Thermodynamic analysis of the hydrocarbon-fuelled air-turborocket engine with complete-combustion gas generator

The Air-Turborocket (ATR) engine can work at Mach 0∼4 or even higher speed, which is considered one of the best low Mach number propulsion systems for reusable hypersonic vehicles. However, because the hydrocarbon-fuelled ATR engine uses a fuel-rich gas generator, the combustion product contains a large amount of C(gr) that can cause coking in the turbine in a few minutes. To solve this problem, an ATR engine cycle with a complete-combustion gas generator (ATR-CCGG) was proposed. The performance of this cycle has been analysed through the thermodynamic model, and the influence factors and the sensitivity study of the cycle performance have been investigated. The results show that when the equivalence ratio is 1, the cycle can get more than 700 s of specific impulse and 1000 m/s of specific thrust at supersonic speed. Although the performance at subsonic speed is lower than that of the LOX/Kerosene ATR engine, the gas generator without C(gr) can ensure the engine to work for hours without coking in the turbine at different Mach numbers, which can be used in reusable hypersonic vehicles or single-stage-to-orbit missions.

Study on the Improvement of Theoretical and Electric Field Simulation Methods for the Accurate Prediction of FEEP Thruster Performance

In this study, we investigate and propose an improved theoretical method to more accurately predict the performance of a field-emission electric propulsion (FEEP) thruster with its complex configuration. We identify critical flaws in the previous theoretical methods and derive corrected equations. Additionally, we define and implement the overall half angle of the Taylor cone to account for variations in the Taylor cone’s half angle depending on the applied voltage. Next, we also establish an improved method of the electric filed simulation in a three-dimensional domain to accurately predict a trajectory of extracted ions and a resulting spatial beam distribution of the FEEP thruster by incorporating a configuration of the Taylor cone with the estimated overall half angle from the results of the present theoretical method. Through comparison with the experimental measurements, we found that the present improved methods for theoretical and electric field simulations can yield more accurate predictions than those of the previous methods, especially for higher V and Iem regimes, which correspond to the actual operating conditions of the FEEP thruster. Consequently, we anticipate that the proposed methods can enhance the reliability and efficiency of the design process by accurately predicting performance when developing the new FEEP thruster with its non-symmetric complex configuration to match specific thrust or spatial beam requirements.

Molecular dynamics study of ionic liquid electrospray: Revealing the effect of types of electric fields

The microscopic behaviors and characteristics of BMIM+BF4− ionic liquid electrospray in the presence of the axial electric field and axial-radial electric field were investigated by molecular dynamics method. The formation and fragmentation of the jet, the structure characteristics of the ions, the interactions between the ions and the energy properties were analyzed. Since ionic liquid is commonly used in spacecraft thruster, their specific impulse and thrust are also evaluated. It was found that the occurrence of ionic liquid electrospray was determined by the axial voltage, not less than 0.8V/nm. The flow rate and radial voltage mainly affect the jet instability and the secondary disintegration. When the jet is broken, it is almost ejected in the form of (BMIM+BF4−)1(BMIM+) and BMIM+, but finally it presents four main clusters, including BF4−, BMIM+, (BMIM+BF4−)1 and (BMIM+BF4−)1(BMIM+). For the unbroken jet at the capillary outlet, BMIM+ is concentrated on the surface of the jet, and BF4− is concentrated inside. The number of clusters will occasionally decrease owing to coalesce. The internal bonds of BMIM+ and BF4− is likely to break, and the bond angle may also vary. In the presence of the axial-radial electric field with an axial electric field intensity of 1.2V/nm, the electrospray ionization effect is the best at the radial electric field intensity of 0.001V/nm, and the specific impulse and thrust can be obtained by adjusting the axial electric field intensity. This work provides a new insight into the microscopic mechanisms of BMIM+BF4− ionic liquid electrospray.

Synergetic and Performance Characteristics of A High-Speed Pre-Cooled Propulsion Concept

Abstract Pre-cooled air-breathing cycles are promising candidates to power future high-speed flight as well as Single-Stage-To-Orbit vehicles, due to their increased efficiency over contemporary propulsion systems and launch vehicles. These concepts usually feature complex interactions in the synergy of their thermodynamic cycles. In this study, a performance model of such a cycle is developed for its air-breathing mode of operation. One-dimensional thermodynamic modeling is employed within a component-level approach, to evaluate the performance and operation of the cycle under investigation in the range of 1.35 = ≤ 8 = 5 and conditions of up to 26 kilometers altitude. The model is validated quantitatively and qualitatively for both design and off-design conditions. The specific impulse Isp and specific thrust, as predicted by the model, agree within less than 5% for both design and off-design point conditions, while it captures the trend of Isp for the range modeled. Moreover the maximum gross thrust point is predicted correctly at M∞ = 4. The fundamental operating principles and synergetic characteristics of the engine at design and off-design conditions are investigated and reported. A model which does not feature a bypass duct is created and compared for the same inflow conditions and mission profile. It is found that the engine without the bypass duct exhibits reduced specific impulse up to 32% lower at off-design conditions while the overall trend of engine efficiency cannot be properly captured without modeling of the bypass duct, especially at the region of M∞ < 3.5.

Installed nacelle aerodynamics at cruise and windmilling conditions

PurposeThe decrease in specific thrust achieved by Ultra-High Bypass Ratio (UHBPR) aero-engines allows for a reduction in specific fuel consumption. However, the typical associated larger fan size might increase the nacelle drag, weight and the detrimental interference effects with the airframe. Consequently, the benefits from the new UHBPR aero-engine cycle may be eroded. This paper aims to evaluate the potential improvement in the aerodynamic performance of compact nacelles for installed aero-engine configuration.Design/methodology/approachDrooped and scarfed non-axisymmetric compact and conventional nacelle designs were down selected from a multi-point CFD-based optimisation. These were computationally assessed at a set of installation positions on a contemporary wide-body, twin-engine transonic aircraft. Both cruise and off-design conditions were evaluated. A thrust and drag accounting method was applied to evaluate different aircraft, powerplant and nacelle performance metrics.FindingsThe aircraft with the compact nacelle configuration installed at a typical installation position provided a reduction in aircraft cruise fuel consumption of 0.44% relative to the conventional architecture. However, at the same installation position, the compact design exhibits a large flow separation at windmilling conditions that is translated into an overall aircraft drag penalty of approximately 5.6% of the standard cruise net thrust. Additionally, the interference effects of a compact nacelle are more sensitive to deviations in mass flow capture ratio (MFCR) from the nominal windmilling diversion condition.Originality/valueThis work provides a comprehensive analysis of not only the performance but also the aerodynamics at an aircraft level of compact nacelles compared to conventional configurations for a range of installations positions at cruise. Additionally, the engine-airframe integration aerodynamics is assessed at an off-design windmilling condition which constitutes a key novelty of this paper.

EVALUATION OF THE SCATTER OF LIQUID LAUNCH VEHICLE POGO OSCILLATION AMPLITUDES DUE TO THE INFLUENCE OF THE SCATTER OF INTERNAL FACTORS

Almost all liquid launch vehicle developers faced the problem of ensuring stability in relation to POGO oscillations. The level of POGO amplitudes oscillations of the launch vehicle can be significantly affected by the scatter of internal factors. The study aims to create a mathematical model that can determine the range of POGO amplitudes in liquid launch vehicles. This will be demonstrated through the example of the Dnipro launch vehicle, which is affected by a variety of internal factors that cause its POGO amplitudes to vary. We developed the non-linear non-stationary mathematical model of POGO oscillations of the prototype of the Dnipro space launch vehicle. The model is built by taking into account the two lower vibration modes of the LV structure, two lower oscillation modes of the oxidizer feedline, and the first oscillation mode of the fuel feedline of the propulsion system. Modeling of dynamic processes was conducted in a combination of four liquid rocket engines based on the schematic of the staged rocket engine. The simulation takes into account cavitation phenomena in the engine pumps and delay times in the gas generators’ chambers. We have developed a method for determining the scatter of the POGO oscillations caused by the action of internal factors, which is based on the use of the LP uniformly distributed sequences. As internal factors, the frequencies, decrements, and shapes of LV structural oscillation modes, the values of pressurization of the propellant tanks, and the engines’ specific thrust impulses were considered. Based on the results of the calculations, the dependence of the POGO amplitudes in two regions of LV instability was determined, and the lower and upper enveloping curves for the POGO amplitudes were constructed. It is shown that the maximum POGO amplitudes oscillations in the first region of instability lie in the range from 0.23 g to 0.72 g and in the second region of instability — from 0 g to 0.60 g. Variants of combinations of internal factors, which provided the largest and smallest values of POGO amplitudes, were analyzed. This made it possible to determine the internal factors, the scatter of which has the greatest effect on the POGO amplitudes scatter: frequency, decrement, shape coefficients of oscillations of the oxidizer feedlines and the LV 1st mode structural longitudinal oscillations in the payload cross-section.

Analysis of carbon nano particle variant as the propellant fuel to increase specific impulses of rockets.

This study compares propellant fuels' specific thrust and impulse parameters using nanocarbon variant fuels from coconut shells and coal. Specific impulses and impulses are essential parameters that determine rocket performance. The specific thrust and impulses are influenced by fuel type, material composition, heat flow, and burning time parameters. The characteristics of nanocarbon as a fuel are proven to have high-value features and a long combustion time. This parameter is essential to add value to specific thrusts and impulses. The method to prove the quality of solid fuel material was done experimentally. The composition of the propellant fuel variant mixed with Ammonium Perchlorate (NH 4ClO 4) as an oxidizer through the printing process and sample testing was carried out using scanning electron microscope (SEM), X-ray diffraction analysis (XRD), combustion time, and specific impulse tests. Test results with the composition of CBK2 (Coconut Shell + NH 4ClO 4 + Binder + Al = 15: 70: 10: 5) produced a specific impulse of 267 seconds or an increase of 37 seconds without carbon. CBB2 (Coal + NH 4ClO 4 + Binder + Al = 15: 70: 10: 5) produced a specific impulse of 261 seconds or an increase of 31 seconds. The composition of nanocarbon as a solid fuel propellant has been proven to increase a rocket's thrust and specific impulses. After all the material variants have been tested for thrust and specific impulse, the best rocket propellant is CBK2.

Performance analysis of pulse detonation ramjet

Abstract A Pulse Detonation Ramjet (PDR) performance evaluation model is established based on the component-based method. The influence law of working parameters of pulse detonation combustor (PDC), PDR geometry parameters, flight speed and altitude on the performance is analyzed. Results are as follows. (1) With the increase of the equivalent ratio, the PDR specific thrust first increases and then decreases, reaching a maximum value around 1.2, but the fuel specific impulse always decreases; (2) The change of inlet throat cross-sectional area has little effect on specific thrust and fuel specific impulse, but the increase of nozzle throat cross-sectional area will have a negative effect; (3) PDR specific thrust and fuel specific impulse increase first and then decrease with the increase of flight Mach number and altitude; (4) Compared with a ramjet engine, the advantage range of PDR can reach 18.04 %, but gradually decreases with the increase of Mach number.

Dynamic Testing of a Hybrid-Propellant Water-Breathing Ram Rocket in Underwater Cruise Conditions

High-speed submerged marine vehicles, such as torpedoes, traveling at velocities of an order of 100 m/s and above, require powerful propulsion to overcome the tremendous hydrodynamic drag. This paper aims to investigate a marine hybrid-propellant water-breathing ram rocket (marine ramjet or ducted rocket) under various underwater cruise conditions. At high underwater cruise speeds, the ram rocket outperforms regular rocket motors, substantially increasing its specific impulse and thrust. This investigation utilized a unique test facility capable of dynamically testing the marine ramjet. Over 20 dynamic experiments have been conducted, revealing the submerged motor characteristics at different cruise speeds, water-to-propellant mass ratios, and oxidizer-to-fuel mass ratios, thereby creating a performance map of the marine ramjet. The results were compared with static firing data and theoretical calculations, showing a good agreement with standard specific impulse improvement of about 55% compared to a regular hybrid rocket, reaching a maximum value of 380 s. The significant increase in performance demonstrates the potential of the water-breathing ramjet for propelling high-speed underwater vehicles.

Air-breathing rotating detonation engine supplied with liquid kerosene: propulsive performance and combustion stability

Experimental results are presented for a rotating detonation engine supplied with liquid kerosene and preheated air without liquid or gaseous additions to the propellant mixture. Various combustion modes for the generic combustor geometry design were observed—from deflagration, through pulsed combustion and high-frequency instabilities, to stable detonation propagation. Attention was paid to detonation stability (if present), its characteristics, and the propulsive performance of the combustor with a focus on specific thrust and pressure gain through thrust and outlet total pressure measurement. These parameters measured for the observed modes were compared. The stability of the detonation combustion proved not to be critical to achieve high performance of the combustion chamber. For example, high performance was achieved for combustion modes with high-frequency instabilities.

Insights into thermodynamic performance of a hypersonic precooled air-breathing engine with a complicated multi-branch closed cycle

An advanced precooled airbreathing engine with a closed Brayton cycle is a promising solution for high-speed propulsion, of which the Synergetic Air Breathing Rocket Engine (SABRE) is a representative configuration. The performance of the latest SABRE-4 cycle was analyzed in this paper. Firstly, a relatively complete engine performance model that considers the characteristics of turbomachinery and heat exchangers was developed. Then, Sobol’ global sensitivity analysis of key performance parameters was carried out to identify the most influential design variables. Optimal specific impulses under different target specific thrusts were obtained by particle swarm optimization, of which the thermodynamic parameters corresponding to a specific thrust of 1.12 kN·s·kg−1 and a specific impulse of 3163 s were chosen as the design values. Four different control laws were analyzed in contrast, and the charge control method had the strongest ability of thrust regulation as well as maintaining a favorable specific impulse performance. Finally, working characteristics under the charge control and over a typical flight envelope were calculated, in which the average value of the maximum specific impulse was as high as 5315 s. This study would help to deepen the understanding of SABRE-4 thermodynamic characteristics and other precooled airbreathing engine cycles with similar layouts.

Design & Development of Hybrid Electric Jet Engine

Hybrid electric jet engines without turbines, in which the compressors are powered by fuel cells rather than the turbines, are a potential and evolving technology for unmanned aerial vehicles. The hybrid electric jet engine combines the benefits of fuel cells and turbo-electric engines. The hybrid engine produces high specific thrust and excellent thermal efficiency.

Investigation of overall performance of turbofan engine with pulse detonation combustor in bypass duct

The performance model was established to study the performance of mixed exhaust turbofan engine with pulse detonation combustor (PDC) in the bypass duct. The effects of the total pressure loss of isolator and bypass circulation parameters on the characteristic of PDC and performance of engine were analyzed. Then, the sensitivity analysis of performance parameters was carried out on the influence of the components parameters. The performance of mixed exhaust turbofan engine with PDC in the bypass duct was compared with that of the conventional afterburner turbofan engine at the same design circulation parameters. The results show that to increase the total pressure loss of isolator is beneficial to the improvement of performance because of the improvement of pressure ratio; when fan pressure ratio is constant, with the increasing of bypass ratio, the specific fuel consumption and specific thrust increase; with the increasing of fan pressure ratio, when bypass ratio is 0.2 to 0.4, specific thrust increases first and then decreases. When bypass ratio is 0.4 to 0.5, specific thrust increases first and then remains constant. When bypass ratio is 0.5 to 0.9, specific thrust increases, but the rate of increase reduces. At different bypass ratio, specific fuel consumption decreases with increasing of fan pressure ratio; the performance of engine is more sensitive to the component parameters that directly affect the flow state parameters of the bypass duct; because of supercharging performance of PDC, PDC in the bypass duct only uses part of air flow to organize combustion that can make specific thrust of turbofan engine with PDC in the bypass duct and conventional afterburner turbofan engine are equivalent. Meanwhile, specific fuel consumption can be reduced by 27.7% at design point and 12.8%-26.8% at off-design points.

Www.ijeit.misuratau.edu.ly ISSN 2410- Paper ID: EN A Comparison Study For the Performance of the Mixed and Unmixed Flow Two Spool Turbofan Engines at Variable Parameters

The following paper is to investigate the influence of some operational and design parameters on the performance of a turbofan engine with two spools and for mixed and unmixed flow. Because there are many influential variables on the performance, here a computer program with VISUAL BASIC has been developed to carry out this study. The input variables taken in consideration are the flight speed, compressor pressure ratio, fan pressure ratio, the altitude and the combustion temperature. The output obtained at these variables, which are the thrust specific fuel consumption, specific thrust, propulsive efficiency and thermal efficiency are depicted and discussed for the two types of flow.The results show that there is a great influence of these variables on the performance parameters.

Preliminary design of lunar high-resolution remote sensing micro-nano satellite

The development and utilization of lunar resources and the construction of lunar scientific research stations are important parts of China’s future lunar exploration plan. Aiming at the problems of China’s micro-nano satellites with low radiation resistance, lack of autonomous Earth-Moon transfer capability, and limited optical system volume and quality, a preliminary design scheme of a lunar remote sensing micro-nano satellite with autonomous Earth-Moon transfer capability is proposed. Using the analytical orbit average technology and transcription and collocation method to calculate the satellite’s Earth-Moon transfer orbit, the satellite can use the Hall thruster with large specific impulse and low thrust to realize the autonomous Earth-Moon transfer. The passive + active method is used to optimize the radiation resistance of the system, and the total radiation dose of some single machines can reach more than 25krad. A compact high-resolution multispectral camera is designed, and imaging simulation and verification are carried out. In the end, the satellite can achieve high-resolution imaging of the moon with a resolution of 0.2 m at an orbital height of 100 km to achieve the purpose of fine exploration of information such as lunar surface topography and geological structure.

INVERTED BRAYTON CYCLE ENGINE OPTIMIZATION FOR HYPERSONIC FLIGHT

Optimization of inverted Brayton cycle engine is made by particle swarm optimization method in this study. The optimum specific thrust value is reached by staying within the optimization constraints. When the total temperature of the cooling section is examined, it is seen that a temperature above the freezing temperature of the air is obtained. A very high specific thrust value, 451 N.s/kg is obtained at the hypersonic flight Mach Number (5 Mach) as a result of optimization. Temperature decrease with increasing altitude effect performance dramatically as positive. Low preburner exit total temperature and turbine total pressure ratio values and high afterburner exit total temperature values are desirable in terms of performance but values of total temperature at cooling section effect performance conversely in terms of specific thrust and specific fuel consumption