Specific Thrust Research Articles

Sustainable consumer behaviour of Indian millennials: Some evidence

Despite policy thrusts and initiatives, driving consumer behaviour towards sustainable consumption is a challenging task. The paper in this context carries out a detailed analysis of consumer behaviour in the Indian context with a specific thrust on sustainable consumer behaviour. The paper builds on the theoretical conception of the ‘Value Belief Norm (VBN)’ framework developed by Stern et al. (1999) and adapts and expands the same in the Indian context by employing exploratory factor analysis. Based on consumption behaviour patterns, the study classifies Indian millennials into five different categories such as aware consumers, conscious consumers, consumers under transition, unwilling consumers, and rejecters respectively, which help to map the heterogeneities present among the Indian millennials in terms of their consumer behaviour. In addition, it emerges from the study that Indian millennials are more concerned about the environmental dimension of consumption than social and economic dimensions, which has clear policy implications in terms of creating awareness and sensitising people.

Multiple ionization characteristics of hall thruster with large height–radius ratio and their effects

Increasing the specific impulse has become one of the main trends in the development of Hall thrusters and may be achieved by increasing the proportion of multiply charged ions in the plume. In this study, the multiple ionization characteristics of Hall thrusters with large height–radius ratios and their effects on the specific impulse and other performance parameters were examined via numerical simulations. The simulation results show that the proportion of multiply charged ions in a Hall thruster with a large height–radius ratio increases by 7.8 percentage points, compared with that of the traditional Hall thruster, and the multiply charged ions can obtain a higher average jet velocity under an accelerating voltage, thereby improving the specific impulse and thrust of the thruster. Further analyses show that a structure with a large height–radius ratio reduces the cooling effect of the channel wall on the electrons, and the increase in electron temperature effectively promotes the occurrence of the multiple ionization process. In addition, the smaller wall area and the larger magnetic field intensity in the near-wall region inhibit the electron near-wall conduction and prevent the potential negative effect of the increase in electron current caused by multiply charged ionization, which ensures a high level of efficiency. It was proved that multiple ionizations contribute positively to the specific impulse and efficiency of Hall thrusters with a large height–radius ratio. This serves as a reference for the high specific impulse design of Hall thrusters with large height–radius ratios and provides a new optimization perspective for improving the specific impulse of Hall thrusters.

Numerical Study on the Effect of Natural Gas on Aircraft Turbofan Engine Propulsion

The world has been concerned and worried about the depletion of the liquid oil fuel, besides new environmental rules are to be followed to reduce pollution hazards and global warming. The utilization of natural gas (as a near term fuel) and hydrogen (as long-term fuel) are receiving great attention, because they have less pollution effects. Since the aviation has a great deal in environmental pollution effects due to the cruise flight in the upper troposphere (supersonic aircraft) or in the lower stratosphere (subsonic aircraft) where most of the ozone concentrate, which helps in protecting the earth form ultra violet radiation. Therefore, the use of alternate fuel has a great attention in aviation. In the present study, the thrust specific fuel consumption and specific thrust for the aircraft during aircraft flight profile are predicted, when using aviation fuel and natural gas. The P&W JT9D –7R-turbofan jet engine is taken as a base line engine propelling the Boeing 747-200 aircraft as a base line aircraft with four engine nacelles mounted on wings. The model engine fuel-air cycle representation is carried out for design point calculations based on sea level static conditions and variable specific heats along engine components. The predicted engine performance results compared very well with the reported values by the manufacture. Predictions carried out using aviation fuel and natural gas show an increase in the specific thrust by 3% and decrease in the thrust specific fuel consumption by 14% and fuel to air ratio by 11%, when using natural gas.

Additive-manufactured single-piece thin multi-layer tungsten heater for an electrothermal thruster.

In this study, a novel single-piece thin multi-layer tungsten resistive heater was successfully fabricated using additive manufacturing and tested as an electrothermal thruster. The heater has 12 resistive layers, with each layer having a thickness and height of 0.15 and 81mm, respectively, and can provide high heating efficiency. A single-piece or monolithic heater was manufactured via additive manufacturing technique, which drastically improved its reliability and decreased its manufacturing cost. In the heating and thrust measurement tests that used nitrogen gas as a propellant, the heater reached a gas temperature of ∼2000K at a 140-A heater current without experiencing any failure. The tungsten-heater resistance linearly increased with an increase in temperature due to the temperature dependence of tungsten's resistivity. The specific impulse and thrust increased with the heater temperature in accordance with the theoretical prediction. Even including a voltage drop due to a contact resistance, the achieved heater efficiency reached 63% at a 100-A heater current even without a thermal insulation around the thruster. The heater efficiency decreased with an increase in the heater temperature due to heat loss to the surroundings. The heat-loss analysis indicated that both thermal conduction and radiation heat losses were crucial for improving the heater performance at a high-temperature operation of over 2000K.

Laser–Accelerated Plasma–Propulsion System

In this paper, the laser-accelerated plasma–propulsion system (LAPPS) for a spacecraft is revisited. Starting from the general properties of relativistic propellants, the relations between specific impulse, engine thrust and rocket dynamics have been obtained. The specific impulse is defined in terms of the relativistic velocity of the propellant using the Walter’s parameterization, which is a suitable and general formalism for closed–cycle engines. Finally, the laser-driven acceleration of light ions via Target Normal Sheath Acceleration (TNSA) is discussed as a thruster. We find that LAPPS is capable of an impressive specific impulse Isp in the 105 s range for a laser intensity I0≃1021W/cm2. The limit of Isp≲104 s, which characterizes most of the other plasma-based space electric propulsion systems, can be obtained with a relatively low laser intensity of I0≳1019W/cm2. Finally, at fixed laser energy, the engine thrust can be larger by a factor 102 with respect to previous estimates, making the LAPPS potentially capable of thrust-power ratios in the N/MW range.

Dual mode operation of a hydromagnetic plasma thruster to achieve tunable thrust and specific impulse

We report here on initial studies of a pulsed hydromagnetic plasma gun that can operate in either a pre-filled or a gas-puff mode on demand. These modes enable agile and responsive performance through tunable thrust and specific impulse. Operation with a molecular nitrogen propellant is demonstrated to show that the hydromagnetic thruster is a candidate technology for air-harvesting and drag compensation in the very low Earth orbit. A dual mode operation is achieved by leveraging propellant gasdynamics to change the fill fraction and flow collisionality within the thruster. This results in the formation of distinct modes that are characterized by the current-driven hydromagnetic waves that they allow, namely, magneto-deflagration and magneto-detonation, respectively. These modes form the basis of using gasdynamics to enable responsive thruster performance. Using time-of-flight emission diagnostics to characterize near-field flow velocities, we find that a relatively dramatic transition occurs between modes as gas is allowed to expand in the thruster, with exhaust velocities ranging from 10 to 55 km/s in the deflagration and detonation regimes, respectively. Simulations of the processed mass bit offer the first glimpse into possible thruster performance and trade-offs between specific impulse and thrust. An impulse bit tunability of ∼22% is predicted, with differing propellant fill fractions when operating in a burst mode.

Charge transport characterization of the alternative low power hybrid ion engine (alphie) with particle-in-cell simulations

The Alternative Low Power Ion Engine (alphie) is a high specific impulse plasma thruster where both electrons and ions flow combined through the open spaces of its two-grids system. Ionizing electrons from its external cathode travel in toward the ionization chamber accelerated by the potential drop established between the grids, whereas ions exit toward the opposite direction and are later neutralized by electrons emitted from the cathode. In this configuration, the ion current is not space-charge limited, contrary to conventional gridded ion engines where only ions are transported through the grids. Since previous approaches that consider only one charged species are not valid for the alphie two-grids system, the characterization of geometrical parameters is studied using a fully particle-in-cell simulation. This numerical scheme calculates the velocities and positions of electrons and ions within a structured axial-symmetric mesh, which mimics one hole of the grids. The charged particle densities, currents, and their axial and radial mean velocities along the simulation domain are evaluated. Additionally, important reductions in the computational time have been obtained using different time scales for ions and electrons. Numerical results show that moderate changes in the thickness and separation of the grids with respect to the values employed in the prototype currently tested in the laboratory do not significantly affect the exhaust ion current or the ion beam collimation. Since the average velocity of the axial ions is essentially insensitive to small modifications of the geometric parameters, so will the specific impulse and thrust delivered.

Molecular dynamics simulation of ionic liquid electrospray: Microscopic presentation of the effects of mixed ionic liquids

Micro-propulsion devices based on electrospray have advantages of high specific impulse, high precision and compact structure, and are able to meet the requirements of gravitational wave detection and other drag-free spacecrafts for micro-propulsion system. For electrospray thrusters, ionic liquid is the preferred propellant because of its low volatility, high electrical conductivity and thermal stability. Most of the studies presented today focus on single ionic liquid, while the properties of single ionic liquid are certain, which leads to a limitation of electrospray characteristics and hence propulsion performance. Mixing different kinds of ionic liquids broadens the properties of ionic liquids, and thus adjusts the characteristics of electrospray to meet the requirements of diversified space missions. In this paper, the effects of mixing two typical ionic liquids, 1-ethyltrimethylimidazole dicyanamide (EMIM-N(CN)2) and 1‑butyl‑trimethylimidazolium hexafluorophosphate (BMIM-PF6), on propellant properties and electrospray characteristics are investigated by molecular dynamics simulation and validated by comparing with experimental data. The result shows that the mixed ionic liquid changes the emission mode of electrospray thruster. Compared with BMIM-PF6, the specific impulse and thrust for the 50:50 (by mass) mixture increases by more than half of the difference between the two before mixing. In addition, the analysis shows that the range of extrusion voltage is widened by adjusting the mixing ratios to meet the mission requirements for certain specific impulse and thrust.

The aerothermodynamic cycle optimal design of a turbofan engine

Abstract To ensure that the aerothermodynamic cycle design of a turbofan engine is more accurate, efficient, and provide a reliable decision-making basis for engine designers, the multi-objective particle swarm optimization (MOPSO) method was used to optimize the aerothermodynamic performance parameters of the turbofan engine at multiple design points (MDPs). Fuel consumption rate and the specific thrust were considered as optimization targets. The thrust requirements and cycle parameter constraints under each working state were comprehensively considered to obtain the optimal performance boundary of the engine, the corresponding cycle parameters, and the correlations between different requirements and constraints. The results showed that the MOPSO algorithm could accurately and completely obtain the optimal performance boundary surface of the engine in the feasible region and the corresponding cycle parameter value. The feasible region obtained by the aerothermodynamic cycle design at MDPs was more accurate and effective than the design at a single design point.

Impact of installation on a civil large turbofan exhaust system at idle descent conditions

Recent trends in civil aero-engine design aim at lowering specific thrust and improving propulsive efficiency by increasing the bypass ratio and therefore, usually also the fan diameter. The integration of these larger diameter engines with the airframe is critical to exhaust performance, and it is important to include these effects in engine performance analysis. The discharge coefficient of the bypass and core nozzles of a high-bypass ratio aero-engine at idle descent conditions is investigated numerically for an aero-engine in isolation and installed on an airframe. The discharge coefficients influence the engine operating conditions and turbomachinery re-matching at these off-design conditions. The maximum difference in the bypass nozzle discharge coefficient between the installed and isolated aero-engine across the descent phase is ≃1.6%. The differences in the core nozzle discharge coefficient between the installed and the reference isolated configuration are ≃43% and ≃−5.4% at the start and the end of the descent phase, respectively. The nozzle discharge coefficients depend on flight Mach number, incidence angle, and the nozzle pressure ratios of the fan and core nozzles. Multiple competing flow mechanisms govern the static pressure on the core nozzle base, which influences the core nozzle discharge coefficient. A novel reduced-order model is developed to estimate the core nozzle discharge coefficient for the installed configuration in idle descent conditions. This approach is based on the effective nozzle pressure ratio and can be implemented in engine performance simulations.

A history of jet noise research at the National Aeronautics and Space Administration.

This paper reviews jet noise research conducted at the National Aeronautics and Space Administration (NASA) from the early 1950s to the present day. Research conducted by NASA's predecessor, the National Advisory Committee for Aeronautics (NACA), and during the early years of NASA focused on turbojet noise, where a common approach for reducing jet noise was to limit the jet exit velocity to speeds that provided acceptable noise levels. Suppressors tested during this time resulted in thrust losses that were too severe to be implemented. With the introduction of turbofan engines in the 1960s, NASA shifted research to programs for both subsonic and supersonic aircraft applications with specific noise reduction goals. Subsonic research focused on increasing the bypass ratio of the engine to reduce the jet exit velocity of the core exhaust and adding mixers to the dual exhaust streams. Advances in computational methods improved aerodynamic designs and jet noise prediction tools. Supersonic applications proved to be more troublesome as programs aimed at large commercial transports required higher specific thrust engines. Changing the engine cycle to reduce jet noise was not compatible with mission range and speed requirements. Research for supersonic commercial aircraft remains an area of interest today at NASA.

A novel full-electric aircraft propulsion based on the DEA compressor

This article introduces a novel full-electric aircraft propulsion designed based on the DEA compressor, named DEAThruster that is supplied by a PEMFC. The thruster designed and modeled in this study is a compact DEAThruster to operate in altitudes up to 15,000 m and in the subsonic/transonic region with a specific thrust around 8300 N s/kg. The results show the thruster can satisfy all expectations, and it can generate up to 8300 N s/kg specific thrust for flight conditions encompassing static condition at sea level up to flight Mach number 0.95 in altitude 15,000 m. The DEAThruster can potentially be a practical alternative for gas turbine propulsions in all aspects when all available options for full-electric propulsions are not competitive for conventional aircraft propulsions in performance, size, and weight.

Performance evaluation for scramjet based on ground direct-connected test: A method investigation

Scramjet has become one of the most efficient hypersonic vehicles because of its broad application potentiality. The ground direct-connected test is a relatively easy and economical way to test the performance of the scramjet under “real” condition. In this paper, an efficient performance evaluation method based on ground direct-connected test was developed. It can determine the performance parameters of the scramjet combustor (internal resistance, combustion efficiency, and total pressure recovery coefficient) and that (specific impulse and thrust) of the overall scramjet during the working time under the assumptions of adiabatic inlet and adiabatic-isentropic nozzle. Compared with the previous evaluation methods, using this method not only greatly reduces the time and money cost, but also provides more accurate boundary conditions for unsteady numerical simulation, and reference values for free jet and flight test furtherly. In the end, a case analysis was conducted and the real-time performance parameters were obtained. The results show that the internal resistance, combustion efficiency, and total pressure recovery of combustor are about 280 N, 0.69, and 0.21, respectively. The specific impulse and thrust of the overall scramjet reach about 6000 Nskg−1 and 370 Nskg−1, correspondingly.

Low-topography deep-seated gravitational slope deformation: Slope instability of flysch thrust fronts (Outer Western Carpathians)

Unlike high-relief mountain areas, low-relief hilly landscapes are usually rarely affected by deep-seated gravitational slope deformations (DSGSDs). However, low-topography flysch thrust fronts can create suitable structural conditions for DSGSDs. The study area of the Kavalčanky ridge represents a relatively low-lying (<120 m of local relief) DSGSD-affected ridge situated at the thrust front in the flysch Outer Western Carpathians (Czech Republic). With the aim of revealing the main controlling factors and temporal constraints of mass-movement activity, a multidisciplinary investigation of DSGSDs was performed. Typical DSGSD landforms were mapped using high-resolution LiDAR-based mapping. Structural analysis revealed the presence of specific flysch thrust structural conditions with competent sandstone units overthrusted on a weak tectonically disrupted claystone basement. Geophysical measurement with the use of electrical resistivity tomography (ERT) and ground penetrating radar (GPR) profiling confirmed the deep reach (>50 m) of the studied DSGSD. Together with slope stability finite element modelling, geophysics suggests that the shear zone of DSGSDs is represented by one of the thrust faults, whereas lateral limits are formed by a set of conjugate strike-slip faults. Radiocarbon dating of bogs within the DSGSD body showed two phases of mass-movement activity corresponding to the Late Glacial-Holocene transition and Middle Holocene. The recent activity was excluded by dendrogeomorphic analysis. We concluded that the structural conditions involving high lithological complexity, tectonic weakening and the presence of major tectonic contacts might create conditions prone to DSGSDs even in relatively low-topography settings with mass movement activity, especially during humid and warmer Late Quaternary periods.

Effects of isentropic efficiency of turbomachinery components on entropy production for small turbojet engine

PurposeUsage of gas turbine engines has increased by day due to rising demand for military and civil applications. This case results in investigating diverse topics related to energy efficiency and irreversibility of these systems. The purpose of this paper is to perform a detailed entropy assessment of turbojet engines for different flight conditions.Design/methodology/approachIn this study, for small turbojet engines used in unmanned aerial vehicles, parametric cycle analysis is carried out at (sea level-zero Mach (hereinafter phase-I)) and (altitude of 9,000 m- Mach of 0.7 (hereinafter phase-II)). Based on this analysis, variation of performance and thermodynamic parameters with respect to change in isentropic efficiency of the compressor (CIE) and turbine (TIE) is examined at both phases. In this context, the examined ranges for CIE is between 0.78 and 0.88 whereas TIE is between 0.85 and 0.95.FindingsIncreasing isentropic efficiency decreases entropy production of the small turbojet engine. Moreover, the highest entropy production occurs in the combustor in the comparison of other components. Namely, it decreases from 2.81 to 2.69 kW/K at phase-I and decreases from 1.44 to 1.39 kW/K at phase-II owing to rising CIE.Practical implicationsIt is thought that this study helps in understanding the relationship between entropy production and the efficiency of components. Namely, the approach used in the current analysis could help decision-makers or designers to determine the optimum value of design variables.Originality/valueDue to rising isentropic efficiencies of both components, it is observed that specific fuel consumption (SFC) decreases whereas specific thrust (ST) increases. Also, the isentropic efficiency of a compressor affects relatively SFC and ST higher than that of the turbine.

Turbofan engine performances from aviation, thermodynamic and environmental perspectives

In this paper, JT15D turbofan engine and its main subcomponents are assessed with the aviation, energy, exergy, environmental, and sustainability analyses. In the first stage, the system's specific thrust and specific fuel consumption of the engine are found as 315.9 N s/kg and 15.8 g/kN.s, respectively. Then, system's energetic efficiency is estimated as 21.15% and exergetic efficiency is accounted to be 19.919%. The system's exergetic improvement potential rate, productivity lack ratio and fuel exergy waste ratio are determined as 1573.535 kW, 402.024% and 80.081%, respectively. The ecological and environmental effect factors of the turbofan system are computed to be 5.020 and 4.020, respectively, while the sustainable efficiency factor and exergetic sustainability index rates are found as 1.249 and 0.249, respectively. Finally, among the system parts, the Combustion Chamber (CC) has minimum rates of sustainable efficiency factor, exergetic efficiency and sustainability index, while it has utmost rates of ecological and environmental effect factors, fuel exergy waste ratio, irreversibility and productivity lack ratios. As a general conclusion, the combustion chamber and low pressure compressor components should be optimized for better performance of the system.

Performance assessment of a solid oxide fuel cell turbine-less jet hybrid engine integrated with a fan and afterburners

Reducing carbon emissions in the field of aviation is important with the rising prosperity of air transport. Electric aircraft play an essential role in the process. Hybrid engines, composed of a solid oxide fuel cell, a fan, a compressor, and an afterburner are proposed. Fuel cells provide energy to the compressor and the fan. Internal duct air and external duct air are mixed at the inlet of the nozzle. Propulsion power is output by the nozzle. In order to study the effects of operating parameters on the performance of the hybrid engine, thermodynamic performance models are built. The main conclusions are as follows: (1) The specific thrust and specific impulse of the engine are dramatically increased with increasing bypass ratios. Meanwhile, fuel cell power rises slowly, which indicates that the engine with fans has more advantages than the one without fans in the respect of thrust-weight ratios. (2) Fuel cell power first increases then decreases with increasing fan pressure ratio under the conditions of constant overall pressure ratios. (3) Increasing afterburner temperature is the most effective way to improve the thrust of the engine. When the temperature is 2217 K, the specific thrust is up to 3085 N/(kg⋅s−1). Meanwhile, the specific impulse and thermal efficiency are 2693 s and 60.5%, respectively.

Thermodynamic parameters optimization of an aviation three-shaft turbofan engine with an intercooler and a recuperator under flight condition

The use of heat recovery in turbofan engines to improve their efficiency is not so obvious. Installing a recuperator behind the turbine of the main engine duct to heat the air at the entrance to the combustion chamber, on the one hand, reduces fuel consumption, but on the other hand, leads to a decrease in the temperature at the entrance to the jet nozzle of the main duct, thereby reducing the gas flow rate from the nozzle and, consequently, the specific thrust and thrust of the engine as a whole. The installation of an intercooler for intermediate cooling of the air between compressors allows to reduce the net-work and power of the high-pressure compressor and increase the gas pressure behind the turbine, and, consequently, the speed, specific thrust, and thrust of the engine. However, the amount of possible gain in the efficiency of a turbofan engine due to heat recovery will depend on how well the parameters of the working process of the engine and heat exchanger are selected. And this can be done only by optimizing these parameters in terms of evaluating the efficiency of the engine in the aircraft system. The article presents the results of optimizing the parameters of the working process of three-shaft turbofan engines with an intercooler and a recuperator in the system of a passenger aircraft of the Boeing 737 MAX 7 according to such criteria as the total weight of the power plant and the fuel required for the flight, and the specific fuel consumption of the aircraft per ton-kilometer of the transported commercial load.

Comparative study of PTFE filled different dopants as propellants for laser-electric hybrid thruster

This study presents a broad comparison of samples, which are made of polytetrafluoroethylene (PTFE) filled with different dopants according to the same preparation steps, as propellants on a laser-electric hybrid thruster with a common design base. The impulse bit, specific impulse, efficiency and discharge success rate of the propellant samples were measured and analyzed under the same operation. The experimental results illustrated that the dopants of alloy could improve the impulse bit, that of salt could promote the stability of the discharge process, and the samples doped with oxide performed dramatic specific impulse and thrust efficiency. And three representative samples were adopted to investigate the dependence of propulsion characteristics on different working parameters in the respective doping ratios. It is found that the sample doped with TiO2 in lower doping rate was expected to rank maximal specific impulse and efficiency at most operations, especially under the conditions of low laser energy, long discharge width and high initial voltage. And the samples doped with Co–Cr alloy in low doping ratio appeared better performance at low discharge width and laser energy, while those in high doping ratio performed greater performance at high discharge width and laser energy.

Thermodynamic performance analysis of the rotating detonative airbreathing combined cycle engine

A configuration is presented for the rotating detonative airbreathing combined cycle engine with compatibility between the turbomachinery and rotating detonation combustor being realized. Two propulsion units are merged to achieve three different operating modes, which provides a favorable propulsion choice for advanced full range and trans-aerosphere vehicles. For the vehicle taking off from the ground, the engine operates at turbojet mode until the thrust demand could not be satisfied. Then the ramjet unit ignites and the operating mode turns into transition mode, during which the ratio of the thrust generated by ramjet unit to the total thrust increases with an increase in Ma0. When the ratio reaches 1.0, the engine transforms into the ramjet mode. Thereafter, a mode transition strategy with the thrust and freestream mass flowrate remaining constant is proposed for the combined cycle engine, based on the steady state characteristics of turbine and ramjet units. Furthermore, the parametric performance simulation model of the engine is developed to verify the potential performance improvements generated by the application of rotating detonation technology, and analyze the application characteristics of the mode transition strategy formulated in this study. It is shown the lower initial Mach numbers of mode transition result in higher values of equivalent specific thrust but lower values of equivalent specific impulse during the transition mode. For the detonative combined cycle engine, the minimum value of initial Mach number of mode transition is below 2.0, which is much lower than that of the conventional turbine based combined cycle engine. Therefore, the performance during transition mode is expected to be improved with the application of rotating detonation technology.