Thrust Ratio Research Articles

Analysis of the Principle of Jet Engine and State-of-art Application

Abstract. In 1903, the Wright Brother made the dream of flight come true. Contemporarily, as the development of modern technology, aviation has become a major type of daily transportation. The jet engines are the magic that makes the airplanes soar high in the blue sky. As the jet engine is a major invention in human history that brought humans to the sky, it is meaningful to gig into the principle behind the engines. In this article, the principle of 5 jet engines are explained. There are benefits but also impacts due to the human beings. Environmental problems are caused by the exhausted gases, healthy of people who work and live near the airstrips and airports are threatened. Different types of jet engine produced by different companies latest are enumerated and stated with detailed data such as thrust and bypass ratio. According to the analysis, the current limitations and challenges of the development with the future prospect are stated.

Research on layout optimization for propellers of a submarine operation vehicle

The submarine operation vehicle was an important equipment for exploring the marine environment and performing underwater tasks. In this paper, the calculation and analysis of hydrodynamic performance and the layout optimization of the propellers under different layouts were completed. Based on the left-right symmetrical layout of the four ducted propellers, the two ducted propellers were taken as the research object to optimize its vector layout, and an approximate model was constructed according to the numerical calculation results under different combinations of propeller spacing and deflection angle. The simulation calculation and verification are carried out, and the optimal layout of the four ducted propeller were determined considering the interference between the ducted propellers on both sides. Finally, an experimental platform for the thrust test of the two ducted propellers were built. It was found that the axial thrust ratio of the rear propeller and the front propeller in experimental results were in good agreement with the simulation results, which verified the feasibility of using the numerical simulation method to study the mutual interference between the ducted propellers.

A novel simplified simulation method for the precooler based on the virtual source term: modeling, validation and application

Precooling technology is one of the promising approaches for expanding the flight envelope of conventional turbine engines and enhancing flight performance at high Mach numbers. Within this domain, a crucial challenge is to rapidly and accurately obtain the flow and heat transfer characteristics of the precooler. Currently, full-size simulation of the precooler is impractical, while simplified and fast simulation methods employed in previous research have exhibited certain deficiencies, including excessive simplification and unconvincing validation. In this study, a novel simplified simulation method for the annular precooler based on the virtual source term model was proposed. This method comprehensively considered the axial and radial non-uniformity of airflow caused by the coupling effect between the precooler and the inlet channel in the actual flow field. Subsequently, systematic validations of the reliability and superiority of this method were conducted, along with its application in the overall performance calculation of the precooled engine. The results demonstrate that the simplified method proposed in this paper showcases higher accuracy than the methods employed in previous research. Compared with the real model, the relative errors of the overall parameters are within ± 1 %, and the relative errors of the distortion indices for total temperature and total pressure are 5.1 % and −1.5 %, respectively. Meanwhile, the number of computational mesh cells is reduced by more than 95 %. Moreover, the first application of the simplified method in the overall performance calculation of the precooled engine reveals that the traditional low-dimensional model underestimates both the total pressure loss of the precooler and the fuel consumption for precooling, resulting in larger calculation deviations in the thrust and equivalence ratio, respectively. The values of thrust calculated with the high-dimensional virtual model show an average decrease of 2.01 % compared to those derived from the low-dimensional model. In addition, a higher Mach number results in lower effectiveness, higher air inlet temperature, and reduced resistance of the precooler to heat exchange imbalance caused by flow non-uniformity. These factors lead to an escalating relative deviation of the equivalence ratio between the two models, reaching 26.45 % at Mach 5.

Wing deformation improves aerodynamic performance of forward flight of bluebottle flies flying in a flight mill.

Insect wings are flexible structures that exhibit deformations of complex spatiotemporal patterns. Existing studies on wing deformation underscore the indispensable role of wing deformation in enhancing aerodynamic performance. Here, we investigated forward flight in bluebottle flies, flying semi-freely in a magnetic flight mill; we quantified wing surface deformation using high-speed videography and marker-less surface reconstruction and studied the effects on aerodynamic forces, power and efficiency using computational fluid dynamics. The results showed that flies' wings exhibited substantial camber near the wing root and twisted along the wingspan, as they were coupled effects of deflection primarily about the claval flexion line. Such deflection was more substantial for supination during the upstroke when most thrust was produced. Compared with deformed wings, the undeformed wings generated 59-98% of thrust and 54-87% of thrust efficiency (i.e. ratio of thrust and power). Wing twist moved the aerodynamic centre of pressure proximally and posteriorly, likely improving aerodynamic efficiency.

Tuning the Ballistic Performance of a Single-Burning-Rate Grain Solid Rocket Motor via New Discontinuous Embedded Metal Wires

This work proposes a new effective method to realize variable thrust through discontinuous embedded metal wires in the solid rocket motor (SRM). We aimed to study the influence of discontinuous embedded metal wires on the performance of an SRM with a single-burning-rate grain. A model based on convection heat transfer, heat conduction, and heat radiation was established to calculate the heat transfer in the discontinuous embedded metal wires in the grain, to then obtain the burning rate ratio. Most importantly, a solid rocket motor was designed to verify the feasibility of variable thrust and of the present model prediction, with the embedded silver–nickel alloy wire divided into two segments in the grain. According to the SRM ignition experiment, the silver–nickel alloy wires raised the burning rate of the grain. The pressure varied regularly with changes in the discontinuous embedded metal wires. The theoretical burning rate ratio matched the experimental result well. Based on the verified model, the effects of the burning rate, pressure exponent, burning rate ratio, and number of wires on thrust were investigated. Burning rate, burning rate ratio, and pressure exponent were found to be positively correlated with thrust ratio. The thrust ratio could reach 12.5 when the burning rate ratio was 5. The ability to adjust thrust tended to increase with an increase in the number of wires. This study also provided a method to assess whether the consecutive embedded metal wires had been broken or not. The method using discontinuous embedded metal wires in the grain was proven to be feasible to realize multi-thrusts of single-burning-rate grain, which is a new idea for the design of a multi-thrust SRM.

3D relativistic MHD simulations of the gamma-ray binaries

Abstract In gamma-ray binaries neutron star is orbiting a companion that produces a strong stellar wind. We demonstrate that observed properties of ‘stellar wind’–‘pulsar wind’ interaction depend both on the overall wind thrust ratio, as well as more subtle geometrical factors: the relative direction of the pulsar’s spin, the plane of the orbit, the direction of motion, and the instantaneous line of sight. Using fully 3D relativistic magnetohydrodynamical simulations we find that the resulting intrinsic morphologies can be significantly orbital phase-dependent: a given system may change from tailward-open to tailward-closed shapes. As a result, the region of unshocked pulsar wind can change by an order of magnitude over a quarter of the orbit. We calculate radiation maps and synthetic light curves for synchrotron (X-ray) and inverse-Compton emission (GeV-TeV), taking into account $\gamma $ – $\gamma $ absorption. Our modelled light curves are in agreement with the phase-dependent observed light curves of LS5039.

Detailed Gaseous and Particulate Emissions of an Allison 250-C20B Turboshaft Engine

Abstract Aviation is known to be one of the most significant contributors to air pollutants. This includes gaseous emissions, like carbon dioxide (CO2) and nitrogen oxides (NOx), and also particulate matter (PM), especially in the form of soot. This study conducted emission measurements on an Allison 250-C20B turboshaft engine operating on Jet A-1 fuel with a focus on gaseous compounds (e.g., ozone precursors) and PM. The different engine loading points were chosen based on the percentage thrust ratios of the International Civil Aviation Organization LTO-Cycle. A standard FTIR/O2/FID system to measure general gaseous combustion compounds, e.g., CO2, carbon monoxide (CO), unburned hydrocarbons (UHC), and NOx. For the investigation of the volatile organic compounds (VOC), which are known to act as ozone precursors, a gas chromatograph was applied. Different measurement methods were used to characterize the PM emissions. For the particle size distribution (PSD), we used two types of electrical mobility analyzers and an aerodynamic aerosol classifier. All measurement systems yielded comparable PSD results, indicating reliable results. The particle measurement methods all show increasing aerosol diameter modes (electrical and aerodynamic) with increased engine loading. The aerosol diameter modes were shifting from 29 nm to 65 nm. The size and shape of different individual particles were evaluated with a scanning electron microscope. A correlation between the injection system and the particle formation was established. Gaseous turboshaft engine emissions show high CO and UHC values at Ground Idle level. NOx levels were the highest at Take-Off conditions. Acetylene and ethylene were the most significant contributors to ozone formation.

Analytical Model Development for Rotors Hovering Above Heaving Surfaces

A potential flow model was developed for rotors hovering in-ground-effect above a heaving surface using a similar methodology to classical, analytical, static ground effect models. Experimental performance measurements for rotors hovering above a surface undergoing single degree-of-freedom heaving motion were used for model validation, and potential mission trends were generated for a representative naval helicopter. Unlike prior empirical models, the current model was able to capture the effect of rotor hub height as well as ground motion parameters. A new thrust ratio was proposed that compared the thrust produced in-ground-effect above a heaving surface to that produced above a static surface. Better agreement was found to occur at the higher collective pitch settings and hub heights, with the model predicting the classical thrust ratio within 5% of the measured value for approximately 75% of the test points. The thrust ratio relative to a static surface was found to produce better agreement, and approximately 80% of the experimental tests points were predicted within 5% of the measured value.

Fluidic Thrust Vector Control of Aerospace Vehicles: State-of-the-Art Review and Future Prospects

Abstract An efficient propulsion system holds the key to the smooth operation of any aerospace vehicle over different flight regimes. Apart from generating the necessary thrust, emphasis has also been laid on vectoring the direction of thrust. The primitive modes of thrust vectoring chiefly focused on mechanical means such as the use of gimbals or hinges. The current state-of-the-art technologies demand more efficient methods for thrust vectoring, which minimize the use of mechanical components. These methods termed fluidic thrust vector control methods, employ secondary jets for achieving the required attitude, and trajectory of the aerospace vehicles such as aircraft, rockets, and missiles. Such methods have greatly helped in reducing vehicle weight, vehicle maintenance requirements, and enhancement of stealth characteristics of such vehicles. This work presents a review of the various fluidic thrust vectoring systems, starting with a brief overview of traditional thrust vectoring systems, followed by a discussion on the various aspects of fluidic thrust vectoring systems. It also highlights the effect of the various geometrical and operating conditions on the performance parameters of the thrust vectoring system such as the thrust vector angle, system thrust ratio, and thrust vectoring efficiency among others. For ensuring the comprehensive character of this work, synthetic jet vectoring techniques have also been included due to their nonmechanical nature and similarities with purely fluidic thrust vectoring techniques.

Investigation on dynamic performance of wind turbines using different scaling methods in wind tunnel tests

The relationship between rotor thrust (T) and tip speed ratio (TSR) directly influences the response of the scaled offshore wind turbine in the wind tunnel test. In this paper, a series of 1:100 scaled wind tunnel tests based on NREL 5 MW wind turbine are designed to study the dynamic performance of the wind turbine based on two different T-TSR scaling methods: (1) geometrically scaled blade with increased wind speed (GB-IW) and (2) redesigned blade with scaled wind speed (RB-SW). Test result indicates that the GB-IW method can meet the requirement of T, but unavoidably result in the distortion of TSR (40.2 % to 61.7 % lower than the target). However, using the RB-SW method, the T-TSR curve can be simulated well under different operational conditions. The amplitudes of structural response including base moment, nacelle acceleration and displacement from the GB-IW method are much lower than those from the RB-SW method, indicating that the GB-IW method leads to an overestimation of damping ratio but underestimation of the fluctuation of both natural frequency and structural response. The RB-SW method can provide a more reasonable estimation of structural dynamic performance, compared with the GB-IW method. This study confirms the importance of selecting reasonable scaling method for wind tunnel test, which is critical to the understanding of the complex aerodynamic and structural dynamic performance of offshore wind turbines.

Characterization of pintle engine performance for GO2/kerosene propellants

Liquid rocket engine with a wide range of thrust variations has drawn much attention in recent years. A GO2/kerosene deep-throttling variable thrust engine using a liquid–gas pintle injector is presented and tested with aim of characterizing engine performance, regulating capability, and heat flux distribution. Three-dimensional numerical simulations are conducted to investigate the flame appearance and spray characteristics of the engine. The SST k-omega model is used for modeling turbulence, the Wave model is used for modeling the secondary breakup, and a multi-step quasi-global combustion mechanism and eddy dissipation concept model are used for modeling combustion. The GO2/kerosene pintle engine achieves stable combustion at 0.18 ∼ 4.35 MPa with a thrust of 21.30 ∼ 864.70 N. The thrust ratio of the pintle engine is 40.6:1 indicating a deep-throttling variable thrust capability. The thrust regulation of the pintle engine using the closed-loop control has high accuracy and response speed, as well as high disturbance suppression capability. The effects of combustor pressure, spray characteristics, and reactants concentration distribution on flame appearance and heat flux distribution of the engine are studied in detail. The engine equipped with a liquid–gas pintle injector could ensure a good atomization effect and high combustion efficiency by actively adjusting the propellant injection pressure drop coefficient. This work will provide a reference for the design of the pintle injector and thermal protection system of deep-throttling variable thrust rocket engines.

Investigation of dynamic mixing combustion characteristics in variable thrust hybrid rocket motors

The dynamic mixing combustion characteristics in variable thrust hybrid rocket motors (HRMs) have been investigated by using a newly developed two-dimensional transient numerical model. This model takes into account the heat transfer process in the solid fuel grain coupled with dynamic combustion flow in HRMs. Fire experiments were conducted to validate the developed model by using an optical-access slab burner and a lab-scaled HRM. It shows that the model developed in the present work can predict accurately the regression behavior of fuel grains in HRMs. Based on this model, simulations of two variable thrust processes with different thrust ratio were performed. The results show that the recirculation zones located at pre- and post-combustion chamber directly affects the inhomogeneity of the temperature distribution of the inner surface of the fuel grain, leading to a fuel regression rate difference along axial direction. In thrust variation process, the fluctuation duration of the fuel regression rate is longer than the regulation time of oxidizer flow. The fluctuation range of regression rate is influenced by the amplitude and proportion in oxidizer flow regulation. In addition, a decrease in the characteristic velocity and combustion efficiency (from 97.1% to 72.63% and 70.64%, respectively) was observed during the thrust variation process. A slight thrust loss was observed at all stages of the variable thrust process and reaches to maximum value at an oxidizer flow rate of 20 g/s for the lab-scaled HRM studied in the present work.

Dynamic Characteristics of LOX/Kerosene Variable Thrust Liquid Rocket Engine Test System Based on General Modular Simulation Method

For the research demand of reusable LOX/kerosene variable thrust liquid rocket engine, a test system with electric displacement pumps is designed and a multidisciplinary modular dynamic simulation method based on AMESim platform is used to analyze the system. The method comprehensively considers the characteristics of complex components in the engine and realizes the fast module assembly and variable step size solution. Considering the combustion model of thrust chamber, the positive displacement pump model with complex leakage channels, and the cooling jacket heat transfer model, the component dynamic equations are deduced and the final model simulation results reveal that the system has a smooth ignition, stage turning, and shutdown process. The thrust can reach 6900 N in high working condition and the variable thrust ratio is 5 : 1. The dynamic characteristics of the system show that the performance error of main pump components is less than 5%, the maximum average temperature rise of the thrust chamber coolant is about 28°C, and the time of stage adjustment is within 300 ms, which mean the overall design of the system is reasonable. Although the accumulation of LOX before kerosene injection can adversely affect the temperature of the thrust chamber, large pressure pulses do not occur due to the ignitor’s duty flame. Moreover, the pintle injectors based on PID control can effectively stabilize the pressure drop at lower conditions. The system and the simulation method provide important support for the actual engine test and the general LRE dynamic characteristics analysis.

지면효과 속에 놓인 이중로터의 공력 특성에 대한 수치적 연구

Multi-copters are used in various fields as unmanned aerial vehicles. Consequently, the stability of the operation conditions of the unmanned aerial vehicle and the improvement of aerodynamic characteristics are required. Therefore, this study is intended to examine the effect of the wake interaction between rotors in ground effect when a set of double rotors is positioned adjacent to the ground through an unsteady numerical analysis. For this reason, the rotor is positioned adjacent to the ground, and the thrust ratio is measured while changing the distance between the rotors which are rotating at a fixed rotational speed. With the ground height as z and the rotor radius as R, the smaller the height ratio(z/R), the higher the thrust. As the distance between the rotors became smaller, the collision and interaction between wing tip vortices of the rotor blades got stronger, and the flow was not properly propagated into the direction of the wake. It is shown that when the rotor gets closer to the ground, the flow could not be fed back in after flowing along the ground, so the thrust ratio due to the ground effect tends to decrease. With these results and analysis, the effect of wake interference between rotors on the ground effect could be confirmed.

Computational Investigation of Effects of Side-Injection Geometry on Thrust-Vectoring Performance in a Fuel-Injected Dual Throat Nozzle

Dual-throat Nozzle (DTN) is known as one of the most effective approaches of fluidic thrust-vectoring. It is gradually flourishing into a promising technology to implement supersonic and hypersonic thrust-vector control in aircrafts. The main objective of the present study is numerical investigation of the effects of secondary injection geometry on the performance of a fuel-injected planar dual throat thrust-vectoring nozzle. The main contributions of the study is to consider slot and circular geometries as injector cross-sections for injecting four different fuels; moreover, the impact of center-to-center distance of injection holes for circular injector is examined. Three-dimensional compressible reacting simulations have been conducted in order to resolve the flowfield in a dual throat nozzle with pressure ratio of 4.0. Favre-averaged momentum, energy and species equations are solved along with the standard k- ε model for the turbulence closure, and the eddy dissipation model (EDM) for the combustion modelling. Second-order upwind numerical scheme is employed to discretize and solve governing equations. Different assessment parameters such as discharge coefficient, thrust ratio, thrust-vector angle and thrust-vectoring efficiency are invoked to analyze the nozzle performance. Computationally predicted data are agreed well with experimental measurements of previous studies. Results reveal that a maximum vector angle of 17.1 degrees is achieved via slot injection of methane fuel at a secondary injection rate equal to 9% of primary flow rate. Slot injection is performing better in terms of discharge coefficient, thrust-vector angle and thrust-vectoring efficiency, whereas circular injection provides higher thrust ratio. At 2% secondary injection for methane fuel, vector angle and vectoring efficiency obtained by slot injector is 8% and 34% higher than the circular injector, respectively. Findings suggest that light fuels offer higher thrust ratio, vector angle and vectoring efficiency, while heavy fuels have better discharge coefficient. Increasing center-to-center distance of injector holes improves thrust ratio, while having a negative effect on discharge coefficient, vector angle and vectoring efficiency. A comparison between fuel injectant of current study and inert injectant in the previous studies indicates that fuel reaction could exhibit substantial positive effects on vectoring performance. Secondary-to-primary momentum flux ratio is found to play a crucial role in nozzle performance.

Study on the effect of mist flux on water mist fire extinguishing

A series of combustion experiments were carried out by varying the pressure of water mist and the power of the fire source in order to quantitatively study the fire-extinguishing conditions of water mist. The main characteristics of an oil pool fire under the action of water mist are revealed through the analysis of experimental parameters, and a fire extinguishing model is established. The results showed that the thick flammable vapor layer on the oil surface is the source of water mist-induced flame expansion. Besides that, the splashing of oil droplets is one of the reasons for the flame's enhancement. After simplifying the combustion model, the heat and mass transfer approach was utilized to quantitatively study the water mist flux model for fire suppression. When the thrust ratio of fire versus spray (α) exceeds 5, the plume thrust has a significant impact on the water mist droplet trajectory, and the fire extinguishing fails. In other words, the water mist's cooling action is the primary mechanism for putting out the pool fire. The failure to extinguish the fire was mostly due to the mist flux being insufficient to absorb enough heat from the fuel surface to break the combustion equilibrium conditions.

Low-Order Aerodynamic Model for Interference in Multirotor Systems

The paper proposes an aerodynamic model capable of capturing interference effects in arbitrary multirotor systems. The current analysis is limited for nonoverlapping rotors. The model is based on a combined inflow theory approach, where self-induced and influence-induced velocities are obtained using different inflow models. The generalized dynamic wake model was used to obtain the self-induced velocity distribution, whereas the influence-induced velocity field was computed using Beddoes’s generalized wake model. The results of the combined inflow model were compared against free-vortex wake simulations of a twin-rotor system in edgewise flight. Furthermore, the study examined the effects of tip-to-tip separation distances and freestream orientation angles on overall system performance in terms of thrust and power ratios. The comparison indicated that the results of the proposed model agreed within around 5% with the free-vortex wake analysis over all operating conditions. The computational time of the combined inflow model was two to three orders of magnitude lower than the free-vortex wake approach, depending on the code architecture and domain discretization. The analysis shows that the proposed approach is adequate to model primary aerodynamic interference effects in multirotor systems operating at advance ratios greater than or equal to 0.15.

Analysis of the Efficiency of the Electrohydrodynamic Propulsion Based on the Biefeld-Brown Effect for Manned and Unmanned Aircrafts

The phenomenon known as the Biefeld-Brown effect or ionic wind has been the subject of study since its discovery with the aim of being applied to aircraft propulsion. Electrohydrodynamic (EHD) thrusters have been developed in a thorough and detailed way since 2000. Subsequently, some attempts have been made to extend them to electric-powered atmospheric aircrafts. This paper is framed in this field by performing a rigorous and detailed analysis of the energy efficiency, measured as the ratio of thrust to energy consumed, for both EHD systems and conventional electric propulsion. This is done using simple EHD thruster configurations together with accurate measurements of the current consumed under controlled environmental conditions. The results are not only compared with the performance of an electric propulsion motor-propeller set, but also consider future research, summarizing the issues that arise to scale them. Future designs of low weight, high performance thrusters might allow this technology to be applied to manned or remotely piloted aircrafts.

Experimental study on the interaction of water mist spray with two buoyant non-premixed flames

Single-fire and double-fire extinguishments by water mist were conducted to assess the similarities and differences. The double-fire extinguishment requires higher applied volume flux of the water mist and spray momentum, while it is achieved faster than the single-fire extinguishment, owing to the smaller size of the single flames of the double-fire. Both the air entrainment restriction among the two flames of the double-fire and the typical re-ignition phenomenon have a significant effect on the double-fire extinguishment. A double-fire merging from the burner surface is more difficult to be extinguished in terms of the critical operating pressure since the air entrainment restriction among the two flames of the double-fire is more intense. A double-fire with a smaller flame spacing requires a longer extinguishing time since re-ignition happens more easily. The equivalent diameter (De) is applied to calculate the plume–spray thrust ratio of the double-fire and evaluate the competition between the double-fire and the water mist quantitatively.

Phase Relationships in the Carbon–Titanium–Uranium System for Ultra-High Temperature Nuclear Fuels

Carbides of uranium have attracted interest as fuels for nuclear thermal propulsion (NTP) to drive deep-space exploration owing to their attractive thermal and neutronic properties. Optimization of NTP technology, however, requires ultra-high temperature reactor environments to maximize the ratio of thrust to propellant to achieve peak rocket engine efficiency. Incorporation of transition metals into uranium carbides offers a pathway to increase the melting point of carbide fuels to address the operational challenges posed by NTP. A thermodynamic model has been developed to examine the phase relationships in the C–Ti–U system at NTP conditions. Calculated phase relationships at ultra-high temperatures predict a stable (U, Ti)C solid solution. Additionally, experimental work on C–Ti–U synthesis via arc melting has been carried out, providing data on phase constitution and compositions that can be used to further refine the thermodynamic model that has been developed.