Thrust Augmentation Research Articles

A reliable combustion strategy for a throttling-assisted supersonic combustor with flight Mach 7

As the flight Mach number increases, scramjet engines struggle to achieve stabilized combustion. In the present investigation, a reliable combustion strategy is developed for a kerosene-fueled supersonic combustor. Assisted by air throttling, this enables the combustor to operate efficiently at a flight Mach number of 7. Various experimental measurement techniques are used to capture the combustion process and flow characteristics. A three-dimensional numerical method based on the Reynolds-averaged Navier–Stokes equations is employed to analyse the effects of air throttling on the non-reacting and reacting flow fields. The whole combustion test is effectively divided into five processes covering the establishment of non-reacting flow, ignition, flame stabilization, and flameout. Analysis of the experimental and numerical flow fields indicates that, as the origin of the initial flame in the combustor, the aerodynamic compression induced by air throttling promotes the kerosene/air sub-mixing region. The main mixing process is enhanced by the arrangement of ramps and cavity, which contribute to effective multi-channel injection. Two combustion modes are observed, namely combined cavity shear-layer/recirculation stabilized combustion and jet-wake stabilized combustion. In the reacting flow field, the additional injection of throttling gas improves the thrust augmentation at the cost of reduced combustion intensity. The outlet combustion efficiency and total pressure recovery coefficient are found to decrease by 8.49% and 28.79%, respectively.

Minimum Condition Requirements in Laser Driven ICF for Advanced Space Propulsion Applications

Inertial Confinement Fusion (ICF) is one of the candidate methods by which spacecraft can be propelled onto interplanetary and interstellar trajectories. In this paper we discuss the application of laser driven ICF to advanced space propulsion and present fundamental equations which relate the capsule mass and nozzle jet efficiency to propulsion parameters. This is also shown for the scenario of using expellant propellant for the purpose of thrust augmentation. Keywords: Inertial Confinement Fusion, Space Propulsion

A constant pressure differential valve and its controlling characteristics for aero engine

No matter in traditional hydraulic mechanical or newly numerical controlling system, pressure differential valve (PDV) always plays an important role in main fuel oil system and thrust augmentation fuel system. A throughout study on PDV is required to gain a more precise control and effective design for aero engine. Therefore, this paper presents a PDV and establishes the mathematical model to analyze its working characteristics, which shows that PDV performance is greatly affected by return oil inlet orifice and spring stiffness. The simulation and experiment are carried out to analyze the effect of PDV performance on system outlet volume flow rate and their results agree well with a maximum error of approximately 5 %. According to the study, it is found that there is a maximum increase of 9.99 % in system volume flow rate with pump rotating velocity from 300 to 900 rpm, but a maximum decrease of 3.4 % with outlet load pressure from 0.5 to 1.6 MPa. In addition, calculated curve slope reduces with return oil inlet orifice number increase and spring stiffness decrease, which means that PDV shows a faster response to working condition change and a better dynamic characteristic in controlling process. It is excepted that the research can be helpful in a further understanding of aero engine controlling system and a meaningful instruction on system optimization as well as fuel efficiency improvement.

Drag reduction characteristics of recirculation flow at rocket base in an RBCC engine under ramjet/scramjet mode

To reduce the drag generated by the recirculation flow at the rocket base in a Rocket-Based Combined Cycle (RBCC) engine operating in the ramjet/scramjet mode, a novel annular rocket RBCC engine based on a central plug cone was proposed. The performance loss mechanism caused by the recirculation flow at the rocket base and the influence of the plug cone configuration on the thrust performance were studied. Results indicated that the recirculation flow at the rocket base extended through the entire combustor, which creates an extensive range of the “low-kinetic-energy zone” at the center and leads to an engine thrust loss. The plug cone serving as a surface structure had a restrictive effect on the internal flow of the engine, making it smoothly transit at the position of the large separation zone. The model RBCC engine could achieve a maximum thrust augmentation of 37.6% with a long plug cone that was twice diameter of the inner isolator. However, a shorter plug cone that was half diameter of the inner isolator proved less effective at reducing the recirculation flow for a supersonic flow and induced an undesirable flow fraction that diminished the thrust performance. Furthermore, the effectiveness of the plug cone increased with the flight Mach number, indicating that it could further broaden the operating speed range of the scramjet mode.

Trade Study of Engine Installation Position for a Rear-mount Aircraft

The meticulous analysis of large bypass ratio engine installation constitutes a vital foundation for the integrated design of civil aircraft engines. To holistically evaluate the performance implications between the engine and the airframe, a trade design concerning the engine installation position for a rear-mounted regional aircraft is conducted by utilizing computational fluid dynamics, incorporating a powered nacelle to account for the engine’s intake and exhaust ramifications. The engine’s longitudinal, spanwise, and vertical positioning impacts on aircraft drag and installed net thrust were discerned. The nacelle’s longitudinal displacement exerts the most substantial influence on the aggregate aircraft gain. The optimal configuration, achieved by advancing the engine 500 millimeters forward, yields a drag advantage of 2.24 counts and a thrust augmentation of 0.49 counts, culminating in a combined gain of 2.73 counts.

Thrust performance of the rocket-based combined-cycle engine under ejector mode

The Rocket-Based Combined-Cycle (RBCC) engine integrates the preponderance of ramjet and rocket engines. It can perform excellently at a lower Mach number than a ramjet while consuming less fuel than the rocket. The higher specific impulse under lower flight Mach conditions guarantees the competitiveness and application prospect of an RBCC engine for reusable space transportation and hypersonic cruise vehicles. With a wide range of Mach numbers, the flow choking between primary and secondary streams in the inner flow passage of the engine becomes complicated. The flow choking not only affects the mass flow ratio between the air stream and rocket plume but also determines the thrust performance of the engine. However, the relationship between flow-choking states and thrust performance has not been revealed. This investigation aimed to provide a better design reference for the RBCC trajectory design and a basis for the RBCC engine geometry design so that the thrust performance under different flow choking states was studied. The findings indicate that the engine is more favorable in the supersonic regime status than others. At the premise of the lower total pressure air stream condition, the thrust of the RBCC engine is attributed to the rocket pressurization effect. On the contrary, its thrust is ascribed to the air stream when the air stream total pressure is higher. Besides, the thrust augmentation of the RBCC engine is essentially due to the larger entrainment ratio. There is a rocket mass flow rate range in different mixer diameters, named region B in this paper, in which the air mass flow rate can be boosted with the rocket mass flow rate and the specific impulse ratio is higher than the rocket specific impulse. Significantly, the RBCC design state should be maintained in region B to meet the operational requirements of engines under the precondition of an insufficient flight Mach number.

Design of thrust augmentation control schedule during mode transition for turbo-ramjet engine

A thrust augmentation control schedule was designed for turbo-ramjet engine during mode transition process to achieve high thrust performance. First, a mathematical model of turbo-ramjet engine was established, analyzing the reason for the existence of an insufficient thrust working range under conventional mode transition control schedule. On this basis, a thrust augmentation control schedule was devised for the thrust insufficient range considering the matching between the air inlet available airflow and the engine demand airflow. The control schedule could increase the flow capacity of the engine and reduce the overflow drag of air inlet by opening the ramjet bypass while keeping the turbofan components working at the maximum state, which could improve the maximum installation thrust performance of the engine and achieve a better air inlet/engine matching. By further designing the optimal regulation rule for the rear variable area bypass injector (RVABI), the installed thrust of the engine could be increased by more than 37.93% at most, while simultaneously reducing the fuel consumption by 0.16%. Finally, the turbo-ramjet engine was simulated along the isobaric flight path. The simulation results revealed that under conventional mode transition control schedule, the thrust insufficient occurred when Mach number is above 2.7. The designed thrust augmentation control schedule could effectively improve the maximum thrust of the thrust drop zone and provide sufficient thrust margin for engine during the whole working process. At the same time, considering the smooth thrust and thrust augmentation requirements, a compromise mode transition control schedule between the thrust augmentation control schedule and the conventional mode transition schedule is proposed. This control schedule could address the problem of insufficient thrust at high Mach numbers while the smoothness transition of engine total thrust could be guaranteed.

Effect of aspect ratio on the propulsive performance of tandem flapping foils

In this work, we describe the impact of aspect ratio ( $AR$ ) on the performance of optimally phased, identical flapping flippers in a tandem configuration. Three-dimensional simulations are performed for seven sets of single and tandem finite foils at a moderate Reynolds number, with thrust producing, heave-to-pitch coupled kinematics. Increasing slenderness (or $AR$ ) is found to improve thrust coefficients and thrust augmentation but the benefits level off towards higher values of $AR$ . However, the propulsive efficiency shows no significant change with increasing $AR$ , while the hind foil outperforms the single by a small margin. Further analysis of the spanwise development and propagation of vortical structures allows us to gain some insights into the mechanisms of these wake interactions and provide valuable information for the design of novel biomimetic propulsion systems.

Thrust Augmentation and its Control Methods: Short Overview

Augmentation plays a crucial role in many aircrafts and jet engines to elevate the thrust. Thrust is oe of the four fundamental forces of an aircraft. Augmentation methods are used in certain cases to increase the thrust to perform better landing and take-off. There is various augmentation methods in which afterburner is mostly used over other augmentation methods because of its advantages. In this paper an overview has been provided for thrust augmentation and its different control methods.

Preliminary Study on Mixing Mechanism of Three-Dimensional Vortices in the Wake of Lobe-Type Rear Variable Area Bypass Injector

In this work, an innovative lobe-type rear variable area bypass injector is proposed, which can take into account both mixing exhaust performance to increase thrust at a high Bypass Ratio (BR), and bypass throttling at low BR. In order to explore the aerodynamic characteristics of the variable injector, the unsteady motion characteristics and the congenerous thermal transport effect of three-dimensional wake vortices within the BR range of 0.18 to 0.68 are numerically analyzed. The results demonstrate that the injector can adjust the sidewall of the collapsible lobe based on the injection of bypass flow, and significantly strengthen streamwise vortices induced by lobe peak (SV-LP), which can effectively drive the migration of the heterogeneous fluid element, so as to compensate for the inefficient mixing ability of Kelvin-Helmholtz (K-H) vortices. When the BR is raised to 0.36, the increasing SV-LP will gradually replace the low-disturbance K-H vortices in the dominant position of the quasi-ordered flow field by virtue of a strong entrainment effect. Therefore, the injector can maintain a relatively high thrust augmentation in a wide working envelope.

A Comparative Performance Analysis of the Novel TurboAux Engine with a Turbojet Engine, and a Low-Bypass Ratio Turbofan Engine with an Afterburner

Presented herein is a comparative performance analysis of a novel turbofan engine with an auxiliary combustion chamber, nicknamed the TurboAux engine, against a turbojet engine, and a low bypass ratio turbofan engine with an afterburner is presented. The TurboAux engine is an adaption of the low-bypass ratio turbofan engine, but with secondary combustion in an auxiliary bypass annular combustion chamber for thrust augmentation. The TurboAux engine is envisioned with the desire to facilitate clean secondary burning of fuel at temperatures higher than in the main combustion chamber with air exiting the low-pressure compressor. The comparative study starts by analyzing the turbojet engine and its performance with and without an afterburner segment attached. In parallel, the conventional turbofan and its mixing counterpart are analyzed, also with and without an afterburner segment. A simple optimization analysis of a conventional turbofan is performed to identify optimal ‘fan’ pressure ratios for a series of low-bypass ratios (0.1 to 1.5). The optimal fan pressure ratios and their corresponding bypass ratios are adapted to demonstrate the comparative performance of the varying configurations of the TurboAux engine. The formulation and results are an attempt to make a case for charter aircrafts and efficient close-air-support aircrafts. The results yielded increased performance in thrust augmentation, but at the cost of a spike in fuel consumption. This trade-off requires more in-depth investigation to further ascertain the TurboAux’s utility.

УВЕЛИЧЕНИЕ ТЯГИ ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ ПУТЕМ МОДЕРНИЗАЦИИ СОПЛОВОГО БЛОКА

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.

Oxidation Resistant Nickel Aluminide Coating on Niobium Alloy C103 for High Temperature Aerospace Applications

Owing to its fairly higher melting point (2468°C) and superior thermal conductivity, Niobium alloys are rendered as candidate material for high temperature applications, where state of the art Nickel base super alloys cannot be used. Its applications possess wide range of aerospace components mainly encompassing aircraft exhaust chambers, thrust augmenters, rocket engines and other industries such as chemical and petrochemical. The only limitation with niobium alloys is their inferior oxidation resistance, which can be tackled by developing a high temperature oxidation resistant coating. In this regard, Nickel Aluminide (NiAl) having a melting point of 1638°C embraces a paramount importance amongst the coating categories. In this study, NiAl was synthesized using vacuum melting and strip casting followed by ball milling. The powder was coated on C-103 Niobium alloy specimens using air plasma technique and evaluated for high temperature stability. It was deduced that the coating was exceptionally stable up to 1234 °C in ambient conditions.

Aerodynamic performance of hovering micro revolving wings in ground and ceiling effects at low Reynolds number

Numerous investigations have been conducted to understand the wall effects on rotors. The purpose of this study is to further investigate the aerodynamic performance of revolving wings, especially when it is very close to the ground and ceiling (i.e., less than half the wingspan) at low Reynolds numbers. Hence, the ground and ceiling effect for hovering micro revolving wings at low Reynolds numbers are investigated by improving the theoretical models. The theoretical model for the ground effect is established based on the wall-jet assumption, and that for the ceiling effect is improved by considering the uneven spanwise distribution of induced velocity. These two models are validated by comparing the results of experiments and CFD simulations with the Lattice-Boltzmann Method (LBM). Both ground and ceiling effects are found helpful to enhance the thrust, especially with small wing-wall distances, by making a difference to the induced velocity and the pressure distribution. By comparing the thrust generation and aerodynamic efficiency between the ground and ceiling effects, the former is found more helpful to the thrust augmentation, and the latter is more beneficial for the aerodynamic efficiency promotion.

Investigations of entrainment characteristics and shear-layer vortices evolution in an axisymmetric rear variable area bypass injector

In this study, firstly, for the axisymmetric RVABI, the change-rule of adverse pressure gradient caused by radial velocity during the transition of internal flow mode in variable geometry is summarized, and a Bypass Ratio (BR) iterative algorithm based on the empirical correlation of non-equilibrium pressure is proposed. The algorithm can estimate the nonlinear relationship between area ratio and BR, with an error range falling below 6.5%. Then, we discuss the favorable effect of uniform mixing on the thrust augmentation of mixed exhaust under variable BR conditions. From this point of view, the characteristics of vortices evolution in different shear strength jets are compared, to clarify the effect of variable cycle parameters on jet mixing. As the results suggest, when Λ is as low as 0.22, the K-H disturbance is of high-frequency wavelet property, and it is difficult to induce large-scale spanwise vortices. The macro migrations of fluid elements in spanwise vortices and the diffusion effect caused by edge tearing is weak, which is not conducive to the energy exchange between the two streams. However, the low Λ jet will also correspondingly weaken the viscous dissipation effect of vortices. It is concluded that the dissipation level is proportional to the 2.31 power of the Λ.

Turbojet Thrust Augmentation through a Variable Exhaust Nozzle with Active Disturbance Rejection Control

Turbojets require variable exhaust nozzles to fit high-demanding applications; however, few reports on nozzle control are available. The purpose of this paper is to investigate the possible advantages of an exhaust gas control through a variable exhaust nozzle. The control design method combines successful linear active disturbance rejection control (LADRC) capabilities with a loop shaping controller (LSC) to: (i) allow designing the closed-loop characteristics in terms of gain margin, phase margin and bandwidth, and (ii) increase the LSC disturbance rejection capabilities with an extended state observer. A representation of the nozzle dynamics is obtained from first principles and adapted to achieve a stream-velocity-based control loop. The results show that the resulting controller allows improving the expansion of the exhaust gas to the ambient pressure for the whole operating range of the turbojet, increasing the estimated thrust by 14.23% during the tests with experimental data.

Thrust Augmentation of Micro-Resistojets by Steady Micro-Jet Blowing into Planar Micro-Nozzle

The present work investigates the impact of steady micro-jet blowing on the performance of a planar micro-nozzle designed for both liquid micro-thrusters and nitrogen cold-gas micro-resistojets. Two micro-injectors have been placed into the divergent region along the sidewalls, injecting a secondary flow of propellant perpendicularly to the wall where they have been located. The micro-jet actuator configuration is characterized by the dimensionless momentum coefficient cμ. The best performance improvement is retrieved at the maximum cμ for both water vapor (Δ%T,jet = +22.6% and Δ%Isp,Tjet = +2.9% at cμ = 0.168) and nitrogen gaseous flows (Δ%T,jet = +36.1% and Δ%Isp,Tjet = +9.1% at cμ = 0.297). The fields of the Mach number and the Schlieren computations, in combination with the streamline visualization, reveal the formation of two vortical structures in the proximity of secondary jets, which energize the core flow and enhance the expansion process downstream secondary jets. The compressible momentum thickness along the width-wise direction θxy in presence of secondary injection reduces as a function of cμ. In particular, it becomes smaller than the one computed for the baseline configuration at cμ > 0.1, decreasing up to about and -57% for the water vapor flow at cμ = 0.168, and -64% for the nitrogen gaseous flow at cμ = 0.297.

Thrust Augmentation of Single Cycle Pulse Detonation Engine using Nozzles

Pulse detonation propulsion systems have the potential to provide better performance with additional advantages such as considerably light in weight, cost effective and reduced complexity in comparison with other propulsion systems which are currently in use. These improvements are due to the high thermodynamic efficiency obtained because of constant-volume combustion. Pulse detonation cycle can be used for both air-breathing and rocket based systems. Present study investigates the effect of nozzles in various configurations. They are straight nozzle, conical and bell-shaped nozzles with varying length, divergent angles and area ratios on the thrust augmentation of Pulse Detonation Engine (PDE) test rig which was developed by research team at Punjab Engg. College (PEC), Chandigarh. It was found from the experiments that the conical nozzle with high divergent angle of 20° and high nozzle area ratio of around 23 increased the thrust to 14%. The bell shaped nozzle, with 20° angle of divergence and a nozzle area ratio of just around 7, produced 59.5% more thrust in comparison with baseline engine. The augmentation in thrust was found to be as high as 55% in comparison with straight nozzle. Divergent nozzles produced negative thrust with less divergent angle but gave an increment of 11.28% with high angle of divergence in comparison with a straight nozzle.

Computational Investigation of Multiple Oscillating Hydrofoils and its Effect on Thrust Augmentation

Previous investigations of flapping hydrofoils for the purpose of thrust production have been limited to one or two in tandem. Tandem foils were found to have superior performance because the performance of the aft foil was augmented by the vortices shed from the fore foil. It is however not clear if increasing the number of foils will continue to have increased performance or if there exist an optimal number after which the overall performance either stagnates or reduces. A 2D computational study was conducted to investigate the effect of increasing the number of hydrofoils to four at a Reynolds number of 8000 flapping in-phase and out-of-phase. Optimal and sub-optimal conditions found previously with tandem hydrofoils were found also be applicable to three and four hydrofoils.

Experimental Study on the Effect of Water Injection on a Micro Turbojet Engine

Due to increasing emissions in the 21st century, aviation industry faces a rising pressure to reduce its contribution to air pollution. A possible way to mitigate such harmful effects on the environment is to use water injection as a thrust augmentation method which was developed in the 1960s. Since the emergence of modern engines, the need for the performance-enhancing effect of water injection has become abandoned. However, considering its beneficial impact on both emissions and engine structure, it can still be used as an alternative nowadays as well. As a result of the cooling effect of water injection, the service life of engine components extends which could also lead to a reduction in maintenance and operating costs. This paper presents how water injection on TKT-1 experimental jet engine affects performance and emissions. The modifications implemented on the test bench enabled us to obtain a measure of the water injection operation with the nozzle placed in the intake duct. Results reveal that such method has no visible impact on the increase of thrust due to the experimental environmental conditions and the limitations of the system. Nevertheless, water injection proved to be effective in reducing emissions. In addition to the beneficial effect of lower gas temperature throughout the turbine section of the engine, the emissions of NOx gases also fall significantly (i.e. 30%) compared to dry operation.