Thrust Augmentation Research Articles

Assessment of Mechanical Design Parameters for an Aero Gas Turbine Engine Jet Pipe Casing using Finite Element Analysis

Aero Gas Turbine engines power aircrafts for civil transport application as well as for military fighter jets. Jet pipe casing assembly is one of the critical components of such an Aero Gas Turbine engine. The objective of the casing is to carry out the required aerodynamic performance with a simultaneous structural performance. The Jet pipe casing assembly located in the rear end of the engine would, in case of fighter jet, consist of an After Burner also called as reheater which is used for thrust augmentation to meet the critical additional thrust requirement as demanded by the combat environment in the war field. The combustion volume for the After burner operation together with the aerodynamic conditions in terms of pressure, temperature and optimum air velocity is provided by the Jet pipe casing. While meeting the aerodynamic requirements, the casing is also expected to meet the structural requirements. The casing carries a Convergent-Divergent Nozzle in the downstream side (at the rear end) and in the upstream side the casing is attached with a rear mount ring which is an interface between engine and the airframe. The mechanical design parameters involving Strength reserve factors, Fatigue Life, Natural Frequencies along with buckling strength margins are assessed while the Jet pipe casing delivers the aerodynamic outputs during the engine operation. A three dimensional non linear Finite Element analysis of the Jet pipe casing assembly is carried out, considering the up & down stream aerodynamics together with the mechanical boundary conditions in order to assess the Mechanical design parameters

Analytical Model for Ring-Wing Propulsor Thrust Augmentation

A simple, accurate, reduced-order model is developed for predicting the maximum available thrust performance of subsonic ring-wing propulsors varying in shroud airfoil length, thickness, camber, and/or angle of attack relative to the centerline. The method, applicable to drones, aircraft, ships, and more, identifies the potential upper limit to the thrust augmentation available from hover to cruise, as well as those dominant geometric features contributing to its achievement and those factors limiting its attainment. It employs classic thin airfoil aerodynamic principles and is an extension and enhancement of two previous reduced-order models. Sensitivity studies are conducted of impacts on thrust augmentation from shroud-induced viscous effects, radial loading variations induced by the propeller, and losses due to propeller-induced swirl. The model’s efficacy is assessed using over 40 independently published experimental and state-of-the-art inviscid and viscous computational data sets for more than 20 different configurations. It is shown that the model’s algebraic solutions provide very good engineering estimates at all operating conditions from hover to cruise, thereby enabling rapid preliminary design of ring-wing-based systems. Additionally, the new model provides a means to establish maps of the practical thrust augmentation limits for ring wings, providing rational starting points for formal optimization studies.

Influence of water injection on performance of scramjet engine

In order to investigate the influence of water injection on performance of scramjet engine, a thermodynamic model with the constant cross sectional area heating process was developed to analyze the effects of water content and energy on performance of scramjet engine. The influence of water injection on the maximum fuel equivalence ratio of scramjet engine was also studied. The maximum specific thrust was obtained by optimizing the compression ratio of the inlet, the fuel equivalence ratio and the water content. Results show that the specific thrust of scramjet engine can be augmented by water injection at the cost of the decrease of fuel impulse. The maximum fuel equivalence ratio of scramjet engine can be increased by water injection at high flight Mach numbers, while slightly decreased at low flight Mach numbers. So thrust augmentation by water injection is more suitable for scramjet engines operated at high flight Mach numbers, which is further confirmed by optimization results of the maximum specific thrust of scramjet engine. Optimization results of the maximum specific thrust also show that improving the isentropic efficiency of the inlet and the maximum combustion temperature is beneficial to the augmentation of the maximum specific thrust of scramjet engine.

Гиперфорсированные воздушно-реактивные двигатели

This paper introduces a thrust augmentation method for super- and hypersonic jet engines by means of applying water at the engine intake. This method expands the use of jet engines with subsonic combustion, allowing velocities up to Mach 8 and altitude up to 45 km. At velocities higher than 3–4 Mach, stagnation temperature of the air is getting higher than the critical temperature of water, which makes the existence of water at the gas turbine engine intake impossible. Water vapour as a working medium of a jet engine creates the so-called inner thermodynamic circle. This phenomenon defines the physics of the thrust augmentation method proposed. The author discusses three variants of hyper afterburner application: hyper afterburner turbojet, hyper afterburner ramjet, and hyper afterburner turbo ejecting engine. The presented basic specifications of the hyper afterburner engines qualitatively differ from those of their prototypes (engines without the hyper afterburner thrust augmentation function). The proposed thrust augmentation method of jet engines is of a special interest for the aerospace field, particularly, for creating air launch systems. It is shown that the application of hyper afterburner in turbo ejecting engines can increase velocity and altitude of the launch aircraft up to Mach 7 and 40 km respectively, thus opening new avenues in space exploration.

Experimental and Numerical Studies of Secondary Injection in Nozzle Divergence for Thrust Augmentation

Experimental and Numerical Studies of Secondary Injection in Nozzle Divergence for Thrust Augmentation

Computational research on a combined undulating-motion pattern considering undulations of both the ribbon fin and fish body

In this work, the combined undulating-motion pattern (CUMP) of the undulatory ribbon-fin propulsion is numerically investigated, inspired by the swimming of the large-mouthed catfish (Silurus soldatovi meridionalis Chen). The swimming of a juvenile large-mouthed catfish is observed and recorded to study the undulation of its elongated anal fin, commonly known as the “ribbon fin”. Compared with the knifefish, which undulates the ribbon fin to swim while maintaining a relatively rigid fish body, the large-mouthed catfish performs a CUMP propulsion, which comprises the undulation of the fish body and the active undulation of the ribbon fin. Accordingly, a fin–body fish model is used to study the effect of the CUMP propulsion on the swimming performance. Simulation results show that the phase-angle difference between the undulations of the fish body and the ribbon fin has a significant impact on the thrust and heave forces. More specifically, an in-phase CUMP propulsion, which is commonly found in the swimming of actual large-mouthed catfish, can dramatically improve the thrust. In addition, the effects of the angular amplitude of ribbon fin and the swimming frequency are considered as well. Finally, the reasons for the thrust augmentation in the in-phase CUMP propulsion are detailed and analyzed.

Independent caudal fin actuation enables high energy extraction and control in two-dimensional fish-like group swimming

We study through numerical simulation the optimal hydrodynamic interactions and basic vorticity control mechanisms for two fish-like bodies swimming in tandem. We show that for a fish swimming in the wake of an upstream fish, using independent pitch control of its caudal fin, in addition to optimized body motion, results in reduction of the energy needed for self-propulsion by more than 50 %, providing a quasi-propulsive efficiency of 90 %, up from 60 % without independent caudal fin control. Such high efficiency is found over a narrow parametric range and is possible only when the caudal fin is allowed to pitch independently from the motion of the main body. We identify the vorticity control mechanisms employed by the body and tail to achieve this remarkable performance through thrust augmentation and destructive interference with the upstream fish-generated vortices. A high sensitivity of the propulsive performance to small variations in caudal fin parameters is found, underlying the importance of accurate flow sensing and feedback control. We further demonstrate that using lateral line-like flow measurements to drive an unscented Kalman filter, the near-field vortices can be localized within 1 % of the body length, and be used with a phase-lock controller to drive the body and tail undulation of a self-propelled fish, moving within the wake of an upstream fish, to stably reach the optimal gait and fully achieve maximum energy extraction.

Investigation of the Rotor–Obstacle Aerodynamic Interaction in Hovering Flight

In this paper, a comprehensive experimental survey of the aerodynamic interaction of a hovering rotor in the proximity of a ground obstacle is described, taking advantage of multiple experimental techniques. Load measurements on the rotor were carried out to assess the change in the rotor performance for different positions with respect to the cubic obstacle, thus simulating a set of possible hovering flight conditions around the obstacle. Laser Doppler anemometry measurements of the rotor inflow were used to investigate how the aerodynamic interaction affected the rotor performance. Stereoscopic particle image velocimetry measurements in the region between the rotor and the obstacle were carried out to gain a better insight of the interacting flow field. The investigation showed two main regions of interest. The first region is the one above the edge of the obstacle, where the rotor experiences a gradual ground effect as it is positioned over the obstacle. The second region, probably of more interest, is the one just beside the obstacle where a recirculation region between the rotor and the obstacle develops, causing both a significant reduction in the thrust augmentation experienced in unobstructed hover in ground effect and significant pitching and rolling moments, due to the nonsymmetrical inflow pattern on the rotor.

Aero engine compressor cooling by water injection - Part 1: Evaporative compressor model

The need for more fuel-efficient turbofan engines has led to a rise in compressor pressure ratio and turbine inlet temperature respectively. The latter has been possible with advancements in turbine blade technology. Nevertheless, this higher temperature during combustion increases the production of thermal Nitrogen Oxides. Contrary to this high-pressure, high-temperature aero-engine design trend, regulations are pushing towards capping or reducing emissions. Injecting atomised water into a jet engine is an alternative to mitigate Nitrogen Oxides that is applied extensively to stationary gas turbines. The application for jet engines is very limited and dates back to the early Boeing 707 and 747 for thrust augmentation. The focus of this study is to investigate the performance benefits of water injection when applied to 2 and 3-spool compressors, under a wide range of different environmental conditions, and for different injection properties. In this first paper, a thermo-analytical compressor model with water droplet investigations in the Lagrangian frame of reference is explored. The methodology is applied to two different engine architectures, representative of modern turbofan engines. This water injection study focuses on cooling the core and shows that the percent reduction in compressor discharge temperature is promising over a wider range of ambient conditions than expected. The effect of droplet sizes or quantity utilised were seen to be more influential. The 3-spool compressor also appears to benefit more from water injection, mainly due to the higher operating pressures and temperatures found on the Intermediate Pressure Compressor which enables more efficient evaporation, as compared to a booster compressor. Given the design of this compressor, two locations of injections were considered. Reductions in Compressor Discharge Temperature of 60 and 80 K were seen for the 2 and 3-spool engines, for a 2% injection ratio, accompanied by reductions in specific compressor work of 16 and 17%. Part 2 of this study has considered boundary conditions obtained here, to investigate the performance and emissions of complete jet engines.

Emulsified jet fuel flame characterization with image processing

This work describes a RGB digital image processing approach of emulsified jet fuel flame, which allows the characterization of the combustion phenomenon in the case of new fuels through color chemiluminiscence measurements. By applying RGB techniques, the image processing of the flame reveals useful parameters in an effective and cost-efficient technique for the determination of relevant chemical species, such as CH* and C2*, equivalence ratio, and temperature estimation.Second generation emulsified aviation fuels containing water-jet fuel have been a challenge for simultaneous thrust augmentation and pollution diminution, with subsequent cost reduction and fossil fuel dependence. Testing new fuels would normally require expensive equipment and reliable investigation techniques, while image processing proved to be a reliable method for the estimation of combustion chemical species and temperature in the case of classic fuels.For the combustion behavior of emulsified jet fuel, a co-annular spray burner was used, allowing the complex investigation with UV-VIS spectrometer and flame photography. RGB image processing techniques showed good agreement with more complicated diagnosis tools, such as spectrometers.

Numerical analysis and design optimization of supersonic after-burning with strut fuel injectors for scramjet engines

Scramjets are a class of hypersonic airbreathing engine that offer promise for economical, reliable and high-speed access-to-space and atmospheric transport. The expanding flow in the scramjet nozzle comprises of unburned hydrogen. An after-burning scheme can be used to effectively utilize the remaining hydrogen by supplying additional oxygen into the nozzle, aiming to augment the thrust. This paper presents the results of a single-objective design optimization for a strut fuel injection scheme considering four design variables with the objective of maximizing thrust augmentation. Thrust is found to be augmented significantly owing to a combination of contributions from aerodynamic and combustion effects. Further understanding and physical insights have been gained by performing variance-based global sensitivity analysis, scrutinizing the nozzle flowfields, analyzing the distributions and contributions of the forces acting on the nozzle wall, and examining the combustion efficiency.

Thrust Augmentation by Airframe-Integrated Linear-Spike Nozzle Concept for High-Speed Aircraft

The airframe-integrated linear-spike nozzle concept applied to an external nozzle for high-speed aircraft was evaluated with regard to the thrust augmentation capability and the trim balance. The main focus was on the vehicle aftbody. The baseline airframe geometry was first premised to be a hypersonic waverider design. The baseline aftbody case had an external nozzle comprised of a simple divergent nozzle and was hypothetically replaced with linear-spike external nozzle configurations. Performance evaluation was mainly conducted by considering the nozzle thrust generated by the pressure distribution on the external nozzle surface at the aftbody portion calculated by computer simulation at a given cruise condition with zero angle of attack. The thrust performance showed that the proposed linear-spike external nozzle concept was beneficial in thrust enhancement compared to the baseline geometry because the design of the proposed concept had a compression wall for the exhaust flow, which resulted in increasing the wall pressure. The configuration with the boattail and the angled inner nozzle exhibited further improvement in thrust performance. The trim balance evaluation showed that the aerodynamic center location appeared as acceptable. Thus, benefits were obtained by employing the airframe-integrated linear-spike external nozzle concept.

Numerical Studies on Thrust Augmentation in High Area Ratio Rocket Nozzles by Secondary Injection

Single stage to orbit propulsion devices are being developed as part of low cost access to space endeavors. Sea level operation of high area ratio rocket nozzle used in rocket engines leads to an overexpanded flow condition resulting in high side loads. Secondary injection of propellants in high area ratio nozzle is an attractive option to overcome the inefficiency of operation of such nozzles in sea level conditions in addition to the augmentation of thrust. A numerical study on thrust augmentation in high area ratio nozzle by secondary injection of propellants is presented here. The turbulent compressible reacting flow in rocket nozzle with auxiliary injection is simulated using conservation equations for chemical species based on finite rate chemistry model and compressible Navier-Stokes equations with AUSM+-up upwind scheme based unstuctured finite volume solver. An optimized eight step, six species reduced H2-O2finite chemistry reaction model is used to model the supersonic combustion. The indigenously developed solver has an efficient rescaling algorithm to alleviate the effect of stiffness in conventional explicit algorithm for simultaneous solution of reacting flow. The code is validated using the wall pressure and hydrogen concentration values reported for the similar high area ratio rocket nozzle. Accurate prediction of nozzle performance is possible with present turbulent reacting flow simulation as it take care of all losses in nozzle flow. Extensive computations have been performed for the performance estimation of high area ratio rocket nozzle for various prospective auxiliary injection options.

Enhanced thrust performance of a two dimensional elliptic airfoil at high flapping frequency in a forward flight

This study is motivated by our earlier investigation Lua et al. (2016) which shows that a two-dimensional elliptic airfoil undergoing sinusoidal flapping motion experiences thrust deterioration when the flapping frequency exceeds a certain critical value (or critical Strouhal number, Stcr). To alleviate this unfavorable thrust generation condition, we propose two novel effective angle of attack profiles, namely smooth trapezoid (STEA) profile and elliptic trapezoid (ETEA) profile. These profiles are designed to ensure that the airfoil still experiences the same effective angle of attack amplitude as the sinusoidal flapping while concurrently reducing the detrimental effect of high rotation rate. The effectiveness of these two proposed profiles is confirmed by our numerical and experimental studies. In particular, our results show that at low to moderate flapping frequency, thrust generation is almost invariant to the type of angle of attack profiles applied, but at high flapping frequency (St>Stcr), the proposed profiles produce higher thrust than the sinusoidal flapping. The thrust augmentation can be attributed to the suppression of the adverse low pressure region in the vicinity of the airfoil, which is a consequence of the reduced rotation rate. Of the two proposed profiles, the ETEA with its steep acceleration phase produces a higher time-average thrust than the STEA near the stroke reversal. Also, in line with our previous finding, thrust augmentation for both STEA and ETEA is a function of the base length of the trapezoid, with a broader one producing a better thrust performance.

Characterizing and modeling the enhancement of lift and payload capacity resulting from thrust augmentation in a propeller-assisted flapping wing air vehicle

Biologically-inspired flapping wing flight is attractive at low Reynolds numbers and at high angles of attack, where fixed wing flight performance declines precipitously. While the merits of flapping propulsion have been intensely investigated, enhancing flapping propulsion has proven challenging because of hardware constraints and the complexity of the design space. For example, increasing the size of wings generates aerodynamic forces that exceed the limits of actuators used to drive the wings, reducing flapping amplitude at higher frequencies and causing thrust to taper off. Therefore, augmentation of aerodynamic force production from alternative propulsion modes can potentially enhance biologically-inspired flight. In this paper, we explore the use of auxiliary propellers on Robo Raven, an existing flapping wing air vehicle (FWAV), to augment thrust without altering wing design or flapping mechanics. Designing such a platform poses two major challenges. First, potential for negative interaction between the flapping and propeller airflow reducing thrust generation. Second, adding propellers to an existing platform increases platform weight and requires additional power from heavier energy sources for comparable flight time. In this paper, three major findings are reported addressing these challenges. First, locating the propellers behind the flapping wings (i.e. in the wake) exhibits minimal coupling without positional sensitivity for the propeller placement at or below the platform centerline. Second, the additional thrust generated by the platform does increase aerodynamic lift. Third, the increase in aerodynamic lift offsets the higher weight of the platform, significantly improving payload capacity. The effect of varying operational payload and flight time for different mixed mode operating conditions was predicted, and the trade-off between the operational payload and operating conditions for mixed mode propulsion was characterized. Flight tests revealed the improved agility of the platform when used with static placement of the wings for various aerobatic maneuvers, such as gliding, diving, or loops.

CFD modeling for flow field characterization and performance analysis of HGITVC

The present study aims at numerically investigating the performance of the Hot Gas Injection Thrust Vector Control (HGITVC) system within the divergent part of a supersonic nozzle. For this purpose, the following measures were taken: considering the flow to be compressible and turbulent, using k-ω-SST model for turbulence modeling, simulating and comparing single and dual injections using a three dimensional and multi block in-house code and validating simulation results using experimental data. The findings indicate that the strength of secondary jet rises as the injection area and injector to nozzle pressure ratio increases. Moreover, it is proved that increase in strength of injection leads to decrease in the amplification factor and increase in the deflection angle and the axial thrust. In addition, a sharp drop in the performance of HGITVC is observed not within acquired permissible range for strength of injection which reduces the system performance. These findings lead us to conclude that in comparison with single injection, the dual injection cannot increase the thrust ratio which is the main objective of this study. However, it helps to inject more gas without shock impingement phenomenon and also improves the Axial Thrust Augmentation (ATA).

Investigation of Thrust Augmentation and Acoustic Performance by Ejectors on PDE

AbstractThrust augmentation and acoustic performance of a Pulse Detonation Engine (PDE) with ejector system is experimentally investigated. For these tests the

Optimization of Combined Rocket and Ramjet/Scramjet Ballistic Missile Designs

A preliminary design optimization of a revolutionary concept involving integrated airbreathing propulsive assist is presented. The conceptual design for the proposed airbreathing thrust augmenter is an annular ramjet/scramjet integrated into a liquid-oxygen-rocket propellant rocket core. The airbreathing engine makes use of the same tankage and turbopump hardware as the rocket engine for supplying fuel. The aeropropulsive environment is simulated using an axisymmetric Euler solver with a decoupled semiempirical viscous model. The flight performance is simulated using a seventh- to eight-order Runge–Kutta integrator for the six-degree-of-freedom dynamics. The optimizer used for design iterations is a binary-encoded genetic algorithm with a proven track record for addressing complex flight vehicle design problems. Results show the marked improvement in performance achieved by the use of such thrust augmenters on a military-dominated niche of the overall spectrum of missile designs.

Thrust augmentation optimization through supersonic after-burning in scramjet engine nozzles via surrogate-assisted evolutionary algorithms

Scramjets are a class of hypersonic airbreathing engine that are associated with realizing the technology required for economical, reliable and high-speed access-to-space and atmospheric transport. The expanding flow in the scramjet nozzle comprises of unburned hydrogen which under ideal conditions, can be utilized to introduce an after-burning scheme. After-burning augments the thrust produced by the scramjet nozzle and creates a more robust nozzle design. This paper presents a single-objective design optimization considering three design variables with the objective of producing maximum thrust augmentation. It is found that significant levels of thrust augmentation are produced based upon contributions from increased pressure, mass flow and energy in the nozzle. Further understanding of the phenomenon by which thrust augmentation is being produced is provided in the form of variance-based global sensitivity analysis, force contribution breakdowns, analysis of the nozzle flowfields, analysis of the surface pressure and shear stress distributions acting on the nozzle wall and analysis of the combustion efficiency.

Performance investigation of a pulse detonation turbine engine

In order to obtain the performance of a pulse detonation turbine engine (PDTE) and compare the PDTE performance with the traditional turbine engine performance, analytical and experimental investigations were conducted in the present work. Ideal thermodynamic cycle analysis was firstly carried out to obtain the upper performance limit of both the PDTE cycle and the Brayton cycle-based engines. A simplified analytical model and a liquid-fueled dual tube PDTE system were built up in order to study the real performance of the PDTE. Thrust performance of the PDTE system was obtained under a self-aspirated mode. The calculated thrust performance based on the simplified analytical model was compared with the experimental data to verify the reliability of the simplified analytical model. Additionally, the specific thrust and specific fuel consumption (SFC) of the PDTE system were compared with the ideal performance of the PDTE cycle and the Baryton cycle-based engines. The results indicate that the ideal PDTE performance is much better than the traditional turbine engine performance. The dual tube PDTE system can be successfully operated at frequencies from 10 Hz to 20 Hz under the self-aspirated mode, indicating the feasibility of the PDTE concept. The simplified analytical model could estimate the real performance of the PDTE if proper losses were being taking into consideration. The specific thrust of the PDTE system is higher than that of the ideal Brayton cycle-based engines, while the SFC is much lower. When an ejector was used for thrust augmentation, the performance could be largely increased. It proves the performance advantages of the PDTE over the traditional turbine engine experimentally for the first time.