Thrust Ratio Research Articles

Dual-Thrust Rocket Motor with an intermediate Nozzle at the condition of higher boosting pressure than the sustaining pressure

Boost-sustain solid propellant rocket motors have been used to meet the demand for short and medium range tactical missiles. These motors obtain a wide variety of thrust-time profiles if the propellant characteristics and grain geometry are properly selected. These Dual Thrust Rocket Motors (DTRM) can have various designs; the DTRM with an intermediate nozzle is one design that yields a high boost-sustain thrust ratio with a stable operation. Under certain conditions, the flow through the intermediate nozzle at the start of motor operation can be reversed. The present paper is intended to understand the phenomena and determine the operating parameters of DTRM with the reverse flow intermediate nozzle. In this study, a mathematical model describing the performance is derived from the main governing equations of internal ballistics (IB). Also, a set of experiments are conducted so as to inspect the details of the physics of such configuration DTRM. The experimental findings are utilized to validate the mathematical model. Results of the theoretical and experimental investigation explain most of the IB characteristics that were not discussed before in the open literature.

Analysis of thrust-induced drilling in composite materials using a hemispherical drill

Although composite parts are produced to a near-net shape, drilling is frequently required to make precise holes for assembly. A hemispherical drill has a quadratic distributed load over a central circular area of the radius to produce a concave, a hemicave, or a hole in structural parts. However, a hemispherical drill is characterized by a relatively high thrust force in drilling composites due to the negative rake and null cutting speed at the near-tool center, which is similar to cutting phenomenon in the chisel edge of a twist drill. An analysis of thrust-induced delamination when hemispherical drilling composite materials is presented in this study. The theoretical analysis shows that the thrust force for a hemispherical drill that produces delamination increases as the thrust ratio (α) decreases. This induces delamination in composite materials. The maximum theoretical thrust force for a hemispherical drill occurs at α = 2, which is about 30 % of the value for a twist drill. In other words, a hemispherical drill produces lower bearing stresses than a twist drill, which can cause separation of the plies at the exit, as the interlaminar bonding yields. The results agree with industrial experience, which shows that a hemispherical drill causes more delamination during drilling. The proposed model shows that α must be selected when a hemispherical drilling is used for composites.

A study on the thrust vector control using a bypass flow passage

Bypass actuators are both indispensable and critical components for thrust vector control of supersonic vehicles. Detailed flow features concerned with bypass flow play a key role in thrust vectoring systems. An analytical model was created in supersonic nozzle with bypass flow passage. Meanwhile, a computational fluid dynamics-based study was conducted for studying this phenomenon. To validate the analytical and numerical models, comparisons with experimental result were performed. Good agreement was observed in overall region. Analytical and numerical pressure distributions under different injection positions were compared. The distance between separation point and injection position becomes smaller as the secondary injection position moves upstream. This is because of the reducing of bypass flow rate. The effects of expanded ratio and mass flow rate were adequately researched and tested. Significant variation of system thrust ratio and specific impulse takes place in the over-expanded conditions. It is indicated that the secondary flow produces great effects for over-expanded conditions. The separation point of boundary layer moves upstream as bypass flow rate increases. System thrust ratio and specific impulse will reduce if bypass flow rate increases.

Analytical and experimental approach of the high density transverse flux linear stepper motor

Detailed analytical and experimental approaches were performed to calculating the ratio of the pull-in thrust to the input power of High Density Transverse Flux linear Stepper Motor (HDTFLSM) with considerations of the electrical time constant, the mechanical time constant and the friction. The analytical approach involves the use of Fourier series analysis to derive the equation for the pull-in thrust to input power ratio. The experimentation was performed under different load conditions until the motor come to the standstill value. A chopped voltage was applied to motor at different speeds and the obtained results were compared with those for normal voltage. Comparison between analytical and experimental results shows good agreement, for instance at 0.22 m/s the difference percentage on normal voltage and chopped voltage are 6.1% and 6.7%, respectively. The chopped voltage has higher ratio of the pull-in thrust to the input power compare to normal voltage so that the machine performance can be improvised to a greater extent. The maximum improvement percentage of 18.8% on pull-in thrust to input power ratio for motor with chopped voltage was obtained at 0.22 m/s.

Numerical simulation of X-wing type biplane flapping wings in 3D using the immersed boundary method

The numerical simulation of an insect-sized ‘X-wing’ type biplane flapping wing configuration is performed in 3D using an immersed boundary method solver at Reynolds numbers equal to 1000 (1 k) and 5 k, based on the wing's root chord length. This X-wing type flapping configuration draws its inspiration from Delfly, a bio-inspired ornithopter MAV which has two pairs of wings flapping in anti-phase in a biplane configuration. The objective of the present investigation is to assess the aerodynamic performance when the original Delfly flapping wing micro-aerial vehicle (FMAV) is reduced to the size of an insect. Results show that the X-wing configuration gives more than twice the average thrust compared with only flapping the upper pair of wings of the X-wing. However, the X-wing's average thrust is only 40% that of the upper wing flapping at twice the stroke angle. Despite this, the increased stability which results from the smaller lift and moment variation of the X-wing configuration makes it more suited for sharp image capture and recognition. These advantages make the X-wing configuration an attractive alternative design for insect-sized FMAVS compared to the single wing configuration. In the Reynolds number comparison, the vorticity iso-surface plot at a Reynolds number of 5 k revealed smaller, finer vortical structures compared to the simulation at 1 k, due to vortices’ breakup. In comparison, the force output difference is much smaller between Re = 1 k and 5 k. Increasing the body inclination angle generates a uniform leading edge vortex instead of a conical one along the wingspan, giving higher lift. Understanding the force variation as the body inclination angle increases will allow FMAV designers to optimize the thrust and lift ratio for higher efficiency under different operational requirements. Lastly, increasing the spanwise flexibility of the wings increases the thrust slightly but decreases the efficiency. The thrust result is similar to one of the spanwise studies, but the efficiency result contradicts it, indicating that other flapping parameters are involved as well. Results from this study provide a deeper understanding of the underlying aerodynamics of the X-wing type, which will help to improve the performance of insect-sized FMAVs using this unique configuration.

Collar Length on the Performance of a Nozzle Using Fluidic Counterflow for Thrust Vectoring

Numerical simulation is applied to study internal flow structure and performance of counterflow thrust vectoring nozzle with collar length in zero attack angle and subsonic conditions. The changes of thrust vectoring angle, resultant thrust ratio and secondary mass flow ratios are obtained. Results indicate that the thrust vectoring angle firstly increases and then decreases with increasing suction collar length, the max of which is 5.5o; but resultant thrust ratio decreases, and its range is 0.76-0.78;During the procedure of increasing suction collar length, suction flow shifts from counterflow to coflow, and mass flow ratio decreases.

Two-dimensional quasi-double-layers in two-electron-temperature, current-free plasmas

The expansion of a plasma with two disparate electron populations into vacuum and channeled by a divergent magnetic nozzle is analyzed with an axisymmetric model. The purpose is to study the formation and two-dimensional shape of a current-free double-layer in the case when the electric potential steepening can still be treated within the quasineutral approximation. The properties of this quasi-double-layer are investigated in terms of the relative fraction of the high-energy electron population, its radial distribution when injected into the nozzle, and the geometry and intensity of the applied magnetic field. The two-dimensional double layer presents a curved shape, which is dependent on the natural curvature of the equipotential lines in a magnetically expanded plasma and the particular radial distribution of high-energy electrons at injection. The double layer curvature increases the higher the nozzle divergence is, the lower the magnetic strength is, and the more peripherally hot electrons are injected. A central application of the study is the operation of a helicon plasma thruster in space. To this respect, it is shown that the curvature of the double layer does not increment the thrust, it does not modify appreciably the downstream divergence of the plasma beam, but it increases the magnetic-to-pressure thrust ratio. The present study does not attempt to cover current-free double layers involving plasmas with multiple populations of positive ions.

Experimental study on the thrust and the ratio of thrust to power of multi-points/grid ionic wind exciter

The thrust of ionic wind exciters in 19 points-grid and 31 points-grid electrodes configurations was measured in this paper by means of air suspension test platform without friction. Meanwhile, a control method was also provided for multipoint-grid ionic wind exciter. And then the study was done on the factors, e.g. the gap of point-grid electrodes, the spacing between points, the curvature radius of the point, and the voltage, current of corona discharge, all of which may affect the thrust character; so some methods for getting high thrust at low power were also proposed in order to improve the thrust and the thrust per unit power at the same time.

Design performance evaluation and vortex structure investigation of different S-shaped intermediate turbine ducts

The demand of further increasing bypass ratio of aeroengine will lead to low pressure turbines with higher diameter. Therefore, it is necessary to design a duct to guide the hot gas flow which is expelled from the upstream high pressure (HP) turbine stage to the downstream low pressure (LP) turbine stage. Named by its position, this kind of duct is always called intermediate turbine ducts (ITDs). Due to the pursuit of higher thrust ratio of the aeroengine, this kind of ITDs has to beas short as possible which leads to aggressive (high diffusion) S-shaped ITDs' geometry. In this paper, two different schemes of high diffusion separation-free S-shaped ITDs were studied with the aid of three-dimensional CFD programs. Although these two ITDs have the same area ratios (AR), because of the different duct length, they have totally different area as well as area change rates. With the detailed calculation results, comparisons were made to investigate the underneath physical mechanisms. Additionally, a direct estimation of the ITDs' loss is given at the end of this paper and some ITDs' novel design idea is proposed to initiate some further discussions.

Design and Optimization of 3D Wagon Wheel Grain for Dual Thrust Solid Rocket Motors

AbstractThe combination of various design and performance parameters in dual thrust solid rocket motors coupled with constraints may lead to a multitude of configurations. It is difficult to initially determine a suitable design without optimization. In this article, a methodology for the design of a 3D wagon wheel that provides two levels of thrust, namely boost phase thrust and sustain phase thrust, is investigated. The proposed framework consists of an integrated approach of grain design, computer aided design and hybrid optimization. A hybrid genetic algorithm and simulated annealing are used for optimization due to their robustness and capability to find global optimal solution. Several benchmark test functions are used to validate the performance of the optimization algorithm and solution. Maximization of the average thrust ratio and the total impulse are the objective functions under the given constraints.

The influence of engine thrust behaviour on the aerodynamics of engine airframe integration

The wind tunnel is still an important pillar for the aerodynamic investigations of engine installation effects, such as the strong aerodynamic interferences between an underwing mounted engine and the wing itself. To consider real engine effects on the aerodynamics in the wind tunnel environment the turbine powered simulator (TPS) technique is one of the widely applied technologies. Especially for underwing installed engines thrust affects the aerodynamic drag. Reflecting the fact that in modern aircraft design engine installation improvements on drag counts in small numbers, and considering that the TPS technology cannot match the real engine in all parameters, the questions arises if no matching parameters exist what is their influence on drag. Usually fan Mach numbers and total thrust coefficients of the real engine can be correctly realized by TPS. In the present contribution the essential parameters of engine simulation are determined and in particular the thrust ratios, i.e. turbine thrust to fan thrust of real engines and simulators at different bypass ratios (BPR) in the range of 6–17 at cruise conditions will be evaluated. Considerable differences of the thrust ratios between the engines and the simulators exist in particular at high freestream Mach numbers. Therefore, the aim of this contribution will be to clarify the influence of the thrust ratio on the drag behaviour of an aircraft configuration. First, the theoretical background on the drag determination theory is derived. Based on this theory the analysis of the thrust ratio is done using mainly experimental results. Additional numerical results by solving the Reynolds-averaged Navier–Stokes equations are used to support the analysis.

Active Dual Thrust Modulation of a Solid Rocket Motor

Based on a combustion characteristic that some propellants cannot burn in an intermediate pressure range while they can burn at lower and higher pressure, an active thrust modulation system was developed. The motor changed its thrust in dual combustion mode. It chose alternately high and low thrusts during combustion. The transitions from the low mode to the high mode were attained by a secondary ignition system, and those from the high mode to the low mode were done by a brief gas release system. Up to three times mode transitions were successfully demonstrated. Required time for the mode transitions and the effects of the thrust modulations on the thrust performance were evaluated. On the present motor scale, the transition time from the low mode to the high mode ranged from 0.12 s to 0.23 s, and that from the high mode to the low mode was from 0.19 s to 0.38 s. The thrust variable range was adjustable by the throat area. The ratio of the low-mode thrust to the high mode thrust was variable from 3 to 5. Specific impulse was decreased from 195 s to 175 s, when the number of thrust modulations was increased.

Uncertainty-based computational robust design optimisation of dual-thrust propulsion system

This paper proposes a robust design optimisation methodology for performance evaluation of high-fidelity propulsion systems under both aleatory and epistemic uncertainties. In this paper, uncertainty-based robust design methodology is applied using first-order orthogonal design to estimate worst case of the uncertainties. The effectiveness and computational efficiency of proposed formulation is obtained by decoupling computation of both types of uncertainties in the optimisation process loop. The system robustness is measured by directly minimising the variance of target mean and standard deviation and adhering to performance constraints. A parametric sensitivity analysis of performance parameters due to input parameters variations is also performed to identify the most significant design variables. A hybrid genetic algorithm and simulated annealing approach are used as an optimiser. By using the proposed method, the objective function to maximise the booster phase to sustain phase thrust ratio and total impulse could be achieved while adhering to the motor performance constraints to provide an insensitive solution.

Development of preliminary design program for combustor of regenerative cooled liquid rocket engine

An integrated program was established to design a combustor for a liquid rocket engine and to analyze regenerative cooling results on a preliminary design level. Properties of burnt gas from a kerosene-LOx mixture in the combustor and rocket performance were calculated from CEA which is the code for the calculation of chemical equilibrium. The heat transfer of regenerative cooling was analyzed by using SUPERTRAPP code for coolant properties and by one-dimensional correlations of the heat transfer coefficient from the combustor liner to the coolant. Profiles of the combustors of F-1 and RS-27A engines were designed from similar input data and the present results were compared to actual data for validation. Finally, the combustors of 30 tonf class, 75 tonf class and 150 tonf class were designed from the required thrust, combustion chamber, exit pressure and mixture ratio of propellants. The wall temperature, heat flux and pressure drop were calculated for heat transfer analysis of regenerative cooling using the profiles.

MACHINING OF CORTICAL BONE: SIMULATIONS OF CHIP FORMATION MECHANICS USING METAL MACHINING MODELS

This paper investigates chip formation in the machining of cortical bone and the application of isotropic elastic-plastic material models with a pressure dependent yield stress and a strain path dependent failure strain law to finite element calculations to predict observed behaviour. It is shown that a range of models can be created that result in segmented chip formations and a range of specific cutting forces similar to those observed experimentally. Results from the simulations provide an explanation for differences in the ratio of thrust to cutting forces observed between previous experimental studies, namely that the cutting tools used may have had different edge sharpness or degree of damage induced by the material removal process. Measurements of edge profiles from one of these studies support that explanation and emphasize the importance of tool toughness in maintaining efficient cutting of bone.

Robust Design Optimization of an Aerospace Vehicle Prolusion System

This paper proposes a robust design optimization methodology under design uncertainties of an aerospace vehicle propulsion system. The approach consists of 3D geometric design coupled with complex internal ballistics, hybrid optimization, worst‐case deviation, and efficient statistical approach. The uncertainties are propagated through worst‐case deviation using first‐order orthogonal design matrices. The robustness assessment is measured using the framework of mean‐variance and percentile difference approach. A parametric sensitivity analysis is carried out to analyze the effects of design variables variation on performance parameters. A hybrid simulated annealing and pattern search approach is used as an optimizer. The results show the objective function of optimizing the mean performance and minimizing the variation of performance parameters in terms of thrust ratio and total impulse could be achieved while adhering to the system constraints.

Influence of Design Parameters on End Effect in Long Primary Double-Sided Linear Induction Motor

A 1-D analytical technique for predicting the airgap magnetic field distribution and the time-averaged thrust of long primary double-sided linear induction motor (LP-DSLIM), taking account both fundamental and entry end effect component, has been developed. The analytical results are compared and validated by the predictions from corresponding finite element analyses. The end effect-related attenuating constant and thrust ratio factor are investigated as the functions of the motor design parameters, which involve the airgap length, secondary thickness and conductivity, frequency, and pole pair numbers. The corresponding measures to alleviate the end effect at high speed are given. The method allows the assessment of the end effect and helps guide the design procedure for LP-DSLIM.

Numerical Calculation Analysis of Structural Parameters of Water-Lubricated Thrust Bearing

The affect of the thrust pad inclination and ratio of length to breadth (L/B) on the lubricity of a water-lubricated thrust bearing is studied in numerical method. The research results indicate that when the elastic deformation is considered, the distribution of both water film thickness and pressure presents to be paraboloid. With the increase of thrust pad inclination, both maximum water-film pressure and the friction coefficient increase, and the minimum water-film thickness decreases first and then increases; when the L/B increases in the range of 1~2, both maximum water-film pressure and the friction coefficient increase, but the minimum water-film thickness decreases.

Three-Dimensional Analysis of a Supersonic CombustorCoupled to Innovative Inward-Turning Inlets

Three-dimensional simulations are employed to examine the effect of inlet distortion and model fidelity with Reynolds averaged Navier-Stokes or large eddy simulation approaches for a generic circular cross-sectional supersonic combustor at a flight condition of Mach 6 and an altitude of about 24.2 km. To examine inlet distortion effects on combustion, frozen and finite rate chemistry simulations are performed on combustors connected to two different types of streamline-traced inlets (denoted and Jaws) with the Wilcox k-ω turbulence model. For comparison, uniform inflow boundary condition to the combustor is also simulated. The metrics employed include qualitative assessments related to flow structure as well as quantitative values offuel combustion efficiency and thrust ratios. The results indicate a complex overall effect of distortion due to inlet design. For Jaws, the increased pressure loss associated with distortion is mitigated slightly by improved combustion efficiency and better thrust performance. The Scoop inlet has lower distortion and better recovery, but the combustion coefficient is lower than Jaws. In the second part ofthis study, finite rate chemistry results with unsteady Reynolds averaged Navier-Stokes are compared with those from large eddy simulation with an uniform inflow profile. Differences in transient processes include the manner in which large- and small-scale structures originate and evolve in the cavity recirculation and downstream regions. The finer details observed with the large eddy simulation model have significant consequences on the overall field, including the unsteady position of the shock structure arising from the interaction of the incoming flow with the cavity shear layer, combustion processes, and injection jet interactions.

Coseismic fault slip of the 2008 Mw 7.9 Wenchuan earthquake estimated from InSAR and GPS measurements

We infer co‐seismic fault slip during the 2008 Mw 7.9 Wenchuan earthquake from interferometric synthetic aperture radar (InSAR) and GPS observations of ground deformation. We use ALOS/PALSAR data from ascending orbits on six tracks, and we do not use data that are strongly affected by ionospheric perturbations. We use a fault model composed of three planar fault segments of the Beichuan fault, and one planar segment representing the parallel Pengguan fault. Maximum thrust‐slip is up to 6.7 m near the surface, and occurs in two locations, near Yingxiu in the south and Beichuan in the center of the rupture. Maximum strike‐slip is over 4 m, and occurs near Pingtong and Nanba along the northern end of the rupture. We find that the ratio of coseismic thrust‐ to strike‐slip on the Beichuan fault decreases from 1.5 to 0.7 from the SW to the NE.