Experimental evaluation and numerical simulation of performance of the bypass dual throat nozzle

2020 ◽  
pp. 095441002095988
Author(s):  
Mohammad Hadi Hamedi-Estakhrsar ◽  
Hossein Mahdavy-Moghaddam
Keyword(s):  
Numerical Simulation ◽  
Discharge Coefficient ◽  
Deflection Angle ◽  
Pressure Ratio ◽  
Modern Concept ◽  
Thrust Coefficient ◽  
Thrust Vector Control ◽  
Nozzle Pressure ◽  
Thrust Vector ◽  
Turbulent Air Flow

Bypass dual throat nozzle (BDTN) is a modern concept of fluidic thrust vector control. This method able to solve the problem of thrust loss without need the secondary mass flow from other part of engine. Internal nozzle performance and thrust vector angles have been measured in the BDTN experimentally and numerically. A new simple approach is proposed to detect the thrust deflection angle. Numerical simulation of 3-D turbulent air flow is carried out by using the RNG k-e turbulence model. The obtained results of thrust coefficient, discharge coefficient and thrust deflection angle have been validated by comparing with measured experimental data. The results show that for nozzle pressure ratio of 1–4 the tested nozzle able to deflect the thrust vector of 26.5°-19°. By increasing NPR from 2 up to 4, the thrust coefficient values will change in the range of 0.85-0.93. Also the effect of different positions of the bypass channel on the BDTN performance parameters has been investigated numerically. The predicted results show that the BDTN configuration with bypass duct on the first nozzle throat has the highest value of thrust deflection angle over the range of NPRs.

Theoretical and Numerical Studies on the Performance of Shock Vector Control

2019 ◽  
Author(s):  
Kexin Wu ◽  
HeuyDong Kim
Keyword(s):  
Experimental Data ◽  
Vector Control ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Symmetry Plane ◽  
Control Effect ◽  
Thrust Coefficient ◽  
Thrust Vector Control ◽  
Pressure Distributions ◽  
Thrust Vector

Abstract The transverse injection into a supersonic flow is a significant application that appeared in numerous aerodynamic applications, such as drag reduction and fluidic thrust vectoring control. Nowadays, fluidic thrust vector control is gradually replacing mechanical thrust vector control to redirect various air vehicles. Shock vector control is very popular in fluidic thrust vector control field due to lots of advantages, such as simple structure, more integrated control effect, and quick vectoring response. In present works, numerical simulations and theoretical analyses were conducted to investigate the shock vectoring performance in a three-dimensional rectangular nozzle. To validate the reliability and accuracy of the present numerical methodology, static pressure distributions along upper and lower nozzle surfaces in the symmetry plane were compared with experimental data published by NASA. It was evident that present numerical results present great approximations with experimental data. Control variables of the slot injector were studied, which not only include slot length and slot width but also contain uniform mass flow ratio and injection pressure ratio. Performance variations were illustrated clearly, such as static pressure distributions along upper and lower nozzle surfaces, deflection angle, resultant thrust coefficient, and thrust efficiency. Useful conclusions were obtained for further investigations on shock vector control.

Fluidic thrust vector control based on counter-flow concept

2018 ◽  
Vol 233 (4) ◽  
pp. 1412-1422 ◽  
Author(s):  
Kexin Wu ◽  
Heuy Dong Kim ◽  
Yingzi Jin
Keyword(s):  
Vector Control ◽  
Mass Flow ◽  
Deflection Angle ◽  
Pressure Ratio ◽  
Supersonic Nozzle ◽  
Flow Ratio ◽  
Counter Flow ◽  
Thrust Vector Control ◽  
Thrust Vector ◽  
Mass Flow Ratio

Computational studies are conducted on the supersonic nozzle to investigate the possibility of utilizing counter-flow in fluidic thrust vector control. In this work, the design Mach number of the symmetric supersonic nozzle is set to be 2.5. For the validation of methodology, numerical results are compared with experimental data referred from the literature. Two-dimensional numerical simulations are based on well-assessed standard k–ɛ turbulence model with standard wall functions. Second-order accuracy is ensured to reveal more details of flow field. The system thrust ratio, deflection angle, and secondary mass flow ratio were studied for a wide range of nozzle pressure ratios and secondary pressure ratios. The results indicate that deflection angle and secondary mass flow ratio are found to be decreased with increasing nozzle pressure ratio as well as system thrust ratio. The secondary mass flow ratio and deflection angle decrease with the increase of secondary pressure ratio, and system thrust ratio increases with the increasing of secondary pressure ratio. The secondary mass flow rate remains under 2.4% of the primary flow to obtain efficient thrust vector control at high Mach number.

An Investigation of Flow Characteristics and Parameter Effects for a New Concept of Hybrid SVC Nozzle

2018 ◽  
Author(s):  
F. Song ◽  
J. W. Shi ◽  
L. Zhou ◽  
Z. X. Wang ◽  
X. B. Zhang
Keyword(s):  
Static Pressure ◽  
Pressure Ratio ◽  
Flow Characteristics ◽  
Exhaust System ◽  
Aircraft Engine ◽  
Thrust Coefficient ◽  
Thrust Vectoring ◽  
Static Pressure Distribution ◽  
Thrust Vector ◽  
Vector Angle

Lighter weight, simpler structure, higher vectoring efficiency and faster vector response are recent trends in development of aircraft engine exhaust system. To meet these new challenges, a concept of hybrid SVC nozzle was proposed in this work to achieve thrust vectoring by adopting a rotatable valve and by introducing a secondary flow injection. In this paper, we numerically investigated the flow mechanism of the hybrid SVC nozzle. Nozzle performance (e.g. the thrust vector angle and the thrust coefficient) was studied with consideration of the influence of aerodynamic and geometric parameters, such as the nozzle pressure ratio (NPR), the secondary pressure ratio (SPR) and the deflection angle of the rotatable valve (θ). The numerical results indicate that the introductions of the rotatable valve and the secondary injection induce an asymmetrically distributed static pressure to nozzle internal walls. Such static pressure distribution generates a side force on the primary flow, thereby achieving thrust vectoring. Both the thrust vector angle and vectoring efficiency can be enhanced by reducing NPR or by increasing θ. A maximum vector angle of 16.7 ° is attained while NPR is 3 and the corresponding vectoring efficiency is 6.33 °/%. The vector angle first increases and then decreases along with the elevation of SPR, and there exists an optimum value of SPR for maximum thrust vector angle. The effects of θ and SPR on the thrust coefficient were found to be insignificant. The rotatable valve can be utilized to improve vectoring efficiency and to control the vector angle as expected.

Numerical Simulation of Supersonic Jet Control by Tabs with Slanted Perforation

2019 ◽  
Vol 0 (0) ◽  
Author(s):  
G. Ezhilmaran ◽  
Suresh Chandra Khandai ◽  
Yogesh Kumar Sinha ◽  
S. Thanigaiarasu
Keyword(s):  
Numerical Simulation ◽  
Mach Number ◽  
Near Field ◽  
Pressure Ratio ◽  
Supersonic Jet ◽  
Pressure Gradients ◽  
Nozzle Pressure Ratio ◽  
Nozzle Pressure ◽  
Jet Control ◽  
Different Diameters

Abstract This paper presents the numerical simulation of Mach 1.5 supersonic jet with perforated tabs. The jet with straight perforation tab was compared with jets having slanted perforated tabs of different diameters. The perforation angles were kept as 0° and 10° with respect to the axis of the nozzle. The blockage areas of the tabs were 4.9 %, 4.9 % and 2.4 % for straight perforation, 10° slanted perforation ( {{{\Phi }}_{\ }} = 1.3 mm) and 10° slanted perforation ( {{{\Phi }}_{\ }} = 1.65 mm) respectively. The 3-D numerical simulations were carried out using the software. The mixing enhancements caused by these tabs were studied in the presence of adverse and favourable pressure gradients, corresponding to nozzle pressure ratio (NPR) of 3, 3.7 and 5. For Mach number 1.5 jet, NPR 3 corresponds to 18.92 % adverse pressure gradients and NPR 5 corresponds to 35.13 % favourable pressure gradients. The centerline Mach number of the jet with slanted perforations is found to decay at a faster rate than uncontrolled nozzle and jet with straight perforation tab. Mach number plots were obtained at both near-field and far field downstream locations. There is 25 % and 65 % reduction in jet core length were observed for the 0° and 10° perforated tabs respectively in comparison to uncontrolled jet.

Demonstration of Fluidic Throat Skewing for Thrust Vectoring in Structurally Fixed Nozzles

2001 ◽  
Vol 123 (3) ◽  
pp. 502-507 ◽  
Author(s):  
P. J. Yagle ◽  
D. N. Miller ◽  
K. B. Ginn ◽  
J. W. Hamstra
Keyword(s):  
Conceptual Development ◽  
Discharge Coefficient ◽  
Companion Paper ◽  
Nozzle Throat ◽  
Thrust Coefficient ◽  
Thrust Vectoring ◽  
Flow Rates ◽  
Critical Design ◽  
Design Variables ◽  
Nozzle Pressure

The experimental demonstration of a fluidic, multiaxis thrust vectoring (MATV) scheme is presented for a structurally fixed, afterburning nozzle referred to as the conformal fluidic nozzle (CFN). This concept for jet flow control features symmetric injection around the nozzle throat to provide throttling for jet area control, and asymmetric injection to subsonically skew the sonic plane for jet vectoring. The conceptual development of the CFN was presented in a companion paper (Miller et al. [1]). In that study, critical design variables were shown to be the flap length and expansion area ratio of the nozzle, and the location, angle, and distribution of injected flow. Measures of merit were vectoring capability, gross thrust coefficient, and discharge coefficient. A demonstration of MATV was conducted on a 20 percent scale CFN test article across a range of nozzle pressure ratios (NPR), injector flow rates, and flow distributions. Both yaw and pitch vector angles of greater than 8 deg were obtained at NPR of 5.5. Yaw vector angles greater than 10 deg were achieved at lower NPR. Values of thrust coefficient for the CFN generally exceeded published measurements of shock-based vectoring methods. In terms of vectoring effectiveness (ratio of vector angle to percent injected flow), fluidic throat skewing was found to be comparable to shock-based vectoring methods.

The Formula of Discharge Coefficient of the Cooling Hole in the Combustion Chamber for a Uniform Penetration Boundary Method

2020 ◽  
Author(s):  
Sen Wang ◽  
Donghai Jin ◽  
Xingmin Gui ◽  
Xiaoheng Liu
Keyword(s):  
Numerical Simulation ◽  
Mach Number ◽  
Combustion Chamber ◽  
Discharge Coefficient ◽  
Pressure Ratio ◽  
Diameter Ratio ◽  
Geometric Parameters ◽  
Cooling Flow ◽  
Cooling Hole ◽  
Aerodynamic Parameters

Abstract At present, researches on the discharge coefficient of the combustion chamber cooling holes are mostly based on the experimental study of the porous plates considering various geometric structures under different flow conditions. In this method, the average discharge coefficients of multiple holes are obtained. Since the discharge coefficient will be applied to the numerical simulation, it is important to obtain an accurate formula for each cooling hole. Therefore, the discharge coefficient will be associated with some local aerodynamic parameters around the cooling holes and geometric parameters of the cooling holes. In this paper, the geometric parameters consist of length-to-diameter ratio (L/d) and inclination angle (α). The aerodynamic parameters cover the Mach number of cooling flow and mainstream (Mac, Mam), characterizing the flow feature, and the pressure ratio (π) which associates the cooling flow with the mainstream. The purpose of this paper is to study the discharge coefficient of a single circular hole with variable geometries under a cold condition of a combustion chamber, which has low crossflow and low-pressure ratio on both sides of the hole. In this environment, according to the research results, the discharge coefficient is sensitive to the cooling flow Mach number, length-to-diameter ratio and pressure ratio (π ≈ 1.05). Discharge coefficient decreases with L/d linearly, conforms the quadratic function with Mac and changes complexly at π ≈ 1.05. Other parameters have little effect on the discharge coefficient. The data for discharge coefficient of the cylindrical hole considering different parameters is obtained through numerical simulation and the correlations summarized by these data are valid for the following ranges: L/d = 3∼12, α = 20°∼45°, Mac = 0.05∼0.15, Mam = 0∼0.1, π = 1.05∼1.15. Compared with the CFD data, the prediction formula has a maximum error of less than 3% and a mean absolute error of 0.78%.

Dynamic Experimental Investigations of a Bypass Dual Throat Nozzle

2015 ◽  
Vol 137 (8) ◽  
Author(s):  
Rui Gu ◽  
Jinglei Xu
Keyword(s):  
Experimental Study ◽  
Pressure Ratio ◽  
The Other ◽  
Experimental Investigations ◽  
Vector State ◽  
Nozzle Pressure Ratio ◽  
Dynamic Hysteresis ◽  
Nozzle Pressure ◽  
Thrust Vector ◽  
Secondary Injection

The bypass dual throat nozzle (BDTN) does not consume any secondary injection from the other part of the engine, while it can produce steady and efficient vectoring deflection similar to the conventional dual throat nozzle (DTN). A BDTN model has been designed and subjected to dynamic experimental study. The main results show that: (1) The frequency spectrums of the dynamic pressures are different between each thrust vector state. (2) The variation rates of dynamic vector of the new BDTN can reach as high as 50 deg/s, 40 deg/s, and 34 deg/s under nozzle pressure ratio NPR = 3, 5, and 10, separately. (3)The dynamic hysteresis time is less than 1 ms.

Numerical investigation of separation pattern and separation pattern transition in overexpanded single expansion ramp nozzle

2014 ◽  
Vol 118 (1202) ◽  
pp. 399-424 ◽  
Author(s):  
Y. Yu ◽  
J. Xu ◽  
J. Mo ◽  
M. Wang
Keyword(s):  
Numerical Simulation ◽  
Pressure Ratio ◽  
Hypersonic Vehicle ◽  
Small Range ◽  
Nozzle Pressure Ratio ◽  
Rocket Nozzle ◽  
Nozzle Pressure ◽  
Separation Pattern ◽  
Separation Results ◽  
Free Shock

Abstract Flow separation results in many problems to single expansion ramp nozzle (SERN) and hypersonic vehicle. However, little research has been conducted on the separation patterns and their effects on SERN’s performance. In the present paper, the numerical simulation is adopted to get the intuitive results and help to analyse the separation phenomena in SERN thoroughly. The main separation pattern is the restricted shock separation (RSS) in SERN, and the free shock separation (FSS) only appears in a small range of the nozzle pressure ratio (NPR), which is much different from the axisymmetric rocket nozzle. Further CFD results show that the separation pattern transition makes great effects on the performance of SERN, especially the lift. Moreover, the performance of SERN has an extreme in the separation pattern transition because of the main jet impinging on the expansion ramp. The transitions occur in both the startup and shutdown processes but the critical nozzle pressure ratios of the separation pattern transitions are different, which leads to a hysteresis loop of SERN performance.

Demonstration of Fluidic Throat Skewing for Thrust Vectoring in Structurally Fixed Nozzles

2000 ◽  
Author(s):  
Patrick J. Yagle ◽  
Daniel N. Miller ◽  
K. Brant Ginn ◽  
Jeffrey W. Hamstra
Keyword(s):  
Conceptual Development ◽  
Discharge Coefficient ◽  
Companion Paper ◽  
Nozzle Throat ◽  
Thrust Coefficient ◽  
Thrust Vectoring ◽  
Flow Rates ◽  
Critical Design ◽  
Design Variables ◽  
Nozzle Pressure

The experimental demonstration of a fluidic, multi-axis thrust vectoring (MATV) scheme is presented for a structurally fixed, afterburning nozzle referred to as the conformal fluidic nozzle (CFN). This concept for jet flow control features symmetric injection around the nozzle throat to provide throttling for jet area control, and asymmetric injection to subsonically skew the sonic plane for jet vectoring. The conceptual development of the CFN was presented in a companion paper (Miller et al., 1999). In that study, critical design variables were shown to be the flap length and expansion area ratio of the nozzle, and the location, angle, and distribution of injected flow. Measures of merit were vectoring capability, gross thrust coefficient, and discharge coefficient. A demonstration of MATV was conducted on a 20%-scale CFN test article across a range of nozzle pressure ratios (NPR), injector flow rates, and flow distributions. Both yaw and pitch vector angles of greater than 8° were obtained at NPR of 5.5. Yaw vector angles greater than 10° were achieved at lower NPR. Values of thrust-coefficient for the CFN generally exceeded published measurements of shock-based, vectoring methods. In terms of vectoring effectiveness (ratio of vector angle to percent injected flow), fluidic throat skewing was found to be comparable to shock-based vectoring methods.

scholarly journals Multiparameter Optimization of Thrust Vector Control with Transverse Injection of a Supersonic Underexpanded Gas Jet into a Convergent Divergent Nozzle

Energies ◽  
2021 ◽  
Vol 14 (14) ◽  
pp. 4359
Author(s):  
Vladislav Emelyanov ◽  
Mikhail Yakovchuk ◽  
Konstantin Volkov
Keyword(s):  
Vector Control ◽  
Automatic Generation ◽  
Injection System ◽  
Gas Jet ◽  
Outlet Section ◽  
Thrust Coefficient ◽  
Thrust Vector Control ◽  
Injection Nozzle ◽  
Thrust Vector ◽  
Multiparameter Optimization

The optimal design of the thrust vector control system of solid rocket motors (SRMs) is discussed. The injection of a supersonic underexpanded gas jet into the diverging part of the rocket engine nozzle is considered, and multiparameter optimization of the geometric shape of the injection nozzle and the parameters of jet injection into a supersonic flow is developed. The turbulent flow of viscous compressible gas in the main nozzle and injection system is simulated with the Reynolds-averaged Navier–Stokes (RANS) equations and shear stress transport (SST) turbulence model. An optimization procedure with the automatic generation of a block-structured mesh and conjugate gradient method is applied to find the optimal parameters of the problem of interest. Optimization parameters include the pressure ratio of the injected jet, the angle of inclination of the injection nozzle to the axis of the main nozzle, the distance of the injection nozzle from the throat of the main nozzle and the shape of the outlet section of the injection nozzle. The location of injection nozzle varies from 0.1 to 0.9 with respect to the length of the supersonic part of the nozzle; the angle of injection varies from 30 to 160 degrees; and the shape of the outlet section of the injection nozzle is an ellipse with an aspect ratio that varies from 0.1 to 1. The computed fluid flow in the combustion chamber is compared with experimental and computational results. The dependence of the thrust as a function of the injection parameters is obtained, and conclusions are made about the effects of the input parameters of the problem on the thrust coefficient.

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