CFD Analysis of Aircraft Curved C-D Nozzle: Technical Note

2019 ◽  
Vol 11 (3) ◽  
Author(s):  
F. Ferdaus ◽  
S. Sivaganesan ◽  
C. Dhanasekaran ◽  
G. Sathishkumar ◽  
S. Sivabalan
Keyword(s):  
Mach Number ◽  
Technical Note ◽  
Cfd Analysis ◽  
Flow Conditions ◽  
Flow Properties ◽  
Subsonic Flows ◽  
Gas Dynamic ◽  
Sonic Flow ◽  
Exit Mach Number ◽  
Gas Dynamic Equations

A nozzle for the aircraft can be designed by considering the exit Mach number. In order to get a premeditated Mach number, we need to convert pressure energy into kinetic energy by using a nozzle. Convergent nozzles are utilized for subsonic flows while Convergent-Divergent (C-D) nozzle is utilized for supersonic flows. Curved nozzle flow is accelerated from low subsonic to sonic velocity at the throat and further expanded to supersonic velocities at the exit, in a C-D nozzle. This paper details the relevancies on designing a curved nozzle to attain super-sonic flow and maximizing the optimal thrust and devoid of flow separation due to shock waves. The navigation of the flow must be parallel to the axis of the nozzle for achieving extreme thrust and proficiency. Based on the fundamental gas dynamic equations, this study aims to develop a theoretical approach for the calculation of the flow properties along the axis of the C-D Nozzle. The flow conditions were selected in consideration of the pressure, temperature and gases accessible at the exit of the combustion chamber.

Computational Analysis of Straight Nozzle: Technical Note

2019 ◽  
Vol 11 (2) ◽  
Author(s):  
F. Ferdaus ◽  
R. Sridhar ◽  
G. Sathishkumar ◽  
S. Sivabalan
Keyword(s):  
Flow Separation ◽  
Supersonic Flows ◽  
Technical Note ◽  
Flow Conditions ◽  
Convergent Nozzle ◽  
Turbine Engine ◽  
Sonic Flow ◽  
Induced Flow ◽  
Exit Mach Number ◽  
Maximum Thrust

Most of the modern aircraft and military aircraft are powered by the modern gas turbine engine. They have nozzles to produce the required speed. Depending upon the required exit Mach number, a nozzle can be designed to be used for subsonic and supersonic flows. For the sonic flows, the convergent nozzle is used and for supersonic flows a convergent–divergent (CD) nozzle is used. In a CD nozzle, a straight nozzle flow is accelerated from low subsonic to sonic velocity at the throat and further expanded to supersonic velocities at the exit. This paper focuses on designing a straight nozzle to attain super-sonic flow and optimizing it to achieve maximum thrust without flow separation due to shock waves. This research also confirms that at which angle of deflection on the divergent portion produces more speed. The flow conditions were selected in view of the pressure, temperature and gases that are accessible at the exit of the combustion chamber. At the exit of the nozzle, the shock induced flow separation due to, over, under and optimum expansion conditions were studied.

Experimental Investigation of Gas Turbine Axial Diffuser Performance: Part I — Parametric Analysis of Influential Variables

2020 ◽  
Author(s):  
Kenneth Brown ◽  
Stephen Guillot ◽  
Wing Ng ◽  
Lee Iksang ◽  
Kim Dongil ◽  
...  
Keyword(s):  
Experimental Investigation ◽  
Mach Number ◽  
Gas Turbines ◽  
Axial Flow ◽  
Flow Distribution ◽  
Initial Study ◽  
Flow Conditions ◽  
Inlet Flow ◽  
Viscous Loss ◽  
Exit Mach Number

Abstract An experimental investigation of the effect of inlet flow conditions and improved geometries on the performance of modern axial exhaust diffusers of gas turbines has been completed. As the first of a two-part series, this article concentrates on characterizing diffuser sensitivity to parametric variations in internal geometry and inlet flow conditions. Full-factorial experiments were carried out on five parameters including the inlet Mach distribution, shape of the support struts, shape of the oil-drain strut, diffuser hade angle, and the hubcap configuration. To enable an efficient sweep of the design space, experiments were performed in this initial study at a down-scaled turbine exit Reynolds number (ReH roughly 3% of the value for an H-class diffuser) and at a full-scale turbine exit Mach number. The study was accomplished in a continuous, cold-flow wind tunnel circuit, and tailored distributions of Mach number, swirl velocity, and radial velocity derived from on-design conditions of an industry diffuser were generated. Measurements included 5-hole probe traverses at planes of interest. Diffuser performance was most sensitive to the inlet Mach distribution with losses of 0.081 points of pressure recovery due to a nonuniform Mach distribution with higher velocity near the hub versus a uniform one. Detailed comparisons of axial flow variation for a top-performing configuration versus related configurations shed physical insight regarding the evolution of kinetic energy distortion into viscous loss in the wake, as well as highlight the benefit of uniform inlet profiles in practice despite the lower theoretical recovery of such cases. The results presented here isolate the inlet flow distribution as a parameter of high interest for further study which is carried out for both on- and off-design conditions in the companion article [1].

Computational and Experimental Testing of Aircraft Bell Nozzle: Technical Note

2019 ◽  
Vol 11 (5) ◽  
Author(s):  
F. Ferdaus ◽  
R. Ganapathi
Keyword(s):  
Mach Number ◽  
Supersonic Flow ◽  
Experimental Testing ◽  
Technical Note ◽  
Aviation Industry ◽  
Raspberry Pi ◽  
Cfd Analysis ◽  
Nozzle Design ◽  
Hypersonic Speed ◽  
Nozzle Pressure

In order to gain the supersonic and hypersonic speed, the nozzle plays a major role in the aviation industry. Based on the nozzle design, the required mach number can be calculated. In this paper, the convergent divergent bell nozzle which is basically used for the supersonic flow is analysed and designed using CATIA Software. The mesh of the designed nozzle is carried out in ANSA and then analysed using CFD. Different nozzle designs are assessed through of CFD analysis to choose the best performing nozzle that can be manufactured for experiments. For the experimental test Raspberry Pi, pressure sensor and Python coding was developed to test bell nozzle pressure.

Numerical Modeling of Heat Transfer and Pressure Losses for Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

2007 ◽  
Author(s):  
Lamyaa A. El-Gabry
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Computational Study ◽  
Numerical Models ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Pressure Losses ◽  
Blade Tip ◽  
Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. CFD predictions of blade tip heat transfer are compared to test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; they are flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25, 2.0, and 2.75% of blade span. The tip heat transfer results of the numerical models agree fairly well with the data and are comparable to other CFD predictions in the open literature.

scholarly journals Aerodynamic Losses in Turbines with and without Film Cooling, as Influenced by Mainstream Turbulence, Surface Roughness, Airfoil Shape, and Mach Number

2012 ◽  
Vol 2012 ◽  
pp. 1-28 ◽  
Author(s):  
Phil Ligrani
Keyword(s):  
Surface Roughness ◽  
Mach Number ◽  
Film Cooling ◽  
Total Pressure ◽  
High Speed ◽  
Density Ratio ◽  
Aerodynamic Loss ◽  
Pressure Losses ◽  
Loss Coefficients ◽  
Exit Mach Number

The influences of a variety of different physical phenomena are described as they affect the aerodynamic performance of turbine airfoils in compressible, high-speed flows with either subsonic or transonic Mach number distributions. The presented experimental and numerically predicted results are from a series of investigations which have taken place over the past 32 years. Considered are (i) symmetric airfoils with no film cooling, (ii) symmetric airfoils with film cooling, (iii) cambered vanes with no film cooling, and (iv) cambered vanes with film cooling. When no film cooling is employed on the symmetric airfoils and cambered vanes, experimentally measured and numerically predicted variations of freestream turbulence intensity, surface roughness, exit Mach number, and airfoil camber are considered as they influence local and integrated total pressure losses, deficits of local kinetic energy, Mach number deficits, area-averaged loss coefficients, mass-averaged total pressure loss coefficients, omega loss coefficients, second law loss parameters, and distributions of integrated aerodynamic loss. Similar quantities are measured, and similar parameters are considered when film-cooling is employed on airfoil suction surfaces, along with film cooling density ratio, blowing ratio, Mach number ratio, hole orientation, hole shape, and number of rows of holes.

Investigation of the 3D Unsteady Rotor Pressure Field in a HP Turbine Stage

2002 ◽  
Author(s):  
E. Valenti ◽  
J. Halama ◽  
R. De´nos ◽  
T. Arts
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Pressure Field ◽  
Pressure Ratio ◽  
Turbine Stage ◽  
Time Resolved ◽  
Mach Number Distribution ◽  
Rotor Surface ◽  
Exit Mach Number ◽  
Unsteady Pressure Measurements

This paper presents steady and unsteady pressure measurements at three span locations (15, 50 and 85%) on the rotor surface of a transonic turbine stage. The data are compared with the results of a 3D unsteady Euler stage calculation. The overall agreement between the measurements and the prediction is satisfactory. The effects of pressure ratio and Reynolds number are discussed. The rotor time-averaged Mach number distribution is very sensitive to the pressure ratio of the stage since the incidence of the flow changes as well as the rotor exit Mach number. The time-resolved pressure field is dominated by the vane trailing edge shock waves. The incidence and intensity of the shock strongly varies from hub to tip due to the radial equilibrium of the flow at the vane exit. The decrease of the pressure ratio attenuates significantly the amplitude of the fluctuations. An increase of the pressure ratio has less significant effect since the change in the vane exit Mach number is small. The effect of the Reynolds number is weak for both the time-averaged and the time-resolved rotor static pressure at mid-span, while it causes an increase of the pressure amplitudes at the two other spans.

Prediction of the Flow and Force Characteristics of Safety Relief Valves

2006 ◽  
Author(s):  
W. Dempster ◽  
C. K. Lee ◽  
J. Deans
Keyword(s):  
Operating Conditions ◽  
Cfd Analysis ◽  
Flow Conditions ◽  
Relief Valve ◽  
Good Correspondence ◽  
Geometry Modeling ◽  
Valve Opening ◽  
Satisfactory Prediction ◽  
Force Characteristics

The design of safety relief valves depends on knowledge of the expected force-lift and flow-lift characteristics at the desired operating conditions of the valve. During valve opening the flow conditions change from seal-leakage type flows to combinations of sub-sonic and supersonic flows It is these highly compressible flow conditions that control the force and flow lift characteristics. This paper reports the use of computational fluid dynamics techniques to investigate the valve characteristics for a conventional spring operated 1/4” safety relief valve designed for gases operating between 10 and 30 bar. The force and flow magnitudes are highly dependent on the lift and geometry of the valve and these characteristics are explained with the aid of the detailed information available from the CFD analysis. Experimental determination of the force and flow lift conditions has also been carried out and a comparison indicates good correspondence between the predictions and the experiment. However, attention requires to be paid to specific aspects of the geometry modeling including corner radii and edge chamfers to ensure satisfactory prediction.

Thin Film Coatings of WO3 by Cold Gas Dynamic Spray: A Technical Note

2005 ◽  
Vol 14 (2) ◽  
pp. 183-186 ◽  
Author(s):  
Ha Yong Lee ◽  
Young Ho Yu ◽  
Young Cheol Lee ◽  
Young Pyo Hong ◽  
Kyung Hyun Ko
Keyword(s):  
Thin Film ◽  
Technical Note ◽  
Film Coatings ◽  
Thin Film Coatings ◽  
Gas Dynamic ◽  
Cold Gas Dynamic Spray ◽  
Gas Dynamic Spray ◽  
Cold Gas

Loss Generation in Transonic Turbine Blading

2017 ◽  
Author(s):  
Penghao Duan ◽  
Choon S. Tan ◽  
Andrew Scribner ◽  
Anthony Malandra
Keyword(s):  
Experimental Data ◽  
Boundary Layer ◽  
Mach Number ◽  
High Speed ◽  
Turbine Blades ◽  
Surface Boundary ◽  
Loss Model ◽  
Exit Mach Number ◽  
Loss Characteristic ◽  
Shock Loss

The measured loss characteristic in a high-speed cascade tunnel of two turbine blades of different designs showed distinctly different trend with exit Mach number ranging from 0.8 to 1.4. Assessments using steady RANS computation of the flow in the two turbine blades, complemented with control volume analyses and loss modelling, elucidate why the measured loss characteristic looks the way it is. The loss model categorizes the total loss in terms of boundary layer loss, trailing edge loss and shock loss; it yields results in good agreement with the experimental data as well as steady RANS computed results. Thus RANS is an adequate tool for determining the loss variations with exit isentropic Mach number and the loss model serves as an effective tool to interpret both the computational and experimental data. The measured loss plateau in Blade 1 for exit Mach number of 1 to 1.4 is due to a balance between a decrease of blade surface boundary layer loss and an increase in the attendant shock loss with Mach number; this plateau is absent in Blade 2 due to a greater rate in shock loss increase than the corresponding decrease in boundary layer loss. For exit Mach number from 0.85 to 1, the higher loss associated with shock system in Blade 1 is due to the larger divergent angle downstream of the throat than that in Blade 2. However when exit Mach number is between 1.00 and 1.30, Blade 2 has higher shock loss. For exit Mach number above around 1.4, the shock loss for the two blades is similar as the flow downstream of the throat is completely supersonic. In the transonic to supersonic flow regime, the turbine design can be tailored to yield a shock pattern the loss of which can be mitigated in near equal amount of that from the boundary layer with increasing exit Mach number, hence yielding a loss plateau in transonic-supersonic regime.

Mathematical Analysis for Off-Design Performance of Cryogenic Turboexpander

2011 ◽  
Vol 133 (3) ◽  
Author(s):  
Subrata K. Ghosh ◽  
R. K. Sahoo ◽  
Sunil K. Sarangi
Keyword(s):  
Pressure Ratio ◽  
Expansion Ratio ◽  
Flow Conditions ◽  
Flow Properties ◽  
Dimensional Solution ◽  
One Dimensional ◽  
Design Performance ◽  
Fluid Properties ◽  
The Mean ◽  
Inlet Conditions

A study has been conducted to determine the off-design performance of cryogenic turboexpander. A theoretical model to predict the losses in the components of the turboexpander along the fluid flow path has been developed. The model uses a one-dimensional solution of flow conditions through the turbine along the mean streamline. In this analysis, the changes of fluid and flow properties between different components of turboexpander have been considered. Overall, turbine geometry, pressure ratio, and mass flow rate are input information. The output includes performance and velocity diagram parameters for any number of given speeds over a range of turbine pressure ratio. The procedure allows any arbitrary combination of fluid species, inlet conditions, and expansion ratio since the fluid properties are properly taken care of in the relevant equations. The computational process is illustrated with an example.

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