Stagnation temperature effect on the supersonic flow around pointed airfoils with application for air

2018 ◽  
Vol 19 (3) ◽  
pp. 312
Author(s):  
Rahima Takhnouni ◽  
Toufik Zebbiche ◽  
Abderrazak Allali
Keyword(s):  
High Temperature ◽  
Mach Number ◽  
Supersonic Flow ◽  
Incidence Angle ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Oblique Shock Wave ◽  
Stagnation Temperature ◽  
Perfect Gas ◽  
Application Limit

The aim of this work is to develop a new numerical calculation program to determine the effect of the stagnation temperature on the calculation of the supersonic flow around a pointed airfoils using the equations for oblique shock wave and the Prandtl Meyer expansion, under the model at high temperature, calorically imperfect and thermally perfect gas, lower than the dissociation threshold of the molecules. The specific heat at constant pressure does not remain constant and varies with the temperature. The new model allows making corrections to the perfect gas model designed for low stagnation temperature, low Mach number, low incidence angle and low airfoil thickness. The stagnation temperature is an important parameter in our model. The airfoil should be pointed at the leading edge to allow an attached shock solution to be seen. The airfoil is discretized into several panels on the extrados and the intrados, placed one adjacent to the other. The distribution of the flow on the panel in question gives a compression or an expansion according to the deviation of the flow with respect to the old adjacent panel. The program determines all the aerodynamic characteristics of the flow and in particular the aerodynamic coefficients. The calculation accuracy depends on the number of panels considered on the airfoil. The application is made for high values of stagnation temperature, Mach number and airfoil thickness. A comparison between our high temperature model and the perfect gas model is presented, in order to determine an application limit of the latter. The application is for air.

Experimental Investigation on the Annular Sector Cascade of a High Endwall-Angle Turbine

2016 ◽  
Author(s):  
Weiliang Fu ◽  
Jie Gao ◽  
Chen Liang ◽  
Fukai Wang ◽  
Qun Zheng ◽  
...  
Keyword(s):  
Mach Number ◽  
Static Pressure ◽  
Incidence Angle ◽  
Pressure Sensors ◽  
Leading Edge ◽  
Control Program ◽  
Turbine Cascade ◽  
Cascade Flow ◽  
Annular Sector ◽  
The Way

The flow in high endwall-angle turbine is complex, and it is different from the ordinary turbine flow in characteristics. In order to study the flow field characteristics of high endwall-angle turbines, the annular sector cascade experimental study of high endwall-angle turbines is carried out. The blade is studied experimentally in the form of annular sector cascade. The cascade includes 7 blades, and makes up 6 flow passages, in order to simulate full cascade flow. The experimental Mach number is adjusted by the way of changing inlet total pressure, and the Mach number influence (0.7, 0.8 and 0.9) on annular sector cascade flow is studied. Based on it, the inlet incidence angle (−15°, −7.5°, 0°, 7.5° and 15° )is changed with the way of changing sector straight pipes upstream of the cascade, and its influence on turbine flow fields is studied at the Mach number of 0.8. Here, five-hole probes are used to measure aerodynamic parameters distributions downstream of the cascade, and static pressure taps are positioned on the blade surface to measure surface static pressure distribution. The auto-traversing system and pressure sensors were operated by a self-compiled program based control program. The results indicate that there are two passage vortices inside the turbine cascade flow passage under the high Mach number condition, and the passage vortex near the high endwall-angle region is bigger. As Mach number increases, the passage vortices inside turbine cascade passage will become strong, and moves towards the blade mid-span. Besides, it is shown that the way of changing sector straight pipes can achieve the variation of inlet incidence angles. And, the blade profile with big leading-edge radius has good design and off-design performance. Detailed results and analyses are presented in the paper.

scholarly journals Unsteady Pressure Measurements on a Biconvex Airfoil in a Transonic Oscillating Cascade

1986 ◽  
Vol 108 (1) ◽  
pp. 53-59 ◽  
Author(s):  
L. M. Shaw ◽  
D. R. Boldman ◽  
A. E. Buggele ◽  
D. H. Buffum
Keyword(s):  
Mach Number ◽  
Incidence Angle ◽  
Dynamic Pressure ◽  
Leading Edge ◽  
Pressure Transducers ◽  
High Incidence ◽  
Shock Motion ◽  
Plate Analysis ◽  
Phase Angles ◽  
Unsteady Pressure Measurements

Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfoils were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence angle of 0.0 deg and a Mach number of 0.65 in order to obtain results in a shock-free compressible flow field. Subsequent tests were performed at an angle of attack of 7.0 deg and a Mach number of 0.80 in order to observe the surface pressure response with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and −90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semichord reduced frequencies up to about 0.5 at a Mach number of 0.80). Results from the zero-incidence cascade are compared with a classical unsteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.

scholarly journals Calculation of aerodynamic characteristics of flying objects using Prodas and Fluent environments

Mechanik ◽  
2017 ◽  
Vol 90 (7) ◽  
pp. 591-593
Author(s):  
Leszek Baranowski ◽  
Michał Frant
Keyword(s):  
Mach Number ◽  
Experimental Method ◽  
Drag Force ◽  
Lift Force ◽  
Incidence Angle ◽  
Theoretical Method ◽  
Aerodynamic Characteristics ◽  
Pitching Moment ◽  
Force Coefficient ◽  
Moment Coefficient

The article presents the methodology of determining the basic aerodynamic characteristics using the Fluent theoretical method and the theoretical and experimental method using the Prodas program. Presented calculations were made for a 122 mm non-guided missile. In order to compare both methods, the results of calculations of coefficient of drag force, lift force coefficient and pitching moment coefficient as a function of incidence angle of attack and Mach number are shown in graphs.

scholarly journals Aerodynamic Study of the NASA’s X-43A Hypersonic Aircraft

2020 ◽  
Vol 10 (22) ◽  
pp. 8211
Author(s):  
Àlex Navó ◽  
Josep M. Bergada
Keyword(s):  
Angle Of Attack ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Oblique Shock ◽  
Force Balance ◽  
Oblique Shock Wave ◽  
Vertical Force ◽  
Pitching Moment ◽  
Cfd Model ◽  
Hypersonic Aircraft

A 2D aerodynamic study of the NASA’s X-43A hypersonic aircraft is developed using two different approaches. The first one is analytical and based on the resolution of the oblique shock wave and Prandtl–Meyer expansion wave theories supported by an in-house program and considering a simplified aircraft’s design. The second approach involves the use of a Computational Fluid Dynamics (CFD) package, OpenFOAM and the real shape of the aircraft. The aerodynamic characteristics defined as the lift and drag coefficients, the aerodynamic efficiency and the pitching moment coefficient are calculated for different angles of attack. Evaluations are made for an incident Mach number of 7 and an altitude of 30 km. For both methodologies, the required angles of attack to achieve a Vertical Force Balance (VFB) and a completely zero pitching moment conditions are considered. In addition, an analysis to optimise the nose configuration of the aircraft is performed. The mass flow rate throughout the scramjet as a function of the angle of attack is also presented in the CFD model in addition to the pressure, density, temperature and Mach fields. Before presenting the corresponding results, a comparison between the aerodynamic coefficients in terms of the angle of attack of both models is carried out in order to properly validate the CFD model. The paper clarifies the requirements needed to make sure that both oblique shock waves originating from the leading edge meet just at the scramjet inlet clarifying the advantages of fulfilling such condition.

Iterative Solutions of the Equations Jor Plane Oblique Shock Waves

1959 ◽  
Vol 63 (587) ◽  
pp. 669-672 ◽  
Author(s):  
A. R. Collar
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Supersonic Flow ◽  
Free Stream ◽  
Oblique Shock ◽  
Oblique Shock Wave ◽  
Two Dimensional ◽  
Iterative Solutions ◽  
Flow Through ◽  
Image Position

If a plane oblique shock wave, inclined to the free stream at the angle ε, is produced in two-dimensional supersonic flow of Mach number M by (for example) a wedge which deflects the flow through an angle δ, the equation connecting these quantities may be writtenIn this form, δ is given explicitly when M, ε are fixed. Similarly, we may obtain M explicitly when ε, δ are fixed; equation (1) may be written (see, for example, Liepmann and Puckett, Equation 4.27)

The Effects of Inlet Reynolds Number, Exit Mach Number and Incidence Angle on Leading Edge Film Cooling Effectiveness of a Turbine Blade in a Linear Transonic Cascade

2015 ◽  
Author(s):  
Cong Liu ◽  
Hui-ren Zhu ◽  
Zhong-yi Fu ◽  
Run-hong Xu
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Film Cooling ◽  
Incidence Angle ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Exit Mach Number

This paper experimentally investigates the film cooling performance of a leading edge with three rows of film holes on an enlarged turbine blade in a linear cascade. The effects of blowing ratio, inlet Reynolds number, isentropic exit Mach number and off-design incidence angle (i<0°) are considered. Experiments were conducted in a short-duration transonic wind tunnel which can model realistic engine aerodynamic conditions and adjust inlet Reynolds number and exit Mach number independently. The surface film cooling measurements were made at the midspan of the blade using thermocouples based on transient heat transfer measurement method. The changing of blowing ratio from 1.7 to 3.3 leads to film cooling effectiveness increasing on both pressure side and suction side. The Mach number or Reynolds number has no effect on the film cooling effectiveness on pressure side nearly, while increasing these two factors has opposite effect on film cooling performance on suction side. The increasing Mach number decreases the film cooling effectiveness at the rear region mainly, while at higher Reynolds number condition, the whole suction surface has significantly higher film cooling effectiveness because of the increasing cooling air mass flow rate. When changing the incidence angle from −15° to 0°, the film cooling effectiveness of pressure side decreases, and it presents the opposite trend on suction side. At off-design incidence of −15° and −10°, there is a low peak following the leading edge on the pressure side caused by the separation bubble, but it disappears with the incidence and blowing ratio increased.

Cavity Flame Holding for High Speed Reacting Flows

2010 ◽  
Author(s):  
Onur Tuncer
Keyword(s):  
Mach Number ◽  
High Speed ◽  
Fuel Injection ◽  
Transport Model ◽  
Leading Edge ◽  
Reacting Flows ◽  
Stagnation Temperature ◽  
Detailed Chemical Kinetics ◽  
Flame Holder ◽  
Shear Stress Transport Model

Combustion phenomena in a ramjet combustor with cavity flame-holder is studied numerically. Combustor follows a constant area isolator and comprises of hydrogen fuel injected sonically upstream of the cavity. Secondary fuel injection is performed at the cavity backwall. A diverging section follows the cavity to prevent thermal choking. These concepts are also utilized in practice. Calculations were performed for an entrance Mach number of 1.4. Stagnation temperature is 702 K, corresponding to a flight Mach number of 3.3 at an altitude of 12.5 km. Detailed chemical kinetics are taken into account with a reaction mechanism comprising of 9 species and 25 reaction steps. Turbulence is modeled using Menter’s k–ω shear stress transport model, which is appropriate for high speed internal flows. It is observed that flame anchors at the leading edge of the cavity, and the flame is stabilized in the cavity mode rather than the jet-wake mode. Numerical simulation captures all the essential features of the reacting flowfield.

scholarly journals Plasma-Assisted Control of Supersonic Flow over a Compression Ramp

Aerospace ◽  
2019 ◽  
Vol 6 (3) ◽  
pp. 35 ◽  
Author(s):  
Yasumasa Watanabe ◽  
Alec Houpt ◽  
Sergey Leonov
Keyword(s):  
Supersonic Flow ◽  
High Speed ◽  
Leading Edge ◽  
Pressure Change ◽  
Discharge Voltage ◽  
Stagnation Temperature ◽  
Plasma Power ◽  
Supersonic Wind Tunnel ◽  
Compression Ramp ◽  
Plasma Control

This study considers the effect of an electric discharge on the flow structure near a 19.4° compression ramp in Mach-2 supersonic flow. The experiments were conducted in the supersonic wind tunnel SBR-50 at the University of Notre Dame. The stagnation temperature and pressure were varied in a range of 294–600 K and 1–3 bar, respectively, to attain various Reynolds numbers ranging from 5.3 × 105 to 3.4 × 106 based on the distance between the exit of the Mach-2 nozzle and the leading edge of the ramp. Surface pressure measurements, schlieren visualization, discharge voltage and current measurements, and plasma imaging with a high-speed camera were used to evaluate the plasma control authority on the ramp pressure distribution. The plasma being generated in front of the compression ramp shifted the shock position from the ramp corner to the electrode location, forming a flow separation zone ahead of the ramp. It was found that the pressure on the compression surface reduced almost linearly with the plasma power. The ratio of pressure change to flow stagnation pressure was also an increasing function of the ratio of plasma power to enthalpy flux, indicating that the task-related plasma control effectiveness ranged from 17.5 to 25.

Assessment of the Academic CFD Code Galatea-I With the DLR-F11 Model in High Lift Configuration

2016 ◽  
Author(s):  
Sotirios S. Sarakinos ◽  
Georgios N. Lygidakis ◽  
Ioannis K. Nikolos
Keyword(s):  
Mach Number ◽  
Incidence Angle ◽  
Three Dimensional ◽  
Leading Edge ◽  
Navier Stokes ◽  
Test Case ◽  
High Lift ◽  
Low Mach Number ◽  
Flow Phenomena ◽  
Number Flow

In this study the development and assessment of an academic CFD (Computational Fluid Dynamics) code, named Galatea-I, is reported. The proposed solver employs the RANS (Reynolds-Averaged Navier-Stokes) approach, modified by the artificial compressibility method, along with the SST (Shear Stress Transport) turbulence model to predict steady or unsteady turbulent incompressible flow phenomena on three-dimensional unstructured hybrid grids, composed of prismatic, tetrahedral and pyramidal elements. Parallel processing and an agglomeration multigrid method have been included for the acceleration of the solver’s methodologies. Galatea-I is evaluated against a test case of the HiLiftPW-2 (Second High Lift Prediction Workshop). In particular, the low Mach number flow at 7° incidence angle over the DLR-F11 aircraft configuration of Case 1 of the aforementioned workshop was examined; it considers a three-element wing with a leading edge slat and a trailing edge flap attached on a body pod, without including though any of the support brackets used in the wind tunnel experiments. The obtained results are close to the available experimental data, as well as the numerical results of other reference solvers, indicating the proposed methodology’s potential to predict accurately such low Mach number flows over complex geometries.

The representation of GT component maps using Mach numbers

2007 ◽  
Vol 111 (1121) ◽  
pp. 453-460
Author(s):  
V. E. Kyritsis ◽  
P. Pilidis ◽  
K. Ramsden
Keyword(s):  
Chemical Composition ◽  
Mach Number ◽  
Mass Flow ◽  
Rotational Speed ◽  
Pressure Ratio ◽  
Working Fluid ◽  
Stagnation Temperature ◽  
Perfect Gas ◽  
Full Account ◽  
Mach Numbers

Abstract Component maps are produced under certain environmental conditions using air as the working fluid during static ground operation. Any changes of the component characteristics when operating under different temperature conditions and/or with different working fluid are partially taken into account, because of the existence of the gas constant and the ratio of the specific heats in the non-dimensional mass flow and rotational speed. This provides a second order correction for the component characteristics, which may be adequate for the initial modeling of engines. However, for rigorous performance calculations correction factors are applied to the non-dimensional mass flow, rotational speed and pressure ratio distributions of a map, when deviations from the reference conditions under which it was extracted, are experienced. In the current study, a different approach is considered in order to eliminate the inaccuracies caused by the varying temperature and chemical composition. It makes direct use of inlet and circumferential Mach numbers based upon stagnation temperature in conjunction with dimensionless enthalpy variation. A sensitivity analysis against gas property variations is conducted to quantify the benefits gained in precision. Generally, the well-known relationships correlating the Mach number with total and static properties are based on the assumption of perfect gas and constant gas properties. Introducing dependency on temperature and/or chemical composition for the caloric properties of the semi-perfect gas, proper mean values are defined and some theoretical corrections are provided for the well-known equations. The mass flow compatibility equation is then based on the ‘corrected’ expression correlating dimensionless mass flow and Mach number and takes full account of gas property variations.

Sign in / Sign up

Export Citation Format

Share Document