Abstract: In this paper multi-objective genetic algorithms are
employed for Pareto approach optimization of ideal Turboshaft
engines. In the multi-objective optimization a number of conflicting
objective functions are to be optimized simultaneously. The
important objective functions that have been considered for
optimization are specific thrust (F/m& 0), specific fuel consumption
( P S ), output shaft power 0 (& /&) shaft W m and overall efficiency( ) O
η .
These objectives are usually conflicting with each other. The design
variables consist of thermodynamic parameters (compressor pressure
ratio, turbine temperature ratio and Mach number).
At the first stage single objective optimization has been
investigated and the method of NSGA-II has been used for multiobjective
optimization. Optimization procedures are performed for
two and four objective functions and the results are compared for
ideal Turboshaft engine. In order to investigate the optimal
thermodynamic behavior of two objectives, different set, each
including two objectives of output parameters, are considered
individually. For each set Pareto front are depicted. The sets of
selected decision variables based on this Pareto front, will cause the
best possible combination of corresponding objective functions.
There is no superiority for the points on the Pareto front figure,
but they are superior to any other point. In the case of four objective
optimization the results are given in tables.
Abstract: LES with mixed subgrid-scale model has been used to
simulate aerodynamic performance of hypersonic configuration. The
simulation was conducted to replicate conditions and geometry of a
model which has been previously tested. LES Model has been
successful in predict pressure coefficient with the max error 1.5%
besides afterbody. But in the high Mach number condition, it is poor in
predict ability and product 12.5% error. The calculation error are
mainly conducted by the distribution swirling. The fact of poor ability
in the high Mach number and afterbody region indicated that the
mixed subgrid-scale model should be improved in large eddied
especially in hypersonic separate region. In the condition of attach and
sideslip flight, the calculation results have waves. LES are successful
in the prediction the pressure wave in hypersonic flow.
Abstract: The aerodynamics characteristics of a blended-wing body (BWB) aircraft were obtained in Universiti Teknologi MARA low speed wind tunnel. The scaled-down of BWB model consisted of a canard as its horizontal stabilizer. There were four canards with different aspect ratio used in the experiments. Canard setting angles were varied from -20q to 20q. All tests were conducted at velocity of 35 m/s, with Mach number 0.1. At low angles of attacks, the increment of lift slope for various canards aspect ratio is small and almost constant. Higher canard aspect ratio will cause higher drag. However, canard has a high effect to the moment at zero lift, CM,0.The visualization using mini tuff was performed to observe the airflow at the upper surface of canard. KeywordsAerodynamics,blended-wing body, canard, wind tunnel.
Abstract: The aim of this paper is to develop a new two dimensional time accurate Euler solver for shock tube applications. The solver was developed to study the performance of a newly built short-duration hypersonic test facility at Universiti Tenaga Nasional “UNITEN" in Malaysia. The facility has been designed, built, and commissioned for different values of diaphragm pressure ratios in order to get wide range of Mach number. The developed solver uses second order accurate cell-vertex finite volume spatial discretization and forth order accurate Runge-Kutta temporal integration and it is designed to simulate the flow process for similar driver/driven gases (e.g. air-air as working fluids). The solver is validated against analytical solution and experimental measurements in the high speed flow test facility. Further investigations were made on the flow process inside the shock tube by using the solver. The shock wave motion, reflection and interaction were investigated and their influence on the performance of the shock tube was determined. The results provide very good estimates for both shock speed and shock pressure obtained after diaphragm rupture. Also detailed information on the gasdynamic processes over the full length of the facility is available. The agreements obtained have been reasonable.
Abstract: In the present paper, a numerical investigation has
been carried out to classify and clarify the effects of paramount
parameters on turbulent impinging slot jets. The effects of nozzle-s
exit turbulent intensity, distance between nozzle and impinging plate
are studied at Reynolds number 5000 and 20000. In addition, the
effect of Mach number that is varied between 0.3-0.8 at a constant
Reynolds number 133000 is investigated to elucidate the effect of
compressibility in impinging jet upon a flat plate. The wall that is
located at the same level with nozzle-s exit confines the flow. A
compressible finite volume solver is implemented for simulation the
flow behavior. One equation Spalart-Allmaras turbulent model is
used to simulate turbulent flow at this study. Assessment of the
Spalart-Allmaras turbulent model at high nozzle to plate distance,
and giving enough insights to characterize the effect of Mach number
at high Reynolds number for the complex impinging jet flow are the
remarkable results of this study.
Abstract: A full six degrees of freedom (6-DOF) flight dynamics
model is proposed for the accurate prediction of short and long-range
trajectories of high spin and fin-stabilized projectiles via atmospheric
flight to final impact point. The projectiles is assumed to be both rigid
(non-flexible), and rotationally symmetric about its spin axis launched
at low and high pitch angles. The mathematical model is based on the
full equations of motion set up in the no-roll body reference frame and
is integrated numerically from given initial conditions at the firing
site. The projectiles maneuvering motion depends on the most
significant force and moment variations, in addition to wind and
gravity. The computational flight analysis takes into consideration the
Mach number and total angle of attack effects by means of the
variable aerodynamic coefficients. For the purposes of the present
work, linear interpolation has been applied from the tabulated database
of McCoy-s book. The developed computational method gives
satisfactory agreement with published data of verified experiments and
computational codes on atmospheric projectile trajectory analysis for
various initial firing flight conditions.
Abstract: This research paper presents the CFD analysis of
oscillating airfoil during pitch cycle. Unsteady subsonic flow is
simulated for pitching airfoil at Mach number 0.283 and Reynolds
number 3.45 millions. Turbulent effects are also considered for this
study by using K-ω SST turbulent model. Two-dimensional unsteady
compressible Navier-Stokes code including two-equation turbulence
model and PISO pressure velocity coupling is used. Pressure based
implicit solver with first order implicit unsteady formulation is used.
The simulated pitch cycle results are compared with the available
experimental data. The results have a good agreement with the
experimental data. Aerodynamic characteristics during pitch cycles
have been studied and validated.
Abstract: The present work describes a computational study of
aerodynamic characteristics of GLC305 airfoil clean and with 16.7
min ice shape (rime 212) and 22.5 min ice shape (glaze 944).The
performance of turbulence models SA, Kε, Kω Std, and Kω SST
model are observed against experimental flow fields at different
Mach numbers 0.12, 0.21, 0.28 in a range of Reynolds numbers
3x106, 6x106, and 10.5x106 on clean and iced aircraft airfoil
GLC305. Numerical predictions include lift, drag and pitching
moment coefficients at different Mach numbers and at different angle
of attacks were done. Accuracy of solutions with respect to the
effects of turbulence models, variation of Mach number, initial
conditions, grid resolution and grid spacing near the wall made the
study much sensitive. Navier Stokes equation based computational
technique is used. Results are very close to the experimental results.
It has seen that SA and SST models are more efficient than Kε and
Kω standard in under study problem.