مجله مکانیک سیالات و آیرودینامیک
سال دهم شماره 1 (بهار و تابستان 1400)

The precise method for aerodynamic parameters measurement in aeroballistics labs, is the free flight method. It can be used to test different models at a wide range of flight speeds including subsonic, transonic, supersonic and hypersonic speeds, and in the millimetric to metric range dimensions. It can be utilized to measure the path-equation related parameters and the aerodynamic and aeroelastic parameters. Moreover, it has the capability for checking the performance of designed control systems, testing material temperature tolerance, testing structural strength and addressing high-speed collision and re-entry problems. In this lab, stations for measuring flight model data are located at certain distances that extract data using imaging and telemetry techniques. The main goal is to provide a suitable mathematical model with precise accuracy in order to extract aerodynamic coefficients from raw data measured according to the number of measurement stations and their intervals. Therefore, for the purpose of using imaging data, an overview of aerodynamic coefficients extraction in the aeroballistics test is carried out. The results show that static and dynamic coefficients are not independent in extraction and should be calculated simultaneously. Dynamic coefficients cannot be ignored if the magnitudes are significant. The measurement accuracy required when the dynamic coefficients cause the damping of the oscillations is about 0.1 degree with 20 stations, while the accuracy of 0.1 degree with 10 stations is inadequate and higher than that is required.

The purpose of this work is to investigate the effect of magnetic field direction on heat transfer of Newtonian and non-Newtonian fluids in both uniform and non-uniform forms, with the power-law model by using the multiple relaxation time lattice Boltzmann method (MRT-LBM) with written computer code by Fortran language. The natural convection is created in the two-dimensional cavity with lozenge barrier and is examined in three different temperature boundary conditions. The cold wall of the cavity is investigated in three modes: smooth, curved and diagonal. The results show that increasing the Rayleigh number and decreasing the power-law index and the Hatmann number increase the strength of fluid flow and heat transfer. The smooth design of the wall increases the average Nusselt number by about 30%. Placing the barrier at a constant cold temperature increases the average Nusselt number by 20% on average. The effect of the magnetic field is highest for the smooth wall and lowest for the diagonal wall and this effect decreases with increasing the power-law index. In general, an applied non-uniform magnetic field increases the average Nusselt number by about 10% and increases the flow strength. The effect of wall shape and type of magnetic field applied on shear thickening fluid is negligible. Further reduction of flow strength and average Nusselt number is observed by applying a magnetic field horizontally.

In hydrodynamic applications, accurately predicting fluid flows with cavitation is very important. In this regard, prediction of the cavity dimensions and the pressure distribution and the flow dynamics, inside and around the cavity, specifically at the closing point has frequently been under consideration. In this study, cavitating flow around cylindrical projectiles with flat or hemispherical heads is considered numerically. To this end, four turbulence models of k-ε-Realizeable, k-ω Standard, k-ω SST, and GEKO, in combination with the Zwart cavitation model are considered using the Fluent software. Flows with a vast range of cavitation numbers (0.1-1.8) are considered in comparison with the experimental and numerical results of other researchers. Our results show that the last turbulence model proposed by Menter, namely the GEnelarilized-KOmega )GEKO( model, in which two extra equations are solved, predicts the results much better, particularly for higher cavitation numbers.

The main challenges facing the designers of hypersonic flying objects are the significant amount of pressure drag and aerodynamic heating. However, blunt noses are preferred for better heat distribution, but they produce a lot of drag force. Spikes and aero disks are effective tools for reducing drag and heating. In this study, the effects of aerospike geometry on the drag reduction of three types of hemispherical, Trident and HB1 noses, in a hypersonic wind tunnel have been evaluated. Experiments were performed on blunt noses in the two conditions of with and without aerospike at Mach number of 6.4 by measuring the drag force and observing the shock waves using the Schlieren technique. For this study, two hemispherical aerospikes with length to diameter ratios of 1 and 2 were considered. The results show that in all three noses, the aerospike converts a strong bow shock into a weak shock, and followed by a rotational zone or dead air zone, reduces the pressure in front of the nose and hence, leads to less drag exertion. The main cause of drag reduction is the aerospike nose and the optimal spike length to diameter ratio is with the blunt nose geometry. The highest drag reduction of 74.8% is observed for an aero disk with the length to diameter ratio of 2 in the hemispherical blunt nose.

Density flows are one of the most important factors in reducing the useful life of water structures, especially reservoir dams. Therefore, researchers have always been investigating solutions to eliminate these flows and increase the useful life of dams. One of the most practical methods known is to build an obstacle in the path of these currents. In this laboratory-numerical study, the effect of trapezoidal permeable obstacle (filled with 0.5 cm diameter grains) on the head of density flow is investigated and the effects of variables such as discharge, concentration, slope and height of the obstacle are studied. Based on the percentage reduction of the density flow head obtained from the experiments, the concentrated salt flow head is modeled using one of the soft computing methods known as the adaptive neural-fuzzy inference system (Anfis). Then, by comparing the results with the classical multivariable regression method, the performance of these two methods is compared. The error of training, validation and test data for the Anfis method are shown to be 0.07, 0.033 and 0.03, respectively, while for the multivariable regression method the mentioned errors are shown to be 0.12, 0.199 and 0.1084, respectively. Also, regression values for the training and test data for the Anfis method, are found to be 0.9954 and 0.9652 respectively whilst for the multivariable regression method, the mentioned parameters are shown to be 0.93108 and 0.90396 respectively which demonstrates the superiority and the efficiency of the adaptive neural-fuzzy inference system in modeling the head data.

The cargoes in a fighter aircraft should be released in such a way so as not to hit the fuselage, but to hit the targets accurately. In order to identify and eliminate the mentioned problems and achieve the optimal conditions, the computational fluid dynamics (CFD) plays an important role. In this study, the trajectory of the projectile and its angular variations have been investigated such that in the process of the simulation, one of the wings of the Su-22 aircraft with a NACA-64a210 delta-shaped cross-section, along with the pylon, a rectangular plate with two elliptical sides and a cargo with NACA-0008 cross-section with delta-shaped blades have been considered. The study has been carried out for different flight conditions, including three altitudes of 5, 10, and 12 km above the ground, various Mach numbers of 0.5, 0.8, and 1.2, different attack angles (AOA) of 0, 2, 4, and 8 degrees and lateral wind with speeds of 0, 40, and 60 m/s. The purpose of this study is to model and investigate the safe and unsafe release of cargo by the wings of Sukhoi-22 aircraft in a different flight package. This is done in order to find carefully the physical releasing conditions in turbulent compressible flows. Finally, by examining the results, the effect of each of the parameters of flight altitude change, Mach, angle of attack, and lateral wind speed on the projectile movement path in the release conditions without initial force is assessed, to identify all safe and unsafe releases.

In this paper, the thermodynamic modeling and analysis of various components of a turboprop engine and their relation in the steady-state design and off-design performances are addressed. An algorithm is presented and a computer code is developed for the analysis of the cycle in design and off-design conditions. Then off-design performance curves are plotted for different turbine inlet temperatures, flight elevations and Mach numbers and their results are interpreted. The results are compared and validated with the results of GasTurb software. In addition, the effects of the combustion chamber defects on the performance of the engine are investigated. The results of this paper can be used as the first step in the study of the performance and upgrading of these engines. The results of this study show that the specific fuel consumption will rise by 14% for a 10% increase in the combustion chamber pressure drop. Furthermore, the specific fuel consumption increases by 13% and the engine power reduces by 17.5% for a 10% reduction in the combustion efficiency.

In this research the aerodynamic performance of a discrete wing capable of morphing, has been studied based on the numerical techniques. In the proposed wing the chord-wise strips have been implemented to resemble a bird wing’s feathers. The morphing mechanism composed of changing the wrist angle which is equal to the partially sweep angle and the wing planform area. Here it is shown that by utilizing this proposed simple mechanism, the MAV/UAV could enhance its aerodynamic performance and also produce control power in the longitudinal and lateral maneuvers. Firstly, the benchmark geometry and a suitable numerical scheme are introduced. The results are then compared and validated against the reference available data. Consequently, the parametric study is performed by selecting the morphing wing tip sweep angle and the angle of attack as the variables and the force coefficients and efficiency as the performance indices. This kind of morphing could make the efficiency increase, up to 13 percent. Flow simulations around the wing are depicted for various morphing angles and for two angles of attack, both in the steady and unsteady manner. The results are also compared and analyzed. Based on the current research, one may continue to estimate the control forces produced by this simple mechanism to expedite the longitudinal and lateral maneuvers in a specified MAV.

Turbocharging is a prevalent method for the promotion of an UAV flight level without having to make major changes in its engine. It depends on the correct selection and precise control of the turbocharger. Today, the mathematic simulation of the engine cyclic processes as a strong tool to estimate performance and reduce costs and testing time is taken into consideration. In this research multi-zones thermodynamic modeling according to the following steps is performed. At the first step, a geometrical model of the Wankel engine is developed and the geometrical characteristics of the equivalent reciprocating engine is achieved. At the second step, the equivalent reciprocating engine is simulated by GT-Power commercial software. Then, using the empirical results of dynamometer tests the developed model is vitrificated and by this means the effect of altitude conditions on the engine performance is studied. At the last step, according to the matching theories, a proper turbocharger is selected and by using appropriate control mechanisms, the performance of the turbocharged engine at the desired altitude is evaluated. Results indicate that for the operating engine speed and full load condition, turbocharging leads to 41% power increment and 5% specific fuel consumption reduction at the target altitude compared to the naturally aspirated engine at designed working altitude.

In the present study, the thermal performance of the air-side compact heat exchanger of an automotive condenser in the three-dimensional model has been numerically simulated based on the design of experiment (DOE). For this purpose, the effective geometrical parameters such as fin pitch, louvre pitch, louvre angle, fin height and fin width and their effects on condenser thermal performance are studied and investigated. Finally, for parameters effective on the compact heat exchanger operation, such as air-side heat transfer coefficient, static pressure drop and other thermal parameters, relationships are extracted based on the design of experiment which decrease the computational cost with the least error. The results show quantitatively that by increasing the fin length, louver angle, louver pitch and decreasing the fin pitch, the heat transfer rate increases by about 42%. Also, the results show that the heat dissipation, heat transfer coefficient and pressure drop, have improved by 8.1%, 5% and 10.5% respectively compared to recent numerical results. Eventually parametric studies including the effect of ambient temperature, condenser wall temperature and vehicle speed on thermal performance and pressure drop are presented.

One of the combustion methods that has been developed in the last two decades, but has not yet become fully operational and commercial, is the use of a continuous detonation wave with only one initial start-up in the combustion chamber. This method is called rotating detonation combustion or RDC for short. In this article, while introducing rotating detonation combustion, this process is also simulated in two waves. It also presents a method that by considering the physics of the problem, any number of waves can be simulated in the combustion chamber. In this method, by considering the periodic boundary conditions along the wave propagation in two-dimensional mode, the detonation wave becomes continuous and rotational. By setting the appropriate code (UDF: User Defined Function) to apply the boundary condition at the input, we manage the injection into the domain, in such a way that there is always fresh reactant in front of the wave so that the wave continues to move and the injection stops when the wave passes and the pressure increases, Then the injection is resumed after the wave is gone and the pressure drops. In order to propagate the detonation unilaterally before reaching the continuous propagation conditions, the wall boundary condition at the periodic boundaries of the continuous state is used. By applying the appropriate initial conditions, the transition from deflagration to detonation is eliminated and the cost of calculations is reduced. Despite the mentioned simplifications, the model presented in this article demonstrates single-digit error percentage in predicting the detonation wave velocity.

Hydrazine monopropellant thrusters are most widely used for navigation and control systems of re-entry and manned payloads. In this paper, the effect of the internal geometry of the injector on the characteristics of the outlet liquid sheet, such as the liquid sheet thickness, the spray cone angle, the average output velocity, and its mass flow rate, has been studied. The injector chosen for the case study is the fuel injector of a 10N monopropellant hydrazine thruster. This injector was designed in such a way to achieve a medium spray angle and a very small sheet thickness, which is suitable considering the limited length of the catalytic chamber and leads to finer atomization. For this purpose, simulation of the internal flow based on the computational fluid dynamics is performed to predict the output flow characteristics, and then parametric studies are conducted to investigate the effects of geometry. The results of these studies show that the internal structure of the injector has a great influence on the control of the characteristics of the liquid sheet, and the ratios of fluid swirl radius to the nozzle radius/inlet duct radius have the greatest effect on the output spray characteristics, so that minimum average speed is achieved at larger ratios and maximum average speed is obtained at smaller ratios.