Chapter 3 Results and Discussion
3.4 Aerodynamics Simulation with Different Angle of Attack
3.4.1 Flow Conditions and Simulation Conditions
Table 6 is the flow conditions choosing from the 3DOF simulation results, Figure
3.1. When the rocket is flying with subsonic speed, the momentum of the axial
direction is not so big, comparing with the force resulting from the side wind. We have to simulate larger angle of attack in the subsonic cases. When the rocket is stably flying with super sonic speed, we could just simulate the small angle of attack. In cases of Ma=0.2, we simulate cases of A. Ao. .=0°,2°. In cases of Ma=0.5, we simulate cases of A. Ao. .=0°,2°,4°. In transonic cases of Ma=0.9 and 1.1, we simulate two angle of attack, A. Ao. .=0°,2°. In supersonic cases of Ma=1.5 and 2.5, we simulate angle of attack, A. Ao. .=0°,1°.
3.4.2 Results in Difference Angle of Attack
3.4.2.1 Density, Pressure and Mach Number Distributions
The density, pressure and Mach number distributions are shown in Figure 3.8 to
Figure 3.21. First, we can easily observe the shock wave in cases of Ma=2.5 and 1.5.
In cases of A.o.A.=0 degree, the density, pressure and Mach number distributions near the rocket surface are axial symmetry. In cases with A.o.A. equal to 1, 2 and 4 degrees, there are higher density, pressure and lower velocity in the windward side of wall than the leeward of wall.
3.4.2.2 Axial-Force Coefficients
We can observe that there are small differences between difference attack angles because the cases with less attack angle have less effect on axial-force. In supersonic
cases, the difference between the axial-force coefficients using turbulent flow model and the reference axial-force coefficients are less than 3%, this can be use to validate our numerical code. The numerical results with difference angle of attack have not large difference because our angle of attack is small. As the attack angles become large, the axial-force coefficients become larger but not a big mount. In transonic cases, the difference between the axial-force coefficients using turbulent flow model and the reference axial-force coefficients are in the range of 0% to 50%. The transonic flow is hard to predict so the difference is acceptable. In the cases of Ma=0.5, H=0 and 5000 meter, the difference between the axial-force coefficients using turbulent flow model and the reference axial-force coefficients are less than 20%. As the attack angles become large, the axial-force coefficients become larger, too. In the cases of Ma=0.2, H=24000 meter, the difference is much bigger the others because the Reynolds number of this cases are less than Re , but the difference cr using laminar flow model is still mot small, and so as the cases of Ma=0.2. I think it could have something wrong with my setting to simulate these cases.
3.4.2.3 Normal-Force Coefficients
The sounding rocket is plane symmetry. If the angle of attack of flow field is zero, the normal-force acting on the body surface should be zero. In the case of Ma=2.5,
is less than 1%. This is a very accurate simulation. In the cases of Ma=1.5 and 1.1, the difference between numerical result using turbulent flow model and reference data is about 15%, but it is still acceptable. In the case of Ma=0.9, the difference is much bigger, because the transonic cases are hard to predict. In the cases of subsonic, the differences between numerical results using turbulent flow model and reference data are all less than 8%, but the differences between numerical results using laminar flow model and reference data are in the range of 15% to 30%.
3.4.2.4 Pitching-Moment Coefficients
The sounding rocket is plane symmetry. If the angle of attack of flow field is zero, the pitching acting on the body surface should be zero. In the supersonic cases, the difference between pitching-moment coefficient using turbulent flow model and the reference data are in the range of 5% to 15%. In the cases of transonic cases, because it is hard to predict the flow field in transonic cases, the difference between pitching-moment coefficient using turbulent flow model and the reference data are in the range of 15% to 40%. In the cases of subsonic cases, the difference between pitching-moment coefficient using turbulent flow model and the reference data are in the range of 0% to about 10%, but the difference between pitching-moment coefficient using laminar flow model and the reference data are in the range of 15% to 30%.
3.4.2.5 Location of Pressure Center
The location of pressure center determines the stability of the flying sounding rocket. We get it by taking the top of the nose of sounding rocket as the original point.
In case of Ma=2.5, the location of pressure center is nearer the top of nose than other cases. The difference between the location of pressure center using turbulent flow model with the reference data is 5.48%. In cases of Ma=0.9, 1.1 and 1.5, the difference between the location of pressure center using turbulent flow model with the reference data are less than 4%. In the cases of supersonic and transonic using turbulent flow model, when the velocity becomes larger, the location of pressure center becomes smaller. It means that the rocket is more stable in transonic than in supersonic in the regime of Ma=0.9 to 2.5. In cases of subsonic, the difference between the location of pressure center using both flow model with the reference data are less than 5%. Because Xcp =Cm Cn, and the normal-force coefficients and the pitching-moment coefficients we discuss previously are not very close to the reference data, that the differences are less than 5% is not very reliable.