You're confusing static, dynamic and total pressures. Static pressure is the pressure of a fluid on a body when the body is at rest relative to the fluid. Dynamic pressure is the velocity created pressure. Total pressure is the sum of the two, and is what is used in the ideal gas law. To compute lift force static pressure is what is integrated around the wing surface. Total pressure remains constant in the fluid flow for low Mach numbers. Static pressure can and absolutely does change significantly as it accelerates through a streamline such as in low-speed aerodynamics. I understand the semantics on the different kinds of pressure can be confusing. But you should know that when aerodynamics refers to "pressure" as it applies to lift generation, they are referring to static pressure.
I should also mention that this pressure absolutely does change a lot over the flow field, and is commonly used to experimentally and mathematically quantify lift. The following is an image of the pressure distribution of a NACA 2412 airfoil at low speeds.
Just to explain the chart a little bit, in aerodynamics, pressure is usually simplified to a Pressure Coefficient (CP) value. A CP of 0 is when static pressure equals atmosphere. A CP value of 1 occurs at the stagnation point (where velocity is 0, therefore static pressure equals total pressure). Note how this type of chart has an inverted y-axis (a common convention so that the wing upper surface is at the top). Notice how the static pressure on the lower surface is roughly atmospheric, while the upper surface pressure suction peak is high. In this case roughly equal in magnitude to the dynamic pressure. This is a typical pressure distribution for most airfoils, with the suction peak increasing in magnitude as angle of attack increases.
This plot can be obtained mathematically using some sort of potential flow scheme (see: XFOIL for 2D airfoils), or experimentally using pressure taps on a wind tunnel model. The area between the upper and lower surface curves is directly proportional to lift. The larger the difference between upper and lower surfaces, the more lift.
I should also mention that this pressure absolutely does change a lot over the flow field, and is commonly used to experimentally and mathematically quantify lift. The following is an image of the pressure distribution of a NACA 2412 airfoil at low speeds.
https://www.chegg.com/homework-help/questions-and-answers/n-...
Just to explain the chart a little bit, in aerodynamics, pressure is usually simplified to a Pressure Coefficient (CP) value. A CP of 0 is when static pressure equals atmosphere. A CP value of 1 occurs at the stagnation point (where velocity is 0, therefore static pressure equals total pressure). Note how this type of chart has an inverted y-axis (a common convention so that the wing upper surface is at the top). Notice how the static pressure on the lower surface is roughly atmospheric, while the upper surface pressure suction peak is high. In this case roughly equal in magnitude to the dynamic pressure. This is a typical pressure distribution for most airfoils, with the suction peak increasing in magnitude as angle of attack increases.
This plot can be obtained mathematically using some sort of potential flow scheme (see: XFOIL for 2D airfoils), or experimentally using pressure taps on a wind tunnel model. The area between the upper and lower surface curves is directly proportional to lift. The larger the difference between upper and lower surfaces, the more lift.