A method is discussed which accounts for the effect of changes in Reynolds number on the pressure distributions of transonic wings. The major Reynolds number effect is a change in shock location which accompanies a change in trailing-edge separation. The method discussed here depends on several characteristics of separating transonic flows, each of which is supported by experimental data from a variety of transonic wing tests: (1) Pressure distributions forward of the shock are not affected by trailing-edge separation. (2) When trailing-edge separation occurs, the shock moves forward. The relationship between shock location and trailing-edge pressure is the same whether changes are caused by changes in Reynolds number, changes in transition location or use of separation control devices (BLC, vortex generators, etc.). (3) The combined effects of changes in angle of attack and Mach number on trailing-edge separation are shown to be a function solely of the transonic similarity parameter previously presented by Melnick and Grossman. (4) When correlated in terms of the M-G similarity parameter, the trailing-edge separation data are shown to follow a simple and universal variation with Reynolds number.The method discussed here applies to cases where there is no separation at the shock and scale effects are dominated by conditions at the trailing edge. Additional study is required for those cases where separations exist at both the shock and the trailing edge.