This paper deals with the high altitude simulation and testing of higher area ratio rocket nozzles at sea level, by adopting the Second Throat supersonic Exhaust Diffuser (STED) configuration. To evaluate the performance of such nozzles, the low-pressure environment of the flight situation has to be simulated in the ground testing installation, using supersonic exhaust diffusers. The diffuser uses the kinetic energy of rocket exhaust itself to maintain lower nozzle back pressure through a benign lambda shock structure, which isolates the vacuum chamber from the ambient. Cold flow experiments were conducted with a nozzle area ratio of 70, using the Second Throat supersonic Exhaust Diffuser (STED) to simulate the high altitude condition. Cases with small and large vacuum chambers were considered, in order to highlight the effect of vacuum chamber during the motor transient start-up period. The effects of varying diffuser parameters were also studied with reference to achieving nozzle and diffuser full flow conditions. Based on the cold flow test results, a diffuser was designed for testing a solid rocket motor with 70 area ratio nozzle. The main objective here was to determine the minimum starting pressure value corresponding to nozzle full flow condition. Hot flow tests were conducted with small and large vacuum chamber using STED, similar to the cold flow tests. The diffuser performance characteristics during the starting transient for the cases with small and large vacuum chamber were studied and the results were analyzed. In order to counter the longer ignition transients, tests were conducted with a nozzle closure arrangement, which would burst near about the diffuser starting pressure and aid in the quick-start of the diffuser. Numerical simulations have also been carried out for visualizing the flow field in the nozzle and diffuser regions. In the numerical study, the time-averaged Navier-Stokes equations have been solved with the Spalart-Allmaras turbulence model, using commercial CFD software. The detailed shock structure and the pressure recovery pattern along the diffuser duct are vividly captured by the numerical predictions. The predicted minimum starting pressure values are also in good agreement with the experimental results obtained in this study. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.