This article discusses the developmental challenges of the low-thrust, long-duration solid rocket motor for the launch of the experimental Reusable Launch Vehicle-Technological Demonstrator (RLV-TD). The main challenges were: (1) developing a motor case and subsystems with low inert mass; (2) design of an optimum nozzle such that the motor can have maximum specific impulse at atmospheric conditions, but with no flow separation at low operation pressures; (3) developing a slow-burning propellant (3 mm/s) to meet the mission requirements; (4) design of propellant grain for the motor so that it has long burning time, the vehicle experiences low dynamic pressure at the transonic regime, and the motor is without combustion instability; (5) developing necessary thermal protection system to take care of long-duration operations, and (6) developing the igniter to ensure the ignition of the motor, especially when easy ignition is difficult with slow-burning propellants and that there should be sufficient overlap of igniter functioning with motor initiation. Performance of the motor in flight indicated that the design met all the required criteria within the expected tolerance.

Crew escape system (CES) is one of the most critical subsystems in a human-rated launch vehicle. CES has four different types of solid motors. One is the high-altitude escape motor (HEM). This motor has the scarfed nozzle region at the divergent aft-end side which is different from conventional nozzle to accommodate the nozzle inside the envelope of the crew module shroud. The composite ablative phenolic liners are bonded to the nozzle metallic hardware by means of adhesive. The structural integrity of this scarfed nozzle region due to temperature and pressure plays an important role for the success of the motor. The nozzle hardware of this motor comprises of three subassemblies connected together by flanged joint. Composite ablative liners are primarily designed to satisfy the thermal and internal ballistics constraint requirements. Though metallic nozzle hardware is designed to bear the complete internal pressure loads, liners share a substantial part during operation due to its stiffness which is comparable with that of the metallic backup. In addition, the high thermal gradient also results in stresses near the inner surface of the liner. These stresses cannot be estimated by closed-form solutions considering the complexity in geometry, direction-dependent material property, and arbitrary temperature distribution arising during operation. The temperature of the liner at its inner surface is the highest due to its direct contact with hot gases. The temperature within the liner decreases across thickness. It is required to be ensured that the temperature at the liner-hardware interface does not exceed the safe-operating temperature. Thermo-structural analysis of composite ablative liners is essential to estimate the complete stress state in liners and to arrive at the minimum available structural margins. The temperature and scarfed geometry make the analysis all the more complicated. Challenges in modelling liners with the 3D contact surface elements are highlighted. Varying pressure load along the inner surface of the liners is simulated in an exclusive load step, and in addition, temperature data estimated from transient thermal analysis are applied in another load step. Temperature-dependent thermo-physical and structural properties are used for the analysis. The temperature distribution across the liner thickness and stresses in the liner and metallic nozzle are reported at different cross sections and their criticality is analysed. This paper highlights the integrated 3D finite-element modelling and analysis of composite liners of solid rocket scarfed nozzle and also covers the comparison of the mechanical strain and thermal parameters with post-static test values.

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