This article presents the results of research on materials used in the nozzle section of a rocket engine. The study analyzed the properties of high-strength metal alloys, heat-resistant, and thermal insulation materials. A review of materials proposed for use in nozzle constructions was conducted. Three different materials were selected for experimental studies: structural steel 30KhGSA, textolite (grade A), and pressed graphite. To assess the thermal and erosion resistance of these materials, fire tests were conducted on a 70 mm caliber engine test stand. As a result of the tests, the level of erosion wear was determined in different nozzle zones: the converging section, the throat, and the diverging section. Changes in the geometric dimensions (diameter and angles) of nozzle components were monitored. The graphite material demonstrated the highest thermal and erosion resistance in the throat region. While steel materials are effective in terms of mechanical strength, and textolite is suitable as a thermal insulator, it was found to be significantly worn under high temperatures. Overall, the results of the conducted research can serve as a scientific and practical basis for designing and selecting materials for rocket engine nozzle blocks. The work provides recommendations for the proper selection of materials for small- and medium-caliber solid-fuel rocket engines.

Pintle injectors offer variable thrust capability and combustion stability advantages for liquid rocket engines. This study presents experimental and numerical investigation of spray characteristics for a liquid–liquid pintle injector using water as simulant. Ten cold flow tests covering total momentum ratio (TMR) from 0.36 to 2.76 captured spray angle variations from 26° to 80°. Computational fluid dynamics (CFD) simulations using Ansys Fluent 2025 R1 with the Volume of Fluid method and dispersed interface modeling showed good agreement with experimental spray angles for TMR > 0.74 (error < 8%), but demonstrated increasing discrepancy at lower TMR values (up to 62% error at TMR = 0.36). This deviation indicates limitations of steady-state RANS models in capturing unsteady, fuel-dominated flow regimes. The experimental dataset provides validation benchmarks for CFD modeling and contributes to injector design optimization for sounding rocket applications.

The design of small upper stages must meet many propulsion requirements: repeated restarts, stable operation at low mass flow rate, strict limits on mass and volume, and compatibility with limited ground infrastructure. The central engineering task is a justified selection of the liquid propellant rocket engine feed architecture for a given mission profile and propellant pair, since this choice affects specific impulse, service life, risk, and ground processing effort. This work systematizes propellant feed schemes for a small upper stage and proposes a unified analytical framework that links engine cycle and feed method with propellant selection and tank pressurization modes. Pressure fed and turbopump schemes are treated as mature solutions; pump cycles are used mainly with cryogenic propellants and in high energy demand cases that require high efficiency and compact hardware. In parallel, electro pump schemes are actively studied as a promising direction due to lower mechanical complexity, precise control, and straightforward integration with modern control systems. The analysis shows how chamber pressure, allowable throttling range, restart capability, and the mass and volume metrics of tanks and hardware bound the rational domain of each scheme. The outcome is a set of criteria for early design that maps mission requirements to feed system architecture and supports a technologically feasible choice for small upper stages. Keywords: pressure-fed system, turbopump-fed supply system, small upper stages, rocket engines, liquid rocket engine.