<div class="htmlview paragraph">Significant improvements in engine readiness with reductions in maintenance costs and turnaround times can be achieved with an engine condition monitoring systems (CMS). The CMS provides health status of critical engine components, without disassembly, through monitoring with advanced sensors. Engine failure reports over 35 years were categorized into 20 different modes of failure. Rotor bearings and turbine blades were determined to be the most critical in limiting turbopump life. Measurement technologies were matched to each of the failure modes identified. Three were selected to monitor the rotor bearings and turbine blades: the isotope wear detector and fiberoptic deflectometer (bearings), and the fiberoptic pyrometer (blades). Signal processing algorithms were evaluated for their ability to provide useful health data to maintenance personnel. Design modifications to the Space Shuttle Main Engine (SSME) high pressure turbopumps were developed to incorporate the sensors. Laboratory test fixtures have been designed for monitoring the rotor bearings and turbine blades in simulated turbopump operating conditions.</div>

To obtain high specific work output with small mass flow rate, high-pressure ratios across the turbines are required in liquid rocket engine turbopumps. An impulse-type supersonic turbine can achieve this. To prevent losses due to low blade aspect ratio and issues related to manufacturing and industrial problems, partial admission configuration is adopted. Partial entry in a turbine is achieved by adjusting the extent of the nozzle arc of admission, leading to a strong unsteady circumferential asymmetry of flow parameters in the rotor passage, and degradation in efficiency. The pressing need of aerodynamic design of supersonic partial admission turbines to improve their efficiency demands an investigation of the viscous fluid dynamic of the turbine flow field. This work reports the aerothermodynamic steady state CFD analysis to obtain the performance parameters of a three-dimensional partial admission turbine for LOX booster turbopump in a semicryogenic engine using ANSYS® CFX. The areas of steady loss have been identified through entropy generation contours, and the effects associated with aerodynamic loss structures like secondary flow, shock location, recirculation with additional pumping and mixing losses have been investigated for designed operating condition corresponding to 100% nominal thrust.

<div class="htmlview paragraph">This technical paper discusses the evolution of turbopumps designed for U.S. rocket engines, starting with the Thor, Jupiter, and Atlas missiles, which were developed in the 1950s; the F-1 and J-2 engines, which were developed in the 1960s for the Saturn V vehicle to put man on the moon; and the Space Shuttle Main Engine (SSME), which was developed in the 1970s for the United States' first reusable space vehicle. Technology development in the 1980s that will influence the design of turbopumps for the National Launch System (NLS), the National Aero-Space Plane (NASP), and Orbital Transport Vehicles (OTV) are also discussed.</div>

Experimental study on stability of axial direction of Balance Piston mechanism Vibration problems often occur in development of rocket engine turbopump, and an axial vibration is one of these problems. Rocket turbopumps often use balance piston mechanism (BP) which is a self-balancing system in order to balance the large axial thrust derived from high discharged pressure. Turbopump rotor with BP is movable in axial direction and supported by fluid stiffness generated by BP. Although BP has a stable characteristic statistically, it becomes dynamically unstable in some cases because of compressibility of the fluid in BP chamber. Stability of BP has been studied in analysis approaches, but experimental approach has not been conducted yet. In this paper experimental research is conducted to confirm the occurrence of self-excited oscillation and compare the experimental results with one dimensional theoretical analysis. From the experimental results, effect of BP parameters on stability is revealed and it is concluded that there is contradictory relationship between BP static and dynamic characteristics. Published in Nihon Kikai Gakkai ronbunshu

This study evaluated a rocket engine system for a suborbital vehicle, utilizing 70 MPa hydrogen gas as fuel and 5 MPa liquid nitrous oxide as oxidizer. The engine produced no carbon dioxide emissions, eliminating a major greenhouse gas. The use of high-pressure hydrogen gas removed the need for a turbopump, simplifying the design. Lightweight composite materials were employed for the propellant tanks. Two-dimensional flight simulations were conducted to estimate payload capacity, which was then compared with that of a hybrid rocket engine system. The nitrous oxide tank was assumed to maintain constant pressure throughout operation. The hydrogen–nitrous oxide engine demonstrated payload capabilities on par with the hybrid system. Regenerative cooling was implemented, with high-pressure hydrogen gas absorbing engine heat despite temperature rise from the Joule–Thomson effect. The recovered heat was used to vaporize liquid nitrous oxide, helping sustain the required 5 MPa tank pressure. To remain within dynamic pressure limits during flight, the engine operated in a throttled mode. Maximum payload was achieved at an equivalence ratio of 1.04. Although nitrous oxide is a greenhouse gas, it was effectively consumed during combustion. Overall, the hydrogen–nitrous oxide engine system shows strong potential as an economical, efficient, and environmentally safer solution for suborbital flight applications.

methane-rich turbopumps. Before 2014, only two full-flow staged-combustion rocket engine designs had advanced enough to undergo testing: the Soviet RD-270 Starship is a two-stage, fully reusable, super heavy-lift launch vehicle under development by American aerospace company SpaceX. Currently built and launched from Starbase in Texas, it is intended as the successor to the company's Falcon 9 and Falcon Heavy rockets, and is part of SpaceX's broader reusable launch system development program. If completed as designed, Starship would be the first full The Block 2 version of Starship is 52.1 m (171 ft) tall, 9 m (30 ft) wide, and is composed of four general sections: the engine bay, the oxygen tank, the fuel tank, and the payload bay. The retired Block 1 was constructed in a similar manner, though it was only 50.3 m (165 ft) tall. Elon Musk stated in 2021 that the vehicle has a dry mass of roughly 100 t (220,000 lb). The windward side is protected by a heat shield, which is composed of eighteen thousand hexagonal black tiles that can withstand temperatures of 1,400 °C (2,600 °F). It is designed to protect the vehicle during atmospheric entry and to be used multiple times with minimal maintenance between flights. The silica-based tiles are attached to Starship with pins, and had small gaps in between to allow for heat expansion. After flight test 4, SpaceX added a secondary ablative layer under the primary heat shield, though this was only added to the flaps of the flight test 6 vehicle. This ablative layer is likely composed of pyron, which is similar in composition to carbon composites. The total mass of the heat shield and ablative layer of a Block 1 ship is 10.5 t (23,000 lb). After flight test 10, SpaceX added a felt, called "crunch wrap," between the gaps in between the tiles to prevent heat seeping in. R… In July 2022, Booster 7 tested the liquid oxygen turbopumps on all 33 Raptor engines, resulting in an explosion at the vehicle's base, which destroyed a pressure pipe and caused minor damage to the launchpad. By the end of November, Ship 24 had performed 2 static test fires, while Booster 7 had performed 6 static test fires and finally on February 9, 2023, a static fire with 31 engines at 50% throttle. In January 2023, the whole Starship stack underwent a full wet dress rehearsal. After a launch attempt aborted on April 17, 2023, Booster 7 and Ship 24 lifted off on April 20 at 13:33 UTC in the first orbital flight test. Three engines were disabled during the launch sequence and several more failed during the flight. The booster later lost thrust vectoring control of the Raptor engines, which led to the rocket spinning out of control. The vehicle reached a maximum altitude of 24 mi (39 km). Approximately 3…

Rocket engine nozzle blocks operate under extreme thermal and oxidative loads, requiring materials with high temperature resistance, dimensional stability, and a predictable lifetime without active cooling. This review provides a comparative overview of multimatrix composite materials-including C/C, C/SiC, SiC/SiC, MMC, and polymer-based ablative systems-representing the full spectrum of materials used in non-cooled rocket nozzles. The study highlights the evolutionary continuum from polymeric ablative systems to carbon, ceramic, and metallic matrices, demonstrating how each class extends operational limits in temperature capability, reusability, and structural integrity. Polymer and ablative composites serve as the foundation of thermal protection through controlled ablation and insulation, while carbon- and ceramic-based systems ensure long-term performance at ultra-high temperatures (>1600 °C). MMCs bridge these classes by combining strength, impact toughness, and thermal conductivity in transition zones. Particular attention is given to manufacturing technologies such as PIP, CVI, LPI, RS, powder metallurgy, casting, diffusion bonding, and filament winding, emphasizing their effect on microstructure, porosity, and lifetime. A practical selection matrix linking nozzle zones, mission profiles, and composite types is proposed, outlining trade-offs among performance, mass, lifetime, and manufacturability, and guiding the design of next-generation thermal protection and propulsion systems based on the multimatrix concept.