RM 4000
- Type:
Liquid-fuel engine- Propellant:
LOX/RP-1- Mixture Ratio:
2.7:1- Cycle:
Oxidizer-rich staged combustion- Length:
3,800 mm- Diameter:
1,900 mm- Dry Weight:
5,400 kg- Nozzle Ratio:
25:1- Thrust (vac):
5,500 kN- Thrust (sl):
5,100 kN- Throttle Range:
100% to 40%- Thrust-to-weight ratio:
97:1- Chamber Pressure:
26.0 MPa- Isp (vac):
330 seconds- Isp (sl):
305 seconds
Overview:
The SDI RM 4000 is a reusable, high-performance, oxidizer-rich staged combustion cycle engineburning liquid oxygen and RP-1 kerosene designed by SDI Rocket Engines. The RM 4000 engine is designed to produce 5,100 kN of sea-level thrust at a chamber pressure of approximately 26 MPa, with a specific impulse of approximately 305 seconds at sea level and 330 seconds in vacuum.
Design:
The RM 4000 employs an oxidizer-rich staged combustion (ORSC) cycle where the entire liquid oxygen flow and a small bleed of RP-1 fuel are combusted in a single preburner to produce a hot, oxygen-rich gas that drives both the oxidizer and fuel turbopumps on a common shaft. The oxidizer rich turbine exhaust is then injected into the main combustion chamber where itserves as the oxidizer for the primary combustion process with the bulk of the RP-1 fuel. Liquid oxygen enters the engine at the low-pressure oxidizer turbopump (LPOTP) inlet at approximately 0.45 MPa. The LPOTP, an axial-flow inducer driven by a hydraulic turbine tapped from the high-pressure pump discharge, raises the LOX pressure to approximately 2.8 MPa, providing sufficient net positive suction pressure to suppress cavitation at the high-pressure oxidizer turbopump (HPOTP) inlet. The HPOTP is a single-stage centrifugal pump and raises the LOX pressure to approximately 38 MPa. From the HPOTP discharge the LOX flow splits, approximately 98% is directed to the preburner where it is combusted with a small RP-1 bleed while the remaining ~2% bypasses the preburner and is injected directly into the MCC as liquid film coolant along the chamber wall. The preburner's hot, oxygen-rich exhaust gas then passes through the single-shaft turbine, exits at approximately 28.5 MPa, and enters the MCC injector as the primary oxidizer stream. Fuel (RP-1) meanwhile enters the engine at the low-pressure fuel turbopump (LPFTP) inlet at approximately 0.35 MPa. The LPFTP raises the pressure to approximately 3.5 MPa. The high-pressure fuel turbopump (HPFTP), a single-stage centrifugal pump on the same shaft as the HPOTP, raises the RP-1 pressure to approximately 42 MPa. From the HPFTP discharge, the RP-1 flow splits three ways: approximately 3–5% is routed to the preburner as the fuel component, approximately 60–65% passes through the regenerative cooling jacket of the MCC and nozzle before entering the MCC injector as warm, partially cracked kerosene; and the remaining 30–35% enters the MCC injector directly as subcooled liquid through dedicated fuel injectors.
The RM 4000 uses a single-shaft turbopump configuration in which the high-pressure oxidizer pump, the high-pressure fuel pump, and the single-stage impulse turbine are all mounted on a common rotor supported by two angular contact ball bearings and one roller bearing. This arrangement simplifies the mechanical design, eliminates the need for inter-pump gearing or separate turbine drive shafts, and provides inherent speed synchronization between the two pumps. The High-Pressure Oxidizer Turbopump (HPOTP) is a single-stage centrifugal pump with a fully shrouded impeller machined from a single forging of Monel K-500 nickel-copper alloy (selected for its compatibility with high-pressure liquid and gaseous oxygen). The pump operates at approximately 12,800 RPM and produces a pressure rise of approximately 35.2 MPa at a flow rate of 14,700 kg/s of LOX. The volute and diffuser are cast from Inconel 718 alloy. The High-Pressure Fuel Turbopump (HPFTP) is also a single-stage centrifugal pump with a fully shrouded impeller, machined from a single forging of Ti-5Al-2.5Sn ELI (Extra Low Interstitials) Grade 6, alpha-phase titanium alloy optimized for cryogenic temperatures. The HPFTP operates at the same 12,800 RPM as the HPOTP (on a common shared shaft) and produces a pressure rise of approximately 38.5 MPa at a flow rate of 5,450 kg/s of RP-1. The higher pressure rise compared to the HPOTP compensates for the additional pressure losses in the regenerative cooling circuit that the fuel must traverse before reaching the MCC. The turbine is a two-row velocity-compounded (Curtis) impulse stage, designed to extract the required power at a low pressure ratio (1.137) and moderate rotational speed. The first-row rotor blades are investment-cast single-crystal CMSX-4 nickel superalloy with internal convection cooling passages and an external yttria-stabilized zirconia (YSZ) thermal barrier coating approximately 150 μm thick. The second-row rotor blades are uncooled IN718. The nozzle/stator vanes are also single-crystal CMSX-4. Total blade count is 89 in the first row and 107 in the second row, with a nozzle vane count of 46. The turbine tip speed is approximately 370 m/s, well within the structural capability of the single-crystal blades. The design life of the turbine section is 36,000 seconds of cumulative hot-fire time, corresponding to approximately 100 nominal mission duty cycles of 155 seconds each with margin for engine start and shutdown transients, ground testing, and acceptance test firings. The high-pressure turbopump's rolling-element bearings are the primary life-limiting components in the turbomachinery. The RM 4000 employs hybrid ceramic bearings consisting of silicon nitride (Si₃N₄) rolling elements in M50 steel races that offer significantly longer service life than all-steel bearings due to the lower density of the ceramic balls (reducing centrifugal loading), their superior surface finish (reducing wear), and their self-lubricating properties in the presence of trace LOX or kerosene films. Bearing coolant is provided by filtered, temperature-controlled RP-1 tapped from the HPFTP discharge. Interpropellant seals between the LOX-side and RP-1-side of the common shaft are critical safety components that prevent mixing of oxidizer and fuel within the turbopump cavity. The RM 4000 uses a triple-seal arrangement consisting of a primary LOX-side labyrinth seal, a helium-purged intermediate cavity, and a primary fuel-side labyrinth seal. The helium purge cavity is maintained at a slightly higher pressure than either propellant side, ensuring that any leakage is helium into the propellant streams rather than propellant into the purge cavity. Seal wear is monitored in real time by measuring helium flow rate into the purge cavity an increase above a threshold value triggers an EHM caution flag. Each propellant feed line of the engine further includes a low-pressure turbopump (boost pump) upstream of the high-pressure turbopump. These boost pumps are designed to raise the propellant inlet pressure sufficiently to prevent cavitation at the HPOTP and HPFTP inlets, even with relatively low tank pressurization levels. The Low-Pressure Oxidizer Turbopump (LPOTP) is an axial-flow inducer pump driven by a single-stage hydraulic turbine. The hydraulic turbine is powered by a tap-off of high-pressure LOX from the HPOTP discharge which is throttled through an orifice, expanded through the turbine, and recombined with the main LOX flow downstream of the inducer. The LPOTP operates at approximately 5,200 RPM and produces a pressure rise of approximately 2.35 MPa. The Low-Pressure Fuel Turbopump (LPFTP) is similarly an axial-flow inducer driven by a hydraulic turbine tapped from the HPFTP discharge. It operates at approximately 6,800 RPM and produces a pressure rise of approximately 3.15 MPa. Both low-pressure pumps use hydrostatic journal bearings lubricated by the working fluid (LOX or RP-1), eliminating rolling-element bearing wear as a life-limiting factor.
The RM 4000 preburner is a compact, cylindrical combustion chamber that operates at a pressure of 32.8 MPa and an oxidizer-to-fuel ratio of approximately 56:1 by mass. At this mixture ratio the combustion gas is overwhelmingly composed of gaseous oxygen (~94% by mass), with minor fractions of CO₂, H₂O, CO, and soot-free hydrocarbon fragments. The gas temperature of 810 K is low enough to be compatible with nickel superalloy turbine materials but high enough to deliver the required turbine work. The preburner injector uses 271 coaxial swirl injection elements arranged in a concentric ring pattern on a faceplate manufactured by laser powder bed fusion (LPBF) from Inconel 718. Each element consists of an inner fuel post that delivers a swirling annular film of RP-1 surrounded by an outer oxidizer annulus that delivers LOX. The swirl imparted to the fuel stream promotes rapid atomization and mixing. The faceplate is regeneratively cooled by a fraction of the incoming RP-1 flow that passes through internal cooling channels before being injected through the fuel posts. This cooling scheme is designed to maintain the faceplate temperature well below the oxidation threshold of IN718, even in the aggressive oxygen-rich environment. Combustion stability in the preburner is ensured by acoustic cavities (Helmholtz dampers) machined into the injector faceplate rim and by the inherent stability characteristics of the coaxial swirl injection pattern. The main combustion chamber s a regeneratively cooled, convergent-divergent thrust chamber that operates at a nominal chamber pressure of 26.2 MPa and an overall mixture ratio of 2.70:1 (O/F by mass). The combustion gas temperature at the injector end is approximately 3,680 K, and the characteristic velocity (c*) is 1,805 m/s with a c* efficiency of 99.5%. The MCC liner is fabricated from a CuCrZr alloy that offers the high thermal conductivity necessary for effective regenerative cooling at extreme heat fluxes. The liner is produced as a single-piece near-net-shape forging, after which the regenerative cooling channels are formed by ultrafast laser drilling, a process which creates hundreds of axial cooling channels approximately 1.2 mm wide and 3.5 mm deep along the inner surface of the liner. The channels are then closed out by electrodepositing a structural nickel jacket over the channel tops. The peak hot-wall heat flux occurs at the throat, where it reaches approximately 160 MW/m². RP-1 enters the cooling jacket at the nozzle end of the MCC, flows forward (counterflow to the combustion gas), and exits at the injector end at a temperature of approximately 450 K. At this temperature, the RP-1 has undergone some degree of thermal cracking, which is beneficial as the endothermic cracking reaction absorbs additional heat from the wall. A structural closeout jacket of electrodeposited nickel surrounds the CuCrZr liner, providing the mechanical strength to contain the high chamber pressure (26.0 MPa) while the liner handles the thermal load. The jacket is approximately 8 mm thick and is bonded to the liner during the electrodeposition process. The MCC chamber has a contraction ratio (chamber-to-throat area) of 3.2:1, a cylindrical section length of approximately 380 mm, and a throat diameter of 375 mm. The convergent section uses a smooth, bell-shaped contour that minimizes boundary-layer separation and combustion-gas recirculation. The main injector uses a coaxial swirl element pattern with 547 injection elements arranged in 12 concentric rings. Each element consists of an inner fuel post carrying warm RP-1 (approximately 450 K, partially cracked) and an outer annulus carrying oxidizer-rich gas from the turbine exhaust. The fuel posts impart a tangential swirl component to the RP-1, producing a hollow-cone spray that is sheared and atomized by the surrounding high-velocity oxidizer gas stream. The oxidizer gas enters the elements at approximately 250 m/s and 750 K. The injector faceplate is manufactured by LPBF additive manufacturing in Inconel 718, allowing complex internal manifolding, integrated acoustic damping cavities, and cooling passages to be produced as a single monolithic part. Combustion stability is addressed by four complementary measures: Helmholtz acoustic damper cavities integrated into the faceplate periphery, tuned to the first tangential (1T) acoustic mode of the chamber; acoustic absorber baffles (five radial blades extending approximately 40 mm into the chamber from the injector face); careful matching of injection element impedance to suppress coupling between the propellant feed system and chamber acoustic modes; and the inherent stability of the high-density, high-velocity coaxial swirl injection pattern, which provides strong viscous damping at the injector face.
The RM 4000 ignition system consists of a dual-redundant augmented spark igniter (ASI) system for both the preburner and the main combustion chamber. Each ASI contains a spark plug and independent propellant (GOX/RP-1) feed lines that create a small pilot flame prior to main propellant valve opening. The preburner is ignited first, once preburner combustion is confirmed (by chamber pressure and optical flame detectors) the main oxidizer valve is opened and the turbine then spools up, followed by main fuel valve opening and MCC ignition. The entire start sequence from ignition command to mainstage (full thrust) takes approximately 3.2 seconds. The start sequence is controlled by the engine digital controller and is fully automated, no ground-side commanding is required after the initial ignition signal. The ignition system is designed for unlimited restarts without refurbishment, enabling the multiple-burn mission profiles required for booster landing (boostback, reentry, landing burns). Each ASI spark plug is rated for 10,000 ignition cycles. The engine main oxidizer valve (MOV) is a pneumatically actuated, hydraulically damped ball valve in the main LOX feed line downstream of the HPOTP. The MOV controls engine start and shutdown and provides gross thrust modulation. Actuation time from closed to full open is approximately 350 ms. The Main Fuel Valve (MFV) is an identical ball valve in the main RP-1 feed line located downstream of the HPFTP which is sequenced with the MOV during start and shutdown. The Oxidizer Preburner Valve (OPV) is a variable-position, electromechanically actuated sleeve valve that meters LOX flow to the preburner. The OPV is the primary thrust control mechanism, by adjusting the LOX flow to the preburner the preburner gas temperature and mass flow rate can be varied, which changes the turbine power, turbopump speed, and thus the overall propellant flow rate and chamber pressure. Thrust modulation from 40% to 105% is achieved by varying the OPV position. The fuel preburner Valve (FPV) is a variable-position sleeve valve that meters RP-1 flow to the preburner. The FPV works in concert with the OPV to maintain the desired preburner mixture ratio across the throttle range. Chamber Coolant Valve (CCV): A variable-position valve that controls the fraction of RP-1 routed through the regenerative cooling circuit versus the direct-injection bypass. The CCV is adjusted as a function of thrust level to maintain appropriate cooling jacket temperatures at all operating points. All valve actuators are powered by engine-driven hydraulic and pneumatic systems, eliminating the need for vehicle-level hydraulic power during engine operation. The RM 4000 engine nozzle is a truncated ideal contour (TIC) bell nozzle with an area ratio of 25:1.The nozzle is fabricated in two sections. The upper section, from the throat to approximately 10:1 area ratio, is a regeneratively cooled extension of the MCC liner , CuCrZr alloy with laser-drilled cooling channels carrying RP-1. This region experiences the highest heat flux downstream of the throat (approximately 25 MW/m² at the throat-adjacent section). The lower section, from 10:1 to 25:1 area ratio, is a radiatively cooled structure fabricated from C-103 niobium alloy coated with a disilicide (R512E) oxidation-resistant coating. The radiatively cooled section operates at a wall equilibrium temperature of approximately 1,300–1,500 K, within the capability of coated C-103.
Control:
Each RM 4000 engine incorporates a dedicated engine digital controller (EDC), a self-contained, dual-redundant embedded computer that manages all engine functions: start and shutdown sequencing, thrust and mixture ratio control, valve positioning, ignition system firing, and interface with the engine health management system. The EDC receives high level commands from the vehicle flight computer (e.g., "throttle to 72%," "shutdown," "gimbal to +3.2° pitch") and autonomously executes the required valve, igniter, and actuator sequences. The dual redundant architecture uses two independent processor channels running identical software, with a hardware voter that selects the output of the primary channel unless it disagrees with the secondary, in which case the secondary assumes command. The EDC communicates with the vehicle flight computer via a redundant fiber optic data bus and receives power from the vehicle electrical bus. In the event of a complete loss of communication with the vehicle computer, the EDC can autonomously command a safe engine shutdown. The RRM 4000 EHM system monitors engine parameters including vibration, turbopump speed, temperatures, pressures. and flow rates throughout the engine in real time at sampling rates between 1 kHz and 50 kHz. To measure engine vibration six triaxial accelerometers are mounted on the high-pressure turbopump housing, preburner, MCC, and gimbal block. These sensors capture both broadband vibration levels and narrowband spectral signatures associated with turbopump rotordynamics, bearing defects, and combustion instability. Vibration data is processed on-board by the EDC using fast Fourier transform (FFT) algorithms, and amplitudes at critical frequencies (synchronous, sub-synchronous, and combustion acoustic modes) are compared to redline thresholds. Turbopump speeds are measured by redundant magnetic pickup sensors on both the high-pressure and low-pressure pump shafts which measure rotational speed with an accuracy of ±5 RPM. Speed excursions outside the nominal band are an early indicator of bearing degradation, seal rub, or cavitation. Temperature measurements are taken from over 40 thermocouples and resistance temperature detectors (RTDs) which monitor critical temperatures including turbine inlet, turbine discharge, preburner injector faceplate, MCC throat hot-wall (via embedded thermocouples in the CuCrZr liner), bearing coolant, and nozzle extension skin temperature. More than 25 pressure transducers monitor pressures at the pump inlets and discharges, preburner chamber, MCC chamber (redundant), nozzle base (for separation detection), and the helium interpropellant seal purge cavity. Turbine flowmeters in the main LOX and RP-1 feed lines measure propellant flow rates for mixture ratio control and performance tracking. Additionally two high-bandwidth photodiodes view the preburner combustion zone and the MCC injector face, detecting ignition confirmation and real-time combustion intensity fluctuations. A spectral emission sensor at the turbine exhaust duct monitors for the presence of metallic emission lines (iron, nickel, copper) that indicate turbine blade or bearing material erosion. The EHM system maintains a continuously updated fault model with approximately 150 monitored parameters and associated redline thresholds. After each flight, the full-rate EHM data set (approximately 120 GB per engine per flight) is downloaded from the EDC's solid-state data recorder and analyzed by ground-based machine learning algorithms that compare the engine's performance and signature data to its historical baseline and to the fleet-wide database.
