Space Shuttle Main Engine Research Articles

Numerical modeling for primary atomization of liquid jets

In the proposed numerical model for primary atomization, surface-wave dispersion equations are solved in conjunction with the jet-embedding technique of solving mean flow equations of a liquid jet. Linear and approximate nonlinear models have been considered. In each case, the dispersion equation is solved over the whole wavelength spectrum to predict drop sizes, frequency, and liquid-mass breakup rates without using any empirical constants. The present model has been applied to several low-speed and high-speed jets. For the high-speed case (the LOX/H2 coaxial injector of the Space Shuttle Main Engine Preburner), predicted drop sizes and liquid breakup rates are in good agreement with the results of the CICM code, which have been calibrated against measured data.

Steady-State Analysis of a Nonlinear Rotor-Housing System

The periodic steady-state response of a high-pressure oxygen turbopump (HPOTP) of a space shuttle main engine (SSME), involving a clearance between the bearing and housing carrier, is sought. A harmonic balance method utilizing Fast Fourier Transform (FFT) algorithm is developed for the analysis. An impedance method is used to reduce the number of degrees of freedom to the displacements at the bearing clearance. Harmonic and subharmonic responses to imbalance for various system parameters are studied. The results show that the computational technique developed in this study is an effective and flexible method for determining the stable and unstable periodic response of complex rotor-housing systems with clearance-type nonlinearity.

Analysis of Variable Fluid Properties, Turbulent Annular Seals

A computational analysis for calculation of the dynamic force response in turbulent flow annular pressure seals of arbitrary nonuniform clearance is presented. Cryogenic liquids are considered as barotropic. The fluid motion is described by a bulk flow model and Moody’s friction factor is introduced to accomodate surface roughness effects. The numerical solution scheme is computationally efficient and accurate for seal operation at arbitrary eccentricity ratios. Numerical predictions show good agreement with available experimental results. The effects of eccentricity and liquid properties on the performance of the Space Shuttle Main Engine High Pressure Fuel Turbopump interstage seal at full power level are discussed as a reference case.

Parameter Estimation by Nonlinear Smoothing for Fault Monitoring on Rocket Engines

Abstract Nonlinear parameter estimation algorithms based upon the extended Kalman filter are designed for the problem of estimating some of the parameters that describe the dynamics of a reusable space propulsion system. In this paper, the results of applying these algorithms to simulated data and to test firing data from the Space Shuttle Main Engine (SSME) are presented. The parameters that are estimated are among those likely to change when some common engine degradations occur. The SSME operating at its 100% Rated Power Level is used as the baseline system. All estimator designs are based upon a reduced-order dynamic model of the SSME. The simulated data used to drive the algorithms is generated by a high fidelity transient simulation of the SSME with small magnitude random “dither” signals applied to the inputs and with substantial random noise added to the measured outputs. The results for a healthy SSME indicate good parameter estimation accuracy. However, the algorithms fail to track the correct parameter values when parameter changes representing engine degradations are introduced if the change results in a significant change in the values of dynamic variables that are not included in the reduced order model used to design the estimator. Preliminary results of using hot fire data from the SSME are also presented. These results indicate that the estimator produces valid results when it is driven with real data, but that it must incorporate logic to change the estimator design as the commanded power level of the engine changes.

A Distributed Fault-Detection and Diagnosis System Using On-Line Parameter Estimation

The development of a model-based fault-detection and diagnosis system (FDD) is reviewed. The system can be used as an integral part of an intelligent control system. It determines the faults of a system from comparison of the measurements of the system with a priori information represented by the model of the system. The method of modeling a complex system is described and a description of diagnosis models which include process faults is presented. There are three distinct classes of fault modes covered by the system performance model equation: actuator faults, sensor faults, and performance degradation. A system equation for a complete model that describes all three classes of faults is given. The strategy for detecting the fault and estimating the fault parameters using a distributed on-line parameter identification scheme is presented. A two-step approach is proposed. The first step is composed of a group of hypothesis testing modules, (HTM) in parallel processing to test each class of faults. The second step is the fault diagnosis module which checks all the information obtained from the HTM level, isolates the fault, and determines its magnitude. The proposed FDD system was demonstrated by applying it to detect actuator and sensor faults added to a simulation of the Space Shuttle Main Engine. The simulation results show that the proposed FDD system can adequately detect the faults and estimate their magnitudes.

Dual mixture ratio H2/O2 engine for single stage to orbit application

This paper discusses propulsion requirements for rocket-powered single stage to orbit (SSTO) vehicles and defines a dual mixture ratio Hi/Oi staged combustion engine concept with characteristics matched to those requirements. The objective of this study was to configure an engine within the limits of current technology that when combined with a vehicle based on the material advances of the National Aerospace Plane program could provide an effective SSTO system. This engine uses high mixture ratio to provide high bulk density propellant usage and high thrust for the booster mode and transitions to low mixture ratio for high specific impulse during the upper stage mode. The internal operating characteristics of this engine are within the limits of current technology. Preliminary SSTO system studies indicate significant vehicle weight reductions relative to the Space Shuttle main engine (SSME) powered systems. This engine could be available before the end of this century.

Seal-Rotordynamic-Coefficient Test Results for a Model SSME ATD-HPFTP Turbine Interstage Seal With and Without a Swirl Brake

Test results are presented and compared to theory for a model Space Shuttle Main Engine(SSME) Alternate Turbopump Development(ATD) High-Pressure Fuel Turbopump (HPFTP) with and without swirl brakes. Tests are conducted with supply pressures out to 18.3 bars and speeds out to 16,000 rpm. Seal back pressure is controlled to provide four pressure ratios at all supply pressures. Three inlet guide vanes are used to provide the following three fluid prerotation cases: (a) no pre-rotation, (b) moderate prerotation in the direction of rotation, and (c) high prerotation in the direction of rotation. Test results demonstrate the pronounced favorable influence of the swirl brake in reducing the seal destabilizing forces. Without the swirl brake, the cross-coupled stiffness k increases monotonically with increasing inlet tangential velocity. With the swirl brake, k tends to either be constant or decrease with increasing inlet tangential velocity. Direct damping either increases or remains relatively constant when the swirl brake is introduced. Direct stiffness is relatively unchanged. No measurable differences in leakage were detected for the seal with and without the swirl brake. Comparisons between Scharrer’s (1988) theory and measurements for the seal without a swirl brake indicate that the predictions can be used to provide design guidelines only. Specific predictions for rotordynamic coefficients should be treated cautiously, since systematic differences were observed between theory and experiment due to changes in running speed, supply pressure, and pressure ratio.

A Convolution Approach for the Transient Analysis of Locally Nonlinear Rotor Systems

A computationally efficient convolution method, based on discretized impulse response and transition matrix integral formulations, is developed for the transient analysis of complex linear structures interacting through strong local nonlinearities. In the formulation, the coupling forces due to the nonlinearities are treated as external forces acting on the coupled subsystems. Iteration is utilized to determine their magnitudes at each time increment. The method is applied to a generic rotor-housing model representing a turbopump of a space shuttle main engine (SSME). In that model, the local nonlinearity is due to clearances between the rotor bearing outer races and the carrier attached to the housing. As compared to the fourth-order Runge-Kutta numerical integration methods, the convolution approach proved more efficient and robust for the same accuracy requirement. This is due to the closed-form formulation of the convolution approach which allows for the use of relatively larger time increments and for a reduction in the roundoff errors.

Turbine blading designed for high heat load space propulsion applications

The operating conditions and the propellant transport properties used in Earth-to-orbit (ETO) applications affect the aerothermodynamic design of ETO turbomachinery in a number of ways. This paper discusses some aerodynamic and heat-transfer implications of the low molecular weight fluids and high Reynolds number operating conditions on future ETO turbomachinery. The objective of this work was to examine turbine blading concepts for increasing life and reliability by reducing thermal heat load. Using the current Space Shuttle main engine high-pressure fuel turbine as a baseline, aerothermodynamic comparisons were made for two alternate fuel turbine geometries. The first was a revised first-stage rotor blade designed to reduce peak heat transfer. This alternate design resulted in a 23% reduction in peak heat transfer. The second design concept was a single-stage rotor designed to yield the same power output as the baseline two-stage rotor. Since the rotor tip speed was held constant, the turbine work factor doubled. In this alternate design, the peak heat transfer remained the same as the baseline. Although the efficiency of the single-stage design was 3.1 points less than the baseline two-stage turbine, the design was aerothermodynamically feasible and may be structurally desirable.

The liquid rocket booster as an element of the U.S. national space transportation system

Liquid rocket boosters (LRBs) were first considered for the U.S. Space Transportation System (STS) during the early conceptual phases of the Space Shuttle program. However, solid rocket boosters (SRBs) were ultimately selected for the STS, primarily due to near-term economics. Liquid rocket boosters are once again being considered as a possible future upgrade to the Shuttle. This paper addresses the findings of these studies to date, with emphasis on the feasibility, benefits, and implementation strategy for a LRB program. The principal issue relating to LRB feasibility is their ability to be integrated into the STS with minimal vehicle and facility impacts. Booster size has been shown to have a significant influence on compatibility with the STS. The physical dimensions of the Orbiter and STS support facilities place an inherent limitation on the size of any booster to be used with this system. In addition, excessively large diameter boosters can cause increased airloads to be induced on the Orbiter wings, requiring modification of STS launch trajectory and possible performance losses. However, trajectory and performance analyses have indicated that LRBs can be designed within these sizing constraints and still have sufficient performance to meet Space Shuttle mission requirements. In fact, several configurations have been developed to meet a design goal of providing a 20,000 lb performance improvement to low Earth-orbit (LEO), as compared with current SRBs. Several major system trade studies have been performed to establish a baseline design which is most compatible with the existing Space Transportation System. These trades include propellant selection (storable, hydrogen-oxygen, hydrocarbon-oxygen, and advanced propellants); pump-fed vs pressure-fed propellant feed system design; engine selection (Space Shuttle Main Engine, Titan LR-87, and advanced new engines); number of engines per booster; and reusability vs expendability. In general, it was determined through these trade studies that several options exist for designing a LRB that can be integrated into the STS with manageable impacts on STS facilities and operational procedures. While LRBs offer a potential 40% improvement in Shuttle performance, their most significant benefit is the potential improvements they offer in the area of Shuttle safety. This begins during ground handling operations, where LRBs eliminate the need for large quantities of hazardous solid propellants to be emplaced in the Kennedy Space Center Vehicle Assembly Building. In the pre-launch phase, all LRB engines can be ignited on the launch pad and verified prior to release of the STS. During flight, LRB engines can be shut down on command should the need arise. Further, missions could be aborted safely during the boost phase—an option not available with SRBs. A related benefit of LRBs is their ability to accomplish a mission even if one engine fails, assuming the LRB is designed with sufficient performance margin. An implementation plan has been developed which indicates that LRBs can be operational by 1997. The attractive features of the LRB have prompted NASA to include this booster as a principal element of the agency's long range plan for enhancing STS capabilities through an evolutionary program of block changes. The implementation of LRBs offers an attractive option for developing a safer, more reliable, and better performing STS.

Space shuttle main engine model identification

System identification techniques are used to represent the dynamic behavior of the SSME (space shuttle main engine). The comparison of the responses of a nonlinear simulation with the responses of an identified model indicates very good agreement. The identified model can be used for control design purposes. The identified model does not include valve linkage backlash and valve stiction nonlinearities. These nonlinearities should be added to the identified model, and the validity of the model should be checked by comparing it with the full nonlinear simulation. The identified model is valid for a limited response region at about the 100% power level operating condition.< <ETX xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">&gt;</ETX>

Modelling the thermo-mechanical behavior of a carbide inclusion in a nickel superalloy including residual stress effects with a simple finite-element model

Fractography analyses of the turbine blades in NASA's Space Shuttle Main Engines (SSME) have shown that some types of fractures in the blades can be correlated with inclusions such as carbides, possibly because the inclusion causes a concentration of stress (or strain) in the surrounding matrix. The Lockheed DIAL Finite-Element code was set up to do 2-dimensional axisymmetric modelling of a sphere (MC-type carbide/inclusion) in a cylinder (DS MAR-M-246/matrix) to investigate the basic features of the inclusion problem. The material properties, both thermal and mechanical, were allowed to vary with temperature because the actual turbine blade is exposed to high temperatures during operation. Some analyses also included an initial (or residual) stress distribution as a way of incorporating shotpeening effects that might prevent fracture. The computations showed that stresses large enough to cause fracturing can occur in either the inclusion or the matrix. The residual compressive stresses assumed to occur from the shotpeening can act to reduce the likelihood of fracture.

Space Shuttle guidance for multiple main engine failures during first stage

This paper presents contingency abort guidance schemes recently developed for multiple Space Shuttle main engine failures during the first two minutes of flight (first stage). The ascent and entry guidance schemes greatly improve the possibility of the crew and/or the Orbiter surviving a first stage contingency abort. Both guidance schemes were required to meet certain structural and controllability constraints. In addition, the systems were designed with the flexibility to allow for seasonal variations in the atmosphere and wind. The ascent scheme guides the vehicle to a desirable, lofted state at solid rocket booster burnout while reducing the structural loads on the vehicle. After Orbiter separation from the solid rockets and the external tank, the entry scheme guides the Orbiter through one of two possible entries. If the proper altitude/range/velocity conditions have been met, a return-to-launch-site 'Split-S' maneuver may be attempted. Otherwise, a down-range abort to an equilibrium glide and subsequent crew bailout is performed.

Vortex shedding flowmeters for high velocity liquids

The applicability of vortex shedding flowmeters to measurement in the space shuttle main engine (SSME) ducts is reported here. The liquid oxygen (LOX) flows in the engine ducts exceed by as much as a factor of 10 the maximum flow velocities flow which commercially available meters are designed. The largest Reynolds number is about 3 × 10 7. Tests results show that the vortex shedding flowmeters can measure flow in the SSME ducts and do so without any upstream flow conditioning. The meter measuring element for meter bores up to 59 mm can be introduced through an opposed pair of standard SSME duct instrument ports.

Vibration analysis of rotor systems using reduced subsystem models

A general impedance method using reduced submodels has been developed for the linear dynamic analysis of rotor systems. Formulated in terms of either modal or physical coordinates of the subsystems, the method enables imbalance responses at specific locations of the rotor systems to be efficiently determined from a small number of 'master' degrees of freedom. To demonstrate the capability of this impedance approach, the Space Shuttle Main Engine high-pressure oxygen turbopump has been investigated to determine the bearing loads due to imbalance. Based on the same formulation, an eigenvalue analysis has been performed to study the system stability. A small 5-DOF model has been utilized to illustrate the application of the method to eigenvalue analysis. Because of its inherent characteristics of allowing formulation of reduced submodels, the impedance method can significantly increase the computational speed.

Probabilistic structural analysis methods for improving Space Shuttle engine reliability

Probabilistic structural analysis methods are particularly useful in the design and analysis of critical structural components and systems that operate in very severe and uncertain environments. These methods have recently found application in space propulsion systems to improve the structural reliability of Space Shuttle Main Engine (SSME) components. A computer program, NESSUS, based on a deterministic finite-element program and a method of probabilistic analysis (fast probability integration) provides probabilistic structural analysis for selected SSME components. While computationally efficient, it considers both correlated and nonnormal random variables as well as an implicit functional relationship between independent and dependent variables. The program is used to determine the response of a nickel-based superalloy SSME turbopump blade. Results include blade tip displacement statistics due to the variability in blade thickness, modulus of elasticity, Poisson's ratio or density. Modulus of elasticity significantly contributed to blade tip variability while Poisson's ratio did not. Thus, a rational method for choosing parameters to be modeled as random is provided.

Structural tailoring of Space Shuttle Main Engine turbopump blades SSME/STAEBL

Computerized structural optimization is applied to designing space shuttle main engine (SSME) turbopump blades. Optimization is implemented using a specially developed code, SSME/STAEBL. The capabilities of this code are described with particular attention to those features special to the SSME blade environment. These include severe thermal and pressure loads. Examples of the use of the code are presented. It is shown that incorporation of thermal effects, including temperature-dependent material properties, is important in design optimization, since these effects can add significantly to the weight of the optimal design. The design of single crystal blades is discussed. Crystal axis orientation is found not to play a major role in the optimized design. However, situations are identified in which the orientation could be important.

A New Model for Leakage Prediction in Shrouded-Impeller Turbopumps

The swirling secondary flow in the shroud-to-housing passage of a turbopump stage is analyzed. The flow model is based on the weighted-residual finite-element method to solve the axisymmetric flow field in this passage. A unique feature of the present analysis is the inclusion of two primary flow segments in the definition of the solution domain. This allows full interaction between the primary and secondary flow streams and alleviates what would otherwise be an iterative procedure to calculate the leakage flow rate. Full ellipticity of the flow field throughout the entire domain is maintained for accurate modeling of separation and recirculation zones. Applicability of the computational model to generally shaped secondary flow passages is illustrated through a sample pump that is similar to the boost-impeller stage in the oxidizer turbopump of the U.S. Space Shuttle Main Engine. The secondary flow field in this case is investigated over a range of Reynolds number.

Explicit Runge-Kutta method for unsteady rotor/stator interaction

A quasi-three-dimensional rotor/stator analysis has been developed for blade-to-blade flows in turbomachinery. The analysis solves the unsteady Euler or thin-layer Navier-Stokes equations in a body-fitted coordinate system. It accounts for the effects of rotation, radius change, and stream-surface thickness. The Baldwin-Lomax eddy-viscosity model is used for turbulent flows. The equations are integrated in time using a four-stage Runge-Kutta scheme with a constant timestep. Results are shown for the first stage of the Space Shuttle Main Engine high pressure fuel turbopump. Euler and Navier-Stokes results are compared on the scaled single- and multi-passage machine. The method is relatively fast and the quasi-three-dimensional formulation is applicable to a wide range of turbomachinery geometries.

Improved Rotor Response of the Uprated High Pressure Oxygen Turbopump for the Space Shuttle Main Engine

During development testing of the High Pressure Oxygen Turbopump (HPOTP) of the Space Shuttle Main Engine (SSME) to produce 109 percent of the rated thrust level, subsynchronous rotor whirl was encountered. This whirl was attributed to bearing wear reducing the radial bearing stiffness that caused the rotor second bending mode critical speed to enter the operating speed range. To eliminate this whirl, the pump end bearing loads were reduced to increase bearing life and damping added between the rotor and housing. This was achieved by converting impeller annular seals into “damping” seals that react part of the applied load and also damp the rotor response. Furthermore, the second rotor critical speed was increased by the added stiffness of the seal conversion and stiffening the rotor shaft. The bearing load reduction was verified by strain gaging the pump end bearing support into a load cell. These strain gages also were used to directly measure bearing ball wear during engine tests.