Cavitation dynamics continue to pose a significant risk in the development and operation of launch vehicle (LV) propulsion systems. In addition to generating unsteady loads that can directly damage turbopump hardware, cavitation dynamics often couple with LV fluid feed systems, producing system wide POGO instability that can cause catastrophic failures. Despite its importance, the current understanding of cavitation dynamics, and especially pump transfer matrices, is limited. Given the relatively sparse amount of inducer transfer matrix data available, there is a critical need for more in-depth characterization of the cavitation dynamics in turbopump inducers to avoid POGO instability. This paper defines and validates a new reduced-order approach to infer key parameters such as cavitation compliance, K, and mass flow gain factor, M, from simple, single sensor unsteady pressure measurements during inducer inlet pressure ramps. The utility of this approach is demonstrated for a range of inducer geometries reported in the literature. The results are in agreement with experimental data and the paper provides a new capability supporting the assessment of launch vehicle POGO instability.

Air breathing rocket engines require turbomachinery and ducting that is substantially lighter than that used in ground based or aerospace gas turbines. In order to reduce the weight of the axial compressor, the design of the inter-spool swan neck duct is targeted. In this paper a circumferential splitter blade is used to reduce loading and diffusion on the duct endwalls. The splitter and duct geometry are coupled and optimised together using 2D CFD. A design is selected that is 30% shorter than ducts that are currently used in aerospace gas turbines and the 3D flow features are investigated in further detail using an experimental rig and 3D CFD. This paper shows that the “splittered” duct has 3 benefits over a conventional duct design: First, separation of the endwalls is prevented even at short duct lengths, this will reduce distortion into the downstream compressor. Second, losses generated by corner separations on structural struts can be reduced by 20%, enabling short ducts to achieve high performance. Third, splittered ducts are shown to be twice as robust to uncertain inlet flow conditions as conventional ducts. This allows a designer to target high performance short designs with reduced risk.

The tonal sound production during thermoacoustic instability is detrimental to the components of gas turbine and rocket engines. Identifying the root cause and controlling this oscillatory instability would enable manufacturers to save in costs of power outages and maintenance. An optimal method is to identify the structures in the flow-field that are critical to tonal sound production and perform control measures to disrupt those “critical structures”. Passive control experiments were performed by injecting a secondary micro-jet of air onto the identified regions with critical structures in the flow-field of a bluff-body stabilized, dump, turbulent combustor. Simultaneous measurements such as unsteady pressure, velocity, local and global heat release rate fluctuations are acquired in the regime of thermoacoustic instability before and after control action. The tonal sound production in this combustor is accompanied by a periodic flapping of the shear layer present in the region between the dump plane (backward-facing step) and the leading edge of the bluff-body. We obtain the trajectory of Lagrangian saddle points that dictate the flow and flame dynamics in the shear layer during thermoacoustic instability accurately by computing Lagrangian Coherent Structures. Upon injecting a secondary micro-jet with a mass flow rate of only 4% of the primary flow, nearly 90% suppression in the amplitude of pressure fluctuations are observed. The suppression thus results in sound pressure levels comparable to those obtained during stable operation of the combustor. Using Morlet wavelet transform, we see that the coherence in the dominant frequency of pressure and heat release rate oscillations during thermoacoustic instability is affected by secondary injection. The disruption of saddle point trajectories breaks the positive feedback loop between pressure and heat release rate fluctuations resulting in the observed break of coherence. Wavelet transform of global heat release rate shows a redistribution of energy content from the dominant instability frequency (acoustic time scale) to other time scales.

sCO 2 power cycles offer improved cycle efficiencies compared with traditional steam Rankine cycles. However, the turbomachinery required to support such a cycle does not exist at a commercial scale and requires development. This paper describes a new 10 MWe scale sCO 2 turbine was developed and demonstrated in an sCO 2 closed-loop recompression Brayton cycle. Since this turbine was developed for Concentrating Solar Power (CSP) applications, a target inlet temperature of over 700°C was chosen using funding from the US DOE SunShot initiative and industry partners. However, it can be applied to traditional heat sources such as natural gas, coal, and nuclear power. Traditional Rankine steam cycle thermal efficiencies are typically in the 35–40% range, but can be as high as 45% for advanced ultra-supercritical steam cycles. The sCO 2 cycle can approach 50% thermal efficiency using externally fired heat sources. Furthermore, this cycle is also well suited for bottoming cycle waste heat recovery applications, which typically operate at lower temperatures. The high-power density and lower thermal mass of the sCO 2 cycle results in compact, high-efficiency power blocks that can respond quickly to transient environmental changes and transient operation, a particular advantage for solar, waste heat, and ship-board applications. The power density of the turbine is significantly greater than traditional steam turbines and is comparable to liquid rocket engine turbo pumps. This paper describes the design and construction of the turbine and provides additional testing of the 10 MWe turbine in a 1 MWe test facility including a description of rotordynamics, thermal management, rotor aero and mechanical design, shaft-end and casing seals, bearings, and couplings. Test data for the turbine is included, as it achieves its operational goal of 715°C, 250 bara, and 27,000 rpm.

Turbopumps constitute an essential component of high thrust liquid rocket engines. They are characterized by a compact design, providing a large shaft power at high rotation rates. This is necessary to deliver the propellants at high pressure into the combustion chamber to generate the required engine thrust. Recirculating fluid around the pump has a major influence on axial loads and fluid dynamical losses, impacting turbopump performance and life. Therefore, simplified modelling approaches are required early on for the preliminary design of the pump impeller, and its side cavities and seals. Indeed, past experience within ArianeGroup for pumps and secondary circuits indicates that the coupling between the main flow and the leakage has to be considered at an early stage of the design. The empirical correlations of the flow in the cavities shall be carefully selected, accounting for the particularities of each new configuration. Furthermore, it is also recommended for the impeller design (e.g. for blade leading edge and pressure relieve orifices positioning), that the effects of leakage reinjection into the main flow shall be taken into account. In order to obtain first estimates for early design optimization without the cost of full scale 360° high fidelity computational dynamic simulations (CFD), a reduced model is developed to predict losses and axial thrust on the rotor, including effects of fluid recirculation and reinjection. A two-step approach is followed: Firstly, an empirical model developed by Gülich et al. [1] is applied to characterize leakage loss analytically. Secondly, a reduced numerical model is implemented which features a single passage impeller geometry including seals and side wall gaps. The accuracy of both the analytical model and the simplified numerical model are verified in comparison to high fidelity CFD calculations, evaluating the loss contributions in the leakage path and axial thrust for a range of operating points. In line with expectation, the highest impact on the pump performance are the volumetric losses due to the recirculation of pressurized fluid, with and efficiency decrease of up to 20 % in the investigated cases. The implemented analytical model captures the overall loss mechanisms with a 20 % uncertainty in the design point, disk friction is underpredicted and axial thrust is mostly over-predicted. Due to the simplified numerical model with the single passage impeller geometry including side cavities, the uncertainty can be decreased to about 5 %. At part load operation, the accuracy of both models reduces. It is noted, that thrust prediction is subject to the highest uncertainties. The current work has provided a simplified numerical model that offers the higher flexibility required for the early design phase as compared to a full annulus CFD simulation of the pump, with an increased accuracy as compared to the analytical models.

Over the last two decades, detonation based propulsion has received a great deal of attention as a potential means to achieve significant improvement in the performance of air-breathing and rocket engines. Detonative combustion mode is particularly interesting due to the resulting pressure gain from reactants to products, faster heat release, decreased entropy generation, more available work and higher thrust compared to conventional deflagrative combustion. Rotating detonation engine (RDE) is one such novel combustor concept. Realistic RDE configurations utilize separate fuel and air injection schemes, hence are not perfectly premixed. Moreover, RDE performance is governed by a large number of design parameters and operating conditions. In this context, computational fluid dynamics (CFD) has the potential to enhance the understanding of RDE combustion and aid future development/optimization of this technology. In the present work, a CFD model was developed to simulate a representative non-premixed RDE combustor. Unsteady Reynolds-Averaged Navier-Stokes (RANS) simulations were performed for the full combustor geometry (including the separate fuel and air injection ports), with hydrogen as fuel and air as the oxidizer. Adaptive mesh refinement (AMR) was incorporated to achieve a trade-off between model accuracy and computational expense. A finite-rate chemistry model along with a 10-species detailed kinetic mechanism was employed to describe the H 2 -Air combustion chemistry. Two operating conditions were simulated, corresponding to the same global equivalence ratio of unity but different fuel and air mass flow rates. For both conditions, the capability of the model to capture the essential detonation wave dynamics was assessed. A validation study was performed against experimental data available on detonation wave frequency/height, reactant fill height, oblique shock angle, axial pressure distribution in the channel, and fuel/air plenum pressure. The CFD model predicted the sensitivity of these wave characteristics to the operating conditions with good accuracy, both qualitatively and quantitatively. The present CFD model offers a potential capability to perform rapid design space exploration and/or performance optimization studies for realistic full-scale RDE configurations.

Launch systems development today is heading in two, seemingly divergent, directions. A first direction is towards bigger launchers, designed to carry more than 50 tons into low Earth orbit (LEO). On the other hand, there’s growing interest from government agencies and start-up companies alike in very small vehicles for dedicated launches of small satellites, vehicles that can place just a few hundred, or even a few dozen, kilograms into LEO. At the same time, space programs are starting to look into the possibility of using the existing turbines of rocket engines turbopumps working on classical fuels to work on alternative fuels. The desired characteristics of being a simple, lightweight, high specific work output, low mass flow rate turbine easily translate into supersonic turbines and outweigh the disadvantage of having low efficiency compared to subsonic turbines. Such a turbine needs proof of concept because, at small scales, the flow changes dramatically due to end wall losses which may cause the turbine to choke prematurely and combined with the effect of shock wave losses, characteristic to a supersonic flow, this effect can turn out to be critical. This paper presents a methodology to design such a turbine, taking into account the requirements derived from an application represented by a micro launcher with a maximum payload of 100 kg. As compared to the design of a classical turbine, this methodology is focused on geometrical limitations, to ensure the manufacturability of the turbine, as well as aerodynamic efficiency. The methodology was applied for obtaining the geometry of a turbine for the aforementioned application using a classical fuel as design point. 3D numerical simulations were computed for this geometry, and the efficiency of the turbine was obtained within 8% of the analytical data. To facilitate the use of different fuels, a simple and fast method was also developed for predicting the performance of a turbine of known geometry and performance for an initial working fluid when changing the nature of this working fluid. Benefiting from having the performance already estimated for another working fluid (the design fluid), the Mach numbers similarity criterion can be used to estimate the performance of the same geometry when changing the working fluid, as a known practice in the gas-turbine field. The Mach number appears as a scaling parameter in many of the equations for compressible flows, shock waves, and expansions. Not only is it suited for similarity of a classical turbine, it is also appropriate for a supersonic turbine, due to the fact the conditions behind shock waves are only dependent on the Mach number and the fluid’s properties. Using this method, the performance of the designed turbine was computed for a completely different working fluid than the design one. The largest difference in power output generated by changing the working fluid is of 170%. A new set of numerical simulations was done, and the results confirmed the validity of the method by obtaining a value of the power output within 6.5% for four other working fluids.

An enabling technology for a successful deployment of the sCO 2 closed-loop recompression Brayton cycle is the development of a high temperature turbine not currently available in the marketplace. This turbine was developed under DOE funding for the STEP Pilot Plant development and represents a second generation design of the Sunshot turbine (Moore, et al., 2018). The lower thermal mass and increased power density of the sCO2 cycle, as compared to steam-based systems, enables the development of compact, high-efficiency power blocks that can respond quickly to transient environmental changes and frequent start-up/shut-down operations. The power density of the turbine is significantly greater than traditional steam turbines and is rivaled only by liquid rocket engine turbo pumps, such as those used on the Space Shuttle Main Engines. One key area that presents a design challenge is the radial inlet and exit collector to the axial turbine. Due to the high power density and overall small size of the machine, the available space for this inlet, collectors and transition regions is limited. This paper will take a detailed look at the space constraints and also the balance of aero performance and mechanical constraints in designing optimal flow paths that will improve the overall efficiency of the cycle.

The recent growth of private options in launch vehicles has substantially raised price competition in the space launch market. This has increased the need to deliver reliable launch vehicles at reduced engine development cost, and has led to increased industrial interest in reduced order models. Large-scale liquid rocket engines require high-speed turbopumps to inject cryogenic propellants into the combustion chamber. These pumps can experience cavitation instabilities even when operating near design conditions. Of particular concern is rotating cavitation, which is characterized by an asymmetric cavity rotating at the pump inlet, which can cause severe vibration, breaking of the pump and loss of the mission. Despite much work in the field, there are limited guidelines to avoid rotating cavitation during design and its occurrence is often assessed through costly experimental testing. This paper presents a source term based model for stability assessment of rocket engine turbopumps. The approach utilizes mass and momentum source terms to model cavities and hydrodynamic blockage in inviscid, single-phase numerical calculations, reducing the computational cost of the calculations by an order of magnitude compared to traditional numerical methods. Comparison of the results from the model with experiments and high-fidelity calculations indicates agreement of the head coefficient and cavity blockage within 0.26% and 5% respectively. The computations capture rotating cavitation in a 2D inducer at the expected flow coefficient and cavitation number. The mechanism of formation and propagation of the instability is correctly reproduced.

In this paper we present and validate a shape optimization framework for the design of splitter blades that extends the operative range under cavitation while maintaining the wetted performance of rocket engine turbopumps. For a target turbopump application, the optimization framework allows for independent changes to the blade angle distributions across the span and to the pitchwise position of the splitter blades while preserving the thickness distributions. The optimization is conducted with a surrogate-based gradient method. The geometry is optimized at a fixed cavitation number corresponding to a 5% head coefficient dropoff, while constraints are imposed on the wet pump performance. It is found that this approach, coupled with the optimal design points distribution provided by the Design of Experiment method, reduces the computational cost of the optimization process by minimizing the number of multiphase calculations. The numerical results suggest that the optimized splitter blades successfully increase the pump operative range by 2.2% and increase the head coefficient by 5.3% compared to the baseline case with non-optimized splitters. These results are corroborated by experiments conducted in a closed-loop water test facility. Several pump geometries are tested through rapid prototyping using additive manufacturing. The experimental data validate the optimization framework, demonstrating a 4.7% increase of pump operative range and a 7.6% increase in head coefficient. The calculations are used to gain insight in the physical mechanisms for the performance improvement. The analysis of the results indicates that the improved performance is due to the optimized position and shape of the splitter blades which increase the pump slip factor.

The article describes a refining method for a fuel pump of rocket powerful turbo-pump unit by the joint usage of mathematical optimization software IOSO, meshing complex NUMECA and CFD complex ANSYS CFX. The optimization software was used for automatic change of the geometry of low-pressure impeller, transition duct and high-pressure impeller to find the optimal design. It was mandatory to keep the original variant of the remaining parts of the pump. For this reason, only geometrical parameters of the blades were varied without changing the contours of the pump meridional flow part. The investigated pump consists of five parts: inlet duct, low-pressure screw centrifugal stage, transition duct, high-pressure screw centrifugal stage and volute outlet duct. The pump main parameters with water as the working fluid (based on experiment data) were the following: high-pressure stage rotor speed was 13300 rpm; low-pressure rotor speed was 3617 rpm by gearbox; inlet total pressure was 0.4 MPa; outlet mass flow was 132.6 kg/s at the nominal mode. Creation of vane unit mesh (rotors and stator transition duct) was performed using NUMECA AutoGrid5. Sector models were used for the calculation simplification. The flow around only one blade or screw was considered. Setting up and solution of the task were carried out in the ANSYS CFX solver. Comparison of calculated characteristics of the basic pump with the experimental data was performed before the optimization. The analysis of characteristics for the obtained optimized pump geometry was carried out. It was found that pump with optimized geometry has greater efficiency in comparison with the original pump variant. The obtained reserve can be used to boost the rocket engine, and/or to reduce the loading of the main turbine, which operates in aggressive oxidizing environment.

Boosters are commonly used in liquid propellant rocket engines (LPRE) to allow lower propellant pressures in their storage tanks and, thus, smaller structural masses, contributing to cavitation free operation in the subsequent main turbopumps (TP). Boosters can be identified as key components for the overall performance of large engines, and if their operating requirements are stringent, they can operate under cavitation. Thus, effective design and performance tools are fundamental to design the components of these boosters considering this phenomenon. The simulation techniques based on turbulent and multiphase 3-D Computational Fluid Dynamics (CFD) were used in this work at steady state regime. The simulations were done using the commercial software CFX from ANSYS ® Workbench. The study was conducted analyzing the performance of the first stage of the hydraulic axial turbine of the liquid oxygen (LOX) booster of the Space Shuttle Main Engine (SSME), at various operation points under cavitation, considering 3.0% tip clearance relative to blade height. The results obtained for, the performance parameters of this stage were compared with those obtained through monophase simulation, and the multiphase technique showed results closer to the experimental ones around the design point (DP), with increased simulation times acceptable for the computational resources currently available. Moreover, the results from the current work show the importance of considering the effects of cavitation through multiphase flow in hydraulic turbines.

This study investigates the behavior of a turbopump assembly during critical cavitation of the propellant pumps in the upper rocket engine of the Korea Space Launch Vehicle-II. Turbopumps operate under conditions involving low pressure at the pump inlet and high rotational speeds to allow for a lightweight design. This severe environment can easily cause cavitation to occur in the pump. This cavitation can then cause the pump operation to fail. As the cavitation number in the pump decreases below the critical point, the pump fails to operate. There is concern regarding the behavior of the turbopump assembly arising from pump failure due to cavitation. It is necessary to verify the problems that may occur if the turbopump assembly operates under extreme conditions, such like the critical cavitation. This study performed tests to investigate the breakdown of pumps in the turbopump assembly. Tests were conducted with liquid nitrogen, water, and high-pressure air instead of the mediums used during actual operation of liquid oxygen, kerosene, and hot gas. The turbopump was tested at the design point of 27,000 rpm, while the inlet pressure of each pump was controlled to approach the critical cavitation number. The turbine power output was maintained during the tests. The results show that the breakdown point of the oxidizer pump using liquid nitrogen, which is a cryogenic medium, occurred at a lower cavitation number than during an individual component suction performance test using water. The fuel pump using water, meanwhile, experiences breakdown at similar cavitation numbers in both tests. As the breakdown of the pump occurs, the power required by that pump decreases, and the rotational speed of the turbopump increases. Compared with individual pump suction performance tests, this breakdown test can be used to determine the limit of the propellant inlet pressure of the turbopump and to characterize the behavior of the turbopump assembly when a breakdown occurs. Vibrations were also analyzed for tests at a high cavitation number and at the critical cavitation number. The vibration increased with breakdown and notable frequencies were analyzed.

A modeling framework defined in the time-domain has been developed to characterize the steady state and dynamic behavior of each component of a typical feeding system for liquid rocket engine. A typical water loop for experimental characterization of liquid rocket turbopumps has been modeled according to the modeling framework in order to understand which is the best way to perform forced experiments for the characterization of the transfer matrix of cavitating turbopumps necessary for understanding the POGO instability phenomena that affect rocket launchers. The best results in terms of capability of generating mass flow rate and pressure oscillations at the inlet of the inducer have been obtained by means of a device that produces a volume oscillation located downstream of the pump.

It is well-known that the natural frequencies of structures immersed in heavy liquids will decrease due to the fluid “added-mass” effect. This reduction has not been precisely determined, though, with indications that it is in the 20–40% range for water. In contrast, the mode shapes of these structures have always been assumed to be invariant in liquids. Recent modal testing at NASA/Marshall Space Flight Center of turbomachinery inducer blades in liquid oxygen, which has a density slightly greater than water, indicates that the mode shapes change appreciably, though. This paper presents a study that examines and quantifies the change in mode shapes as well as more accurately defines the natural frequency reduction. A literature survey was initially conducted and test-verified analytical solutions for the natural frequency reductions were found for simple geometries, including a rectangular plate and an annular disk. The ANSYS© fluid/structure coupling methodology was then applied to obtain numerical solutions, which compared favorably with the published results. This initial study indicated that mode shape changes only occur for non-symmetric boundary conditions. Techniques learned from this analysis were then applied to the more complex inducer model. ANSYS numerical results for both natural frequency and mode shape compared well with modal test in air and water. A number of parametric studies were also performed to examine the effect of fluid density on the structural modes, reflecting the differing propellants used in rocket engine turbomachinery. Some important findings were that the numerical order of mode shapes changes with density initially, and then with higher densities the mode shapes themselves warp as well. Valuable results from this study include observations on the causes and types of mode shape alteration and an improved prediction for natural frequency reduction in the range of 30–41% for preliminary design. Increased understanding and accurate prediction of these modal characteristics is critical for assessing resonant response, correlating finite element models to modal test, and performing forced response in turbomachinery.

Due to the critical importance of the turbopump applied in Liquid-Propellant Rocket Engines (LPRE) and the importance in the use of specific engineering software to design and analyze turbomachines, a Project-Based Learning (PBL) methodology was implemented in the undergraduate Turbopumps (TP) discipline at the Aeronautics Institute of Technology (ITA), taught for aerospace engineering students. This methodology was applied, using as a class example, the Liquid Oxygen (LOX) booster turbine of the Space Shuttle Main Engine (SSME), aiming at an enhancement in the discipline’s syllabus, to become the theory and practice closer to the real engineering, and to increase the discipline’s attractiveness. The results obtained with this methodology showed that the students have more interest and attention in the classes in which an engineering problem is evaluated and discussed with details using appropriate examples and engineering software that are used by the academia and industry. Several turbomachines issues as velocity triangles, power, blade geometrical aspects, flow quality, losses and in this case, the importance of tip clearance, could be better understood by the students. About the numerical results, the aim is that the students, after the preliminary project ends, evaluate the results and compare them with experimental data from National Aeronautics and Space Administration (NASA). One of the most important experience in this project is the results evaluation by the students and the discussion around it, as lessons learned, given suggestions to improve the project, if the results are not in the right way what can be done to correct them and understanding all physical phenomena involved. The learning experience was fascinating and effective, as noticed by students and noted by Professors.

Start-transient testing of a hybrid (combined hydrostatic and hydrodynamic action) bearing supplied with air was completed, providing an indication of its performance while operating in a compressible fluid medium. The test start transients were modeled after Rocket Engine Transient Simulation Software (ROCETS) predictions for start-transient behavior of running speed ω( t ) and bearing supply pressure P s ( t ). The top test speed was 15 krpm. The ramp rate, supply pressure P s values at 15 krpm, constant bearing unit load magnitude w 0 , and load orientation (load-on-recess LOR or load-on-land LOL) were varied. Five different load-case combinations were carried out (separately) for LOR and LOL load configurations with ramp rates varying from 2206 rpm/s to 8824 rpm/s. The target pressures at 15 krpm varied from 5.32 bars to 18.25 bars. The tested bearing dimensions were: L = D = 38.1 mm, and C r =.0635 mm. Lift-off occurs due to the increase in P s (ω dependent) and was defined as the point of departure towards the center of the bearing with increasing ω while the rotor remained 0.00254 mm (0.1 mils ) above the bearing surface. This method is limited by the inability to accurately measure an established operating bearing clearance. Evaluation of the lift-off P s versus applied unit load w 0 supports the following conclusions: (1) Lift-off P s is approximately a linear function of w 0 , (2) Changing the ramp rate while keeping constant the specified P s at 15 krpm has no significant impact, (3) Lowering the limit P s at 15 krpm may reduce the lift-off P s value, and (4) The LOR start-transient cases required a higher lift-off speed and lift-off P s values than the corresponding LOL start-transient cases.

Process fluid lubrication of rotating machinery offers advantage of compactness and efficiency while dispensing with complicated oil lubricant supply systems. Prior work in a dedicated test rig demonstrated the performance of water lubricated radial and thrust bearings into high speed and high load conditions. The application related to a high performance rocket engine turbo pump. The test rig was revamped to operate with gas bearings in a program aiming to measure the performance of gas thrust bearings. The gas bearings for lateral support of the rotor are of hybrid type (hydrostatic/hydrodynamic) with flexure pivots and multiple ports for inlet gas pressurization. The paper details the design of the flexure pivot bearings and predictions of the lateral rotordynamics of the rotor supported on the hybrid gas bearings. Troubleshooting operation of the test rotor supported on the novel gas bearings followed with preliminary runs with the bearings supplied with air at 7.9 bar, then 6.5 bar and at 5.1 bar, and shaft speeds to 25 krpm (surface speed=50 m/s). The data recorded showed a very lightly damped system with a critical speed at ∼6 krpm, and susceptible to excite sub synchronous whirl motions when operating above the first critical speed. Ignoring the initial warnings, the operator persisted in operating the rotor to a high speed of 28 krpm while lowering the air supply pressure to 5.1 bar into the bearings. Suddenly, the shaft experienced large amplitude sub synchronous whirl motions, contacted the bearings, and produced a catastrophic failure. The incident produced much damage including a broken coupling, a twisted rotor, sheared covers, and welded pads into the bearing casing. Post-mortem analysis shows the failure is due to a sub synchronous whirl instability of the first rigid body rotor-bearing mode also exacerbated by the rotor approaching second natural frequency of the rotor-bearing system. The rotordynamics model includes the rotor rigidly connected to a long quill shaft and coupling produces results in agreement with the last vibration data set acquired prior to the incident. The experience demonstrates the need for following proper operating procedures while also paying attention to early evidence that could have prevented the mishap.

As global demand for energy increases while environmental regulations tighten, novel power generation cycles are being developed to meet market needs while accommodating green requirements. The Allam cycle is an approach (with high pressure, low pressure ratios, oxygen-fuel combustion and CO2 as a working fluid) that efficiently produces power in a compact plant, avoids NOx emissions, makes efficient use of clean-burning methane (natural gas) and can generate high-pressure carbon dioxide for enhanced oil and gas recovery in the field. The cycle requires oxy-fuel-CO2 combustion at approximately 30MPa and 1150° C turbine inlet temperatures. Due to its relatively compact size (owing to the high operating pressures), the Allam cycle technology can be implemented for a low cost in relative existing power plants, and given the current economic climate, natural gas is the most financially appealing application. Meeting environmental regulations for decades to come, the cycle can dramatically lower the cost of electricity. The critical path for plant implementation lies in demonstrating efficient, stable combustion at 30MPa. Similar pressures have been demonstrated in rocket engine systems, where reusability typically means 3 to 4 hot fires and surviving 8 to 32 minutes lit. The technical challenge here lies in designing durable hardware capable of thousands of hours of continuous operation and efficient combustion over the range of conditions required in a power plant. Production of clean CO2 exhaust (ideally, free of O2 and CO) is another challenge that will drive plant viability.

In connection with the Coal Gasification Program of the U.S. Department of Energy, a need was identified for a high-pressure, high-volume oxygen compressor. At the time, the highest steady operating pressure available from commercial-sized turbocompressors was about 65 bar (950 psia), which was being used in the partial oxidation process for ammonia or methanol synthesis. For advanced coal conversion plants, compressors operating at 115 bar (1667 psia) would be needed. The U.S. Department of Energy (DOE) entered into an agreement with the West German firm of Mannesmann Demag, AG, to develop a high-pressure oxygen compressor. Mannesmann Demag designed and fabricated a third casing centrifugal compressor (to be used in a train with available compressors) having a nominal inlet pressure of 65 bar (950 psia) and a nominal discharge pressure of 115 bar (1667 psia). In exchange, the U.S. DOE agreed to fund the modification and operation of a NASA-owned test facility for testing the compressor. The facility, CTL-V, was modified and operated by the Rocketdyne Division of Rockwell International. Test management and reporting was carried out by DOE’s Energy Technology Engineering Center (ETEC), operated by Rockwell International’s Energy Systems Group. This paper presents a brief description of the compressor test article and test facility. The test program, test results, and analysis are presented in greater detail. The scope of the test program included low- and high-pressure nitrogen shakedown tests, oxygen performance and steady-state tests, and oxygen maximum suction pressure tests. The facility is unique because it is the only one of its kind nationally that can provide the power to drive the compressor and that has liquid oxygen/liquid nitrogen capability (the facility is utilized for development and testing of rocket engine turbopumps). Also, testing and data acquisition, reduction, and analysis were performed in a facility that normally runs tests of few minutes duration (for rocket engine components) but that had to be converted to run endurance tests lasting several hundred hours.