Expander Cycle Engine Research Articles

Comparison between the dynamic characteristics of electric pump fed engine and expander cycle engine

A successful space launch mainly depends on the start-up process and variable operating conditions of rocket propulsion system. With the development of motor and battery technology, new electric pump is gradually used, in place of the traditional turbine pump, in rocket propulsion system. However, it is still unclear about the dynamic characteristics of the two processes in the two schemes (electric pump scheme and turbine pump scheme). In this research, simulation models of the basic components and overall system of an expander cycle hydrogen-oxygen engine were established on the basis of turbine pump and electric pump respectively. In the simulation models, the Centralized Parameter Method and the Stepwise Parameter Method are used for different components. The research results show that the power of hydrogen pump in the electric pump scheme is 25.5% lower than that in the turbine pump scheme, with speed and torque respectively 10.5% and 16.6% lower than those in the latter scheme. In the start-up process of the electric pump scheme, the flow of hydrogen pump underwent an overshoot of 3.7%, and the settling times of hydrogen and oxygen flows were both more than 40% shorter than those in the turbine pump scheme. Under variable operating conditions of turbine pump scheme, the speed of pump experienced an overshoot of about 20%, and the settling times of hydrogen and oxygen flows were 37% and 21% longer than those in the electric pump scheme. These findings are helpful to analyze the said two processes of expander cycle hydrogen-oxygen engine and further understand the difference in dynamic characteristics between the electric pump scheme and the turbine pump scheme.

Preliminary Results for In-Situ Alternative Propellants for Nuclear Thermal Propulsion

Water, ammonia, and other volatiles that can be used for propellant have been found on the Moon, and the technology that will be used to extract them has been laboratory tested. One of the considered propulsion systems for a crewed mission to Mars is nuclear thermal propulsion (NTP). However, current reference missions consider hydrogen as the main propellant, which is technologically difficult to store. Electrolysis units are required to process the lunar water to separate it into oxygen and hydrogen, which is only 1/8 of the mass of water mined. Due to these challenges, a preliminary analysis of alternative propellant nuclear thermal propulsion (A-NTP) expander cycle engines was made. A-NTP engine models that produced 25 000 lbf of thrust, which is comparable to the baseline hydrogen NTP engines, were constructed in Simulink for preliminary analysis, which yielded an Isp of 320.4 s for water and 381.6 s for ammonia. Although this Isp is lower than the most efficient chemical engines, since water and ammonia are used directly and are stored as such, a propellant tank volume decrease of up to 76.1% for water and 69.5% for ammonia is possible. This will decrease the number of launches, given that the tanks are not fully tanked at time of launch and lunar resources are used to fill the tanks completely.

Investigation of a Radial Turbines Compatibility for Small Rocket Engine

In the radial turbine preliminary design for an expander rocket engine, a comparison was made with axial turbine used in Pratt & Whitney RL10 engine. One of the primary requirements of a liquid propellant rocket engine is the generation of a high thrust, which depends on both the mass flow rate of the propellant and the pressure in the thrust chamber. In expander-cycle engines, which are the subject of the present study, the liquid propellant is first compressed using centrifugal turbo-pumps, then it is used to cool the combustion chamber and the nozzle and, once vaporized, it flows through the turbines used to drive the turbo-pumps. The aim was to demonstrate the greater efficiency of the radial turbine with a reduction of the pressure ratio with respect to the axial turbine.

Test facilities for VINCI®

With the replacement of the current upper-stage ESC-A of the Ariane 5 launcher by an enhanced cryogenic upper-stage, ESA’s Ariane 5 Midterm Evolution (A5-ME) program aims to raise the launcher’s payload capacity in geostationary transfer orbit from 10 to 12 tons, an increase of 20 %. Increasing the in-orbit delivery capability of the A5-ME launcher requires a versatile, high-performance, evolved cryogenic upper-stage engine suitable for delivering multiple payloads to all kinds of orbits, ranging from low earth orbit to geostationary transfer orbit with increased perigee. In order to meet these requirements the re-ignitable liquid oxygen/liquid hydrogen expander cycle engine VINCI® currently under development is designated to power the future upper stage, featuring a design performance of 180 kN of thrust and 464 s of specific impulse. Since 2010 development tests for the VINCI® engine have been conducted at the test benches P3.2 and P4.1 at DLR test site in Lampoldshausen under the ESA A5-ME program. For the VINCI® combustion chamber development the P3.2 test facility is used, which is the only European thrust chamber test facility. Originally erected for the development of the thrust chamber of the Vulcain engine, in 2003 the test facility was modified that today it is able to simulate vacuum conditions for the ignition and startup of the VINCI® combustion chamber. To maintain the test operations under vacuum conditions over an entire mission life of the VINCI® engine, including re-ignition following long and short coasting phases, between 2000 and 2005 the test facility P4.1 was completely rebuilt into a new high-altitude simulation facility. During the past two P4.1 test campaigns in 2010 and 2011 a series of important milestones were reached in the development of the VINCI® engine. In preparation for future activities within the frame of ESA’s A5-ME program DLR has already started the engineering of a stage test facility for the prospective upper stage. The new test facility P5.2 is to perform the qualification of the anticipated upper stage with the VINCI® engine. In the past year DLR has started the design phase for these modifications. The main design drivers are the test configuration and operation domain described in the test request.

Parametric Analysis of Cooling Properties of Candidate Expander-Cycle Fuels

Flow evolution and heat transfer capability in the cooling system of liquid rocket engines heavily depend on propellant thermophysical properties. Coolant thermophysical property analysis and modeling is therefore important to study the possibility of relying on a regenerative cooling system, whose performance is crucial to determine feasibility and convenience of pump-fed liquid rocket cycles of the expander type. The aim of the present study is to compare the behavior of different liquid fuels for expander-cycle engines. They are light hydrocarbons, binary mixtures of them, and liquefied natural gas, which is a mixture made basically of methane and minor fractions of other light hydrocarbons and nitrogen. A parametric analysis is carried out by a validated numerical solver to compare temperature increase, pressure loss, and heat transfer evolution for the different fuels along the same straight tube and subjected to assigned heat fluxes. Results show that similar engine performance can be obtained by the different candidate expander-cycle fuels, but significant differences can be seen in the flow evolution through the cooling channels.

Parametric Analysis of Heat Transfer to Supercritical-Pressure Methane

Parametric Analysis of Heat Transfer to Supercritical-Pressure Methane

Measurements of the Effectiveness of Concave Spherical Dimples for Enhancement Heat Transfer

Measurements of the heat transfer enhancement effects of concave spherical dimples have been performed under hydrogen–oxygen hot-fire conditions in a subscale combustion chamber. Seven configurations were tested with a range of area densities from 15–70% and depth-to-diameter ratios of 0.111 to 0.333. Enhancements of 15–82% relative to a smooth surface weremeasured. TheReynolds number was varied by a factor of 5 andwas found to have an insignificant effect on the enhancement factor.Results correlatedwellwith previous studies conductedunder coldflow conditions with similar values of area fraction, depth-to-diameter ratio, and channel height-to-dimple diameter. These results are suitable for use by analysts and designers considering the use of heat transfer enhancement features on hot-gas walls for example in the design of high-performing expander cycle engines. Concave spherical dimples may have advantages over other approaches for enhancing heat transfer, such as protruding ribs with respect to the integrity of the chamberwall and the efficiency of the heat transfer enhancement relative to the associated increase in skin friction.

Stabilization of non-premixed flames in supersonic reactive flows

A Lagrangian framework is set out to describe turbulent non-premixed combustion in high speed coflowing jet flows. The final aim is to provide a robust computational methodology to simulate, in various conditions, the underexpanded GH2/GO2 torch jet that is used to initiate combustion in an expander cycle engine. The proposed approach relies on an early modelling proposal of Borghi and his coworkers. The model is well suited to describe finite rate chemistry effects and its recent extension to high speed flows allows one to take the influence of viscous dissipation phenomena into account. Indeed, since the chemical source terms are highly temperature sensitive, the influence of viscous phenomena on the thermal runaway is likely to be all the more pronounced since the Mach number values are high. The validation of the extended model has been recently performed through the numerical simulation of two distinct well-documented experimental databases. Only a brief summary of this preliminary validation step is provided here. The main purpose of the present work is to proceed with the numerical simulation of geometries that bring together the essential peculiarities of the underexpanded GH2/GO2 torch. The behavior of the corresponding supersonic coflowing jet flames for various conditions is discussed in the light of computational results. To cite this article: J.-F. Izard, A. Mura, C. R. Mecanique 337 (2009).

Recent Advances in LOX / LH2 Propulsion System for Reusable Vehicle Testing

The third-generation vehicle RVT#3 equipped with a pressure-fed engine, which had upgraded in terms of durability enhancement and a LH2 tank of composite material, successfully performed in repeated flight operation tests; and the vehicle reached its maximum flying altitude of 42m in October 2003. The next step for demonstrating entire sequence of full-scale operation is to put a turbopump-fed system into propulsion system. From a result of primary system analysis, we decided to build an expander-cycle engine by diverting a pair of turbopumps, which had built for another research program, to the present study. A combustion chamber with long cylindrical portion adapted to the engine cycle was also newly made. Two captive firing tests have been conducted with two different thrust control methods, following the component tests of combustor and turbopumps separately conducted. A considerable technical issues recognized in the tests were the robustness enhancement of shaft seal design, the adjustment of shaft stiffness, and start-up operation adapted to the specific engine system. Experimental study of GOX/GH2 RCS thrusters have also been started as a part of a conceptual study of the integration of the propulsion system associated with simplification and reliability improvement of the vehicle system.

A feasibility study of a new ATREX engine system of aft-turbine configuration

A feasibility of ATREX (Air-Turbo-Ram Expander cycle) engine with conventional aft-turbine configuration has been studied to be developed in about 10 years, if the development project has started under enough resources. The novel tip-turbine of the original ATREX engine is replaced by a conventional aft-turbine, and the maximum turbine inlet temperature (TTT) is reduced to 1200K, to realize the engine by only using approved metal technologies of modern jet engines. The capability of the performance has been shown by parametric studies by changing components' design parameters. The study shows that the performance of the ATREX engine is not less than that of pre-cooled turbo jet. Some technical issues on developing the new ATREX engine have been addressed. The most important issue would come from the transient total temperature change due to the rapid acceleration from sea level static (SLS) condition (288K) to Mach 6 at 30km of altitude (1680K) in 6 minutes. The deformation due to transient thermal expansion has to be controlled. Especially, the change of the tip clearance and the clearance between rotors and stators are pointed out to be important design issues. The ATREX engine, which has shorter axial length and simpler rotor, has structural advantage over turbo jet.

Development Study of a Precooler for the Air-Turboramjet Expander-Cycle Engine

A review of the development of a precooler for the air-turboramjet expander-cycle (ATREX) engine is given. Three types of precooler for the ATREX engine ground-testmodel were designed, manufactured, and tested under sea-level static conditions. The results suggested two problems affecting the precooler performance, heat transfer rate and air ow pressure drop. One is nonuniformity of the air ow through the tube banks. The other problem is frost formation on the heat transfer surfaces. Concerning nonuniformity of air ow, the shell conŽ guration was modiŽ ed based on analysis by computational  uid dynamics calculation. To improve the precooler performance under frosting condition, a new method to add a condensable gas into the air ow was proposed and examined by experiments on a subscale heat exchanger model. Addition of a small quantity of ethanol can effectively restrain the decline of the precooler performance due to frost formation.

A Computational Model for Thermal Fluid Design Analysis of Nuclear Thermal Rockets

A computational model for simulation and design analysis of nuclear thermal propulsion systems has been developed. The model simulates a full-topping expander cycle engine system and the thermofluid dynamics of the core coolant flow, accounting for the real gas properties of the hydrogen propellant/coolant throughout the system. Core thermofluid studies reveal that near-wall heat transfer models currently available may not be applicable to conditions encountered within some nuclear rocket cores. Additionally, the possibility of a core thermal fluid instability at low mass fluxes and the effects of the core power distribution are investigated. Results indicate that for tubular core coolant channels, thermal fluid instability is not an issue within the possible range of operating conditions in these systems. Findings also show the advantages of having a nonflat centrally peaking axial core power profile from a fluid dynamic standpoint. The effects of rocket operating conditions on system performance are also investigated. Results show that high temperature and low pressure operation is limited by core structural considerations, while low temperature and high pressure operation is limited by system performance constraints. The utility of these programs for finding these operational limits, optimum operating conditions, and thermal fluid effects is demonstrated.

Delta improvement study - Hydrogen upper stage

Introduction T HE Delta commercial launch vehicle has had a long and successful life. It was first launched May 13, 1960 and the mission was unsuccessful. The Echo 1A satellite was carried into orbit on the second Delta mission. Between May 1960 and November 1990, the Delta completed 189 of 201 missions successfully, giving it 94% reliability. The version of the rocket that was considered for upgrade in this study was the Delta II 7925. It first flew on Nov. 26, 1990. The Delta II 7925 consists of three stages and uses nine external solid rocket motors. Propulsion for stage 1 is provided by a liquid core engine burning RP-1 fuel and oxygen. Core stage 2 is a storable liquid core and uses nitrogen tetroxide and A-50 propellants. Stage 3 consists of a solid rocket motor used to insert the payload into geosynchronous transfer orbit (GTO). A payload of 4010 Ib can be placed into GTO using the Delta II 7925 with a 9.5-ft payload fairing. The study reported here is aimed at helping to reduce the use of toxic storable propellants. Special safety steps needed for these chemicals would be eliminated and operations costs would fall as a result. By using a hydrogen and oxygen upper stage, stages 2 and 3 could be eliminated from the Delta II. Figure 1 illustrates this modification. Five engines were examined in the course of this research for use in the hydrogen replacement stage. Included in the research were the following engines: an RL10C, an RL10A-4, an advanced technology low cost (ATLC) engine, an advanced expander cycle engine, and an engine with a simplified injector.

Advanced tube-bundle rocket thrust chambers

An advanced rocket thrust chamber for future space application is described along with an improved method of fabrication. Potential benefits of the concept are improved cyclic life, reusability, and performance. Performance improvements are anticipated because of the enhanced heat transfer into the coolant which will enable higher chamber pressure in expander cycle engines. Cyclic life, reusability and reliability improvements are anticipated because of the enhanced structural compliance inherent in the construction. The method of construction involves the forming of the combustion chamber with a tube-bundle of high conductivity copper or copper alloy tubes, and the bonding of these tubes by an electroforming operation. Further, the method of fabrication reduces chamber complexity by incorporating manifolds, jackets, and structural stiffeners while having the potential for thrust chamber cost and weight reduction.

Status of advanced propulsion for space-based orbital transfer vehicle

This paper discusses the NASA technology program for Advanced Orbital Transfer Vehicle Propulsion. The program requirements and goals for a high performance, multiple restart, variable thrust, orbital transfer propulsion system which could be man-rated, space-based, and compatible with aeroassist maneuvering are presented. Three expander cycle engine concepts and critical technologies are discussed. Progress toward establishing technical feasibility of the concepts is reviewed for trubomachinery, combustion devices, health monitoring, space-based maintenance, and integrated component testing.

Development Study of LOX/LH2, High Pressure Expander Cycle Engine

Development Study of LOX/LH2, High Pressure Expander Cycle Engine

Preliminary experiment on LOX/LH2 high pressure expander cycle engine

The expander cycle has high performance and high reliability. This has been verified by the Pratt & Whitney RL 10 engine that is used to power the Centaur upper stage. Some limitations, however, exist in this cycle on increasing both chamber pressure and thrust because it uses only the heat energy absorbed by the fuel as it cools the thrust chamber and nozzle extension to power the fuel and oxidizer turbopumps. We have proposed the advanced concept chamber with an internal heat exchanger at the 35th IAF congress, Lausanne in order to pick up the large amount heat from combusion energy. This paper describes the brief features of high pressure expander cycle engine using an internal heat exchanger and the results of preliminary tests on the thrust chamber with internal heat exchanger. The average heat flux in heat exchager became about 70% of the value calculated by Bartz's correlation. The combusion performance was not affected significantly by installing a heat exchanger within a combustion chamber.

The study of high pressure expander cycle engine with advanced concept combustion chamber

This paper discusses the feasibility of the high pressure LOX/LH2 expander cycle engine which delivers 147 kN of thrust at 9.8 MPa of chamber pressure with a new concept combustion chamber. The expander cycle has the advantages of high performance and system simplicity, but it has not been applied for high pressure engine due to the existence of an upper limit in the chamber pressure to be achieved. In the expander cycle engine, the maximum chamber pressure is dominated by the amount of heat absorbed by the regenerative coolant. The larger amount of heat exchange requires the longer combustion chamber under the usual idea, which causes a decrease of performance because of reducing expansion ratio under the given engine envelope or increasing vehicle length. In order to achieve the higher chamber pressure without any decrease in performance, higher heat exchangability in the chamber is necessary. We propose the new concept combustion chamber in which higher heat exchangeability is realized by having the additional heat exchanger installed within it. Preliminary tests for the internal heat exchanger and for the expander cycle engine will by carried out in 1985 by ISAS.