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The dual-bell nozzle is a promising altitude adaptive nozzle concept for improving the performance of existing space launchers as well as future reusable launch vehicles. It is characterized by an inflection in the wall geometry which links a base nozzle with a nozzle extension. Two working modes are therefore attained: a low-altitude mode in which the expansion is limited to the first bell and a high-altitude mode in which full expansion is achieved. Even though this capability makes it possible to reduce the non-adaptation losses in engines working from sea level to almost vacuum conditions, the natural transition between the two modes usually takes place prematurely during the ascent trajectory and destructive side-loads may arise. This results in a limitation on the performance gain. In this work, a dual-bell nozzle design capable of upgrading the parallel-staged Ariane 5 launcher is studied. Fluidic control is investigated as a possible solution to limit premature transition and reduce side-loads. The control is obtained by means of a secondary injection in the proximity of the inflection point: different injection strategies are investigated. Reynolds-averaged Navier–Stokes equations at steady state conditions are solved with an in-house CFD tool, to investigate the flow field and determine the required secondary mass flow rate and the range of nozzle pressure ratio for which the control is effective.

A sub-scale test campaign has been conducted in the framework of cooperation between the German Aerospace Center (DLR) and the Japan Aerospace Exploration Agency (JAXA). Both partners had great experience in cold flow dual bell nozzle investigation. The present campaign took place at P6.1 test facility at DLR Lampoldshausen with a common combustion chamber. Two test specimens were tested with similar conditions and measurement systems. The conditions provided at the bench were the combustion of gaseous methane with liquid oxygen at pressure up to 60 bar. The combustion chamber pressure was varied up and down with constant ramps while keeping the mixture ratio constant. Objective of the study was, a part from demonstrating the precision of the design tool, to investigate the sensitivity of the transition to flow property variations (like the one due to the mixture ratio). The transition and retransition conditions, the range of hysteresis and the transition velocity were of particular interest for DLR side. The work of JAXA concentrates on heat flux and combustion instabilities and is presented in the second part of this work.

Single-stage-to-orbit (SSTO) rocket technology offers the potential to substantially reduce launch costs, but is not considered practical for conventional launch vehicles. However, new research in composite propellant tank technology opens the field for reevaluation. One technology that increases the efficiency and feasibility of SSTO flight is an altitude compensating rocket engine nozzle, as opposed to a conventional constant-area, bell nozzle. By implementing an altitude compensation nozzle, such as an aerospike nozzle for inatmosphere flight, the propellant mass fraction (PMF) may be reduced by as much as two percent compared to a conventional rocket engine. In this paper, Optimal Trajectories by Implicit Simulation (OTIS) is used to model SSTO flight trajectories by comparing an aerospike nozzle to a conventional bell nozzle; this includes vacuum thrust, vacuum specific impulse, and nozzle combinations to show that nozzle variability increases the efficiency of SSTO flight through a reduction in PMF. Results suggest that limiting the altitude compensation to lower altitudes further decreases PMF by 0.1-0.2%. Uncertainty estimates suggest that this added benefit is within the range of uncertainty of OTIS, and performs at least as well as the theoretical, infinite nozzle expansion. It is apparent that altitude compensating nozzles outperform conventional bell nozzle designs.

The proliferation of earth orbital missions enabled by nanosatellites, of which CubeSats are the most common, has focused increased attention on low-cost, mobile launch systems. Advanced nozzle design can address the demand for higher booster performance, and has resulted in the investigation of alternate nozzle geometry. A dual-bell nozzle represents a novel approach to this problem by utilizing two theoretically ideal design points over the ascent trajectory instead of just one. The primary focus for this project was to investigate the performance of a representative dual-bell nozzle both computationally and experimentally. First, a dual-bell nozzle was designed which could be used on a representative, nanosatellite launch system with design altitudes of 3 km and 17 km. The nozzle contours were modified to facilitate machining, and incorporated in a supersonic wind tunnel. The dual-bell nozzle fabricated for testing in the wind tunnel had a “first contour” expansion ratio of 1:9.85 corresponding to an ideal design pressure ratio of 134. The downstream, “second contour” had an expansion ratio of 1:19.7 corresponding to an ideal design pressure ratio of 374. The conventional nozzle fabricated for testing in the wind tunnel had an expansion ratio of 1:11.3 corresponding to an ideal design pressure ratio of 102. Imaging of flow structures was achieved using a schlieren optical system. In addition, the computational fluid dynamics program ANSYS Fluent was used to model both a hypothetical, full-scale nozzle as well as the scaled version tested in the wind tunnel. Simulations were performed for full-scale, dual-bell nozzles operating at pressure ratios (chamber to ambient) of 20, 50, 100, 150 and 1000. These results were then compared to a similar conventional nozzle for each case. Preliminary schlieren results indicate that the scale wind tunnel dual-bell contours show flow structures consistent with the Fluent results.

The dual bell nozzle belongs to the family of Altitude Compensating Nozzles (ACN), which constitutes an option of great interest to improve launcher first stage performance. The internal flow of this nozzle adapts to the external pressure by separating at the wall inflection at low altitude (first operating mode) and by full flowing at high altitude (second operating mode). Therefore the dual bell nozzle operates with a separated flow at take off and in the initial portion of the flight trajectory. This characteristic is of important concern since the external flow is neither steady nor axisymmetric, and its coupling with the internal flow separation can cause dangerous side loads. This work presents the results of a time accurate numerical analysis of the effect of unsteady ambient pressure on the separation point and on the shock system inside the dual bell nozzle, operating in the first mode. The external computational domain is characterized by an acoustically open downstream boundary. The numerical simulations show that the oscillation of the separation point is very sensitive to frequencies near the acoustic wave characteristic frequencies.

The paper presents the design of a 30kN multi-thruster annular aerospike engine as candidate for the first stage of a two-stage pressure-fed Nanosat Launch Vehicle (NLV). Detailed description of the injector and layout of single thruster are provided. A major enhancement compared to a classical multi-chamber design with straight thrusters is achieved by introducing nozzle curvature in the circumferential and in the longitudinal direction. This approach allows increasing the available exit area, or pressure ratio, thus improving performance. The design work is supported by numerical analyses of the flow inside the curved thrusters and along the engine including slipstream effects using Reynolds-Averaged Navier-Stokes simulations with realizable k-e turbulence modeling. The computations indicate an improved performance of the dual curved thrusters engine, which is about 2% above the linear design and above an ideal bell nozzle of equal area ratio. Simultaneously, the use of a larger expansion ratio reduces base drag in the transonic flight regime.

The altitude adaption of the dual bell nozzle is achieved by the contour inflection, leading to a symmetrical and controlled separation under sea level conditions. When reaching a certain pressure ratio, the transition to high altitude mode takes place and the nozzle flows full. Former experimental studies have shown the existence of an intermediate flow state before the transition from one operating mode to the other takes place: the sneak transition, characterized by the separation of the flow downstream to the inflection, leading to increased side loads. To minimize this effect, various contour inflection geometries have been investigated. The study was divided into a cold flow and a warm flow test campaign on sub-scale planar models.

Flow instabilities under low-altitude operation mode in dual-bell nozzles are examined. Influences of base nozzle contour are investigated with two types of base nozzle contours: a truncated perfect (TP) nozzle and a compressed truncated perfect (CTP) nozzle. Experimental and numerical results show that the Mach disk oscillation in the axial direction occurs in the dual-bell nozzle with the CTP base nozzle. Such Mach disk oscillation causes higher level wall pressure oscillations than those in the dual bell nozzle with the TP base nozzle. Influences of geometric area ratio of the dual bell nozzle are also examined. As a result, the geometric area ratio does not affect the characteristics of the flow oscillation modes except for the amplitude of the wall pressure fluctuations.

The altitude adaptive dual bell nozzles can increase the overall performance of existing main stage rocket engines like Europe’s Vulcain 2. The Institute of Space propulsion of the German Aerospace Center in Lampoldshausen carried out in the past various cold and warm flow subscale tests to study the flow physics of dual bell nozzles fundamentally. Following this line and in order to increase the technology readiness level hot flow subscale tests were conducted to study the effects of a cooling film on the characteristics of a dual bell nozzle. The experimental results of a first campaign, performed with a representative Vulcain 2 like subscale engine, are presented and discussed.

The dual bell nozzle offers a simple and efficient altitude adaption through its contour inflection, which insures a symmetrical and controlled separation at sea level and a large area ratio at high altitude. The conditions for the flow transition from one operating mode to the other have been intensively investigated in cold flow conditions. A test campaign has been conducted on two dual bell nozzle geometries to verify the theoretical prediction criterion and to investigate the influence of the temperature variations. Therefore, a thin walled axisymmetric and a planar nozzle were designed and tested under hot flow conditions. Wall temperature measurements have been made using both thermocouples and thermal imaging. The region of the contour inflection was of particular interest for the evaluation of thermal behavior in the nozzle wall.

An aerospike nozzle has the ability to perfectly expand the combustion gas to ambient conditions at all altitudes maximizing the efficiency of the nozzle and increasing the performance of the overall system. However, one of the inherent problems with this type of nozzle is the large amount of heat transferred to vital components due to its complex geometry, which could cause performance degradation or even nozzle failure. An aerospike nozzle design was developed that incorporates graphite inserts at the nozzle throat which reduces substantial material erosion at vital areas of the nozzle. This design was integrated with a custom hybrid rocket motor utilizing nitrous oxide and hydroxyl-terminated polybutadiene (HTPB) providing an excellent reusable test bed producing increased performance over that of a solid rocket motor. The unique integration between the hybrid and the aerospike provides a high performance rocket propulsion system capable of consistent experimental testing. A series of hot fire static tests were conducted in order to demonstrate that the aerospike nozzle can not only withstand the intense combustion gas temperatures, but furthermore yield a design that has higher efficiency over a range of design altitudes in comparison to a standard bell nozzle with fixed expansion ratio geometry. The aerospike nozzle was able to outperform an over-expanded bell nozzle by 6.4%, and perform within 0.9% of a perfectly expanded nozzle based on each nozzle’s CF. These tests prove that utilization of an aerospike nozzle has the potential to substantially increase the range or payload capabilities of modern launch vehicles or rockets.

An experimental investigation has been carried out to study the flow characteristics of a dual-bell nozzle during its transition modes. Cold gas tests using gaseous nitrogen were carried out in (i) a horizontal test rig with nozzle exhausting into atmospheric conditions and, (ii) also inside a high altitude simulation chamber with nozzle operation under self-evacuation mode. Transient tests indicate that increasing 0P0 /0t reduces the transition time as well as the amplitude of pressure fluctuations of separation shock at wall inflection point. Sea-level tests show 15-17% decrease in transition NPR during subsequent tests in a single run primarily due to frost formation in the nozzle extension up to the wall inflection location. This reduces the wall inflection angle and hence, the transition NPR. However, tests inside the altitude chamber show nearly constant NPR value during subsequent runs primarily due to decrease in back temperature with decrease in back pressure that prevents any frost formation.

*† ‡ For a dual-bell nozzle to be employed in a future reusable launch vehicle, active control of working mode transition is essential. To establish a technique for such control, the effectiveness of secondary injection was investigated by conducting cold-flow tests. The secondary injection was introduced from the inflection position of the nozzle contour perpendicular to the thrust axis. As a result, the secondary injection successfully changed the transition condition. Based on the results, a system and a sequence to realize the active control of the transition between two working modes of a dual-bell nozzle are herein discussed.

Dual bell nozzles represent a possible solution to improve the performance of large liquid rocket engines for launcher first stages. During the low-altitude mode, the flow is separated and the separation line is located near the inflection of the nozzle wall, in a region characterized by a negative value of the wall pressure gradient and where side loads may occur. At high altitude the flow attaches to the nozzle wall and the full area ratio is used. The flow behavior during the transition between the two operating modes is analyzed by time-accurate simulations carried out on a simplified planar geometry and considering two configurations: one with a wall film cooling and the other without film. Aim of this study is to analyze the effect of the secondary gas injection on the dynamics of the separation point, the transition time and the velocity of the separation front.

The dual bell nozzle is a concept of altitude adaptive nozzles. The flow adapts to the altitude by separation at the wall inflection at low altitude, and full flowing at high altitude. To understand the phenomenology of the flow by the transition from sea level to high altitude mode, a series of tests have been made at the cold flow test bench P6.2 at DLR Lampoldshausen. Three nozzles with different geometries have been tested. Two of them were successively shortened and driven unter the same conditions for each extension length. This study yields the influence of the geometric parameters of the basis and the extension on the transition conditions. Furthermore a transition prediction is given.

In this study we numerically examine thrust production and efficiency of 3D supersonic micronozzles with bell-shaped expanders. Three different expander configurations are simulated (100% ‘full-bell’, 80% and a 60%) for varying nozzle depths (25μm− 400μm) and Reynolds numbers (15− 800) with decomposed hydrogen peroxide as the working gas. Owing to the inherently low Reynolds numbers on the microscale, substantial viscous subsonic ‘boundary’ layers are able to develop along the walls of the nozzle expander. These layers retard the bulk flow field and reduce nozzle performance. It was found that the subsonic layers on the expander walls can grow sufficiently large so as to merge together and revert the entire cross section of the expander flow field to subsonic conditions markedly reducing thrust performance. Thrust production and specific impulse efficiency is computed for the various flow scenarios and nozzle geometries to delineate the impact of viscous forces on nozzle performance. Results are also compared to inviscid theory and previous 3D linear nozzle simulations. It is found that the flow alignment of the bell nozzle comes at the expense of an increased expander length and with that an increase in the corresponding viscous losses via subsonic layer. As such, the 30 ◦ linear nozzle produces more thrust and has a higher efficiency than the bell nozzle at a given Reynolds number.

and increase of launcher reliability and operational efficiency. A common way of slightly improving performance of gas-generatoropen-cycleenginesistheinjectionoftheturbineexhaustgasintothenozzledivergentsection, whichis also used for wall film cooling. The present study focuses on a numerical parametric analysis of the film-cooling efficiency in dual-bell nozzles. The secondary gas injection is made in the first bell, and it is found that the expansion fan originating from the inflection helps the film to better protect the wall. The results of fully-attached-flow simulations are also used to study the influence of film cooling on the expected behavior of nozzle side loads during operation with separated flow in the second bell. Nomenclature A = cross-sectional area a = speed of sound cp = constant-pressure specific heat g = gravity constant l = length M = Mach number _ m = mass flow rate p = pressure R = gas constant Re = Reynolds number r = radius s = slot height T = temperature u = flow velocity V = flight velocity x = abscissa � = change of wall angle at inflection

Dual-bell nozzles permit two different operating modes, which provide higher performance than conventional bell nozzles when applied to rocket engines operating from sea level. During the low-altitude mode, the flow is separated and the separation line is located near the inflection of the nozzle wall, in a region characterized by a negative value of the wall pressure gradient, as in conventional nozzles, and in which side loads may occur. The characteristics of this region in hot full-scale applications are addressed by means of numerical simulations, and scaling laws are sought for cold subscale side-load experiments. Moreover, the flow behavior during the transition between the two operating modes is analyzed by time-accurate simulations.

In this study, we performed a CFD analysis to determine the optimum dual-bell nozzle contour. Especially, we investigated the effect of the deflection angle at the wall inflection on the separation point transition. Based on our present analysis, we found that the deflection angle at the wall inflection should be larger than the angle determined by a simple Prandtl-Meyer expansion. Also, the time to accomplish the separation point transition from the wall inflection to the extension nozzle exit is estimated to be less than 10 ms when our design criterion of the dual-bell nozzle contour is applied to the booster engine of H-2A launch vehicle. This time interval is considered to be essentially instant from the practical point of view of the safe nozzle operation.

The trend toward larger area ratio nozzle extensions for main-stage propulsion means that, whereas previously, nominal sea-level operation has been assuredly in the regime where incipient separation exists at the nozzle exit, new designs approach the boundary where a transition from incipient separation to free shock separation is seen in the nozzle extension. Such separations are not necessarily stable with respect to their location within the nozzle and, as a result of this, it is possible for large side loads to develop. Such behaviour is of immediate concern as the external pressure environment (including buffeting) is far from the steady, axisymmetric ideal that would be desirable and the sideloads generated may interact and excite nozzle structural modes. A clear understanding of the flow behaviour under such conditions is therefore deemed essential with respect to further nozzle extension development. This paper documents the results of an ongoing numerical investigation into the effect of an unsteady ambient pressure environment on the flow issuing from an over expanded separated nozzle. A discussion of the motivation for the research is given first, followed by a review of prior work. The effects of in-phase and outof-phase pressure variations in the far-field are then assessed with respect to conventional and dual bell nozzle configurations. Finally, the results from the numerical work are applied to the design of a new experimental program, the aim of which is to better quantify the effects of an unsteady pressure environment on separated nozzle flow-fields.

Because aerospike nozzle can meet the need of Single-Stage-To-Orbit (SSTO), it has already attracted common attention. To understand the performance of aerospike nozzle, a series of simulation tests on aerospike nozzle unit, which employed solid propellant, were completed. Results indicate that the thrust of aerospike nozzle decreases at beginning and then increases with an increase of backpressure, which is different from traditional bell nozzle. The experimental and numerical results indicate that the resultant force direction of the aerospike nozzle unit may change in a large range during the flight. A full travel optimization method, a parametric optimization essentially, is put forward in this paper, which is performed keeping maximizing total impulse and not restricted to a certain rated altitude. The process of building optimization model including computational procedure of the optimization is presented in detail. Introduction The multi-purpose low-cost reusable SingleStage-To-Orbit (SSTO) earth-orbit transportation system is a strong trend of aerospace technology development. To this system, the aerospike nozzle engine is critical. The aerospike nozzle has already attracted an increasing common attention. That is owing to its outstanding advantages: the automatically compensation ability of altitude performance, which can meet the need of SSTO. Also, in comparison with the conventional bell * PH.D. student, School of Astronautics, BUAA f Professor, School of Astronautics, BUAA Copyright © 2001 by the American Institute of Aeronautics and Astronautics Inc. All right reserved. nozzle it has shorter structure, lighter weight, flexible structure combination, and uniform load distribution, using turbine exhaust gas effectively to form the spike flow. In recent years the aerospike nozzle was widely studied experimentally and theoretically in many countries. Especially in the United States, the aerospike nozzle has been studied since 1960's and the technology level of approaching to practical flights has been reached '. In Europe, an Advanced Rocket Propulsion Technology (ARPT) program is undergoing, under which a number of investigations on aerospike nozzle have been performed \ In Japan, a reusable launcher demonstrator powered by aerospike engines is being studied . All these experimental studies employed the entire aerospike or plug nozzle, either clustered or linear, --' and the performance analysis also was as a whole 1. A whole aerospike nozzle testing is beneficial to gain an understanding of the performance and characteristic of the aerospike nozzle engine. But some characteristics of the nozzle itself, for example, how the thrust vector of the aerospike nozzle unit changes with flight altitude, were ignored owing to the symmetry of aerospike nozzle. In order to facilitate a fully understanding of the nozzle characteristics, in the study reported here an aerospike nozzle unit as shown in Fig.l was employed, and a specially designed testing setup was fabricated. Generally, an aerospike or plug nozzle is designed for a certain rated pressure ratio, which is linked to a certain area ratio. In the case of an ideal aerospike or plug nozzle, to fully determine the plug contour, the definition of the rated pressure ratio or area ratio and mass flow rate should be sufficient. For the portion around the exit lip of internal nozzle America Institute of Aeronautics and Astronautics (c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. it can be done based on a Prandtl-Meyer expansion and the contour followed the streamline 1

Liquid composite wound tanks can weigh three to four times less than a metallic one under the same operating conditions and give a new interest to pressure fed solutions (as a first approach Vacuum Specific Impulse depends mainly on engine area ratio and less on operating pressure) It could lead to a stage easy to operate, reliable, needing no costly solutions (Expander engine, Boost pumps). . On an upper stage with a full diameter tankage system (with or without a common bulkhead),the use of composite tanks is enabling creation of room ,or all around the aft dome of the tank or in length (length could be shorter or longer than an optimized metallic one without a significant mass penalty). Creation of thrust depends mainly on the possibility of implementing nozzles divergent surfaces. A peculiarity of Extendible Exit Cone, Plug (Aerospike ) and of Expansion-Deflection nozzles is to enable very short nozzle length design and so very short stages by comparison with conventional bell nozzle On another side, thermal fluxes depend greatly on the operating pressure: low pressure engine no longer needs regenerative cooling system (ceramic or metallic radiatively cooled engines could be used) Increase of throat area ratio is not a major problem with respect to the available room. Moreover stage without turbomachinery is much less sensitive to the mission profile and become competitive with a multi-boost mission The aim of this presentation is to demonstrate performance competitiveness when use of wound composite tanks will be possible, taking as a main example a conventional cryogenic upper stage and so the interest for R&D dedicated to pressure fed stages .

For an approximate design of aerospike nozzle, some numerical and experimental studies and conclusions are given here. The numerical simulations contain of both the Method of Characteristic and the solving of the Navier-Stokes equations. The experimental system and several kinds of the test aerospike nozzle are showed here. Some tests and calculations curves are presented and explained. Introduction Aerospike nozzle is much complicated in structure than that of the bell nozzle, so the optimized structural design of an aerospike nozzle has more questions need to be answered in ahead. For example, what's the best inclination angle of the side nozzle cell? How much effect to the nozzle performance that the plug base will cause when the supersonic gas flow around the base is closed or unclosed? The study conducted here is trying to answer questions like these. Experimental System The schematic diagram of the experimental system is showed in Figure 1. The system mainly includes the gaseous Oxygen supply system, the pressurized feed system of the fuel ethyl alcohol, the pressurized water circulating cooling system, the high-pressure air source, the vacuum pump and two large cylindrical vacuum exhaust tanks, the amplifier * Professor, School of Astronautics, BUAA f PH.D. Student, School of Astronautics, BUAA * Engineer, School of Astronautics, BUAA Copyright © 2001 by the American Institute of Aeronautics and Astronautics Inc. All right reserved. and computer data acquisition system, manual controlling and monitoring system, the cubic test vacuum room and the test aerospike engine inside it. The measuring parameters are mainly composed of two chamber pressures, chamber temperature, 2-4 oxygen head pressures, 2-4 fuel head pressures, chamber pressure and temperature of the igniter, head oxygen and fuel pressures of the igniter, mass flow rates of oxygen and fuel. Test Aerospike Nozzle After thousands of optimized performance computations have been done by using Method of Characteristic, several kinds of aerospike nozzle configurations have been formed. The configurations composed of bell side nozzle circular-linear composite structure with curved plug contour shown in Figure 2 and Figure 3, rectangular throat linear aerospike nozzle with different side nozzle angles (Fig. 4) and water cooled square shape linear aerospike nozzle (Fig. 5). Most of them will be used to do the experiments to confirm the effectiveness of the design. The model in figure 2 is a practical useful model suggested by us. The benefit of it is that the conventional axisymmetrical thrust chamber can be used with little change, and that the base is always closed both on ground and in a certain flight altitude. The model in figure 3 is a simplified one by using of 14 conical nozzles, 4 thrust chamber heads and one common chamber. The base mass flow rate percent can be adjusted according needs. This model was designed specially to test the optimized base mass flow rate. The model in figure 4 is a two-cell aerospike 1 America Institute of Aeronautics and Astronautics c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. nozzle and is specially used to test the best inclination angle of the primary nozzle or side nozzle. The model in figure 5 is a four-unit aerospike nozzle. Two of them are made up of 5 thrust chamber units, and the others are formed by 2 units. Each unit will be cooled by water so that the model can work in a higher temperature and longer time. It was specially designed to test the effects with or without base closing. Test Results From the tests, some important parameters such as thrust, chamber pressures are measured. Some other interested parameters such as nozzle efficiency, mass flow rate can be calculated by the measured data. Briefly speaking, the results will mainly contained of the multi-tube ignition properties to 14 separated thrust chamber cells, the altitude performance compensating property and its comparison for different nozzle structures, the performance contribution by the plug base and so on. Figure 6 shows the test results of multi-tube ignition. Figure 7 gives the ignition property of the test model in Figure 4. Figure 8 presents the working procedures of the same model. In figure 6, Tj, Pi? P0head and Tc represent the temperature and pressure inside the chamber of the igniter, pressure of the oxygen head position and the temperature at the exit position of several igniter tubes, respectively. Figure 7 still shows the working procedure of the igniter, but it is a procedure accompanying the working procedure of the aerospike engine. Here Pfhead is the fuel head pressure. Figure 8 shows the results of aerospike engine (model in figure 4) of the same test as shown in figure 7. Here F, Pc, PCOhead> Pcftead and Tc represent the thrust, common chamber pressure, one oxygen and one fuel head pressure of the engine, and the common chamber temperature. Numerical simulations Figure 9 shows the optimized results by using Method of Characteristics (MOC). The given conditions are chamber pressure pc =3.0MPa, stagnation temperature T* =3260K, expansion ratio of the primary nozzle 8j =1. From this figure we can see that the best primary nozzle inclination angle 8 is nearly equal to 30 at different backpressure or height. Figure 10 gives the similar results by solving the Navier-Stocks (NS) equations. Figure 11 given the computed results of the cooling at the primary nozzle throat cross section. The calculation conditions are pc =3.0MPa, T* =3260K, the inner wall thickness 8=1.2/ww, the width and deepness of the tunnel 1.7X 1.2. Here, Twg, Twi, Twaterin and Twaterout represent the gaseous sidewall temperature, the liquid sidewall temperature, the water inlet and outlet temperature, respectively. Conclusions For the test research are still undergoing, the complete conclusions can't be given now, except that the best inclination angle of the primary nozzle is probably about 30°. Another conclusion is that the angle will be affected strongly by other parameters . The thrusts of the aerospike nozzle is a function ofpc, Sj, 0, pa, total expansion ration et and the base mass flow rate. This will be also shown later. Reference [1] Dai Wuye, Liu Yu, Zhang Zhengke, Qin Lizi, Wang Yibai, "Numerical investigation on Linear Aerospike Nozzles," Prepared for 37 AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, AIAA2001-3568, Salt Palace, Salt Lake City, UT, 8-11 July, 2001. America Institute of Aeronautics and Astronautics c001 Am eican Intitute of Aenautics & Astnautics or Pulished w ith Perm ision of Auor(s) ad/or Auor(s)' Spsoring O rgazation. A m erica Intitute of A eroautics nd A strautics c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. Figure 2. Circyiar-Mnear composite aerospike oozzle model. Figure 3 Circular-linear composite aerospike nozzles without cooling Figure 4 Rectangular throat linear aerospike nozzles with changeable side nozzle angle Figure 5 Water-cooled square shape linear aerospike nozzles America Institute of Aeronautics and Astronautics c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

The dual-bell nozzle concept has been investigated by means of analytical and experimental work. Based on earlier analytical and numerical work published by the authors, different dual-bell nozzles were designed and experimentally tested, in order to explore this concept regarding its aerodynamic characteristics. This experimental work included cold gas subscale tests and hot gas subscale tests, which were performed within a joint German/Russian Research Programme TEHORA, and complementary in the German National Technology Programme TEKAN. It is shown that depending on the type of nozzle contour used for the dual-bell nozzle extension, a sudden transition from sea-level to altitude mode operation can be achieved. Furthermore, important information from the hot gas tests regarding the wall heat transfer were gained.

Design studies are presented for a 10 ton class high performance expander cycle engine with a chamber pressure above today's steel tube limit providing improved specific impulse for a given fixed bell nozzle design and engine length. Engine requirements are characterized by a 100 kN baseline version with a growth potential to 150 kN, providing a minimum specific impulse of 460 seconds for a maximum engine length of 2.4 m (94.5 inch) and a maximum engine mass of 275 kg (606 Ibm). It was found that the 100 kN (22,500 Ibf) version can be realized with existing or slightly improved technologies, while the 150 kN (33,700 Ibf) growth version needs new combustion chamber technologies with enhanced propellant heating features. Advanced expander chamber technology programs to reach this goal are under way at Dasa. The use of an extendible nozzle offers an additional Isp increase of about 7 seconds, with an associated weight penalty. Cryogenic upper stage engine requirements for Ariane 5 growth versions have recently been defined, calling for an engine of 150 kN thrust with 30% throttling capability. The presented studies are used as a basis for thrust chamber performance optimization and predesign of this new European upper stage engine.

After being tested and characterized in the 1960s and early 1970s, the concept of linear aerospike engines was revived in the last few years as a viable propulsion option for potential Single Stage to Orbit (SSTO) vehicles. In contradiction to the common believe in full altitude adaptation and higher vacuum performance, an evaluation of the published test results shows that this type of nozzle behaves just like any nozzle with the flow separating at too high ambient pressures and performance being limited by geometric and gasdynamic constraints. Altitude adaptation results primarily by the opening of the wake at high ambient pressures and the consequent transfer of pressure information between the ambient environment and the base region of the truncated plug. For even higher ambient pressures, the nozzle flow initially overexpands and finally separates from the wall according to one of the wellknown criteria. As for an aerospike engine the aerodynamic area ratio of the expanding core flow is much smaller than the geometric one, vacuum performance will be lower than for a bell nozzle with an identical overall area ratio. With this consideration the projected vacuum performance of the X-33 and Venture Star engines seems to be quite optimistic.

An experimental study was conducted on two small rockets (110 N thrust class) to directly compare a standard conical nozzle with a bell nozzle optimized for maximum thrust using the Rao method. In large rockets, with throat Reynolds numbers of greater than 1 x 10(exp 5), bell nozzles outperform conical nozzles. In rockets with throat Reynolds numbers below 1 x 10(exp 5), however, test results have been ambiguous. An experimental program was conducted to test two small nozzles at two different fuel film cooling percentages and three different chamber pressures. Test results showed that for the throat Reynolds number range from 2 x 10(exp 4) to 4 x 10(exp 4), the bell nozzle outperformed the conical nozzle. Thrust coefficients for the bell nozzle were approximately 4 to 12 percent higher than those obtained with the conical nozzle. As expected, testing showed that lowering the fuel film cooling increased performance for both nozzle types.

The linear rocket engine is shown to be a viable candidate propulsion system for post-Space Shuttle single-stage-to-orbit systems. The linear engine system has been developed and fired demonstrating high performance and long life with firing durations exceeding 500 seconds. The application of the split or dual combustor to the linear engine permits the uses of two different propellant combinations in a single engine system. The split combustor possesses the advantages of the two position extendible bell nozzle in a fixed nozzle configuration. Engine power cycles and applications to typical vehicles are discussed.