The Compatibility Evaluation Method of the 500N & 120N Japanese Bi

Trans. JSASS Aerospace Tech. Japan Vol. 12, No. ists29, pp. Pg_7-Pg_15, 2014 Original Paper

The Compatibility Evaluation Method of the 500N & 120N Japanese Bi-Propellant Thrusters with the HTV System & Operation Design

1) 1) 1) 1) By Shinichi TAKATA , Hiroshi SASAKI , Tsutomu FUKATSU and Katsuhiko SUGIMORI

1) Japan Aerospace Exploration Agency (JAXA), Sagamihara, Japan

(Received June 12th, 2013)

This paper describes the evaluation method of the new Japanese thrusters design, balancing and verification of interface compatibility with the HTV system and operation design. Thruster performance directly affects the performance of the space craft. Therefore, the thruster replacement from the imported thrusters to the new Japanese thrusters required the evaluation from the view point of high-order system level. It is important to select the appropriate elements, verify their relationships, and create a feedback mechanism which allows modification of each element. This method will facilitate the overall systems design for future inter-orbit transfer vehicles.

Key Words: HTV3, Propulsion, Thrusters, Optimum Design

Nomenclature In conventional spacecraft system design whose propulsion system is not large and/or thrusters firing pattern is not severe, Duty : duty = {X/(X+Y)}*100, % the compatibility between thrusters and high-level systems is On Time : X(ms), Off Time : Y(ms) usually evaluated in one direction, which means thruster's Impulse : impulse bit ratio = X/Y compatibility with high-order system. But HTV is the largest Bit Ratio X : Integrated impulse at one pulse (N*s) vehicle in Japan and it's required to operate in severe pulse Y : Normal thrust*on-time at one pulse (N*s) firing patterns for a long duration especially in rendezvous TWmax : maximum chamber wall temperature at throat with the ISS. And JAXA aimed the unlimited pulse firing TWD : chamber wall temperature at root area thrusters. Therefore the thruster replacement required the fully evaluation of compatibility between thrusters and high-level 1. Introduction systems not in one direction but bi-directionally. This proper feedback mechanism allows modification of the new thrusters The two new types of Japanese bipropellant thrusters in design, evaluation of attitude controllability, revision of the HTV3 (KOUNOTORI 3) demonstrated flawless performance, specification between thrusters and high-order system and and delivered 16 tons of HTV3 to the International Space helps to predict the real profile on orbit. As a results, the Station (ISS). The 500N main engine and the 120N reaction optimum design was achieved. control system (RCS) thrusters were named HBT-5 and This paper describes the new compatibility evaluation HBT-1 respectively (Fig.1). method in large spacecraft system design based on the HTV thrusters development and HTV3 mission results. Fwd. RCS Thruster (HBT-1)

2. HTV System Description

The outline of the HTV system configuration diagram is shown in Fig. 2. The HTV operation team designs a plan by considering the status of ISS, launch vehicle, ground system, Aft. RCS Thruster (HBT-1) and HTV. The plan outlines operation timelines, external thermal environmental conditions, and thruster maneuver tactics. The GN&C system defines the flight control functions. The propulsion system which include the thrusters, generate thrust based on the control command from the GN&C system. The propulsion module configuration diagram is shown in Fig. 3 1).

Main Engine (HBT-5)

Fig. 1. HTV3 view from ISS (C) NASA.

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ISS HTV Operation System Other Systems HTV starts approaching to ISS step by step. During the final approach from 500m point to ISS, the HTV Flight Segment thrusters that locates on +Z axis and face the earth are mainly Pressurized Logistics Carrier ISS : International Space Station used and the severe pulse firing continues for over one hour. GN&C : Guidance, Navigation and Control Un-pressurized Logistics Carrier Exposed Pallet ISS Proximity Phase

Avionics Module GN&C System ISS(International Space Station) Approach Initiation Point Other Systems 10m(Capture Point)

30m Propulsion Module Propulsion System Thrusters Final Approach to ISS 250m Other Systems Other Components 500m (R-bar Initiation Point) Fig. 2. HTV system configuration diagram.

- Generate thrust for orbital change To Earth Main Function - Generate thrust for attitude control RId Maneuver - Store propellants System Regulated/Blowdown Bi-Propellant System Fig. 5. HTV’s final approach to the ISS Fuel : MMH Propellant Oxidizer : MON-3 (NTO) 3. Japanese Thrusters Development Process Pressurization Gas : Helium Main Engine : 500N×4 Thruster RCS(Reaction Control System) : 120N×28 3.1. The goal of HTV thruster development Dry Weight 1380kg The goal of Japanese thruster development was to acquire Maximum MON 1514kg凡MMH 918kg Propellant key technology for inter-orbit transfer system, find injector Size Length : 1273mm× Diameter : 4216mm parameter design know-how to improve the combustion High Pressure Area : 23.1MPa Pressure chamber’s temperature stability during severe pulse firing, Low Pressure Area : 2.75MPa cost reduction, and easy procurement control. Fig. 3. HTV propulsion module configuration. In HTV1 and 2, the imported thrusters that were flight-proven in Apollo and Shuttle missions were used in Almost all components of propulsion system are installed in order to mitigate the component development risks. However, the Aft. propulsion module. Fwd. RCS modules are installed constraints were observed during certain pulse firing in pressurized logistic carrier, so it has the long external tubes conditions. Therefore, key components of the inter-orbit space which feed propellant from tank to Fwd. RCS thrusters (Fig.4). transportation system were domestically developed after And the Fwd. RCS propellant manifold volume is designed to validation of the spaceship’s design technology. be minimized because of mitigating the Extravehicular In the beginning of the Japanese thrusters development, it Activity(EVA) hazard on the propellant leakage failure. As a was planned that the imported thrusters would be simply result, the external tubes and Fwd. RCS manifold tubes are replaced with the Japanese new thrusters. But as shown below, required to be long, but thin. the performance of the Japanese RCS thrusters during pulse operation became different from that of the imported ones. This is because the chamber temperature stability during pulse External Tubes operation with large inlet pressure oscillations was made a top

Fwd. RCS thrusters priority in all of the requirements to thrusters including the thrust performance. But the performance in HTV system level like a rendezvous capability or an attitude controllability could not be changed, then the compatibility of high order system with thruster performance was finally achieved by changing the interface specification between them and making some Propellant Tank Propulsion Module substantial improvements in the thruster injector designs as much as possible. 3.2. Surge pressure evaluation and control Fwd. RCS manifold tubes In general, a large surge pressure occurs in the propellant feed line during initial priming after separation from launch vehicle by opening of either latch valves or pyrotechnic valves.

Its level and behavior must be accurately evaluated by Fig. 4. HTV propulsion system configuration. conducting system level firing test or analysis on the ground in order to prevent a component failure, system rupture or The operation method in ISS proximity phase is shown in leakage. Usually this is the only problem in the initial Fig. 5. After departure of the approach initiation point, HTV activation phase of a spacecraft propulsion subsystem. The conducts a few small maneuvers by RCS thrusters to shift the methodology for analytical prediction of priming water final approach mode (beneath of 500 m from the ISS). Then

Pg_8 S. TAKATA et al.: The Compatibility Evaluation Method of the New Japanese Thrusters with the HTV System Design hammer pressures has been developed and applied to the design and development of a number of spacecraft Operation Plan/GN&C System Design applications. 2) • Generate Mission Duty Cycle using When severe pulse firing operation is required, the pressure the model of limited thrusters fluctuations at inlet of thrusters during pulse operations should performance be fully evaluated. In general, surge pressure is calculated via the following Nikolai E. Zhukovskii’s Eq.(1). It means that Propulsion System Design • Evaluate Surge Pressure Level during operation surge pressure becomes larger as the propellant tube becomes plan pulse firing thinner and the 'v becomes larger. The surge pressure modification reaches maximum in the long propellant tube. 'h 'p/Ug (c/g)'v (1) Check Thrusters Characteristics Where : h = head of surge pressure • Additional Verification if necessary p = surge pressure U = fluid density No Constrain to Operation g = acceleration of gravity c = net fluid speed of sound Fig. 6. General Development Flow of propulsion sub-system. and v = fluid velocity in tube In the development of the MARS Polar Lander descent the thrusters impulse performance during short on-time pulse propulsion system and the Phoenix landing propulsion system, operation must be evaluated accurately in order to conduct the water hammer level was also tested and evaluated because reality-based GN&C analysis. their attitude control systems were required to operate in very 3.3.1. Short on-time pulse operation severe pulse firing mode during the final descent phase. They When the pulse on-time of RCS thrusters is a couple dozen conducted the system level firing test and a set of analyses of millisecond during pulse operation, the thrusters impulse the flight system in order to resolve the unstable thrust levels performance becomes degraded and quite-variable, compared due to the inlet pressure drop. Although their maximum inlet with the large on-time combustion mode or the continuous surge pressure reached more than twice the normal inlet combustion mode. The impulse performance is also affected pressure, they concluded that the propulsion system was by propellant temperature or pulse interval time due to the capable of surviving the large pressure transients without difference of the propellant density or the remained propellant failure or leakage. They also concluded that the affect on the volume in manifold area between thruster valve and injector thrust impulse profile of the pressure transitions was minimum, plate. The lower density with high temperature propellant but the other anomalies like the combustion chamber’s causes the lower impulse performance. And the longer interval thermal instability were not reported. The duration of descent time between each pulse causes the less remained propellant phase which required severe pulse operation was about 40 in manifold area due to vaporization, which delays the seconds and was estimated to be short enough not to cause propellant injection timing of next pulse to the chamber area. other anomalies. 3), 4) Then it’s important to develop the mathematical model of In the development of the Olympus propulsion subsystem, thrusters impulse performance considering the every operating the inlet pressure fluctuations level was fully evaluated and conditions (propellant temperature, inlet pressure, pulse was concluded that it was not significant to thrusters operating on-time, interval time and variability, etc.) and is used for the point or impulse performance. It was estimated that it was design and verification of the GN&C system shown in section because the line pressure oscillation level was not so large 4.3(1). (about 20 % of the normal pressure). 5) 3.3.2. Surge pressure As a result, the surge pressure due to the water-hammer The Fwd. RCS propellant manifold volume is designed to phenomena can be accepted in the propulsion sub-system be minimized as shown in section 2. Then the external tubes design if the propulsion system is kept compact with short and Fwd. RCS manifold tubes which feeds the propellant to feed line, the required pulse patterns are not so severe or the the Fwd. RCS thrusters are required to be long and thin. And duration of the pulse operation is short. In most cases, the it takes about one hour during the HTV’s final approach to the surge pressure level or the thermal tolerance of thrusters are ISS and the HTV RCS thrusters are required to operate in evaluated after the system design is almost completed. The complicated pulse firing mode all that time. 1) general development flow of propulsion sub-system is shown As a result, a large amount of water hammer pressure in Fig.6. oscillations are observed in especially the Fwd. RCS lines due 3.3. Uniqueness of HTV thruster requirements during to the periodic closing of the multiple thrusters valves for a pulse operation long duration shown in Fig.7, whose telemetry is the 1Hz data HTV is the largest vehicle in Japan in order to deliver up to and is much slower than the actual dynamics. 6 tons of pressurized and un-pressurized cargos to the ISS. In In this condition, the transient fluid and heat transfer order to achieve the precise attitude control of HTV especially phenomena like surge pressure or heat soak back phenomena during the ISS proximity phase shown in Fig.5, the thrusters must be considered because it shifts the thrusters’ operating are required to operate in the various pulse firing mode and point, affects the thrusters’ chamber temperature stability and

Pg_9 Trans. JSASS Aerospace Tech. Japan Vol. 12, No. ists29 (2014) changes the thrusters’ impulse performance during the short verified in the real operating conditions because the thrusters on-time pulse operation. Then HTV thrusters must be operated normal operating conditions are sometimes different from the stably even in this condition, and should be better real operating conditions. In this flow, the potential problems performance within the constraints of thermal stability. can be recognized in early phase and be fed back to the Therefore it became very important to develop the HTV thrusters designs. thrusters recognizing the HTV operation plan, the GN&C and the propulsion system characteristics. And the evaluation of Thruster Design Thruster Characteristics thrusters’ compatibility with the HTV system and operation • Injector Design • Impulse Performance Model • Thruster Valve Design 9 At various pulse mode considering design became difficult tasks. • Chamber Design degradation due to propellant temperature • Injector and Chamber Temperature • Performance at off-normal condition

Requirement • Better Impulse Performance • Unlimited Pulse Firing for a Operation Plan Long Duration GN&C System Design • Generate Mission Duty Cycle • Evaluate Controllability

operation plan Requirement modification • Stable Pulse Firing in Large Inlet Propulsion Pressure Oscillations System Design • Evaluate Surge Pressure Level during pulse firing

Optimization of Operation Plan Fig. 8. HTV Thrusters Development Flow.

3.4. Problems in the initial thrusters development phase This was the first development of the 120N and 500N bi-propellant thrusters for complete pulse operation in Japan. Fig. 7. Fwd. line pressure oscillation during the RCS pulse firing They were developed by IHI AEROSPACE CO.,LTD. As (at about 250m point from the ISS) and the continuous firing, 1Hz shown in the section 3.1, JAXA aimed the unlimited pulse telemetry plot. firing thrusters for a long duration in large inlet pressure oscillations. 3.3.3. HTV thrusters development flow During the initial phase of thruster development, it was Due to the requirements to the impulse performance and the difficult to control chamber temperature during pulse firing, large operating shift caused by the surge pressure during the and therefore, took a lot of time to correct the thruster design. pulse operations, the thruster replacement from the imported Fig. 9-1 shows the temperature calculation point in the firing thrusters to the new Japanese thrusters required the evaluation test and Fig. 9-2 shows the test result failures for the 120 N from the view point of high-order system level. Based on RCS thruster in the beginning of the development phase. much experience of the compatibility evaluation tasks in When the thruster designs were not optimized, the chamber system level using the imported thrusters before HTV1, wall temperature at root area began increasing simultaneously analysis tools and know-how were fully used in order to with the transition to pulse mode (after 200 sec), although the conduct it more effectively. temperature was stable during the continuous firing (from 0 to The firing patterns were generated by considering thrusters 200 sec). This phenomenon also occurred during 500N impulse performance, environmental condition on orbit and thruster development. real operation plan. The surge pressure level was estimated by analysis evaluation. These firing patterns and the thrusters Root Area Throat Area operating point shift due to the surge pressure were used as the TWmax (Maximum temperature of TW) thrusters firing test conditions. The GN&C analysis were also TWD1 conducted using thrusters impulse performance. The thruster TW FilmȕǣȫȠ֩ݧᚌ Flow Angle designs, especially injector designs were revised based on the ༓૰ propulsion system and GN&C system analysis results other MMH ȕǡȳᚌFan Angle than the thrusters firing test results. The evaluation results in ᣠ҄д the system level were fed back to the thrusters designs for the MON3 optimization and the thruster designs were optimized step by ǤȳȔȳǸࢲ MMH༓૰ step. The HTV thrusters development flow is shown in Fig.8. The development flow shows that it’s important to evaluate all ȕǣȫȠϬҲྙᲢᲷȕǣȫȠϬҲ්᣽Ჩμ˳්᣽Უ requirements to the thrusters and their priority recognizing the Film Flow TWD2 system performance and real operation conditions in orbit. Fig. 9-1. Temperature calculation point in the firing test. And it’s also important that the thrusters designs must be fully

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1500 ĸ ĺ were also fed back to thruster re-design. The thruster Continuous Firing Pulse Firing (60ms ON䚸30%Duty) design was fixed at the CDR. 1200 (3) In the last stage, thruster qualification test results and verification results to interface specification were

900 checked at the thruster PQR, PQR follow-up, and the Υ TWD1 HTV3 system PQR.

600 Υ TWD2 FY 2006 2007 2008 2009 2010 2011 2012 Υ 䕰䕰䕰 TW max HTV 䕰 HTV1 HTV2 HTV3 Temperature (deg C) (deg Temperature System HTV1 HTV1 Ope. HTV2 Ope. HTV3 Ope. 300 System PQR PQR PQR CDR

Thruster Critical 0 Preliminary Design Qualification Limit Test/Flight Phase Design 0 70 140 210 280 350 Time (sec) 䕰䕰䕰䕰䕰 Thruster PDR CDR PQR PQR Fig. 9-2. Firing test results with inadequate injector design of 120N RCS Development Start Follow-Up GN&C thruster. (HBT-1 Thruster) Analysis Analysis

3.5. Thruster design constraints and thruster (PDR: Preliminary Design Review, CDR: Critical Design development policy Review, PQR: Post Qualification test Review, Ope.: The HTV thrusters development schedule is shown in Operation) Fig.10. It shows that the HTV system design was completed Fig. 10. HTV System and New Thrusters Development Schedule. before the beginning of the HTV thrusters development. It means that the requirements and interface specifications to the 4. System Optimum Design thrusters are fixed and there is little flexibility and a lot of constraints to design the new thrusters. Obviously it was 4.1. Optimum flow (bottom up + feed back) desirable to replace the thrusters without any affect to the As shown in section 3.5, the thruster development approach fixed high-order system design. Furthermore it was aimed to is not a top-down method completely, but one of bottom-up develop the unlimited pulse firing thrusters in order to remove plus feed-back. The top-down method means that the highest the operation constraints of the imported thrusters by level system determines allocation of functions or improving the chamber temperature stability during the pulse requirements to the thrusters and thrusters designs should firing. comply. In this case, new thrusters could be simply replaced But It seemed very difficult to analytically predict with imported thrusters. But the unlimited pulse firing combustion performance precisely during the pulse firing, so capability requirement imposes some restrictions on thruster the primary injector parameters that may affect thrusters design. Therefore, thruster designs, especially the injector performance were selected and numerous combustion tests primary parameters were almost fixed at first, and were conducted in order to evaluate each parameter’s compatibility evaluation with high order system and feedback sensitivity. During this tasks, it was found that the impulse to thruster design were conducted as a type of bottom-up performance was different and degraded in short-on time development (Fig.11-1, 11-2). pulse firing. The new thrusters adopted the 4.2. Extraction of elements multi-injector-elements with film cooling system in order to In order to conduct the compatibility evaluation between achieve the unlimited pulse firing thrusters and it has the the thrusters and high-order system, the primary elements manifold between the thruster valves and injector plate, then were extracted first in thruster level and high-order level. there is the time lag from opening the thrusters valves to As each thruster design variable (Xi) is fixed, each thruster generating the impulse. As a result, it affected the performance (Yi) can be specified based on the thruster firing specifications between thrusters and high-order systems. test results and analysis results. Yi must be appropriately Therefore in the process of the thrusters design, it became modeled considering the system design characteristics and all more important to provide the thrusters modeling and on-orbit operating conditions so that it can be used for evaluated it in high order system and fed-back to thrusters compatibility evaluations with high-level systems. design on a timely basis shown in section 3.3.3. (1) Thruster Design Variable (Xi) Then, the thrusters development policy was set as indicated Xi is the primary thruster design parameter. The injector below. design parameters like film cooling rate, injector elements (1) The thrusters design must be put high priority on the number, injection angle, injection velocity, etc., are chamber temperature stability even in very severe pulse representative examples. firing patterns. The thruster design was almost fixed at The concept of thruster design based on multi-elements the PDR. unlike doublet injector with film cooling system is itself (2) For verification of interface specification to high order conventional. As each Xi was changed, the firing tests with system and operation system, thruster performance was the engineering thruster model were conducted. fully modeled, and this was used for compatibility evaluation in high order system. The evaluation results

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(2) Thruster design results (Yi) Xi : Variable of Thruster Design Thruster Design Yi is thruster performance based on thruster design results, • Material and is modeled from the results of the thruster firing test, • Chamber cooling system • Injector design thermal vacuum test and a lot of analysis. (film cooling rate, injector element number, injection angle, injection velocity, etc.) Specific Impulse, which is usually abbreviated Isp, is a way • Manifold design • Thruster valve design(open/close response, leak tightness, ages, etc.) to describe thruster efficiency during firing, and represents • Chamber and nozzle design force with respect to the amount of propellant used per unit time. The higher the Isp, the lower the propellant flow rate for Yi : Variable of Thruster Design Results (Yi = F(Xi)) a given thrust, but the higher the thermal load to injector and Steady State Performance (Continuous Firing) chamber. • Isp in continuous firing • Temperature profile (injector, chamber, nozzle) Impulse Bit ratio (IB ratio), describes the integrated impulse • Combustion stability in wide operating range (thrust, mixture ratio) at one pulse with respect to the idealized impulse, or the • 䞉䞉䞉 normal thrust times on-time at one pulse. The high IB ratio in Unsteady State Performance (Pulse Firing, Ignition, Cut Off) short on-time is required especially during the maneuvers and • Isp in pulse firing final approach to the ISS. • Impulse bit in pulse firing • Combustion stability on ignition and in pulse firing (3) Design objective (Z) • Effect by large inlet pressure oscillations (caused by water hammer) Zi is the design objective and system level evaluation items • Thruster valve temperature increase in off-time due to heat soak back based on thrusters design results (Yi). In the development of • 䞉䞉䞉 new thrusters, each Zi is verified mainly by analysis. The attitude controllability was evaluated by the GN&C analysis Zi : Design Objective in HTV System Level (Zi = F(Yi)) • Compliance with HTV system Design Requirement tools, which was verified by the GN&C system verification ¾ ISS Safety, Resource, etc. campaign and other satellite experiences before HTV1. The ¾ Mechanical interface (propulsion system) ¾ Avionics interface thermal analysis was conducted using the thermal 9 attitude controllability in GN&C system mathematical model. The thermal mathematical model in ¾ Thermal interface (propulsion system) HTV system level was verified by the system thermal vacuum ¾ Fluid interface (propulsion system) 9 Stability of combustion and performance with thruster operating tests before HTV1 and that of thrusters was verified by the point shift due to transient propellant line pressure change thrusters development campaign. The fluid performance like • No constrain to Operation plan water hammer phenomena was calculated by the transient Fig. 11-2. Extracted elements and each relationship for HTV thruster case. response analysis tools, which was also verified by the results of the HTV propulsion system firing test and priming test before HTV1. 6) These tools were efficiently-utilized in order to evaluate the compatibility between new thrusters and more necessary to model thruster design results (Yi) while existing high-order system considering all operation conditions. This paper especially focuses on the relationship between 1. Analysis of HTV System Design Problems and Operation Plan with new thrusters the 120N RCS thrusters pulse performance and other systems • Elements extraction (Xi, Yi, Zi) as it affects the ISS rendezvous capability, and is the most difficult activity in a HTV mission. 2. Thruster Design (Xi) and Verification feed back (1) Relationship between thruster design(Xi) and • Evaluation of each Xi through a various firing test and analysis modeling(Yi) 3. Modeling of Thruster Performance(Yi) Thruster design parameter (Xi) affects the impulse bit • Define model of each Yi considering every on-orbit operating conditions performance, chamber temperature, or other thruster performance (Yi). 4. Evaluation of Design Objective(Zi) The example of an impulse bit of 60ms (millisecond) on • Identification of Problems or Improvement points based on the evaluation from the view point of the system design and the operation plan. pulse condition with 20 deg C propellant is larger than that of • Change the priority for some Zi (The unlimited pulse firing became more 50 deg C propellant. This is due to increased oxidizer injector important than the high Isp performance in continuous firing, etc.) resistance and decreased flow rate due to vaporization of the No feed back 5. Optimum Design Solution oxidizer at the orifice area of the flow with 50 deg C propellant. This result was fed-back to the flow path design Fig. 11-1. Method for optimum system design solution. (in-depth Fig.8) inside the thruster in order to prevent the oxidizer’s phase transition to gas. 4.3. Clarification of relationship between each element and Also, the impulse bit of 60 ms on-65 ms off pulse condition its modeling is larger than that of 60 ms on-1000 ms off pulse condition. From the lessons learned from HTV1 operation, pressure This is due to a difference of residual propellant in the injector oscillation and its influence of HTV propulsion system shifted between both pulse condition caused by the propellant the imported RCS thruster’s operating point due to long evaporation during off-time. However, if the off-time is very propellant feed line coupling effects, and this caused the short, the remaining propellant in the manifold area between increase of its injector temperature (Fig.13, HTV1 the thruster valves and the chamber area can generate a higher 7) (2009/09/17) Telemetry Data in orbit) . Therefore, it became thrust response. This result was also fed-back to the thruster’s

Pg_12 S. TAKATA et al.: The Compatibility Evaluation Method of the New Japanese Thrusters with the HTV System Design flow volume design. analysis to evaluate controllability in each maneuver and ISS Therefore, detailed modeling of thruster performance based final approach. Countermeasures were discussed and fed back on real conditions in orbit is required using various firing test into the thruster design when inconsistent results were results and understanding of phenomena. The modeling is observed. The thruster team re-designed the thruster, without defined as a function of the on-time, off-time and propellant changing the chamber temperature performance. temperature, which is almost equal to thruster valves In addition, an example of thruster firing pattern is temperature. It is also important to recognize generated through GN&C analysis. The propulsion system thruster-to-thruster variability of performance, which is conducts pressure transient analysis using this pattern in order reflected in the modeling. to recognize the line pressure oscillation level and thrusters Multiple firing tests were conducted with various injection operating shift level as shown in section 4.3 (2-1). conditions using the qualification thrusters in order to Basic policy states that the GN&C control logic cannot be understand the phenomena, or to check the effects of design changed for HTV3. Therefore, the thruster design had to changes. The acceptance test results of the HTV3’s thrusters consider those firing patterns as design constraints as well. using the test conditions shown in Fig.12 were also used for (2-3) Evaluation from operation planning (Zi) evaluation of the thruster-to-thruster variability. The compatibility evaluation with real operational planning is important. The procedure, made by the operation team, is OFFTime(sec) shown in section 2. The propellant tube line temperature, PulseFiringPatterntodefinetheImpulsePerformanceModel 100 maneuver timing, and timeline of final approach to ISS are highly affected by these results. The operation planning is a little bit flexible, and is often coordinated in order to mitigate 10 the load to the thrusters. 4.4. Feedback and optimization

1 As a result of GN&C analysis, it was found that the impulse bit performance for short on-time during the final approach to the ISS should be increased. To achieve this, the thruster’s 0.1 injector was re-designed for a quicker response. One improvement was to decrease the volume from the thrusters’

0.01 valve to the injector face, which made it possible to increase 0 0.02 0.04 0.06 0.08 0.1 the rate of propellant supply to the chamber. ONTime(sec) Analysis and evaluation of the real pulse firing pattern is Fig. 12. 120N RCS thruster pulse firing test condition for impulse important. The operating shift during pulse firing was performance model. evaluated from the results, which were fully attributed to the thruster qualification firing test conditions. (2) Evaluation of design objective It should be added that it is important to follow up the (2-1) Evaluation of propulsion system (Zi) thrusters design and each system’s evaluation status in detail The propulsion system controls the pressure, temperature, and timely. Many companies (JAXA and other contractors) and flow rate of propellant, and supplies the propellant to the were involved in these evaluations, and it was necessary to thrusters. It was developed by Mitsubishi Heavy Industries, clarify the time limit for feedback activities, that needed to be Ltd. as a prime contractor and IHI AEROSPACE CO.,LTD as shared. a sub contractor. Basic policy states that the propulsion system The thrusters’ design was finally approved after the result design cannot be changed for HTV3. Therefore, those newly of continuous feedback and optimization 8), 9). As a result of designed thrusters had to consider those parameters as given HBT-1 thruster design, the primary injector parameters are conditions and design constraints shown in section 3.5. selected so that the thruster temperature shows very stable and Due to the HTV system design, as shown in section 3.3, the has little sensitivity to the operating point shift especially in RCS thruster inlet pressure greatly fluctuates, which causes an the chamber root area. Fig.13 shows the HBT-1 thrusters operating point shift. This phenomenon is prominent in the temperature data in continuous firing test in wide operating final approach to the ISS (Fig.5, Fig.7). In this evaluation, the was used for compatibility evaluation in high order system. operating point shift was evaluated using the analysis tools, The evaluation results were also fed back to thruster re-design. which was correlated well with the HTV propulsion system The thruster design was fixed5. HTV3 Operation Results firing test results held in 2005 before HTV1 as shown in 6) section 4.1(3) . And the evaluated operating point shift was 5. HTV3 Operation Results reflected to the thrusters firing test conditions. In the propulsion system evaluation, using the correlated analysis The new thrusters demonstrated flawless performance 8) tools could remove the extensive system level tests . during their maiden flight in HTV3. Observed performances (2-2) Evaluation of GN&C system (Zi) were within expected values. Based on thruster performance model and other system In HTV1, an increase in injector temperature was observed performance, such as mass property or sensors errors, in the imported RCS thrusters due to the long pulse operation Mitsubishi Electric Corporation conducted the GN&C during the final approach to ISS. This was due to the operating

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6. Accomplishments and Action Assignment TND1 TW(TWmax) Verification of thruster compliance with high-level systems requirements was difficult during the development of HTV thrusters and their replacement in the HTV system. One reason is that it is difficult to confirm the new thruster’s design during unlimited pulse firing and its performance is TWD2 Temp(Υ) different from that of imported thrusters. The other is that the 1200 system and operation design has already been solidified. TWMax:y=4.0336x+399.27,R²=0.9652 HighThrust(135.7N) HighThrust(138.6N) HighMR(1.84) Consequently, it was important to identify all the elements LowMR(1.54) 1000 Normal LowThrust(87.3N) including thruster performance, and define the relationship (122.2N,1.65) LowMR(1.47) with the thruster’s mathematical model in order to evaluate the 800 thruster’s compatibility with high-level systems. This process TND1:y=2.4726x+307.04,R²=0.9903 TWMax 600 within the HTV project facilitates the understanding of the TWD2 mutual systems and their relationships. LowThrust(88.8N) TND1 400 HighMR(1.82) Presently, it is difficult to precisely predict the performance TWD2:y=0.3143x+66.201,R²=0.9889 of each component. Therefore, component design should be 200 performed during development of higher-level systems.

0 80 90 100 110 120 130 140 150 7. Conclusions Thrust(N) Fig. 13. HBT-1 thruster temperature data in continuous firing test 1. The two new types of Japanese bipropellant thrusters in wide operating range. demonstrated flawless performance in HTV3. Thruster performance directly affects the performance of the space craft, then the thruster replacement required the fully shift caused by the propellant line pressure’s transient change evaluation from the view point of high-order system level. (Fig. 7). These results were recreated during the thruster firing test on the ground by simulating conditions in orbit. A similar 2. It is important to define the thrusters performance model thruster firing test was conducted by simulating a worst case considering all the system characteristics and operation scenario (pulse duty, operating point) during the final conditions, and verify their relationship in order to evaluate approach to the ISS using the Japanese RCS thruster (HBT-1) the thruster’s compatibility with high-level systems. It’s also prior to the HTV3 operation. The final operation results important to provide the information to each other timely in (Fig.14, HTV3 (2012/07/27) Telemetry Data in orbit) were order to get time for the evaluation in system-level. within the firing test results (Fig.14, Ground test results of 8, 9) HBT-1 thruster simulating the HTV3 in orbit ). 3. This process also creates a proper feedback mechanism. It allows modification of each element’s design of the new thrusters, and helps set the test conditions on the ground in Temp (deg C) Fwd. RCS Thruster (+Z Axis) Injector Temperature Profile order to predict the real profile on orbit. In some cases the

250 revised. And as a results, the optimum design was achieved. Upper Limit (260 deg C) HTV1(2009/09/17) Telemetry Data in orbit It also helps to get a various data simulating the real

200 Ground test results of Americanthe importedthruster thruster operating conditions in thruster level and system level, and simulating the HTV1 in orbit F : 124.4N (Normal) was fully used to prepare for the real operation. MR : 1.46 (Lowest) 150 4. This method will facilitate the overall systems design for 100 HTV3(2012/07/27) Ground test results of HBT-1 thruster Telemetry Data in orbit simulating the HTV3 in Orbit future inter-orbit transfer vehicles. F : 135.7N (Maximum) MR : 1.46 (Lowest) 50 References

0 㻜㻡㻜㻜㻝㻜㻜㻜㻝㻡㻜㻜㻞㻜㻜㻜㻞㻡㻜㻜㻟㻜㻜㻜㻟㻡㻜㻜㻠㻜㻜㻜 1) S. Takata, T. Fukatsu, T. Imada, K. Matsuyama, S. Matsuo, S. 250m Point Time (sec) Nakai, K. Honda㸸The development of HTV Propulsion System, RId Maneuver R-bar initiation Point (300m in HTV1) Journal of JSASS㸪 58 NO.682㸪November (2010). (in Japanese) 2) Henry C. Hearn : Development and Application of a Priming st Fig. 14. Injector temperature profile during the final approach to ISS in Surge Analysis Methodology, the 41 AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, AIAA 2005-3738, 10-13 HTV1 & HTV3 is shown. Injector temperature profile of the ground test July 2005 results simulating HTV1 & HTV3’s approach to ISS is also shown. 3) Timothy A. Martin, Lawrence E. Rockwell and Carey L. Parish : Test and Modeling of the MARS’98 Lander Descent Propulsion System Waterhammer, AIAA-98-3665

Pg_14 S. TAKATA et al.: The Compatibility Evaluation Method of the New Japanese Thrusters with the HTV System Design

4) J. Greg McAllister and Carey Parish : Phoenix Landing 7) S. Takata, T. Fukatsu, K. Sugimori, K. Matsuyama, S. Matsuo, S. Propulsion System Performance, 45th AIAA/ASME/SAE/ASEE Nakai : The firing test results of HTV RCS thruster simulating the Joint Propulsion Conference & Exhibit, AIAA 2009-5264, 2-5 final approach to the ISS, the 51th Aerospace Propulsion August 2009 Conference of Japan Society for Aeronautical and Space Sciences 5) R. Devey, S. Gaydon and G. Statham : Development and (JSASS), JSASS-2011-0033, March 3-4, 2011. (in Japanese) Pre-Flight Testing of the Olympus Propulsion Subsystem, 8) S. Takata, K. Sugimori, T. Nagata and N. Matsuda : The AIAA/ASME/SAE/ASEE 25th Joint Propulsion Conference, evaluation of thruster design and verification based on the first AIAA 89-2511, 10-12 July 1989 flight of Japanese 120N thruster (HBT-1), the 53rd Aerospace 6) M. Yamamoto, S. Nakai, A. Shisa, H. Mitsuki, H. Mibae, T. Propulsion Conference of Japan Society for Aeronautical and Shiiki, S. Takata, S. Matsuo: Pressure Oscillation and its Space Sciences (JSASS) March 3-4, 2013. (in Japanese) Influence of HTV Propulsion System Due to Long Propellant 9) S. Takata, K. Sugimori, T. Fukatsu : The evaluation of HTV3 Feed Line Coupling Effects, ISTS 2011, 2011-e-35 Propulsion System with New Japanese thruster, the 56th Space Sciences and Technology Conference (in Japanese), 2G02, Nov. 20-22, 2012

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