Preliminary Study for Manned Spacecraft with Escape System and H-IIB Rocket

Trans. JSASS Space Tech. Japan Vol. 7, No. ists26, pp. Tg_35-Tg_44, 2009

By Takane IMADA, Michio ITO, Shinichi TAKATA

Japan Aerospace Exploration Agency ,Tsukuba, Japan

(Received April 30th, 2008)

HTV (H-II Transfer Vehicle) is the first Japanese un-manned service vehicle that will transport several pieces of equipments to ISS (International Space Station) and support human activities on orbit. HTV will be launched by the first H-IIB rocket in September 2009 and JAXA will have the capability to access LEO (Low Earth Orbit) bases with enough volume/weight as the human transport system. This paper is the preliminary study for developing a manned spacecraft from the HTV design and includes clarification of necessary development items. In addition, missing parts in the current HTV design are identified with some analysis, such as LES (Launch Escape System), which is mandatory for all human transport systems.

Keyword: Manned Transportation, H-II, HTV, Escape System

1. Introduction

JAXA announced its long-term vision for the next 20 years This paper uses several data from HTV as an un-manned as "JAXA Vision toward 2025" in April 2005 1) . JAXA but smart transportation vehicle, and launch capability data declared to keep establishing space transportation systems from the H-IIB rocket to estimate as reasonably and with the greatest reliability and competitiveness in the world. realistically as possible. Figure 1 shows an artistic image of Japanese manned spacecraft will be one of the goals of these the launch. This image used a 3-D model that was built reliable transportation systems. Even though development through this preliminary study. has not started yet, JAXA will launch the largest un-manned Through this study, it was identified that current HTV and space vehicle by the most powerful rocket in Japan, HTV H-IIB have several mission items as a manned spacecraft. (H-II Transfer Vehicle) and H-IIB, in the next year. The first Particularly the abort system (LES: Launch Escape System) flight will become an important milestone in JAXA's during power flight, which is unique as a manned system and long-term vision. Then, we will be able to take the next step should be investigated from the early designing process toward developing manned spacecraft. because it affects the transport system design. In previous manned programs in the US and Russia, the mass of Launch Escape Systems were more than half of the crew modules, and the flight path of the rockets were different from un-manned flights to allow the crew to escape in all phases on demand. As the first step of system estimation, we investigated the configuration and size of Launch Escape System with each abort scenario. Abort trajectories were parametrically analyzed, as shown in Section 4 using the sample of a nominal flight path of a manned vehicle with H-IIB and spacecraft, which has a reasonable weight of 14 metric tons. In addition, we tried to make a reasonable concept for this manned spacecraft, which will rendezvous with the spaces station, within the launch capabilities existing within JAXA. We then estimated weight and size of each module of the spacecraft while considering abort capability during each launch phase. In the last half of this paper, we report the analysis result about maximum gravity force to crew during flight and a development plan to demonstrate major functions step by step.

2. Design Baseline Fig. 1 H-IIB manned flight (artist’s image) Figure 2 shows the original HTV configuration. HTV has

logistic carriers for both of pressurized and un-pressurized cargo up to a total of 6 tons. Crew can enter the pressurized Soyuz: 2.9 tons with 3 crew, diameter: 2.2 m 2) section without a space suit to replace several logistics with Apollo Command Module: 5.8 tons, 3 crew, waste. The un-pressurized carrier section has several diameter: 3.9 m 3) mechanics to fix or release the exposed pallet utilized for Orion (planned): 7.3 to 7.7 tons, 6 crew, diameter: un-pressurized cargo exchanging. 5 - 5.5 m 4) JAXA’s Capsule: 5 tons with 4 crew, diameter: 4m

2.1.2. Mission Profile Targets as manned missions were classified into the following points:

- Ballistic Flight (less than 1 revolution) - Low Earth Orbit (a few days) - Round trip to a base in Low Earth Orbit - Round trip to the Lunar Orbit - Round trip and landing to the Lunar Surface

Table 1 Delta-V requirements for missions Fig. 2 HTV (original) configuration

The avionics module has many computers to control not only HTV attitude and position but also all failures in HTV. Basically, HTV has two control strings for nominal operation with failure tolerance and one more contingency string to comply with the same safety requirement as manned vehicles. Figure 3 shows the technical relations between a manned vehicle and HTV.

Table 1 shows necessary Delta-V for each mission. Because a lunar orbit needs plenty of propellant, round trip around the moon is not feasible as the first target within JAXA’s current launch capability. Transportation between the Earth and a human base in LEO (Low Earth Orbit) was selected as a reasonable target for the manned mission in this paper. Also, HTV will demonstrate the rendezvous and berthing capability and become a major milestone for manned flights to LEO. Based on the HTV mission profile, investigated enough in development, rendezvous to the LEO space station needs 2 or 3 days for phase adjusting. In addition, the spacecraft has to have 1- or 2-week survival capability for on-orbit contingency

Fig. 3 Technical relations between HTV and a manned vehicle cases.

2.2. Modules and Functions 2.1. Basic Requirement to Spacecraft Figure 4 shows the construction of the manned spacecraft 2.1.1. Number of Crew investigated in this paper. To estimate the weight and size of The first study to estimate the size of spacecraft is each function, this vehicle was separated into four functional determining the number of crew in it. More is better but the modules, the Launch Escape System, the Re-entry Module, number is restricted by the volume of the manned spacecraft. the Orbital Habitant Module, and the Propulsion Module. This Following are weight and diameter of manned Re-entry configuration was selected to use HTV heritage as much as Modules. The number of crew does not determine the capsule possible, and to enable us to develop and demonstrate each size directly because of the difference of mission, but a crew module separately. As shown in the following sections, the of four was assumed a reasonable number for a 4-meter Propulsion Module and the Orbital Habitant Module can be diameter capsule launched by H-IIB in this paper. developed based on the current design of HTV. But the

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Re-entry Module and the Launch Escape System have to be 2.2.2. Habitant Module developed from the earliest design phase because we have Orbital Habitant Module has several life support only a small technical heritage from previous programs. components and supports the crew staying on orbit. Also, it provides an electrical resource to the other modules on orbit. These functions enable us to shrink the size and weight of the other modules, especially the Re-entry Module. We intend to minimize the environmental control and life support system in the Re-entry Module only for a short duration after separated from other modules. Following are the tentative requirements for the Orbital Habitant Module. They will be modified for each mission, which determines the size of module and on-orbit staying duration.

Supports a maximum of 4 crew staying on orbit up to 3 days (nominal for rendezvous to station) plus 1 week for contingency cases. (Total 10 days) Equips solar paddles to generate 3,000 W (orbital average value) Fig. 4 Manned spacecraft configuration Equips connecting and hatch system with the Re-entry Module which is used for crew going in and out 2.2.1. Propulsion Module Equips mechanism for docking (or capturing as HTV) Required performance of the Propulsion Module is as with sensor/target system follows. Most of them are similar to HTV because both of them use a similar orbit to the space base at LEO. Like the Propulsion Module, the Orbital Habitant Module can be designed based on HTV. Since several functions to

support the crew for 3 - 10 days are not installed in the current Three axes Attitude/Position Control for Docking HTV, the following items should be developed before the first Maximum Acceleration for De-orbit: 0.07 m/s human flight. Full Redundancy, two fail safes Total Delta-V 390 m/s (Orbital transfer: 80 m/s, Rendezvous/Docking: about 200 m/s, De-orbit: about 110 m/s)

As long as total vehicle weight is less than 15 tons, only minor design modifications are required to HTV Propulsion and Avionics Module.

Fig. 6 Orbital habitant module

[Environmental Control/Life Support]

O2/CO 2 Control, Air Circulation Air Temperature Control Waste Management, Food/Water, Sleeping bags, etc.

2.2.3. Re-entry Module The Re-entry Module should be developed from the Fig. 5 Propulsion module preliminary design phase. JAXA has experience with un-manned re-entry vehicles in previous programs. OREX

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(Orbital Re-entry Experiment: Fig. 7) is one such vehicle. analysis because they cause the most uncomfortable gravity OREX conducted de-orbit and atmospheric re-entry force environment on the crew. We have to select a tolerable successfully in 1994. All re-entry vehicles in JAXA flew on thrust and duration for crew carefully with keeping quick ballistic trajectory but a Manned Module should re-enter with separating function enough to depart from additional hazards. controlling lift for both relieving the gravity force on the crew and increasing the accuracy of the splashing point in the sea. Table 2 Component tree of LES Figure 8 shows an artist’s image of the Re-entry Module. It has enough space for seats to absorb shock at splashing down. A conventional revolving shape was selected in this paper as re-entry capsule design, but a more complex and lift effective shape (lifting body) is better to decrease the maximum gravity force during re-entry. Further investigations are required to determine which shape will be used for JAXA's Re-entry Module and it will be the major trade-off point for the manned spacecraft program.

3. Trade-off Study for Spacecraft Configuration

Following are the results of preliminary studies for vehicle configuration. Figure 9 shows an artist’s image of on-orbit flight.

Fig. 7 Orbital re-entry experimental vehicle (OREX)

Fig. 9 Flight configuration (artist’s image)

3.1. Vehicle Construction 3.1.1. Use of Technical Heritage from HTV One of the technical key features of the HTV design is that each function is clearly divided into functional modules. It enables the use of a function as module level if it is required in other vehicles. Fig. 8 Re-entry module (concept image) In this paper, the combination of the Avionics/Propulsion Module in HTV is used as the base design for the Propulsion Module in a manned vehicle. Some equipment in the Avionics 2.2.4. Launch Escape System Module of HTV will be moved to the Orbital Habitant There is no reference system in JAXA for a Launch Escape Module such as electrical power service and environmental System and a suitable responsible section has not been control. Also, most of the structure design of HTV's identified to develop this system. Launch Escape System is Pressurized Carrier will become the baseline of the Orbital not a simple solid rocket booster. It flies with the Re-entry Habitant Module with minor updates. Module in the atmosphere from 0 m/s to hypersonic speed and By succeeding in the module design concept, we can use has to be designed from an aerodynamics point of view. Table much of the HTV design heritage for designing manned 2 shows a component tree of the Launch Escape System. spacecraft. Thrust level and propellant were determined from previous manned programs but to be investigated again in further

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3.1.2. Capability to Support Future Missions 3.2. Weight/Scale Estimation for Each Module In this section, the size and weight of each module were estimated based on HTV and other manned vehicles.

3.2.1. Re-entry Module The weight of the Re-entry Module was roughly estimated by diameter of capsule. As shown in section 2.1, 4 meters was selected as a suitable size to attach at the top of H-IIB launch vehicle with the Propulsion Module/Habitant Module. Its diameter is similar to the Apollo Command Module that weighs 5.8 tons. We estimate the Re-entry Capsule will be lighter than the Apollo Command Module since Apollo has several special components for Lunar missions and a life support system for longer mission duration. Our concept uses an independent Habitant Module for life support on orbit and most of the support equipment for longer stays are moved from the Re-entry Module to the Habitant Module. So, the Re-entry Module in our concept is estimated to have a 5-ton Fig. 10 Module construction weight in this paper.

Figure 10 shows differences in Soyuz, Apollo, and JAXA’s 3.2.2. Propulsion Module concept for module configurations. The Soyuz type is very The Propulsion Module weight and size can be estimated simple for operation but it is not selected in this paper mainly by data from HTV because it is like a subset of the because of the weight penalty for the Launch Escape System Avionics/Propulsion Module in HTV. (LES should also kick the Habitant Module to separate the Figure 11 shows dimensions and weight of the Propulsion Re-entry Module from the rocket). The Apollo type looks Module estimated from the HTV design. better to enhance functions (weight and volume) of the

Habitant Module for longer on-orbit stays. Two major differences between Apollo and JAXA’s concept in this paper are as follows:

Mechanical guidance is used for orbiting connection of the Re-entry Module and the Habitant Module. The Orbital Habitant Module has service functions, such as electrical power supply, crew support, rendezvous and docking.

It is a challenging design concept, but JAXA developed the "Rail & Wheel System" for HTV Exposed pallet in an Fig. 11 Propulsion module size un-pressurized carrier and the experience could be usable for mechanical guidance design of module connection on orbit. 3.2.3. Orbital Habitant Module Service functions were installed into the Orbital Habitant The weight of the primary and secondary structure of the Module while considering the option scenario to re-use the Orbital Habitant Module is about 3 tons as long as the same Orbital Habitant Module on-orbit as one module of the space size of HTV pressurized section is used for the manned station. vehicle. The crew support system has not been estimated exactly yet, but the total weight of the Orbital Habitant 3.1.3. Effective Development Module is roughly set at 5 tons in this paper. Module design in (1) and vehicle configuration in (2) have advantages in a development scenario. Each module can be 3.2.4. Launch Escape System developed by independent steps and schedules. The Figure 12 shows a typical contingency operation scenario Propulsion Module has a minor risk in development because it of the Launch Escape System used in Soyuz. A similar system uses a similar system as HTV. But the Re-entry Module and should be integrated into JAXA's spacecraft for manned the Launch Escape System should be developed through flight. critical milestones and demonstrations. This configuration enables us to start each development independently and minimize the affect of design modification from other modules in the development strategy.

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Fig. 13 Launch escape system Fig. 12 Launch escape system (Soyuz) As shown in Table 3, there is a certain relation between the The Launch Escape System is used if some critical failure maximum air drag and thrust level of the main motor (Abort occurs in the launch system and it is concluded to stop Motor). Each escape system has nearly ten times the ejecting mission by safely separating the Re-entry Module from the weight, but such a large thrust level is not required if LES was rocket. The Launch Escape System should have the following planned to be used only from the ground (smaller thrust is functions for on-demand escape requirements until the launch better to relieve the gravity force on crew). This is a result of vehicle successfully enters its orbit. LES being used under the maximum air drag during boost in the atmosphere. It needs a larger thrust to separate the

LES should work during all critical phases with the Re-entry Module from the rocket. rocket (from pre-launch until orbit) A smaller and heavier Re-entry Module is better to reduce LES should generate impulse to climb to enough thrust of the main motor and relieve the gravity force during altitude to activate the parachute for landing safely launch abort, but a smaller diameter loses lift co-efficient and or splashdown to sea increases the gravity force in the other phase, such as nominal LES should have enough thrust to separate LES re-entry. and the Re-entry Module under the maximum air So, dimension, weight, and shape (lift co-efficient) of the drag Re-entry Module determines thrust, weight, and size of the LES should avoid re-contact after separating from Launch Escape System. Optimizing all of them is the most the rocket important process in the early design phase of manned LES should separate from the Re-entry Module spacecraft. after the end of the necessary phase

As written in a previous section, we have no design 4. Abort Analysis for Boost Phase reference in previous JAXA programs to estimate the Launch Escape System. Weight ratio in Soyuz and Apollo are one of 4.1. Abort from Ground by LES the usable references for estimation. In these programs, LES There are several abort cases during the launch phase, but has almost 60% of the total escape weight. In this paper, two typical scenarios were selected to determine the main escape weight (Re-entry Module) was estimated as 5 tons and parameters of the Launch Escape System. The first scenario is we assumed LES weight as 3 tons. abort from ground. Table 3 compares these Launch Escape Systems. Figure 13 When a time-critical emergency occurs before lift-off, the shows a cutaway of LES with its functions. Launch Escape System is used to separate the crew compartment (usually the Re-entry Module) high enough and far away from the rocket. Table 3 Parameters of launch escape systems Figure 14 shows altitude and velocity in a typical escape case for ground abort. The assumed thrust and weight are written in a previous section. This analysis used a simplified atmosphere model but it shows reasonable results compared with the parameters of Apollo and Soyuz. It takes about 15 seconds to reach to the highest altitude and all sequences prior to deploying the parachute should be completed within this period. The Re-entry Module and the Launch Escape System turn around and the Re-entry Module separates from the Launch Escape System to prepare in deploying the parachute. It is a time Reference data: Apollo 3) , Soyuz 2) critical operation and needs further analyses to cover all abort cases with reasonable performance of the parachute deploying system (how quickly it deploys).

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rockets. Time for re-contact under the maximum SRB acceleration case (near to the end of SRB burn) was estimated and it is more than 11 sec because of the smaller air drag in this case. Time criticality for separation is higher in the maximum air drag case than SRB maximum acceleration case. We believe that characteristics of SRB (can not be cut-off immediately) do not spoil the capability of the escape function.

4.4. Abort Trajectory and G-force Analysis The Launch Escape System does not cover all phases during launch. The Re-entry Module should return safely with the crew any time, even after the Launch Escape System has been separated, and the following part shows the analysis Fig. 14 Altitude/Velocity profile in ground abort result for such cases. It is known that descent from higher altitude with 4.2. Abort at Maximum Dynamic Pressure insufficient velocity causes a very high gravity force and crew Separation under the maximum dynamic pressure is another may be injured. G-forces caused by descent from launch abort reference case in determining the basic performance of the are parametrically analyzed for every situation in this section. Launch Escape System. It needs enough thrust to overtake the Figure 16 and Figure 17 show a sample trajectory for air drag for separating the Re-entry Module from the rocket. manned flight for a nominal case and an abort case with the Also sustaining thrust is required after separation because H-IIB rocket. This flight pattern was designed to keep the the Re-entry Module and LES reduce speed suddenly by air altitude less than 120 km in most of boost phase to avoid drag even though the rocket is much heavier and maintaining steeper re-entry after emergency separation. The original speed. H-IIB has a launch capability up to 16.5 tons into low earth Figure 15 shows the relative distance between the Re-entry orbit but gravity loss spoils it in this flight path and only 14 Module and H-IIB after an emergency separation under the tons can be inserted to orbit. The following G-force analyses maximum dynamic pressure. (Note: 38 kPa is used as the for abort trajectories used this flight path with decreased value. SRBs continue thrusting but main engine is cut-off payload weight in a manned vehicle. during this analysis.)

This graph suggests that it takes only 10 sec for re-contact to rocket. The Launch Escape System should avoid re-contact through its own maneuver with another motor. As written in a previous section, it is preferable to have an active altitude attitude control system for maneuvering not only for turning-around in ground abort cases but also for departing from the rocket under the maximum dynamic pressure.

Fig. 16 H-IIB manned flight path

Fig. 15 Relative distance under maximum dynamic pressure

4.3. Abort at the End of SRB Burn H-IIA and H-IIB rocket have Solid Rocket Boosters (SRB), that are difficult to be immediately cut-off in emergency cases. Abort analysis for maximum acceleration case by SRB is necessary in addition to the maximum air drag case for these Fig. 17 Abort flight pattern (example)

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4.4.1. G-force difference by Abort Timing Figure 18 shows the analysis result with a zero-lift capsule 4.4.3. G-force difference by Re-entry Capsule Weight in ballistic trajectory (Note: a zero-lift type is not Figure 20 shows the result of analysis in which the recommended for manned flight. It was used for this analysis only). Re-entry Module weight changed with the same diameter This graph suggests that the separation point of the 1st/2nd (parametric study for ballistic coefficient). stage (375 sec in H-IIB) is the peak point to cause highest The result shows that maximum G-force during re-entry is G-force during abort. not sensitive to ballistic coefficient. A heavier capsule re-enters lower altitude (higher atmosphere density) before reaching the peak of gravity force and the result shows similar G-force for a lighter capsule.

Fig. 18 Max G-force by abort timing

4.4.2. G-force difference by Abort Altitude Figure 19 shows the result. 1st/2nd stage separation is selected as abort timing since the result in (1) shows the worst case. The result shows lower altitude is better and causes lower gravity force than higher altitude. The ratio is about Fig. 20 Max G-force and altitude by abort weight 0.7 G by 10km. 4.4.4. Effect of Aerodynamics in LES Design As written in Section 3.2, a large thrust in the Launch Escape System is not preferable due to the gravity force caused by the ground abort scenario. A smaller thrust is established if Drag/Weight ratio is decreased. Increasing the escape weight while keeping the same diameter (such as the Soyuz design) enables us to reduce the gravity force and seems a good idea to satisfy both the thrust requirement of air drag and gravity force during ground abort. But aerodynamics (relation between center of mass and air drag) has to be considered. Configuration in the Soyuz type has a Re-entry Module and a Habitant Module in the aft part of the escaping block and the center of mass is located relatively backward compared with other configurations. So, Soyuz has to have four folding drag plates (stabilizer) at the aft end of the escaping block to move aerodynamically and to Fig. 19 Max G-force by abort altitude fly stably from zero to a low mach number. They increase the total air drag area and need higher thrust for separation under the highest air drag condition. In this case, the Launch Escape System has to be re-developed if the Habitant Module adds weight. As a conclusion, in designing a Launch Escape System, we emphasize the importance not only of the thrust level or propellant mass but also aerodynamic and mass balance including future enhancement. (Of course it has been done in previous manned mission as with Soyuz and Apollo.)

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5. Development Steps ・ Recovery Sub-system Analysis

Drag/Parachute Size, Number Higher priority development items for each module are picked up in the following, focusing especially on the Volume Estimation in Re-entry Module Re-entry Module and the Launch Escape System. 5.6. Splash and Recovery 5.1. Whole Vehicle ・ Recovery Area Survey ・ Weight Estimation for Each Module ・ Team Scale Estimation for Sea Recovery Crew Support System Weight in Re-entry Module ・ Time Schedule and Habitant Module ・ Support System Requirement for Recovery Drag/Parachute and Floating Bag Weight, Volume Even though the HTV design and experiences will be 5.2. Launch Phase usable for manned vehicle development, several technically challenging points are still remaining. Especially Re-entry ・ Abort Function with Launch Escape System Module design and operation (including sea recovery) have to Abort Trajectory Analysis be demonstrated prior to human flight. The Launch Escape Abort Demonstration from Ground with Parachute System should also be demonstrated. and Floating Bag In this section, a step-by-step development plan for each module is recommended for cost/time effective development. ・ Abort Analysis during Boost Phase H-IIA has half of the launch capability of H-IIB, and it is Rocket Trajectory Analysis for Manned Vehicle enough to demonstrate functions of the Re-entry Module and Gravity force Analysis for Abort in Boost Phase the Launch Escape System by un-manned demonstration ・ Escape System Function Analysis flight. Table 4 shows demonstration items in each development Nominal Separation Analysis phase. Aerodynamic Analysis after Abort/Separation The following four flight configurations were assumed as Attitude Control Analysis with Pitch Control demonstrations to validate the functions for human flight. Motor (Launch Vehicle Configurations are shown in Fig. 24.) Re-contact Analysis for Maximum Air Drag, Maximum Acceleration by SRB Table 4 Demonstrations in each development step

5.3. On-orbit Phase ・ Re-entry Module/Habitant Module Connecting Mechanism Feasibility Analysis for Rail/Rotation System Development from HTV Pallet System ・ Hatch Design between Re-entry and Habitant Module ・ Solar Paddle Design

5.4. Rendezvous Flight ・ Crew Support System Design for Habitant Module System Tree for Environment Control [Demonstration-1] ・ Rendezvous Flight Control Un-manned Flight Attitude/Trajectory Control, Navigation Demonstrations for Re-entry Vehicle, Recovery in Sea Component Tree Total Weight: 6 ton + Margin i. Re-entry Capsule: 5 ton Propellant Consumption Analysis ii. De-orbit Module: 1 ton

5.5. De-orbit and Re-entry ・ Separation Analysis for Re-entry Module ・ Aerodynamic Analysis for Lift Controlled Re-entry ・ Drag/Parachute Deploy Analysis

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[Demonstration-2] [Demonstration-4] Un-manned, but Manned Flight Path Manned Flight Demonstrations for Launch Escape/Abort System Demonstrations for all Mission Total Weight: 9 ton + Margin Total Weight: 16.8ton + Margin i. Re-entry Capsule: 5 ton i. Re-entry Capsule: 5 ton ii. De-orbit Module: 1 ton ii. Propulsion Module: 1.3 ton iii. Launch Escape System: 3 ton iii. Propellant (full-loaded): 2.5 ton Flight Path and Abort Flight (Re-entry Module iv. Launch Escape System: 3 ton L/D=0.4) is shown in Fig. 21 to Fig. 23. v. Orbital Habitant Module: 5 ton

Fig. 21 Altitude profile for demonstaration-2

Fig. 22 Abort path sample for demonstratino-2 （from left to right: Demonstration 1 - 4）

Fig. 24 Launch vehicle configurations for demo-missions

6. Summary

There are several missing technical parts to build a Japanese manned spacecraft, but the combination of HTV and H-IIB will become the base design of it from their performances to launch and transfer manned-equivalent modules. Technically challenging points especially in the Launch Escape System and the Re-entry module should be Fig. 23 Abort G-load profile for demonstration-2 investigated soon to minimize the risks in development.

[Demonstration-3] References Manned Flight Demonstrations for On-orbit Flight 1) http://www.jaxa.jp/about/2025/index_e.html 2) http://www.russianspaceweb.com Total Weight: 14.3 ton + Margin 3) Anon. : Apollo Operations Handbook Block II Spacecraft, i. Re-entry Capsule: 5 ton SM2A-03-BK-II-(1), 1969 ii. Propulsion Module: 1.3 ton 4) Anon. : NASA's Exploration Systems Architecture Study, iii. Propellant (off-loaded): 1 ton NASA-TM-2005-214062, 2005 iv. Launch Escape System: 3 ton v. Orbital Habitant Module (subset): 4 ton

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