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Manned Spacecraft Development Plan from HTV Technical Heritage

By Takane IMADA

Japan Aerospace Exploration Agency ,Tsukuba, Japan

E-mail: imada.takane@jaxa.jp

This year will be the major milestone for JAXA by launching new heavy rocket H-IIB, and HTV (H-II

Transfer Vehicle) as its first payload. The combination of them will be the infrastructure to transport lots of

logistics, equipments, and experimental modules to LEO station and will be the base design of Manned

Spacecraft in future. We conducted preliminary study especially for LES (Launch Escape System) and

parametric analyses for Abort as shown in our paper for ISTS in 2008, and the missing part of manned flight

with the combination of HTV and H-IIB were selected for this paper. In this ISTS, we will progress the

spacecraft design as much as possible with basing upon the result. Also, the development plan was

investigated before starting concept designing to clarify the items that have not been investigated enough in

past and technical difficulty points.

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

Acronyms

1. Introduction

The first flight of HTV and H-IIB is planned in September

2009 (Fig. 1 shows photo and artistic image for the first flight

model). It will become one of the important milestones in

JAXA's long-term vision. HTV development has been nearly

completed, and the first HTV has already been in

Tanegashima Space Center for launch preparation.

Fig. 1. HTV First Flight Model and Mission Image

HTV is one of the ISS service vehicles and the culmination

of what JAXA has been developing for many years in the area

of launch vehicle, satellite and ISS/JEM program. By

following characteristics, HTV design is expected to be basic

technology for its future projects on orbital transfer vehicles,

free flyer units and manned transportation systems listed in

JAXA Vision 2025i.

- Un-manned rendezvous and berthing function

- Large capacity to add extra-module and components

- Multiply redundant avionics and propulsion systems

- Pressurized section compatible with crew IVA

- Exposed cargo handling capability by robotics arms

We are planning the next program after HTV or ISS.

Manned Spaceship based on HTV design is one of the

candidates. Some dimensions were assumed and some

parameters were determined by analysis in previous paperii.

We have continued feasibility analysis and this report shows

some missing part in previous paper.

2. Vehicle Concept

We estimated the reasonable target as Manned Spacecraft

from HTV technology heritage. (trade-off studies to estimate

it is written in reference document)

Fig.2 shows technical relations between original HTV and

Manned Spacecraft. They have the similar purpose of

operation that transport logistics (or crew) between the ground

and LEO. Lunar missions were not selected because of the

difficulty of enough launch capability with current Japanese

rockets. Multiple launches were not selected to enhance

HTV : H-II Transfer Vehicle

LEO : Low Earth Orbit

LES : Launch Escape System

RM : Re-entry Module

HAB : Habitant Module

SM : Service Module

ISS : International Space Station

JEM : Japanese Experiment Module

IVA : Intra-Vehicular Activity

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launch capability to put manned vehicle to lunar transfer orbit

because of the technical uncertainty of docking and

integrating on LEO with reasonable resource to conduct them.

We consider that transporters between LEO and the earth

will be necessary system especially after the Space Shuttle

retirement in 2010, even though Lunar mission might be

selected as the international target in next decades. Lots of

LEO missions will be required such as maintenance to

un-manned satellites, space telescope. Also, LEO base has the

advantages as the outpost because of the smaller resource to

build, no gravity, and easier to maintain if comparing with

Lunar base.

Fig. 3 shows the artist image about international rescue

mission. We think these kind of international cooperation in

space will be discussed not only for Lunar but also LEO

mission in following years.

Fig. 2. Technical Relations in HTV and Manned Spacecraft

Fig. 3. Spacecraft for International Rescue (Artist Image)

The selected spacecraft as Japanese Manned System as

follows.

[Integrated Vehicle]

- Length: 10m (exclude LES)

- Weight: 16.8 metric ton (operational mission, with LES)

[Re-entry Module]

- Crew: 4

- Diameter: 4m

- Weight: 5 metric ton

[Service Module]

- Delta-V: 390m/s

[Habitant Module]

- Volume: 50 m3

- Crew support system includes environmental control

Fig. 4 shows the combination of each module in launch

configuration.

Every estimated value in preliminary studied one and

should include enough margin to prepare the additional or

un-expected requirements in future. So, these values have

certain margin.

Fig. 4. Manned Spacecraft (Launch Configuration)

3. Subsystems

HTV is preferable base design but we have lots of items to

be developed as "Manned Spacecraft". Especially subsystems

that have interface between modules should be investigated

first to minimize the uncertainties in multiple modules. We

selected propulsion system (Reaction Control System: RCS),

Thermal Control System (TCS), and

Guidance/Navigation/Control (GN&C) System as such kind

of subsystems.

3.1 Propulsion System

Propulsion system in SM is similar to HTV but thrust level

of Main Engine should be three times or more higher (the

reason is shown in section 4.3). Technically it will not have

difficulty for designing.

Thruster system for RM is shown in Fig.5. It keeps attitude

control and de-orbit capability but translational control to thee

axes has been omitted. RM thrusters should withstand high

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heat rate during re-entry flight and all of them do not face to

re-entering direction.

Schematic diagram for RM propulsion system is shown in

Fig.6. Pressuring system was designed based on HTV which

has two failure tolerant for contingency acceleration. From

propellant tanks, every thruster block has independent

isolation valves to prevent failure propagation. This design

may be changed or modified in further designing phase but

philosophy to have "two failure safe" will not be changed.

Fig. 5. Thruster Location in RM

Fig. 6. Schematic Diagram for RM Propulsion System

3.2 Thermal Control System

Thermal Control System (TCS) will be fully different type

from original HTV. HTV has passive radiation and heater to

control temperatures in vehicle, but Manned Spacecraft

should have active radiation and very complicated system.

- RM cannot have radiation surface because it is covered

by thermal protection system to withstand re-entering

heat load

- RM has air circulation for crew support and it needs heat

exchanger to control air temperature

- SM should have radiation surface for thermal control for

both of SM and RM

- RM needs to have component for radiation after

separated from SM (i.e. Water Evaporator)

Fig.7 shows the total fluid system in modules. RM and SM

has fluid interface via quick disconnector for heat transfer

from RM to SM during on-orbit operation. RM has

pressurized section and heat from crew support system

removed via Heat Exchanger. Also, RM has water evaporator

to remove heat after separated from SM.

In this design, HAB has no fluid interface because we

wanted HAB to be used as independent on-orbit module for

an option. It has flexibility for modification or enhancement

easily in future.

Fig. 7. Thermal Control System

3.3 Guidance, Navigation & Control System

HTV has very complicated avionics system to satisfy the

safety requirements as manned vehicle. This technical

heritage can be used for Manned Spacecraft designing also.

Fig. 8 shows the schematic diagram for Manned Spacecraft

and it based on HTV system. The difference is as follows.

- Manned Spacecraft should have two failure operative

function for crew to come back to the earth safety (HTV

is required to safely abort from ISS after two failure as

the final function)

- Re-entry module has the most of failure management in

avionics boxes and it should conduct controlled re-entry

after two failure occurred (We denied to use ballistic

re-entry as an option after a failure. We consider new

manned vehicle should be designed to relief high

G-force stress from crew by ballistic re-entry as the next

generation of manned capsule.)

Fig. 8 also shows the schematic for RM and SM. We are

planning Re-entry module to have the function to conduct the

final de-orbit maneuver by itself to enhance the convenience

as Japanese vehicle. So, both of RM and SM has control

system but SM uses the simplified avionics to control the final

de-orbit maneuver only with 1 failure tolerance.

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Fig. 8. Guidance, Navigation, and Control System

4. Analysis to Determine Parameters

4.1 Launch Pad Abort

We conducted parametric analysis for abort during boost

phase in previous investigation. We determined tentative

parameters of LES such as total weight, dimensions, main

motor thrust, and propellant weight. In this paper, Launch Pad

Abort analysis result is added to select parameters to previous

analysis.

Main motor thrust pattern is defined as fig.9. In previous

analysis, we selected 800kN steady-state thrust as tentative

requirement, but in this paper we defined it as patterned thrust

for more realistic analysis.

Fig. 9. Main Motor Thrust Pattern of LES

To analyze the rotating motion, altitude, and range of

LES+RM during Abort, we made a simplified mass model to

conduct Launch Pad Abort analysis with pitch motor.The

model is shown in Fig.10.

Fig. 10. LES and RM Mass Model for Analysis

Pitch motor firing pattern was determined with considering

LES for other vehicles.

- Pitch Control Motor Thrust: 1500N to 3500N

- Pitch Control Motor Burning Duration: 1 sec

- Pitch Control Start Timing: 0.5 sec after Main Start

Analysis result is shown in Fig. 11 and Table 1. Because

velocity vector is fixed at the end of Main motor burning,

pitch motor should complete all firing before main motor

stops. As shown in Fig. 11, 1,500N looks a little bit

insufficient bending of velocity vector and RM splashdown

point is about 600m from launch pad. We can not assess 600m

is enough safe distance or not from fireball by explosion of

rocket because there is no standard for crew safety in RM

(Probably it is enough for RM, which is covered by Thermal

Protection System). In this paper, we selected 2,500N thrust

because of the RM attitude at the peak point. The attitude

(-172deg) suggests that LES and RM is flying from RM side

and it is preferable attitude for RM/LES separation. 3,500N

case shows that LES and RM rotated too much and RM is not

facing to the flight vector at separation timing.

Fig. 11. Launch Pad Abort Analysis

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Table 1. Parametric Study Result by Pitch Motor Thrust

Thrust of Pitch Motor

1500N 2500N 3500N

Rotation Rate 11.8deg/s 19.7deg/s 27.5deg/s

Max Altitude 1,328m 1,250m 1,143m

Range 577m 897m 1,146m

Attitude at Peak -72deg -172deg -248deg

(-180deg is best for RM Separation)

Fig.12 shows G-force during abort sequence. 2,500N pitch

motor is used for this analysis and time at peak point (LES

and RM separation) is calculated as 14.9sec. Splashdown time

is 33.9sec. This means that LES/RM have to prepare

separation within 15 sec after main motor ignition, and

parachute should completely deploy within following 19 sec.

Certainly this situation is most critical for time, ant these

values will be requirements for separation mechanism and

parachute system as minimum reaction time.

Fig. 12. G-force during Launch Pad Abort

4.2 Rescue Plan after Boost Phase Abort

We conducted boost phase abort analysis and showed in

previous paper but it lacked the following operation scenario.

Fig. 13 shows splashdown point (route) after boost phase

abort.. To rescue astronauts after contingency abort, all line

should be covered by any method.

If we use ships only, six or more number of ships will be

required to wait at determined stand-by points to prepare

picking up astronauts in Pacific Ocean after aborting from

launch vehicle. The dotted red circles show 1,000km radius

from stand-by points. If every ship has 22.5 knot as average

speed, all of the area will be covered within 24 hours.

Usually, the combination of helicopter and carrier ship was

used for sea recovery because helicopter has not enough flight

range from ground. But we selected the different way to

rescue astronauts with shorter time. Fig. 14 is a photo and

performance of ShinMaywa US-2, the latest seaplane

developed in Japan. It is used for Search and Rescue in Pacific

ocean around Japan. It has more than 4,700km range with

Short Take-Off and Landing performance. By using the

combination of two US-2 seaplanes and one ship, we can

cover more than 12,000km range. Blue belt in Fig.13 shows

coverage by two US-2s, one is standing-by at New Caledonia

and another is at Iwakuni Base. It can cover more than five

ships within ten hours. There is one more advantage of

seaplane. It does not consume fuel during picking up

operation at splashdown point (Helicopter consumes fuel for

hovering to pick up astronauts and it decrease operational

range).

Even though further investigation will be required, we think

seaplane will be one of the preferable methods to rescue

astronauts after boost phase abort.

Fig. 13. Abort Area and Rescue Coverage

Fig. 14. ShinMaywa US-2 Seaplaneiii (c) ShinMaywa

note: G-force inverted at 3 sec

because of force changed from

Separation Motor to Air Drag

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4.3 Abort to Orbit

Contingency Orbit Insertion is prepared for the final phase

of launch abort scenario. It allows crew and re-entry module

to escape to orbit by adding delta-V with propulsion system in

Manned Spacecraft itself. The required thrust and delta-V

have directly relation to the range of rescue defined in section

4.2 since it determine the timing to change the abort mode

from "Abort to Ocean" to "Abort to Orbit". If we have longer

range for rescue in sea, the change timing will be later in

launch phase, then smaller thrust and delta-V are required to

put spacecraft to orbit.

Fig.15 shows the nominal flight path for Manned Flight by

H-IIB and Fig. 16 shows abort flight paths by two ways for

aborting at the threshold time to change the method. In this

case, 0.5m/s2 acceleration and 120m/s delta-V are required in

Manned Spacecraft to avoid re-entering to atmosphere. The

original HTV has four Main Engines and total 2,000N to

increase delta-V but new Manned Spacecraft needs more than

6,000N for contingency orbit insertion.

Fig. 15. Nominal Path for H-IIB Manned Flight

Fig. 16. Abort Pattern and Flight Path

4.4 Nominal Recovery Plan

One of the heritages from HTV analysis data is de-orbit and

re-entry analysis. Based on the HTV operation, we selected

"Divided de-orbit maneuvers" (HTV uses three maneuvers for

descent from ISS orbit). Fig. 17 shows multiple de-orbit plan

to allow RM recovery in neighbor from Japan.

Fig. 18 shows splashdown area for SM and RM. Because

SM generated most of delta-V for descent, RM should load

smaller propellant for de-orbiting.

This design enable RM to splashdown to the near point

from Japan (at predetermined point in Recovery Area-1) and

ships can wait at exact the point because SM has already

destructive re-entered to separated ocean as debris.

If some trouble occurred in RM to conduct delta-V,

Recovery Area-2 can be selectable. In this case, RM control

lift and splashdown in further area than debris of SM/HAB

and ship is waiting at the expected RM splash point also.

Fig. 17. Nominal De-orbit Sequence

Fig. 18 Nominal Recovery Area

4.5 Recovery Error

Recovery error was analyzed to estimate necessary resource

to pick up Manned Spacecraft in nominal case. Table 2 shows

re-entry error analysis result from HTV studies. We used very

conservative values to translate these errors into splashing

point error. Table 3 shows the result. Total 289 km is enough

small to be corrected by lift with conventional capsule shape

which has L/D around 0.3.

0

20

40

60

80

100

120

140

160

180

200

0 2000 4000 6000 8000 10000 12000 14000 16000 18000 20000

Distance (km)

Altitud

e (

km)

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So, most of error of splashdown point will come from drift

after parachute deployment. It is estimated as less than 10km

from chasing ship (Ship will chase capsule based on wind data

and current position of capsule). Probably descending RM

will be watched from deck of ship by bare eyes.

Table 2. Re-entry Error at 120km

Interface Error at 120km

HTV Requirement

HTV Analysis Result

I/F

Value

3σ max min 3σσσσ

Inclination Error (deg) 0.03 0.001 -0.001 0.03

Velocity Error （m/s） 1.00 0.200 -0.600 1.00

Location (down range)（km) 100.00 38.749 -75.193 80.0

Location (cross range（km） 4.00 3.398 0.429 4.00

Direction Error （deg） － － － 0.0525

Table 3. Down Range Error by Re-entry Error Item Value Error （km）

Inclination Error （@120km） 0.03 (deg) ±27.9

Velocity Error （@120km） 1.00 (m/s) ±3.4

Location Error （@120km） 80.0 (km) ±80.0

Atmosphere Dispersion ±50 (%) ±119.9

CL Error （nominal：0.2） ±25 (%) ±242.6

CD Error （nominal：1.11） ±25 (%) ±55.8

3σ by RSS ±289.0

5. Overall Development Planiv

Fig.19 shows overall development plan for Manned

Spaceship. It also integrates the operational flights of HTV. In

current plan, total seven HTVs will be launched as a part of

international partnership in ISS program and we intend to use

them as demonstration of important functions for manned

flight by adding some improvements to original HTV.

The development plan is divided into five sections as

follows. All of them are mandatory as manned transport

system.

� Environmental Control and Life Support System

� Manned Spaceship System/HTV Updates

� Manned Re-entry Module Development

� Launch Escape System Development

� Launch Vehicles

We have to select the most effective way to proceed

development each items independently and jointly and we are

proposing to use seven HTV flights as opportunities to verify

some functions for Manned Spacecraft.

Fig. 19. Overall Development Scenario for Manned Spaceship

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5.1 HTV Operational Flights

Seven HTV flights are planned until 2015 as a Japanese

role in ISS program. The contribution by each flight is

assessed by cargo weight that HTV delivered. If we used a

part of HTV performance as transporter, the weight is not

counted as international contribution, and additional flight

may be required. In another word, HTV improvement to

enhance the cargo transport capability is desirable not only

increasing the contribution to ISS program but also

demonstrations Manned Spacecraft by adding some function

without off-loading cargo.

One of the candidates is un-manned capsule demonstration

for re-entry and recovery. This idea will be more reasonable

by the combination with the enhancement of HTV

performance.

5.2 HTV Improvement

We are investigating a few ideas to enhance HTV.

Followings are the examples.

(1) Solar Paddle and Improved Battery System

Original HTV has very complicated electrical power

system. It has four power resources, solar panel, rechargeable

batteries, non-rechargeable batteries, and power line from ISS.

This design was not selected in the early design phase, and

HTV had only non-rechargeable batteries and ISS power. But

we changed the design because of the weight overrun by

additional requirement as manned equivalent system (JAXA

had not have the experience to develop such kind of redundant

system and estimated HTV weight too optimistic).

Solar paddle was not selected because of keeping vehicle

design and interface with launch vehicle. Also,

non-rechargeable batteries were kept in HTV to sustain

insufficient solar power system, then HTV power system

became complex.

We consider that Manned Spacecraft will not have enough

surface to attach body mounted solar panels, and should equip

solar paddle system to provide enough electrical power to

whole vehicle. If HTV takes solar paddle design in advance, it

will give benefit to both of HTV operation and Manned

Spacecraft development. HTV will reduce battery weight (it

occupies 8 - 13% of HTV total hardware weight) and increase

cargo. Also, solar paddle system will be demonstrated

on-orbit before the first Manned Spacecraft flight. Fig. 20

shows advanced HTV which has two solar paddles. Solar

paddle will improve thermal characteristics and decrease the

weight for radiation and heater power.

Fig. 20 HTV Advanced (Solar Paddle Type)

(2) Structure Improvement at Rocket Interface

Current H-IIB rocket uses the same second stage as H-IIA

for minimizing development risk. The stage has only 4 m

diameter and the interface structure with HTV is limited to 3.2

m even though HTV body has 4.2 m diameter.

The second stage enhancement plan in H-IIB rocket will

use larger tanks and expanded interface structure, which will

be desirable for HTV from structural point of view. HTV will

be able to decrease weight or increase the length by the

preferable structural interface with rocket. Fig. 20 shows one

of ideas and it is enhanced 2 m in length and equips re-entry

demonstration capsule with expanded interface ring to 4 m.

Fig. 20 HTV Advanced (Improved Rocket I/F)

(3) H-IIB Enhancement

The second stage enhancement in (2) will give certain

benefit to HTV structure design but we will have larger

benefit by increasing launch capability of H-IIB. We are

conducting the analysis with considering HTV, Manned

Spacecraft, and other future missions such as Lunar inspector

or Lunar lander.

It will become the next target after we complete the

development of H-IIB rocket by test flight.

5.3 Demonstrations by Advanced HTV

As referred in previous section, advanced HTV will

demonstrate some function of Manned Spacecraft. Also,

following demonstrations are investigated as a part of Manned

Spacecraft development.

(1) Re-entry Module Demonstration (Capsule Recovery)v

This type of advanced HTV is most possible one. Many

scientists who conduct experiments in ISS want to get the

their test results, but they have only Soyuz to recover samples

to the Earth after the Space Shuttle retirement. JAXA and

other space organizations recognize the demand and ESA has

expansion plan for ATV to replace logistics carrier to

recovery capsule. JAXA has the similar idea. Un-pressurized

Carrier will be modified to carry a capsule and HTV will keep

cargo transport capability with Pressurized Carrier in this

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scenario.

Fig. 21 shows artist image of this type of HTV with all

improvements in section 5.2. It can demonstrate a part of

function of Manned Spacecraft on orbit and we will have

experience of recovery operation in the Ocean. This is the

final configuration as un-manned spacecraft and will be

succeeded by un-manned demonstrations with Manned

Spacecraft.

Fig. 21 HTV Advanced (Capsule Operation)

(2) Rocket Flight Path for Manned Launch

Manned transportation system cannot be developed by

spacecraft only. We should accomplish crew safety by the

combination of spacecraft and launch vehicle. H-IIB

enhancement in section 5.2 is desirable not only for cargo

transport but also as demonstration of manned flight path.

We consider that current H-IIB does not have desirable

balance in stages as manned launcher. H-IIB should carefully

adjust pitch angles during 1st and 2nd stage boost phase

because of insufficient thrust. We estimate the gravity loss by

current H-IIB design will be almost 15% of total launch

capability and H-IIB 2nd stage enhancement will decrease

gravity loss and satisfy crew safety with easier pitch angle

control.

Manned flight path and rescue in the Pacific Ocean will be

finally demonstrated by un-manned flight with the

combination of Manned Spacecraft and compatible rocket, but

validation flight by enhanced H-IIB and HTV will become

very useful demonstration because JAXA has not have

experience to plan and conduct the manned flight path.

5.4 Demonstrations for Manned Spacecraft

After or in parallel with demonstrations by advanced HTV,

flight demonstrations with Manned Spacecraft will be

necessary to verify each safety function in step by step.

Followings are one of plans to use H-IIA and H-IIB rocket as

efficient as possible. Fig. 22 shows launch configurations for

these missions.

Fig. 23 is artist image for Demonstration-3, which will be

the first manned flight.

[Demonstration-1]

� Un-manned Flight

� H-IIA type202

� Demonstrations for Re-entry Vehicle, Recovery in Sea

� Total Weight: 6 ton + Margin

i. Re-entry Capsule: 5 ton

ii. De-orbit Module: 1 ton

[Demonstration-2]

� Un-manned, but Manned Flight Path

� H-IIA type202 or 204

� Demonstrations for Launch Escape/Abort System

� Total Weight: 9 ton + Margin

i. Re-entry Capsule: 5 ton

ii. De-orbit Module: 1 ton

iii. Launch Escape System: 3 ton

[Demonstration-3]

� Manned Flight

� H-IIB

� Demonstrations for On-orbit Flight

� Total Weight: 14.3 ton + Margin

i. Re-entry Capsule: 5 ton

ii. Propulsion Module: 1.3 ton

iii. Propellant (off-loaded): 1 ton

iv. Launch Escape System: 3 ton

v. Orbital Habitant Module (subset): 4 ton

[Demonstration-4]

� Manned Flight (Enhanced H-IIB)

� Demonstrations for all Mission

� Total Weight: 16.8ton + Margin

i. Re-entry Capsule: 5 ton

ii. Propulsion Module: 1.3 ton

iii. Propellant (full-loaded): 2.5 ton

iv. Launch Escape System: 3 ton

v. Orbital Habitant Module: 5 ton

Fig. 22 Launch Configuration for Demonstration Flights

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Fig. 22 Demonstration-3 (Artist Image)

6. Conclusion

A series of preliminary analyses especially for manned

relate system showed feasibility to develop Manned

Spacecraft from HTV technical heritage. Both of crew support

system and abort system should be developed from the earliest

phase of design but the other systems will be developed based

on HTV design. As follows, we think we will have a major

milestone toward Japanese Manned Spacecraft by HTV

development and successful flight.

� HTV Propulsion/Avionics module will become

Service Module with update

� HTV Carrier System will be structure base of Habitant

Module

� HTV Un-pressurized Carrier will be used for

demonstrations for Manned relate demonstration

� H-IIB has enough launch capability to put Manned

Spacecraft on orbit

� The combination of H-IIB and HTV will be updated

and JAXA will have experiences for manned operation

Japanese Manned Spacecraft has not authorized as JAXA's

program yet, but we should prepare enough to start it

instantaneously when it is authorized. In any cases, JAXA

should complete the first flight of H-IIB and HTV in this

September successfully to go to the next step.

References

i JAXA Vision -JAXA 2025-

(http://www.jaxa.jp/about/2025/index_e.html) ii Takane Imada, M Ito, S Takata: Preliminary Study for Manned

Spacecraft with Escape System and H-IIB Rocket, ISTS., (2008),

2008-g-14

iii The US-2 Amphibian Aircraft - Unrivaled Water Surface

Take-off and Landing Capability | Shinmaywa

http://www.shinmaywa.co.jp/english/guide/museum_us2_02.htm

iv Takane Imada: HTV contribution scenario to Japanese

Human Spaceship Development (2D-07, Japan Society for

Aeronautical and Space Sciences, Space Technology Japan)

v Eiichiro Naoano, Takane Imada, and Yasufumi Wakabayashi:

Preliminary Study of the Recovery Capsule System Derived from

HTV System