Liquid oxygen and liquid methane rocket engines are leading candidates for a variety of space missions, including the critical role of main descent propulsion of planetary landing spacecraft. Particular mission characteristics in this context may dictate the use of propulsion systems with deep-throttling capability and able to deliver good performance over the entire thrust range. A well-known injector type with favourable combustion stability, heat transfer and performance characteristics is the pintle injector. This atomizer has been tested with a variety of propellants, such as monomethyl hydrazine/nitrogen tetroxide, fluorine/methane and oxygen/hydrogen. However, characterization employing liquid oxygen and liquid methane is relatively scarce. In order to obtain relevant parametric performance data at nominal and o-nominal conditions, a miniature cryogenic test stand operating with LO₂/LCH₄ propellants was designed and built. Presented here are the results of a systematic study to determine the uncertainty of measured and calculated parameters defining pintle injector performance. The approach used for the measurement of relevant quantities such as thrust, propellant flow rates, temperature and pressures is discussed and a description of the methodology used in the error analysis is given. Final values of uncertainty in characteristic velocity eciency are presented for dierent pintle injector configurations.

In recent years the cost aspect of the development, production and operation of liquid rocket engines became more and more important with respect to the competitiveness on the commercial satellite launch market. Therefore, one aspect of the development is focused on novel inexpensive hot gas wall materials for inner liners of rocket combustion chambers that can withstand the extreme operational conditions. The copper-base alloy consisting of copper, chromium and zirconium is such an inexpensive material that has the potential to be used as an inner liner material for future rocket combustion chambers. To predict the damage behavior caused by tensile rupture, a viscoplastic model coupled with ductile isotropic damage, crack-closure effect and thermal ageing is defined and implemented as a user-material routine in the commercial finite element package ANSYS. Uniaxial displacement-controlled tensile tests at temperatures up to 1000 K are performed. In contrary to traditional extensometer based measurements, an optical stereo camera system based on digital image correlation is used to determine the local strain distribution in the necking area of an hourglass-shaped test sample. The material model's parameters are then least squares fitted and applied to a two-dimensional thermomechanical finite element analysis of a half-channel model representing a rocket engine combustion chamber wall with a simplified geometry and a representative heat flux, pressure level and maximum temperature. This paper shows that digital image correlation can be used to capture the local strain distribution in the necking area of highly ductile copper materials at elevated temperatures. Indeed, the numerical analysis with tensile test based material parameters results in an estimated life time of 13 cycles but thinning and bulging of the hot gas wall could not be reproduced.

A solar power sail, an extended concept of a pure solar photon sail, is one of the most effective ways of realizing challenging outer deep-space explorations. Thin-film solar cells attached to the surface of the spin-type solar sail membranes generate sufficient electric power to drive highly efficient ion engines even in the outer planetary region. The orbital control capability of the pure solar sail is too small to accomplish the missions within a reasonable period. The ion engines which provide much higher orbital control force than the pure solar sail can realize various outer space missions. JAXA has been preparing for a Jovian Trojan asteroid exploration mission via solar power sail. This paper outlines a solar power sail spacecraft under consideration and discusses the trajectory design method and results for the Jovian Trojan asteroid exploration.

This paper describes how Independent Verification and Validation (IV&V) has grown at the Japan Aerospace Exploration Agency (JAXA) and how the activity was devised over two decades. However, as the development process for aerospace products matured over the decades, the value which IV&V needs to provide has changed currently. Hence, JAXA IV&V refined the concept of IV&V to accommodate mature software products. Definition of IV&V and also the narrow definition of independent, verification and validation are introduced in the paper. The paper also introduces the challenges and aspiration to support the refined concept.

A three-dimensional electrostatic numerical tool has been developed to analyze spacecraft and Hall thruster plume interactions to evaluate spacecraft reliability under operation of the thruster in its designing phase. In the tool, Particle-In-Cell method is adopted to solve trajectories of beam and charge exchange ions under electrostatic field self-consistently. Particle tracking method is also adopted to compute the direct interactions such as the impingements of the plume ions onto the surface of a spacecraft to reduce numerical cost. In this paper, as a first step of the development, specifications of the code are described, and the fundamental functions such as determination of the distribution of the plume, production of charge-exchange ions, and electrostatic potential are introduced. Application analysis of the tool for an actual large-scale satellite, the estimation of the generating torque toward the center of mass of the spacecraft by the impingement of the plume ions, is also shown.

In the current study results from an experimental investigation on an oxygen/methane multi-injector combustion chamber are presented. They provide detailed information about the thermal loads at the hot inner walls of the combustion chamber at representative rocket engine conditions and pressure ranges up to 40 bar. The present study aims to contribute to the understanding of the thermal transfer processes and of the interaction between the injectors and the injector-wall. Furthermore, the test results are used as a test case for the validation of the in-house engineering tool Thermtest. Due to the complex flow phenomena linked to the use of cryogenic propellants, like extreme variation of flow properties and steep temperature gradients, in combination with intensive chemical reactions, the problem has been partially simplified by injecting the propellants in gaseous form.

The similarity between POD (Proper Orthogonal Analysis) and NN (Neural Network) is explained and an example of NN to perform reduced dimension analysis on a combustion oscillation problem is presented. The dimension reduction procedure by Snapshot-POD is shown to be expressed by a three layer AE (Auto-Encoder). Based on this, a DAE (Deep Auto-Encoder), consisting of a four layer encoder and four layer decoder is tested. The encoder has layers of 128-32-8-2 neurons and the decoder has the ones of 2-8-32-128 neurons in its layers. The DAE reduces the dimension of the input data into two, which is the number of the encoder output variables. As a reference, a POD that takes the first and the second mode neglecting higher modes are employed to reduce the dimension into two. A URANS simulated time varying temperature, heat release, and pressure distributions of CVRC (Continuously Variable Resonance Combustor) are analyzed by the POD and DAE. As a result, the 2D data from DAE and POD agreed well. It was confirmed that the dimension reduction performance and the resulting amount of information was almost consistent. By analyzing mode maps, the ability to identify the modes for pumping up the oscillation is demonstrated.

For the demonstration of wide variable thrust range ion thruster, ion beam extraction ON/OFF duty ratio controlling system was developed. We measured the ion beam current trace at downstream of the ion thruster with ON/OFF duty ratio control. In addition, we investigated the effect of this system on the ion beam divergence. The ion beam current changed as the applied voltage changed, that is, we demonstrated the system was useful for extending the thrust range. Beam divergence was almost constant, about 20 degree while changing duty ratio from 1% to 100%; this results shows that variable thrust system with controlling the beam extraction ON/OFF duty ratio affects little on the beam divergence.

This paper studies rendezvous control of spacecraft via constrained optimal control using generating functions. The problem of minimum energy control of the spacecraft transiting between specified states with constraints is formalized into the constrained optimal control problem. The generating function approach is extended to such problems by equipping with penalties. Finally, the developed technique is summarized as algorithms to successfully realize optimal rendezvous control with velocity and thrust bounds.

The accuracy of autonomous orbit determination of Lagrangian navigation constellation will affect the navigation accuracy for the deep space probes. Because of the special dynamical characteristics of Lagrangian navigation satellite, the error caused by different solution technique will cause totally different orbit prediction accuracy. We apply the RKF78 and RK4 to solve the motion equation of Lagrangian navigation satellites. There is no obvious difference when these two methods are used to calculate the orbits around the Earth-Moon triangular libration points. However the calculation error increases when RKF78 and RK4 are used to calculate the orbits around the Earth-Moon collinear libration points. Although the calculation error will be the order of 1×10⁸ meter, it doesn't cause big difference on the AOD with an AOD step of 1 hour. If the AOD step is bigger than 10 hour, the accuracy of autonomous orbit determination using RKF78 is better than the autonomous orbit determination accuracy using RK4.

In the case of three-axis attitude control of spacecraft by control moment gyros (CMG), more than three CMGs are typically used for redundancy. Even when one CMG fails and the redundancy is lost, the attitude control must be maintained. In this paper, by focusing on attitude control with three CMGs, suitable steering laws are considered. In order to use the full angular momentum workspace of three CMGs, singularity problems occur. The problems occur more severely than four CMG case. Two steering laws for the four CMG pyramid configuration with one CMG failure are proposed; Inverse Kinematics Steering Logic (IKSL) and Forward And Backward Reaching Inverse Kinematics (FABRIK). The aim of each steering law is to provide gimbal rates from the calculation of gimbal angles by inverse kinematics of the CMGs. IKSL exactly solves the inverse kinematics of the system, whereas FABRIK uses a heuristic approach to finding an approximate solution of the inverse kinematics problem. Numerical simulations are performed to validate the effectiveness of the proposed steering laws as compared with the Singular Direction Avoidance (SDA) steering law.

A Mars orbiter/lander mission using a micro-satellite (less than 100 kg at Mars arrival) with a deployable membrane aeroshell for the orbit insertion by the aerocapture and the electric propulsion for the trajectory maneuver was considered. The aerodynamic heating environment during the atmospheric flight was investigated solving the thermo-chemical nonequilibrium axi-symmetric Navier-Stokes equations around the spacecraft with the flare-shaped aeroshell. To obtain the appropriate amount of deceleration at the atmospheric pass, the drag modulation technique, in which the aeroshell is timely jettisoned from the spacecraft, was assumed. The hypersonic wind tunnel experiment successfully demonstrated that the backward jettison works well without significant time delay and attitude disturbance. Finally the corridor width for the entry path angle was estimated considering the ability of the orbit insertion, the peak aerodynamic heating, and the capability of the electric propulsion to raise the periapsis altitude. The combination of a low ballistic coefficient flight with the membrane aeroshell and an electric propulsion works well to acquire a finite width of the entry corridor for the aerocapture.

Considering a spacecraft that encounters particle-laden environment, such as dust particles flying up over the regolith by the jet of the landing thruster, high-speed flight of a projectile in particle-laden environment was experimentally simulated by using the ballistic range. In case of high-speed motion, cracking of particles occurs as well as the damage on the projectile surface. To establish the experimental simulation technology and to obtain the fundamental characteristics of such complicated phenomena, a projectile was launched by the ballistic range at the velocity up to 500 m/s and the collective behavior of particles around the body was observed by the high-speed camera. To eliminate the effect of the gas-particle interaction and to focus on the effect of the interaction between the particles and the projectile's surface, the test chamber pressure was evacuated down to 30 Pa. The particles about 100 – 400 μm diameter were scattered in the test chamber uniformly. The projectile was launched into a particle sheet in the tangential direction. The high-speed camera captured both the projectile and particle motion. From the movie, the interaction between the projectile and particle sheet was clarified. The damage on the surface of the projectile recovered after the shot was also observed.

In this study, a water resistojet propulsion system, named AQUARIUS, is proposed and the breadboard model (BBM) test results and the engineering model (EM) design are described. AQUARIUS will be on-boarded the SLS EM-1 CubeSat: EQUULEUS, which is scheduled to be launch in 2019. AQUARIUS uses a storable, safe and non-toxic propellant: water, which allows for the design of all propulsion system below 100 kPa, reduction of dry mass ratio and simplification of feed line routing using soft tubes. AQUARIUS consists of three components as follows: a tank, a vaporization chamber and thruster head. BBM of the vaporization chamber, and the thruster head were designed and tested. The evaporation rate required heat and thrust performance were evaluated. Based on the BBM test results, the engineering model of AQUARIUS was designed.

The number of small satellites in Low Earth Orbit (LEO) is increasing. In LEO, the dominant non-gravitational force arises from aerodynamic effects. These effects must, therefore, be taken into account when determining, predicting, and designing the orbit. This paper presents a detailed analysis of the aerodynamic effects on a small satellite in LEO, using numerical simulation. A sensitivity analysis of the orbital environment and satellite specifications was performed, and the critical parameters were identified. The results were compared with those from analytical models reported in previous studies. Approaches to improving the analytical model were considered. Finally, realistic simulations were performed, to investigate the significance of detailed aerodynamic analysis in predicting the orbital lifetime.

In this study, an attitude control method of spacecraft using only magnetic torquers is proposed. A magnetic torquer can generate torque only in directions perpendicular to the magnetic field of the Earth. Conventionally, a PD feedback control method with a cross product law is used. However, the cross product law has a constraint on the torque output direction, and this constraint restricts the stable region of the attitude control. In order to expand the stable region of the attitude control with magnetic torquers, we propose the control law with the compensation which generates the desired attitude control torque in average through one orbital period. The conventional and proposed control laws for an Earth-oriented spacecraft with magnetic torquers are introduced and their stable regions are analyzed and compared. These control methods are evaluated by numerical simulations for spacecraft attitude motion. The analysis of the stable regions of the conventional and proposed methods is also examined with the numerical simulation results. The results show that the proposed control method with magnetic torquers is effective for the spacecraft attitude control.

Satellite communications are vital in the event of disaster, and fulfill supplementary communications needs, especially in aeronautical applications. The National Institute of Information and Communications Technology (NICT) conducts research on Ka-band satellite communications using the Wideband InterNetworking engineering test and Demonstration Satellite (WINDS). NICT has also developed an aeronautical earth station, which we flight tested in WINDS's Chubu beam. We measured the propagation characteristics of multi-beam antenna mounted on WINDS to measure the variation of Doppler shift for Ka-band airplane-satellite communication. We also measured the tracking performance of the earth station antenna. In practical tests, NICT's aeronautical earth station demonstrated uplink speeds up to 38 Mbps.

JAXA and NEC Corporation developed a state of the art space-borne GPS receiver in 2013. This latest receiver has been reduced in terms of the size and mass of its hardware, along with improved software to achieve high accuracy onboard navigation. In 2016, ASTRO-H, a new X-ray astronomy satellite, was launched with this GPS receiver. The navigation performance of ASTRO-H was evaluated by comparing its onboard GPS receiver data and offline precise orbit determination data. The position error RSS was < 1.7 m (95%) and velocity error RSS was < 11 mm/s (95%). And given this receiver’s specifications of 6 m for position and 30 mm/s for its velocity, it has achieved its design goals. This paper describes how the new GPSR improves the ionosphere-free pseudo-range and carrier phase noise, thereby enabling improved offline precise orbit determination of a satellite. The orbit of ASTRO-H was estimated to be within the precision of a few centimeters, which is among the most accurate for the satellites developed by JAXA to date.