山田 和彦

To create a new flyable detonation propulsion system, a detonation engine system (DES) that can be stowed in sounding rocket S-520-31 has been developed. This paper focused on the first flight demonstration in the space environment of a DES-integrated rotating detonation engine (RDE) using S-520-31. The flight result was compared with ground-test data to validate its performance. In the flight experiment, the stable combustion of the annulus RDE with a plug-shaped inner nozzle was observed by onboard digital and analog cameras. With a time-averaged mass flow of [Formula: see text] and an equivalence ratio of [Formula: see text], the RDE generated a time-averaged thrust of 518 N and a specific impulse of [Formula: see text], which is almost identical to the ideal value of constant pressure combustion. Due to the RDE combustion, the angular velocity increased by [Formula: see text] in total, and the time-averaged torque from the rotational component of the exhaust during 6 s of operation was [Formula: see text]. The high-frequency sampling data identified the detonation frequency during the recorded time as 20 kHz in the flight, which was confirmed by the DES ground test through high-frequency sampling data analysis and high-speed video imaging.

<title>Abstract</title> An arc-heated wind tunnel is one of the most important facilities to reproduce the high-temperature environment during atmospheric entry for plasma studies and spacecraft development. However, the properties of the plasma flow cannot be determined easily, because of the complex physical phenomena, such as arc discharge and supersonic expansion, occurring inside the tunnel. The shock-layer structure should be clarified to evaluate the aerodynamic characteristics, communication conditions, and thermal- protection performance in a high-temperature environment. In this study, shock-layer spectroscopic measurements of a plasma flow in a 1 MW-class arc-heated wind tunnel were performed. The γ-band system spectra of nitric oxide (NO) molecules in the ultraviolet region were measured, and the rotational temperature was determined via spectral fitting through comparison with numerical spectra. The rotational temperature of the NO molecules in the shock layer was 6,620±350 K, whereas that in the free jet was much lower at 770±60 K. This difference is attributed to the increase in translational temperature by flow stagnation across the shock wave, followed by the increase in rotational temperature owing to energy relaxation. A computational science approach revealed the detailed structure of the flow through comparisons with the spectroscopic measurement data. The wind tunnel flow became hypersonic with high temperature and low pressure due to the expansion and acceleration at the nozzle and test chamber. Although the temperature increased across the shock wave, the chemical reaction progressed slowly owing to the low-pressure environment. The rotational temperature in the shock layer increased with the translational temperature; this agrees with the trend of the measurement results. The arc-heated flow was found to be in strong thermo- chemical nonequilibrium in the shock layer. Through this study, a detailed structure of arc-heated flow was revealed and its methodology was also proposed.

<p>Parachute drawing by its cover contributes to simplicity in mechanism of a sample return capsule. Attachment of a band part to suspension lines of the parachute cover is presented to improve attitude stability of the flat plate shaped cover. Aerodynamic characteristics of the cover with the band were evaluated through vertical and horizontal flow wind tunnel tests. The results show that the attachment of the band with an appropriate band perimeter and gap between the band and the cover surface can improve the stability remarkably. Escape route where air flowing inside the band is able to run away is necessary for the stabilization, which is similar as that stability of a parachute relates to air permeation through porosity of the parachute.</p>

<p>An expansion tube is a promising facility to simulate atmospheric entry conditions, although its flow conditions have not been completely characterized mainly owing to its short operation time. In this study, laser absorption spectroscopy was applied to diagnose HEK-X expansion tube flow in the Kakuda Space Center. The target is an absorption line of an oxygen molecule at 763 nm. To increase the sensitivity, optical path length was extended by five times using mirrors. Consequently, an absorption profile with a fractional absorption of 2.4 ± 0.3% was detected at a shock velocity of 7.65 ± 0.05 km/s. The estimated translational temperature from the Voigt fitting was 2750 ± 450 K.</p>

<p>Functionally graded ablative materials with density gradient were newly developed for the thermal protection system of future space exploration missions. The ablating surface was densified to reduce the amount of surface recession, while the density inside the ablator was reduced with expectation of weight reduction and high heat insulation. Typical Bulk specific gravity was found to be about 0.8. Basic thermal characteristics of the developed ablative material were obtained by conducting heating tests. The heating tests were carried out in the arcjet wind tunnel facility in the Chofu Aerospace Center in JAXA for heat flux of 0.9~4.5 MW/m² and impact pressure of 8.6~31 kPa, and in the Sagamihara campus in JAXA for heat flux of 3.6~14.3 MW/m² and impact pressure of 12~54.3 kPa. The amount of surface recession of ablator was successfully obtained during the heating tests. According to X-ray CT inspection conducted after the heating tests, delamination between layers was not observed inside the test piece. In addition, the present study showed the developed ablative material has a potential to reduce the TPS weight by 33.3 % compared with the Hayabusa ablator, although the recession rate of the present ablator is almost same level as the Hayabusa ablator. From a different point of view, the developed ablative material showed a potential to reduce the recession rate in a relatively wide heating environment compared with a medium density ablator, which has the same specific gravity.</p>

The flow enthalpy of an arc-heated wind tunnel is an important parameter for reproducing planetary entry and performing heating tests. However, its distribution is insufficiently clarified, owing to complicated phenomena, such as arc discharge and supersonic expansion. In this paper, the authors assess the enthalpy of an arc-heated flow in a large-scale facility based on measurements and computational results. The flow enthalpy of high-temperature gases, which comprised thermal, chemical, kinetic, and pressure components, was reconstructed based on the measured rotational temperature, heat flux, and impact pressure, in addition to the computational science approach. The rotational temperature of nitric oxide molecules was obtained using emission spectroscopic measurements of band spectra in the near-ultraviolet range. A numerical model was developed and validated based on measured data. The results indicated that the model efficiently reproduced the arc discharge behavior in the heating section and the thermochemical nonequilibrium in the expansion section. It was discovered that the dominant components of the arc-heated flow in the test section were the chemical and kinetic components. The flow enthalpy exhibited a nonuniform distribution in the radial direction. The authors conclude that the flow enthalpy of the core is approximately 28 MJ/kg at the nozzle exit.

The paraglider, a flexible flying vehicle, consists of a parafoil with flexible wings, suspension lines, and a suspended payload. At this time, the suspension lines have several parameters to be designed. Above all, a parameter called Rigging Angle (RA) is sensitive to the aerodynamic characteristics of a paraglider during flight. In this study, the effect of RA is clarified using the two-dimensional stability analysis and a wind tunnel test. The mechanisms about the parafoil-type vehicle stability are clarified through the experimental and analytical approaches as follows. The RA has an allowable range for a stable flight. When the RA is set out of the range, the parafoil cannot fly stably. Furthermore, the behavior of the parafoil wing in the case of lower RA than the allowable range is different from the case of higher RA. The parafoil collapses from the leading edge of the canopy and cannot glide in the case of lower RA.

Attitude control for small satellites is crucial to enable high value missions. Active attitude control is challenging for nanosatellites, due to their small mass and power budgets. On the other hand, the air in low Earth orbit is a promising resource for passive aero-stabilisation of a satellite's orientation. Potential has increased with the development of miniaturised deployable aeroshells for atmospheric entry. The EGG nanosatellite, released into space from the ISS in 2017, is one example of an aeroshell-equipped satellite with no means of active attitude control. Flight data from the EGG mission is a convenient resource to evaluate the concept of passive aero-stabilisation. In this work, a comprehensive coupled atmosphere-orbit-attitude simulation platform for small satellites was developed. The objectives are: (i) to validate the simulation against EGG mission data, and (ii) to evaluate the impact of attitude disturbances on robustness of passive aero-stabilisation. The results provide qualitative validation of the simulation platform, and show that passive attitude control with aeroshell deployed is highly sensitive to initial spin and spacecraft geometric asymmetry. These findings suggest the need for hybrid active control to turn aeroshell-equipped capsules into a viable means of trans-atmospheric transport.

Aerodynamic instability in the attitude of an inflatable re-entry vehicle in the subsonic regime has been observed during suborbital re-entry. This causes significant problems for aerodynamic decelerators using an inflatable aeroshell; thus, mitigating this problem is necessary. In this study, we revealed the instability mechanism using a computational science approach. To reproduce the in-flight oscillation motion in an unsteady turbulent flow field, we adopted a large-eddy simulation approach with a forced-oscillation technique. Computations were performed for two representative cases at transonic and subsonic speeds that were in stable and unstable states, respectively. Pitching moment hysteresis at a cycle in the motion was confirmed for the subsonic case, whereas such hysteresis did not appear for the transonic case. Pressures on the front surface and in the wake of the vehicle were obtained by employing a probe technique in the computations. Pressure phase delays at the surface and in the wake were confirmed as the pitch angle of the vehicle increased (pitch up) and decreased (pitch down), respectively. In particular, we observed that the wake structure formed by a large recirculation behavior significantly affected the pressure phase delay at the rear of the vehicle. The dynamic instability at subsonic speed resulted from flows that could not promptly follow the vehicle motion. Finally, the damping coefficients were evaluated for the design and development of the inflatable vehicle.

<p>The prediction accuracy of arcjet flow using a computer code named ARCFLO3+ is examined by comparing the arc heater operational characteristic data, pitot pressure and cold-wall heat flux data obtained from a segmented constrictor-type arc-heated wind tunnel at JAXA. Results are mainly presented to discuss how the discrepancy between the calculated and measured arc heater operational characteristic data obtained impact the core of the arcjet flow in the test section. Results show that the present computational approach gives a conservative estimation of the arcjet flow core properties within the test section when the mass-averaged enthalpy value obtained through the arc heater is replicated.</p>

An inflatable aerodynamic decelerator with a membrane aeroshell is a promising key technology in the reentry, descent, and landing phases of future space transportation. The membrane aeroshell is generally deformed by the in-flight aerodynamic force; however, the effects of the deformation on the aerodynamic heating are unclear. Here, we investigated aerodynamic heating for an inflatable reentry vehicle, Titans, in the hypersonic regime using flow field simulation coupled with structural analysis. Thermochemical nonequilibrium flows around the Titans with a deformed membrane aeroshell were reproduced numerically for an angle of attack (AoA) values between 0 degrees and 40 degrees. The maximum displacements of the membrane aeroshell by deformation at the AoAs of 0 degrees and 40 degrees were 6.7% and 6.6% of the diameter of the Titans, respectively. The difference in heat fluxes between the deformed and rigid shapes was a remarkable 188.8% for a 0 degrees AoA owing to the considerable changes in the front shock wave shape. Meanwhile, it was indicated that membrane deformation at an AoA of 40 degrees insignificantly affected the peak heat flux value on the inflatable torus because the considerable change in the shock wave shape observed for the case of 0 degrees AoA did not occur. It was found that local wrinkles on the membrane aeroshell were formed by deformation, thus causing the heat flux to increase owing to an increase in local temperature gradient on the surface. (C) 2019 Elsevier Masson SAS. All rights reserved.

The drag coefficient for inflatable reentry vehicles shows discrepancies between a wind tunnel experiment and a flight test in a transonic regime. These discrepancies exist in the drag decrease behavior in the transonic regime and local minimum values of drag above the sonic speed in a wind tunnel. Hence, the present paper focuses on uncovering the reasons and mechanisms behind the same and investigates transonic flowfields around a vehicle by using a transonic wind tunnel and the computational fluid dynamics approach. Several test models with diameters ranging from 56 to 96 mm are used to quantitatively evaluate the effects of scale and a sting attached on the rear. Aerodynamic coefficients, pressures at the rear of the model, and density-gradient distributions are measured for operation conditions of freestream Mach numbers ranging between 0.8 and 1.3. In addition, detailed distributions of the flowfield properties are clarified using the computational fluid dynamics method, which is validated by the experimental data. The results indicate that a sting behind the test models reduces the steep drag decrease at transonic speeds and that shock waves reflected on the test-section walls of the wind tunnel result in local minimum values at supersonic speeds.

Flow features of a supersonic inductively coupled plasma heater that can obtain suitable heat flux for development of membrane material for the flexible aeroshell are numerically examined by means of nonequilibrium magneto-hydrodynamic (MHD) equations. A thermochemical nonequilibrium MHD model was constructed for simulating the radio-frequency discharge of nitrogen from the ICP torch to the conical nozzle, and finally into the ambient test chamber in a uniform manner. The outspread supersonic flow and the thermal nonequilibrium property in the nozzle and in the vacuum chamber were reproduced successfully through the developed numerical model. Due to the effect of the shock wave on the ICP flow, the contours of the translational temperature and Mach number formed separate small areas near the torch outlet in the vacuum chamber. (C) 2019 Author(s).

An inflatable nano-satellite with thin-membrane aeroshell, "EGG", was deployed from the International Space Station at an altitude of approximately 400 km, as part of an orbital deployment mission in 2017. After flight on the low Earth orbit (LEO) for 120 days, EGG successfully reentered Earth's atmosphere and burned out according to mission schedule. During the mission period, surface temperatures of the membrane aeroshell of EGG were measured using thermocouples. The measurement data was sent to the ground station using Iridium short burst data communication. There has not been sufficient investigation of the aerodynamic heating environment of such inflatable vehicles. In this paper, the heat flux on membrane aeroshell was revealed, based on the measured temperature data and a heat conduction simulation technique. In addition, heat flux distributions at an altitude of 120 km were numerically evaluated using the Direct Simulation Monte Carlo (DSMC) method, for cases where the angle of attack was between 0 and 180 degrees. From the reconstructed heat flux history and the DSMC analysis results, it was found that the heat flux on the inflatable torus was higher than that on the membrane aeroshell. Additionally, it was concluded that EGG was in the heating environment of approximately 3-4 kW/m(2) at an altitude of 110 km.

The aerodynamic heating of an inflatable reentry vehicle, which is one of the innovative reentry technologies, was numerically investigated using a tightly coupled approach involving computational fluid dynamics and structure analysis. The fundamentals of a high-enthalpy flow around the inflatable reentry vehicle were clarified. It was found that the flow fields in the shock layer formed in front of the vehicle were strongly in a chemical nonequilibrium state owing to its low-ballistic coefficient trajectory. The heat flux tendencies on the surface of the vehicle were comprehensively investigated for various effects of the vehicle shape, surface catalysis, and turbulence via a parametric study of these parameters. In addition, based on the present results of the computational approach, a new heating-rate method was developed to calculate the heat flux of the nonequilibrium flow. It was demonstrated that the method could well-reproduce the heat flux on the inflatable reentry vehicle.

One of the technologies to make the future planetary exploration more flexible and valuable is a vehicle which can fly freely in the Martian atmosphere. Our group proposes a paraglider which has a flexible and inflatable wing. The paraglider has a large and light wing which can be packed compactly in the launch and cruising phase. To realize this idea, free flight test from high altitude by the Atmospheric balloon must be carried out. However, it is difficult to succeed flight test from beginning. Therefore, we’re going to carry out a relatively simple flight test from low altitude. In this study, result of flight test was reported.

In sample-return missions, the reentry velocity of a sample-return capsule is expected to be approximately 15 km/s however, the reentry velocity of the Hayabusa sample-return capsule was 11.8 km/s. Strong aerodynamic heating caused by a high velocity can damage the capsule during reentry. To overcome this, two designs of highvelocity reentry capsules were proposed. In one design, a rigid flare was attached to decrease the ballistic coefficient by increasing the front projected area. In the other design, the conventional Hayabusa sample-return capsule was used with no modifications. In this study, the aerodynamic heating of the high-velocity reentry capsules and the Hayabusa sample-return capsule was analyzed using numerical simulations. Plasma flow in the shock layer at the front of the capsules and expansion flow in the wake region around the capsules were investigated. The profiles of convective and radiative heat fluxes on the surfaces of these capsules were predicted. The heat fluxes at the stagnation points predicted by the present numerical simulation were in good agreement with that of the empirical models. At the strongest aerodynamic-heating altitude, the total heat fluxes at the rear of the high-velocity reentry capsules and the Hayabusa sample-return capsule were approximately 2% of those in front of the capsules.

Aerodynamic heating around a flare-type membrane inflatable vehicle during Earth atmospheric reentry was investigated using numerical simulation approach. This vehicle, which is mainly composed of the capsule, membrane aeroshell and inflatable torus, has been developed by the Membrane Aeroshell for Atmospheric-entry Capsule (MAAC) group as a one of the innovative reentry systems. Analysis solver for reentry flows around the vehicle was RG-FaSTAR, which is a branch version of JAXA fast aerodynamic routine (FaSTAR). In addition, structure analysis solver also was used for membrane deformation in a loosely-coupled manner with the flow field. In the present research, the effects of angle of attack (AoA) and membrane aeroshell deformation on aerodynamic heating were investigated. The numerical results showed that heat flux distribution drastically varies with the increase in AoA because of changes of flow field, and heat flux value at the stagnation point for case of AoA of 40 degree was 3.09 times as high as that for 0 degree. Moreover, the deformed shapes for case of AoA of 0 and 40 degrees were calculated in the way which the pressure distributions obtained using initial (undeformed) shape were given as the aerodynamic force. The difference of heat fluxes between the deformed and initial shapes on the head capsule part was remarkable as 188.8% for case of AoA of 0 degree. On the other hand, it was indicated that membrane deformation for case of AoA of 40 degree insignificantly affects the peak heat flux value on the inflatable torus such as the case of the AoA of 0 degree.

<p>In recent years, a number of sample return mission and planetary exploration probes have been discussed and proposed. Our group has developed a new atmospheric re-entry vehicle with a membrane aeroshell to increase the variety of these missions. However, there are still several important technical problems to be addressed to apply the membrane aeroshell to an actual mission. One of them is evaluation of the thermal durability of the inflatable structure. The thermal durability of the inflatable structure was evaluated using a 10 kW class inductively coupled plasma (ICP) heater. This ICP heater can produce a plasma flow with a high enthalpy and relatively low heat flux of about 120 kW/m², which is a suitable condition for the heating test of the membrane aeroshell. The tests proved that the inflatable structure, made of polyimide film, silicon rubber adhesive, ZYLON textile, and alumina felt, maintains the gas tight in the plasma flow with a heat flux of 120 kW/m² in 300 s. This layering structure is proposed as a potential candidate for use in actual flight vehicles.</p>

A flight experiment of an inflatable reentry vehicle, equipped with a thin-membrane aeroshell deployed by an inflatable torus structure, was performed using a Japan Aerospace Exploration Agency S-310-41 sounding rocket. The drag coefficient history was evaluated by analyzing the acceleration of the vehicle with the atmospheric density and temperature using a global reference atmospheric model. The vehicle successfully demonstrated deceleration. During the reentry flight, the position, velocity, and acceleration of the vehicle were obtained by using the Global Positioning System. The experimental drag coefficient had an almost constant value of 1.5 in the supersonic region but decreased to 1.0 in the subsonic region. In the transonic region, a steep decrease of the drag coefficient was confirmed. To study the detailed aerodynamics for the reentry vehicle, flowfield simulations were conducted with computational fluid dynamics techniques. The aerodynamic force acting on the vehicle was investigated with the measured data throughout the supersonic and subsonic regions. In the flowfield simulation, the computed result for the drag coefficient showed reasonable agreement with the experimental one. In addition, a compressible effect in front of the vehicle was seen to appear in the supersonic region and a vortex ring at the rear of the vehicle was formed in the subsonic region.