Kazuhiko Yamada

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>

<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>

Balloon Symposium 2018 (November 1-2, 2018. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan

<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>

This study discusses an airbag system as a landing gear of a spacecraft for a planetary exploration. The inflated airbag has the capability to attenuate impact acceleration at the instant of landing and submergence into regolith that covers a planetary surface. Crash tests in this paper verify the attenuation performance of the airbag and specify an issue.