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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 gamma-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 6620 +/- 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 thermochemical nonequilibrium in the shock layer. Through this study, a detailed structure of arc-heated flow was revealed and its methodolo

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 gamma-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 6620 +/- 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 thermochemical nonequilibrium in the shock layer. Through this study, a detailed structure of arc-heated flow was revealed and its methodolo

Over the past decades, engineers have focused on a common goal - to reduce emissions by making industrial processes more efficient and by utilizing renewable energy sources. One possibility to reach this goal can be achieved by employing processes with non-ideal fluids, such as fluids at supercritical conditions in refrigeration, heat pump cycles and power cycles. Most of the flows in industry are turbulent and hence the need to study turbulence in non-ideal fluids arose. Turbulence in non-ideal fluids is extremely challenging since there are many complex effects at play. One such effect is caused by variations in properties, which also occurs in compressible flows, flows with high concentration gradients, or flows in heat exchangers. This thesis presents a review of the existing theory on semi-local scaling for variable property flows with the aim to take it one step further and apply it to turbulent heat flux modeling. Two types of variable property cases are analysed in this thesis; (1) low-Mach number flows with uniform pseudo-heating sources, (2) high-Mach number flows with non-uniform viscous heating. For the former, a Direct Numerical Simulation (DNS) data-base, already available at TU Delft, has been post-processed with the main goal to investigate if the semi-local theory can also be applied to thermal turbulence and its modeling. For the latter, additional DNS simulations of high-Mach number channel flows have been performed to investigate how the viscous heating and its correlations can be accounted for in the semi-local scaling framework. Using a 2-equation heat flux model, we find that modeling thermal turbulence in semi-local scales considerably improves the results for low-Mach number flows. However, for high-Mach number flows, additional unknown (closure) terms arise due to fluctuations in the viscous heating source. A model for the source term in the enthalpy variance equation is successfully proposed. In addition, the DNS stu, Mechanical Engineering | Energy and Process Technology

The purpose of this thesis is investigation of aerodynamic heating and ablation of thermal protection systems for high speed missiles. A one-dimensional solver called AeroheataBS (Aerodynamic Heating and Ablation Simulation) has been developed for this purpose. Time dependent mission profile of a missile, flight velocity, altitude and angle of attack, is used as an input for the aerodynamic heating analysis. Atmospheric properties are used as a function of altitude for the computations. Convectional heating is calculated using external flow correlations and compressibility effects are included through Eckert Reference Temperature method. The flow properties at the boundary-layer edge are computed from shock relations and the surface pressure is evaluated using Modified Newton Theory. The Arrhenius equation is used for the calculation of thermo-chemical decomposition of the insulation material. Effective heat of ablation of the material is used for the surface recession rate. The reduction of convective heating due to injection of the ablation products into the boundary layer is included in the mathematical model. The explicit finite difference method is used for the temporal discretization of the governing equations. Temperature distribution along the nodal points in the body and insulation material is evaluated at each time point. The material properties are computed as a function of temperature. Validation of the methodology developed in the present study is performed through analytical methods where available, flight telemetry data, ground testing and CFD (Computational Fluid Dynamics). The flight data obtained from X-15 and HIFiRE-5 are used for validation but as well as the telemetry data taken from a newly developed missile where the thermal design was carried out by the present methodology, AeroHeataBS. The present predictions are compared with the data that is obtained from the inner surfaces of the missile by the telemetry. The predictions are compared with, Bu tez çalışmasında, yüksek hızlı füzelerin aerodinamik ısınma kestirimi ve aerodinamik ısınma altında ısıl koruma sistemlerinde oluşan aşınmanın incelenmesi amaçlanmıştır. Bu süreçte bir boyutlu AeroheataBS (Aerodynamic Heating and Ablation Simulation) adıyla bir çözücü geliştirilmiştir. Aerodinamik ısınma analizleri için füzeye ait zamana bağlı uçuş hızı, irtifa ve hücum açısı girdi olarak kullanılmıştır. Atmosferik özellikler irtifaya bağlı bir şekilde standart atmosfer modeli kullanılarak çözüme yansıtılmıştır. Aerodinamik ısınmanın hesaplanmasında dış akış bağıntılarından yararlanılmış, yüksek hızlardaki sıkıştırılabilirlik etkileri Eckert Referans Sıcaklık yöntemiyle dikkate alınmıştır. Sınır tabaka kenar çizgisi üzerindeki akış özellikleri şok bağıntılarıyla, yüzey basıncı ise değiştirilmiş Newton kuramı kullanılarak hesaplanmıştır. Malzemenin termo-kimyasal bozunması Arrhenius denklemiyle modellenmiştir. Yüzey aşınması malzemenin etkin bozunma ısısı kullanılarak çözücüye eklenmiştir. Sınır tabakaya salınan bozunma reaksiyonu ürünlerinin, taşınım ısı geçişine etkisi matematiksel olarak modellenmiştir. Bünye denklemleri açık sonlu fark yaklaşımıyla ayrıklaştırılmış ve çözülmüştür. Gövde ve yalıtım üzerinde oluşturulan çözüm ağının her bir düğüm noktasında zamana bağlı sıcaklık hesaplanmıştır. Malzeme özellikleri sıcaklığa bağımlı olarak kullanılmıştır. Çözücünün doğrulama çalışmaları analitik çözümler, uçuşlu test verileri, yer testi ölçümleri ve HAD (Hesaplamalı Akışkanlar Dinamiği) analizleri ile gerçekleştirilmiştir. X-15, HIFiRE-5 sistemlerinin uçuşlu test verilerine ilaveten ısıl tasarımı yapılan bir füzeye ait gövde ile kanard sıcaklıkları da kullanılmıştır. Kıyaslamalarda gövde içerisinden ölçülen ve telemetre ile yer istasyonuna indirilen sıcaklık verileri kullanılmıştır. Açık erişim yayınlarından ulaşılan ve sözü geçen alanda güvenilirliği kanıtlanmış FIAT yazılımına ait sonuçlarla da kıyaslamalı değerlendirmeler yapılmıştır. Sabit aşınma hızı altında

The purpose of this thesis is investigation of aerodynamic heating and ablation of thermal protection systems for high speed missiles. A one-dimensional solver called AeroheataBS (Aerodynamic Heating and Ablation Simulation) has been developed for this purpose. Time dependent mission profile of a missile, flight velocity, altitude and angle of attack, is used as an input for the aerodynamic heating analysis. Atmospheric properties are used as a function of altitude for the computations. Convectional heating is calculated using external flow correlations and compressibility effects are included through Eckert Reference Temperature method. The flow properties at the boundary-layer edge are computed from shock relations and the surface pressure is evaluated using Modified Newton Theory. The Arrhenius equation is used for the calculation of thermo-chemical decomposition of the insulation material. Effective heat of ablation of the material is used for the surface recession rate. The reduction of convective heating due to injection of the ablation products into the boundary layer is included in the mathematical model. The explicit finite difference method is used for the temporal discretization of the governing equations. Temperature distribution along the nodal points in the body and insulation material is evaluated at each time point. The material properties are computed as a function of temperature. Validation of the methodology developed in the present study is performed through analytical methods where available, flight telemetry data, ground testing and CFD (Computational Fluid Dynamics). The flight data obtained from X-15 and HIFiRE-5 are used for validation but as well as the telemetry data taken from a newly developed missile where the thermal design was carried out by the present methodology, AeroHeataBS. The present predictions are compared with the data that is obtained from the inner surfaces of the missile by the telemetry. The predictions are compared with, Bu tez çalışmasında, yüksek hızlı füzelerin aerodinamik ısınma kestirimi ve aerodinamik ısınma altında ısıl koruma sistemlerinde oluşan aşınmanın incelenmesi amaçlanmıştır. Bu süreçte bir boyutlu AeroheataBS (Aerodynamic Heating and Ablation Simulation) adıyla bir çözücü geliştirilmiştir. Aerodinamik ısınma analizleri için füzeye ait zamana bağlı uçuş hızı, irtifa ve hücum açısı girdi olarak kullanılmıştır. Atmosferik özellikler irtifaya bağlı bir şekilde standart atmosfer modeli kullanılarak çözüme yansıtılmıştır. Aerodinamik ısınmanın hesaplanmasında dış akış bağıntılarından yararlanılmış, yüksek hızlardaki sıkıştırılabilirlik etkileri Eckert Referans Sıcaklık yöntemiyle dikkate alınmıştır. Sınır tabaka kenar çizgisi üzerindeki akış özellikleri şok bağıntılarıyla, yüzey basıncı ise değiştirilmiş Newton kuramı kullanılarak hesaplanmıştır. Malzemenin termo-kimyasal bozunması Arrhenius denklemiyle modellenmiştir. Yüzey aşınması malzemenin etkin bozunma ısısı kullanılarak çözücüye eklenmiştir. Sınır tabakaya salınan bozunma reaksiyonu ürünlerinin, taşınım ısı geçişine etkisi matematiksel olarak modellenmiştir. Bünye denklemleri açık sonlu fark yaklaşımıyla ayrıklaştırılmış ve çözülmüştür. Gövde ve yalıtım üzerinde oluşturulan çözüm ağının her bir düğüm noktasında zamana bağlı sıcaklık hesaplanmıştır. Malzeme özellikleri sıcaklığa bağımlı olarak kullanılmıştır. Çözücünün doğrulama çalışmaları analitik çözümler, uçuşlu test verileri, yer testi ölçümleri ve HAD (Hesaplamalı Akışkanlar Dinamiği) analizleri ile gerçekleştirilmiştir. X-15, HIFiRE-5 sistemlerinin uçuşlu test verilerine ilaveten ısıl tasarımı yapılan bir füzeye ait gövde ile kanard sıcaklıkları da kullanılmıştır. Kıyaslamalarda gövde içerisinden ölçülen ve telemetre ile yer istasyonuna indirilen sıcaklık verileri kullanılmıştır. Açık erişim yayınlarından ulaşılan ve sözü geçen alanda güvenilirliği kanıtlanmış FIAT yazılımına ait sonuçlarla da kıyaslamalı değerlendirmeler yapılmıştır. Sabit aşınma hızı altında

Over the past decades, engineers have focused on a common goal - to reduce emissions by making industrial processes more efficient and by utilizing renewable energy sources. One possibility to reach this goal can be achieved by employing processes with non-ideal fluids, such as fluids at supercritical conditions in refrigeration, heat pump cycles and power cycles. Most of the flows in industry are turbulent and hence the need to study turbulence in non-ideal fluids arose. Turbulence in non-ideal fluids is extremely challenging since there are many complex effects at play. One such effect is caused by variations in properties, which also occurs in compressible flows, flows with high concentration gradients, or flows in heat exchangers. This thesis presents a review of the existing theory on semi-local scaling for variable property flows with the aim to take it one step further and apply it to turbulent heat flux modeling. Two types of variable property cases are analysed in this thesis; (1) low-Mach number flows with uniform pseudo-heating sources, (2) high-Mach number flows with non-uniform viscous heating. For the former, a Direct Numerical Simulation (DNS) data-base, already available at TU Delft, has been post-processed with the main goal to investigate if the semi-local theory can also be applied to thermal turbulence and its modeling. For the latter, additional DNS simulations of high-Mach number channel flows have been performed to investigate how the viscous heating and its correlations can be accounted for in the semi-local scaling framework. Using a 2-equation heat flux model, we find that modeling thermal turbulence in semi-local scales considerably improves the results for low-Mach number flows. However, for high-Mach number flows, additional unknown (closure) terms arise due to fluctuations in the viscous heating source. A model for the source term in the enthalpy variance equation is successfully proposed. In addition, the DNS stu, Mechanical Engineering | Energy and Process Technology

Yüksek hızlı hava araçlarının gövde sıcaklıklarını ve aerodinamik ısı akılarını hesaplama yazılımı olan, bir boyutlu aerodinamik ısınma ve ısıl aşınma kestirim yazılımı, AeroheataBS’in doğrulama çalışmaları yürütülmüştür. Şok denklemleri, düz plaka üzeri taşınım ısı transferi yaklaşımları, Eckert’in referans sıcaklık yöntemi ve değiştirilmiş Newton yasası aerodinamik ısınmanın hesaplanmasında kullanılmıştır. Denklemler açık olarak ayrıklaştırılmış ve sayısal olarak çözülmüştür. Işınım ile olan ısı transferi de dikkate alınmıştır. Uçuş hızı, uçuş irtifası ve hücum açısı zamana bağlı olarak tanımlanmıştır. Hesaplamaların doğruluğu literatürde bulunan uçuş verileriyle ve hesaplamalı akışkanlar dinamiği çalışmalarının sonuçlarıyla kıyaslanarak değerlendirilmiştir. Kıyaslamalarda gözlemlenen uyumlu sonuçlar AeroheataBS yazılımında kullanılan yöntemin aerodinamik ısınmanın ve gövde sıcaklıklarının kestiriminde kullanılabileceğini göstermiştir., Validation of one-dimensional aerodynamic heating and ablation prediction program, AeroheataBS to calculate transient skin temperatures and heat fluxes for high-speed vehicles has been performed. In the tool shock relations, flat plate convective heating expressions, Eckert's reference temperature method and modified Newtonian flow theory are utilized to compute local heat transfer coefficients. Corresponding governing equations are discretized explicitly and numerically solved. Time varying flight conditions including velocity, altitude and angle of attack serve as input to the program. In order to examine the accuracy of aerodynamic heating capabilities of AeroheataBS, calculated temperature histories are compared with flight data of the X-15 research vehicle, a modified von-Karman nose shaped body, cone-cylinder-flare configuration and results of conjugate computational fluid dynamics studies. Comparative studies show that computed values are in good agreement with the reference data and prove that methodology established in AeroheataBS is appropriate for estimating aerodynamic heating and structural thermal response.

Yüksek hızlı hava araçlarının gövde sıcaklıklarını ve aerodinamik ısı akılarını hesaplama yazılımı olan, bir boyutlu aerodinamik ısınma ve ısıl aşınma kestirim yazılımı, AeroheataBS’in doğrulama çalışmaları yürütülmüştür. Şok denklemleri, düz plaka üzeri taşınım ısı transferi yaklaşımları, Eckert’in referans sıcaklık yöntemi ve değiştirilmiş Newton yasası aerodinamik ısınmanın hesaplanmasında kullanılmıştır. Denklemler açık olarak ayrıklaştırılmış ve sayısal olarak çözülmüştür. Işınım ile olan ısı transferi de dikkate alınmıştır. Uçuş hızı, uçuş irtifası ve hücum açısı zamana bağlı olarak tanımlanmıştır. Hesaplamaların doğruluğu literatürde bulunan uçuş verileriyle ve hesaplamalı akışkanlar dinamiği çalışmalarının sonuçlarıyla kıyaslanarak değerlendirilmiştir. Kıyaslamalarda gözlemlenen uyumlu sonuçlar AeroheataBS yazılımında kullanılan yöntemin aerodinamik ısınmanın ve gövde sıcaklıklarının kestiriminde kullanılabileceğini göstermiştir., Validation of one-dimensional aerodynamic heating and ablation prediction program, AeroheataBS to calculate transient skin temperatures and heat fluxes for high-speed vehicles has been performed. In the tool shock relations, flat plate convective heating expressions, Eckert's reference temperature method and modified Newtonian flow theory are utilized to compute local heat transfer coefficients. Corresponding governing equations are discretized explicitly and numerically solved. Time varying flight conditions including velocity, altitude and angle of attack serve as input to the program. In order to examine the accuracy of aerodynamic heating capabilities of AeroheataBS, calculated temperature histories are compared with flight data of the X-15 research vehicle, a modified von-Karman nose shaped body, cone-cylinder-flare configuration and results of conjugate computational fluid dynamics studies. Comparative studies show that computed values are in good agreement with the reference data and prove that methodology established in AeroheataBS is appropriate for estimating aerodynamic heating and structural thermal response.

A simplified method is developed to calculate aerodynamic heating, ablation, and structural temperature response for a body traveling at high speeds. Mach number, altitude, and angle of attack are used as a function of time. Compressibility effects are considered by using Eckert's reference temperature approach. Convective aerodynamic heating is calculated using external flow relations. Local pressure values are found through modified Newtonian theory. An approximate recession method that uses the heat of ablation is coupled to the aerodynamic heating. An in-depth solution accounts for material decomposition; however, it does not include the energy absorption of pyrolysis gases through the material. Reduction in the heat transfer coefficient caused by the transpiring gases is estimated. An Arrhenius equation is used to model the density of the ablative material. The method is examined for the validation of six different cases, and predictions are found to be in good agreement with experimental and analytical results. The verification studies indicate that the methodology is well suited for predicting the ablation and thermal response of a thermal protection system. (c) 2019 American Society of Civil Engineers.

A simplified method is developed to calculate aerodynamic heating, ablation, and structural temperature response for a body traveling at high speeds. Mach number, altitude, and angle of attack are used as a function of time. Compressibility effects are considered by using Eckert's reference temperature approach. Convective aerodynamic heating is calculated using external flow relations. Local pressure values are found through modified Newtonian theory. An approximate recession method that uses the heat of ablation is coupled to the aerodynamic heating. An in-depth solution accounts for material decomposition; however, it does not include the energy absorption of pyrolysis gases through the material. Reduction in the heat transfer coefficient caused by the transpiring gases is estimated. An Arrhenius equation is used to model the density of the ablative material. The method is examined for the validation of six different cases, and predictions are found to be in good agreement with experimental and analytical results. The verification studies indicate that the methodology is well suited for predicting the ablation and thermal response of a thermal protection system. (c) 2019 American Society of Civil Engineers.

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.

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.

Materials suitable for the Thermal Protection System (TPS) of the rocket fairing for the Japanese next generation launch vehicle named H-2 were tested using an arc-heated wind tunnel. The main constituents of the candidate materials were epoxy resin and silicon rubber. The heating rate was from 60 to 100 kW/(sq m) in accordance with the maximum heating rate of the nose region of an ascending H-2 rocket. Test results showed that polycondensed silicon rubber was the best candidate as the main component of the TPS, and also that a SiC coating was indispensable. In addition, based on results and a preliminary estimation of the skin temperatures of the fairing nose region, a SiC coating of about two mm in thickness is considered sufficient to achieve thermal protection capability., 日本の次世代打ち上げ用ロケットH-2の頭部フェアリングに用いる熱防御システム(TPS)に適した材料をアーク加熱風洞により試験した。候補材料の主なものはエポキシ樹脂とシリコンゴムである。H-2ロケット上昇時の頭部空力加熱の最大値に合せて試験時加熱率は60〜100kW/平方メートルとした。試験結果によれば、TPS主材としては縮合型シリコンゴムが最適で、また、シリカコーティングは欠かすことのできないものである。また、フェアリング頭部の表面温度に関する試験と推算の結果から、熱防御のためのシリカコーティングの厚さは2mmで十分であると思われる。

Transient thermal response behavior of hypersonic reentry flight test vehicle HYFLEX (Hypersonic Flight Experiment) was investigated. The iterative CFD (Computational Fluid Dynamics)/FEM (Finite Element Method) coupling analysis was carried out. In this study, flow field around the body is computed by the finite volume Navier-Stokes equation solver FIVAD. On the other hand, heat conduction behavior is calculated by the finite element solver THAP. The NWT (Numerical Wind Tunnel) system and CMS (Common work Station) system of NAL are used for simulation, with the data file transfer by LAN (Local Area Network)., 極超音速で大気圏に再突入する実験機HYFLEX(極超音速飛行実験)について、空力加熱解析と構造熱伝導解析を連成させて、非定常熱応答の数値シミュレーションを試みた。対象をHYFLEX機体前胴部とし、空力解析は有限体積法コードFIVADを用い、熱伝導解析は有限要素法コードTHAPを用いた。計算には航空宇宙技術研究所の大規模並列ベクトルコンピュータシステムである数値風洞(NWT)と構造計算用ワークステーションシステム(CMS)を用い、データファイル転送による分散処理を行った。

When a space vehicle reenters onto the atmosphere with high velocity, radiative heating from the shocked air ahead of the vehicle plays an important role on the heat flux to the wall surface as well as convective heating. In the laboratory, the spectroscopic system has been developed for temperature measurement of radiation behind strong shock waves. In this paper, CARS method (Coherent Anti-Stokes Raman Spectroscopy) for radiative flow behind strong shock waves of high Mach number has been developed. The CARS measurement system consists of a YAG laser, a dye laser, optical system, and high-sensitivity I.I.-CCD camera systems with spectroscopy. The preliminary tests have been performed for the purpose of detecting CARS signal from strong shock waves with velocity range over 5 km/s. Finally, both vibrational and rotational temperatures could be estimated by fitting of the CARS spectra obtained from experiment and analytical calculation over Mach 15.

An expansion tube has a unique capability of generating a hypervelocity flow without experiencing a stagnation condition of the test gas. Related to MUSES-C superorbital reentry flow condition, experimental and numerical simulations were conducted in domestic institutions as well as in University of Queensland, Australia. This paper summarizes the characteristics of the facility and results of their investigations., The Institute of Space and Astronautical Science Report SP

Transient thermal response behavior of hypersonic reentry flight test vehicle HYFLEX (Hypersonic Flight Experiment) was investigated. The iterative CFD (Computational Fluid Dynamics)/FEM (Finite Element Method) coupling analysis was carried out. In this study, flow field around the body was computed using the finite volume Navier-Stokes equation solver FIVAD. On the other hand, heat conduction behavior was calculated using the finite element solver THAP. The NWT (Numerical Wind Tunnel) system and CMS (Common Work Station) system of NAL were used for simulation, with the data file transfer by LAN (Local Area Network)., 極超音速で大気圏に再突入するHYFLEX(極超音速飛行実験)実験機について、過渡熱応答挙動を研究した。ここでは反復CFD/FEM(計算流体力学/有限要素法)結合解析を実施した。本研究では機体周りの流れ場をナビエ・ストークス方程式有限体積法コードFIVADを用いて計算した。熱伝導挙動は有限要素法コードTHAPを用いて計算した。シミュレーション計算には航空宇宙技術研究所の数値風洞(NWT)と共通ワークステーションシステム(CMS)を用い、データファイル転送は局地ネットワーク(LAN)を使用した。

This report presents a detailed description and results of force and heat flux measurements conducted at the High-Enthalpy Shock Tunnel (HIEST) of the Japan Aerospace Exploration Agency (JAXA). The HB-2 standard hypersonic ballistic configuration was employed as a model. The force measurement tests used an aluminum alloy model and a three-component aerodynamic balance. An acceleration compensation technique was used to remove vibration components from the original data. Heat flux measurement tests used a chromel model with a total of 28 co-axial thermocouples press-fitted onto the surface. Time histories of temperature data were numerically integrated to determine the surface heat flux by applying one-dimensional heat conduction theory. Both force and heat tests used a conical nozzle, and a total of three enthalpy levels of 4, 8, and 11 MJ/kg and two levels of 4 and 8 MJ/kg were set for the force and heat tests, respectively. Although much higher stagnation enthalpy can be attained, relatively lower enthalpy levels were selected to compare with a blow-down type hypersonic wind tunnel and to alleviate uncertainties inherent in high-enthalpy facilities. The force test results correlated well with other hypersonic wind-tunnel data for the axial force coefficient at the enthalpy level of 4 MJ/kg, while a notable discrepancy was observed at the higher enthalpy levels of 8 and 11 MJ/kg. The real gas effects on aerodynamic characteristics are examined and discussed. In the heat test, a comparison of the heat flux distribution along the surface with data obtained in a blow-down type hypersonic wind tunnel showed good agreement in the nose part, typically within several percent. The present experiment was conducted as a series of comparative test campaigns between two hypersonic facilities in JAXA. Since the tunnel stagnation conditions and the corresponding data are tabulated in detail for each shot, the present test data are believed to serve as a complete database for eva, 本報告は、JAXA高エンタルピ衝撃風洞(HIEST)において実施された、HB-2形状バリスティック標準模型を用いた空気力/空力加熱率測定試験結果に関するものである。力試験はアルミ製模型に十字1列型3分力天秤を装着して実施され、データの振動成分を除去するために加速度補正が施された。熱試験で用いられた模型はクロメル製で計28個の同軸熱電対が装着されてあり、温度時間履歴データを1次元熱伝導の仮定の下で数値的に積分することにより加熱率を求めた。力/熱試験共にコニカルノズルを使用し、力試験についてはエンタルピ4、8、11MJ/kg、熱試験については4、8MJ/kgの試験条件でそれぞれ実施された。力試験結果はエンタルピ4MJ/kg条件の場合、軸力係数に関して他風洞の試験データと良好な傾向の一致を示している一方で、8、11MJ/kgの場合は系統的なずれが観測されており、実在気体効果の影響が可能性として示唆される。一方熱試験における無次元加熱率データは、ブローダウン式コールド極超音速風洞データと模型頭部において数％程度の範囲でよい一致を示した。本試験は、JAXA極超音速関連風洞対応風試の一環として実施されたものであり、また本試験データは高エンタルピ衝撃風洞における基礎データとしても有用である。, JAXA Research and Development Report, 宇宙航空研究開発機構研究開発報告

Hypersonic CFD (Computational Fluid Dynamics) analysis is made for blunt nosed, ramped flat plates. Heat transfer and flow field around compression ramp have been investigated numerically. The present work correspond to the experimental study using the NAL 1.27 m and ONERA (Office National d'Etudes et de Recherches Aerospatiales) S4MA hypersonic wind tunnels, where the infrared thermography measurements are made at M(sub infinity) = 10.0 and angles of attack 0 deg, 20 deg, 30 deg and 35 deg. Heat transfer distributions and shock-shock interaction patterns are compared in detail, changing angles of attack and Reynolds numbers. Comparisons with CFD simulation using NAL and ONERA Navier-Stokes Codes were also made., 極超音速CFD(数値流体力学)解析を鈍頭ノーズ・ランプ平板について行った。圧縮ランプ周りの熱伝達および流れ場を数値的に調べた。本研究はNAL(航空宇宙技術研究所)の1.27m、ONERA(フランス国立航空宇宙技術研究所)のS4MA極超音速風洞を用いた実験研究に対応し、赤外サーモグラフィー計測を主流マッハ数10.0、迎え角0度、20度、30度および35度で行った。迎え角およびレイノルズ数を変えて、熱伝達分布および衝撃波間干渉パターンを詳細に比較した。NALおよびONERAナビエ・ストークスコードを用いたCFDシミュレーションとも比較した。

In the HYFLEX (Hypersonic Flight Experiment), a consecutive hypersonic flight duration of some five minutes is secured. Therefore, it is expected to get the hypersonic aerodynamic/aerothermodynamic performance data good enough in both qualitative and quantitative sense for the region of M greater than 10, where best performance is crucial for the reusable space vehicles. Selection of the measurement items for the onboard measurement system was made among both aerodynamics and aerothermodynamics to get the real gas effects, thermostructural evaluation data, basic data for evaluation of lifting body technology (guidance, navigation and control) etc., consisting of some 250 channels in total. Grounds to select measurement items, the analysis and tests for design and confirmation, and specifications of each items are outlined and discussed in this paper. But mostly measurement mission items are discussed in this paper. Major objective of HYFLEX onboard measurement is to contribute to establishing reliable prediction/design methods for the winged space vehicles. To do this, the comparison between the flight data and the predictions made in advance based on ground tests and analyses is indispensable. For the meaningful comparison, the atmospheric condition along flight path must be known accurately enough. In HYFLEX, the utilization of the remotely sensed atmospheric data by NOAA meteorological satellite was planned to raise the inference accuracy. Data analysis program is briefly presented., 極超音速飛行実験においては、5分間程度の持続的な極超音速飛行時間が確保されるため、再使用型往還機に重要なM＞10領域の極超音速空力、空力加熱特性データが質的および量的ともに充実したレベルで得られることが期待された。機上搭載計測系のための計測項目の選定において、極超音速領域の実在気体効果を得るための空力、空力加熱特性データ、熱構造特性評価用データ、および揚力飛行物体の基本データ(誘導、航法および軌道制御に関するデータ)などを抽出し、約250チャンネルの計測項目計画をたてた。ここでは各計測項目を選定した基準、設計とその確認のための解析・試験、各項目の仕様について概観し、特に具体的議論は主にデータ計測系に限った。HYFLEX機上搭載計測の主目的は有翼宇宙往還機に対する信頼度の高い予測手法や設計手法の確立に資することである。そのためには、地上での事前試験や解析に基づいた予測をフライトデータと比較することが不可欠となる。有意義な比較のためには、飛行経路に沿った大気状態を充分正確に把握する必要がある。NOAA気象衛星のリモートセンシング大気データを使用し、推定精度を上げることを試みている。データ解析計画の概略も述べてある。

The normal emittance of the Black High Emittance (BHE) coating for refractory tiles has been measured for surface temperatures of 473 to 1,473 K. The changes in emittance between new and exposed tiles were evaluated, and no distinct difference was found. The effects of thin metallic and ceramic coatings on BHE were also measured and discussed., OREX(軌道再突入実験機)軌道再突入実験の飛行実験データ解析のため、耐熱タイル黒色コーティングの放射率計測を行った。低温域(473K)での放射率計測には、放射率は同一温度の黒体試料を参照して求めるFT-IR(フーリエ変換赤外分光法)装置(波長範囲0.2〜25マイクロメートル)を使用し、高温域(1,073〜1,473K)では試料移動法により全垂直放射率(波長範囲0.2〜9マイクロメートル)を計測する高温放射率計測装置を使用した。耐熱タイル黒色コーティング供試体については、未使用品およびOREX飛行を模擬したアーク加熱風洞試験済み供試体を計測し、加熱による放射率の変化を観察した。その結果、低温域および高温域とも、両供試体の放射率に顕著な差異のないことが分った。また、耐熱タイル黒色コーティング表面に薄膜コーティングを施して、表面特性を改変した3種類の供試体(SiC、Au、Ir)の計測を行い、薄膜コーティングが計測結果に与える影響などについても討論した。

The determination of the return flight path of a winged reentry vehicle is complicated due to various factors including the flight range and aerodynamic heating. A study of the feasible flight corridor according to the gliding capability of the vehicle is useful, which enables a rough estimate of the return flight path of the vehicle. This research could contribute to HOPE-X (H-2 Orbiting Plane Experimental) development by obtaining a feasible flight corridor. The estimation of the return flight path was studied by means of solving a set of equations of motion for the vehicle. A quasi-steady state was defined as a state in which the equations converge to a certain flight path. It was found that the path which satisfied the quasi-steady state was regarded as a feasible path since it roughly coincided with the dynamic path at a fixed L/D (Lift to Drag ratio). Due to the small reentry angle of the quasi-steady state path, this path was regarded as unrealistic althouth it was feasible., 有翼往還機の帰還飛行経路の決定は、考慮すべき因子がさまざまなため、複雑となっている。機材の滑空能力による実現可能な飛行経路を見つけることが有用であり、これにより、機材の帰還飛行経路のおおまかな評価を行うことが可能になる。この研究によって可能な飛行経路を示すことにより、HOPE-X(宇宙往還技術試験機)の開発に貢献できると期待されている。さまざまなL/D条件の準定常状態の経路、速度制動角、バンク角によって、有翼往還機の飛行経路が決まる。準定常状態経路の再突入速度は、円軌道速度に近く、再突入角が小さいために、再突入フェーズにおいては非現実的な経路であると考えられる。他の方法により、準定常状態の方法によって与えられる経路と接続した再突入経路を見つける必要がある。

JAXA Special Publication, 宇宙航空研究開発機構特別資料

The outline of the mission and initial analysis of the flight data of the recombination heating and ablation sensors installed on the fourth row of the ceramic tile Thermal Protection System (TPS) of Orbit Reently Experiment (OREX) are described., 機体表面での再結合加熱にたいする表面触媒性の効果を観測することを目的とした再結合加熱センサとアブレータによる熱防護特性評価のためのアブレータセンサを製作しOREX第4列タイル上に搭載した。再結合加熱センサは3種類の表面(BHE,SiC,Au)の触媒効果の差異の観測を、アブレータセンサは深さ方向温度履歴データより熱ブロッキング効果を評価する。これらセンサで得られた飛行データの初期解析結果を述べる。

In order to validate real gas CFD (Computational Fluid Dynamics) code and construct appropriate real gas models, analysis of HEK (High Enthalpy shock tunnel at Kakuda Research Center), high enthalpy shock tunnel experiment is made by using one temperature chemically nonequilibrium CFD code. Aerothermodynamic heating distributions on the windward surface of HYFLEX (Hypersonic Flight Experiment) six percent model are compared in detail. Four test conditions in HEK are selected to investigate enthalpy and binary scaling effects on heating characteristics and good agreements are obtained. To extrapolate data from ground to flight, real gas computations are also made for HYFLEX flight conditions at catalytic and non-catalytic wall assumptions. On the other hands, at high altitude and high mach numbers like HOPE-X (H-2 Orbiting Plane Experimental) reentry flight, trim capability of high angle attack is an important design factor. To investigate real gas effects of these aerodynamic characteristics, calculations are made around space shuttle configuration along its trajectory from mach number five to 26. Real gas effects on center of pressure locations are compared well with the shuttle OADB data., 実在気体CFD(数値流体力学)コードを確認して適切な実在気体モデルを構築するために、1温度で化学的に非平衡のCFDコードを使ってHEK(角田宇宙推進技術研究センター高エンタルピー衝撃風洞)実験を解析した。HYFLEX(極超音速飛行実験)6%モデルの風上表面における空力加熱分布を詳細に比較した。HEKについて4テスト条件を選択して、加熱特性に対するエンタルピーと2成分スケーリングの効果を研究し、良い一致を得た。データを地上から飛行へ補外するために、触媒および非触媒壁条件におけるHYFLEX飛行条件について実在気体計算を行った。一方、HOPE-X(宇宙往還技術試験機)再突入飛行のような高高度高マッハ数において、高角度アタックのトリム能力が重要な設計指標となる。これらの空力特性に対する実在気体の効果を研究するために、マッハ数が5から26までのスペースシャトルの軌道に沿った配置周囲で計算を行った。圧力位置の中心に対する実在気体の効果をシャトルのOADBデータと良く対比した。

In the development of the HOPE (H-2 Orbiting Plane) guidance control subsystem, the following technical items are to be studied: Rendezvous-docking phase (RDV), orbital separation and reentry phase, Terminal Area Energy Management phase (TAEM), approaching and landing phase, redundancy techniques, and software. As for RDV, remote navigation RDV and automatic RDV should be combined and its safety and operability should be improved. As for the reentry phase, Lambert rule should be used. After the reentry interface, it would be desirable to use the closed form method, but a sufficient margin should be allowed for the constraints due to the aerodynamic heating rate of HOPE. The flight control system of the TAEM phase should be able to cope with large variations of the aerodynamic parameters in the transonic range. A Microwave Landing guidance System (MLS) should be used at the approaching and landing phase, but the adjustment between the nominal orbit and the flight performance should be very important and arranged to be made. Though the guidance control system had triple redundancy, it would be necessary to study the simplification of redundancy structure, the adequacy of redundancy control rule, and the efficient experimental method. As for software development, the quality assurance and task analysis should be discussed urgently., HOPE(H-2ロケット打上げ型有翼回収機)の誘導制御サブシステム開発には、軌道上・ランデブー(RDV)ドッキングフェーズ、軌道離脱・再突入フェーズ、最終エネルギー調整(TAEM)フェーズ、進入着陸フェーズ、冗長技術、ソフトウェア開発のそれぞれにおける技術課題がある。RDVについては、遠隔操縦RDVと自動RDVを組み合わせ、その安全性と運用性の向上について検討する必要がある。再突入フェーズでは、ランバート誘導則を用いる。再突入インタフェイス以降はクローズドフォーム法を用いるのが良いが、HOPEの場合、空力加熱率による制限が厳しいため、充分なマージンが得られることを検討しなければならない。TAEMフェーズの飛行条件は、主として遷音速の領域で空力パラメタの変化が大きいため、これに適合できるような飛行制御システムを構成する必要がある。進入着陸フェーズではマイクロ波着陸誘導システム(MLS)を使用するが、基準軌道と飛行性能などの調整が重要となり、その検討を行う必要がある。誘導制御システムは3重の冗長構成を基本としているが、冗長構成の簡素化、冗長管理則の妥当性、効率的な実験手法の検討が必要である。ソフトウェア開発については、品質保証の手法や作業分析を早急に検討する必要がある。

Development of the Hypersonic Flight Experiment (HYFLEX) vehicle is described. The HYFLEX vehicle was launched by the first J-1 launch vehicle from Tanegashima Space Center (TNSC) on February 12, 1996. It splashed down in the Pacific Ocean, north east of Bonin Islands, bringing back memorable fruitful results. Here, the progress and the organization of the development, and the feature of the vehicle are described, introducing technological challenges and discussions in the areas of aerodynamics, aerothermodynamics, structures and flight controls which have been of central issues in the development., 極超音速飛行実験機(HYFLEX)の開発について報告する。1996年2月12日に種子島宇宙センターからJ-1ロケットの初号機によって打ち上げられた実験機が日本で初の極超音速揚力面飛行を実施した。小笠原諸島の北東で太平洋に着水した飛行は意義深く成果の多い飛行となった。ここでは開発の経過、開発体制、開発された機体の概要を述べると共に、実験機開発の重点であった空力/空力加熱、構造、飛行制御の分野での技術的な挑戦、議論を重ねた点を紹介する。

With the aim of testing the aerodynamic force and heating in reentry phase of a space transportation vehicle, NAL and NASDA are jointly constructing a large-sized free-position shock tunnel, HIEST (High Enthalpy Shock Tunnel). The concept of the design of this shock tunnel is to achieve efficient energy transport from high-pressure air generated by piston drive to test air. For realizing this shock tunnel, using relatively light piston was proposed and the concept to make piston speed faster when breaking a main diaphragm was developed. In this report, the design concept and the features of the mechanical design as a result of realizing the concept are described., 宇宙輸送機の再突入時の空気力および空力加熱の試験を当面の目的として、航空宇宙技術研究所と宇宙開発事業団は共同で大型の自由ピストン衝撃風洞HIEST(大型高温衝撃風洞)を建設中である。この衝撃風洞の設計思想はピストン駆動高圧空気から試験気体へのエネルギー輸送を効率的に行うことである。これを実現するために、比較的軽いピストンを用い、かつ、主隔膜の破膜時におけるピストン速度を高くする概念を開発した。ここでは、この設計概念とその結果としての機械設計の特徴を述べる。

Numerical study of hypersonic flow around HOPE is performed, using flux-split and flux difference splitting upwind TVD (Total Variation Diminishing) Navier-Stokes codes. New parametric computations are made to determine the basic configuration from two type of HOPE candidate geometry designs. One type is the Double Delta type model (DD model) and the other is the Power Delta type one (PD model). Also, in order to investigate real gas effects, numerical calculations, corresponding to the flow conditions of high enthalpy wind tunnels of DLR, HEG (High Enthalpy Shock Tunnel Goettingen) and Caltech (California Institute of Technology) T-5, are made. In addition, hypersonic perfect gas flow calculations are made for the comparison with experimental data of AEDC (Arnold Engineering Development Center) and ONERA S4MA hypersonic wind tunnel and Calspan's shock Tunnel. Through the present numerical analysis, detailed aerodynamic and aerothermodynamic characteristics of HOPE are investigated. These works have been done as the joint research of NAL and NASDA., HOPE(H-2ロケット打ち上げ型有翼宇宙往環機)周りの極超音速流を流速分割および流速差分分割風上TVD Navier-Stokesコードを用いて数値解析した。2タイプの候補形状設計から基本形態を決めるために新しいパラメータ計算を行った。1つはダブルデルタタイプ(DDモデル)で、他の1つはパワーデルタタイプ(PDモデル)である。また、実在気体効果の調査研究のためDLR,HEG(Goettingenの高エンタルピー衝撃風洞)およびCaltech T-5の高エンタルピー風洞の流れ条件に対応した数値計算を行った。さらに、AEDCおよびONERA S4MAの極超音速風洞とCalspanの衝撃風洞の各試験との比較対応のための数値計算も行った。これらの数値解析を通じて、HOPEの空力・空力加熱特性の詳細を究明できた。本研究はNALとNASDAの共同で実施された。

Numerical Simulation for aerodynamic study of atmospheric space transport vehicles has been conducted at NAL (National Aerospace Laboratory) as the cooperative research work with NASDA (National Space Development Agency). CFD (Computational Fluid Dynamics) technology enlarges its application area with the progress of the computer hardware systems and now CFD has been developed as the strong aerodynamic design tool which covers flight range of space transport vehicle from the launching to the re-entry phase. Here, recent progress of numerical simulations at NAL are introduced., JAXA Special Publication, 宇宙航空研究開発機構特別資料

The following topics were discussed: spaceplane development plan, H-2A rocket, HOPE-X (H-2 Orbiting Plane Experiment), ALFLEX (Automatic Landing Flight Experiment), J-1 rocket, and status of HYFLEX (Hypersonic Flight Experiment) development such as trajectory design, navigation system, flight control, electrical components, aerothermodynamic characteristics, aerodynamics, measurement program, aerodynamic heating, ADS (Air Data Sensor) measurement, gas jet interaction, and thermal protection system., 以下の主題を論じた。宇宙往還機開発計画、H-2Aロケット、HOPE-X(宇宙往還技術試験機)、ALFLEX(小型自動着陸実験)、J-1ロケット、飛行経路設計、航法系、姿勢制御、電気系、熱特性、空力特性、計測計画、空力加熱計測、ADS(大気データセンサ)計測、ガスジェット干渉、および熱防護系などのHYFLEX(極超音速飛行実験)の開発の現況。

In the final phase of the MUSES-C mission, a small capsule with asteroid sample conducts reentry flight directly from the interplanetary transfer orbit at the velocity over 12 km/s. The severe heat flux, the complicated functional requirements, and small weight budget impose several engineering challenges on the designing of the thermal protection system of the capsule. The heat shield is required to function not only as ablator but also as a structural component. The cloth-layered carbon-phenolic ablator, which has higher allowable stress, is developed in newly devised fabric method for avoiding delamination due to the high aerodynamic heating. The ablation analysis code, which takes into account of the effect of pyrolysis gas on the surface recession rate, has been developed and verified in the arc-heating tests in the facility environment of broad range of enthalpy level. The capsule was designed to be ventilated during the reentry flight up to about atmospheric pressure by the time of parachute deployment by being sealed with porous flow-restrict material. The designing of the thermal protection system, the hardware specifications, and the ground-based test programs of MUSES-C capsule are summarized and discussed here in this paper., The Institute of Space and Astronautical Science Report SP

In the perspective of Japanese space projects, the knowledge of high enthalpy and high Mach number flow, occurring during the atmospheric reentry of space vehicles, has regained recently considerable interest. The need to support aerodynamic design of space vehicles under the environments where experiment is very difficult and costly, has driven the use of numerical simulation. Hence, high enthalpy flow workshop was held at NAL in June, 1995 for axisymmetric flow problems. This paper summarizes the computational results of OREX (Orbital Reentry Experiment) real gas flow problem and discusses the detailed comparison of each numerical results., 日本の宇宙プロジェクトを展望すると、宇宙飛行体の大気再突入の間に発生する高エンタルピー、高マッハ数流れに関する知識が最近かなりの関心を呼んできている。実験が極めて困難で費用の掛かる宇宙環境下の宇宙飛行体の空力設計を支援する必要性から、数値シミュレーションの利用が促された。このような状況から、軸対称流問題についての高エンタルピー流れワークショップを1995年6月にNALで開催した。OREX(軌道再突入実験機)実在気体流れ問題の計算結果を要約し、各数値結果の詳細な比較を論じた。

A brief review of the current capabilities and outstanding issues for the efficient analysis, and design of high-speed flight vehicles using unstructured grids is given. It is concluded that all major areas required in the analysis cycle - grid generation, flow solvers and visualization - have seen major advances in recent years. This trend is expected to continue, leading us to believe that turnaround in a matter of hours or minutes for complex geometries and/or physics will become a reality sometime in the next decade., JAXA Special Publication, 宇宙航空研究開発機構特別資料

The forebody flow field of the Orbital Reentry Experiment (OREX) reentry vehicle at the flight condition of the maximum wall heating is studied by using the axisymmetric viscous shock-layer equations with the seven-air-species nonequilibrium chemistry. At this condition, the shock layer is in severe chemical nonequilibrium and the wall heating rate is significantly augmented due to the wall catalycity. The effects of the wall catalycity on the wall heating rate strongly depend on the extent of the chemical nonequilibrium in the shock layer. Consequently, the chemical nonequilibrium analysis is necessary for the prediction of the aerothermodynamic environment of the OREX reentry vehicle. The results show that the viscous shock-layer analysis has sufficient potential to be a useful tool not only for the estimation of the wall heating rate but also for the investigations on the physics of the real gas effects in the shock layer., OREX再突入物体の壁面加熱が最大となるような飛行条件における前面側の流れ場を、7種の化学種の間の非平衡化学反応を伴う軸対称衝撃層方程式によって調べた。この条件では、衝撃層は化学的に甚しく非平衡であり、壁面加熱率は壁面の触媒性によって著しく増大する。壁面触媒性が壁面加熱に与える影響は、衝撃層内における化学的非平衡の程度に強く依存する。したがって、OREX再突入物体の熱空気力学的環境を予測するためには化学的非平衡の解析が必要である。結果は、粘性衝撃層の解析が壁面加熱率の推定のみならず、衝撃層内における実在気体効果の物理に関する研究に役に立つ可能性が十分にあることを示す。

The mechanism of early transition phenomena over an ablating reentry capsule has been examined. A two-equation turbulence model (kappa-epsilon model) coupled with Reynolds averaged Navier-Stokes equations has been employed. Low-Reynolds-number effects on the solid wall were considered by modifying the Chien's correction. As a result, the transition occurred at lower Reynolds numbers with higher ablation rates. The predicted transition-point Reynolds number was 3 x 10(exp 4) at the surface-mass-injection rate of 100 g/sm(sup 2). The principal mechanism of this early transition is thought as follows; the viscosity damping effects are reduced and re-laminarization is prevented in the downstream of the capsule surface, due to the turbulence on the surface and due to the pushing out of near-surface stream-lines from the surface by successive mass injection., The Institute of Space and Astronautical Science Report SP

As a part of the HOPE-X (H-2 Orbiting Plane-Experimental) development, tests were carried out in order to evaluate quantitatively the catalytic activity of the surface of C/C (Carbon/Carbon composite material) thermal protection materials using a plasma wind channel. HF (High Frequency) Plasmatron U-13 at TsNIIMASH (Central Scientific Research Institute of Machine Building) in Russia was used as the plasma wind channel. The temperature of the test objects (thermocouples for the rear side and radiation thermometers for the front side), the gas flow rate, the pressure and the discharge power were obtained at the plasma heating tests under various conditions on different test objects. The important parameters such as the catalytic recombination rate constant, and the ratio of the heating rate of the test object to that of the Cu calorimeter which is the calibration catalyst of a unit, q/qFC were estimated. The value of q/qFC was in the range from 0.36 to 0.46 in both atmospheres of air and nitrogen. The type of the coating, which was used in OREX (Orbital Re-entry Experiment), has thus demonstrated a great effect on suppression of the heating rate against the reference catalyst. Values of the recombination rate constant were determined to be 0.7 to 1.8 m/s, 0.6 to 2.2 m/s, and 0.3 to 2.0 m/s for the air, nitrogen, and oxygen respectively. These low values demonstrate that the test objects have a sufficiently low catalytic activity. In conclusion, the OREX type coating is the material of low catalytic activity equivalent to the SiO2-B2O3 glass coating, which was evaluated as having low catalysts activity in USA and Russia., 本試験では、HOPE-X(宇宙往還技術試験機)の設計のための検討の一部として、C/C(カーボン/カーボン複合材)熱防護材の表面触媒性を定量的に評価するため、プラズマ風洞を用いて試験を実施した。プラズマ風洞には、ロシアTsNIIMASH(機械製作中央研究所)所有のHF(高周波)プラズマトロンU-13を使用した。各種条件で各種供試体のプラズマ加熱試験を実施し、供試体の温度(裏面は熱電対、表面は放射温度計で測定)、ガス流量、圧力、放電出力などを計測し、触媒再結合速度定数、供試体の加熱率と、校正用完全触媒材料であるCuの加熱率との比率q/qFCの見積もりを行った。q/qFCの値は、空気中と窒素中とでほとんど差がなく0.36〜0.46の範囲にあった。この結果、OREX(軌道再突入実験機)タイプコーティングは、完全触媒材料に対する加熱率低減効果が大きいことを示した。再結合の速度定数は、空気で0.7〜1.8m/s、窒素で0.6〜2.2m/s、および酸素で0.3〜2.0m/sとなって、いずれも小さい値を示し、供試体材料は低触媒性材料であることが示された。以上の結果、OREX用コーティングが、アメリカやロシアで評価されたSi2O3-B2O3系ガラス質コーティングと同等の低触媒性材料であることが確認できた。

To obtain the aerodynamic heating data of HOPE-X (H-2 Orbiting Plane Experimental), the design of which is now in progress, experiments placed by NASDA were performed at AEDC (Arnold Engineering Development Center) VKF (Von Karman gas dynamics Facility) tunnel B, C in 1996 and 1997. A comparison test has been performed at the NAL hypersonic wind tunnel and the results are reported here. The model is a 3.6 percent HOPE (H-2 Orbiting Plane) of the reference length 576 mm. In the experiment, the aerodynamic heating is measured by about 400 heating rate sensors embedded on the surface of the model. Although the data are not completely corresponding to the AEDC case, 14 runs of 43 cases of experiments have been performed., 現在設計が進められている有翼宇宙往還機HOPE-Xの空力加熱データ取得の目的でNASDA発注により1996年および1997年にAEDC(アーノルド工学開発センター)VKF(フォン・カルマン空気力学施設)風洞B、Cにおいて試験が実施されていた。今回これらとの対応風試を航技研極超音速風洞において行ったのでそれらの結果の比較について報告する。模型は基準長576mmの3.6%HOPE(H-2ロケット打ち上げ型有翼回収機)であって、今回の風試では表面上のおよそ400個所に埋め込まれた加熱率センサの一部を用いて空力加熱を計測した。AEDCにおいて行われたケースに完全に対応したデータは必ずしも得られていないが、14Run、43ケースの試験を行った。

An integrated numerical method to obtain trajectory-based aerodynamic heating environment for entry probe vehicles with ablation is briefly described. In this method, a thermochemical nonequilibrium CFD code is loosely coupled with an ablation module along the entry trajectory. Radiative heating is estimated by the use of multi-band radiation module. The effect of earlier turbulent transition of the boundary layer due to ablation product gas is included. Calculated results for MUSES-C sample return capsule are shown and are compared with the results of existing analysis., The Institute of Space and Astronautical Science Report SP

The re-entry capsule USERS/REV was successfully recovered in May 2003. REV has a heat shield system to protect inner equipments against the severe heating environment during the re-entry. The heat shield system is mainly made of three kinds of layers, i.e., CFRP ablator, felt type insulation material with thermal anchor and an aluminum shell structure. In the design of the heat shield system, contradictory requirements of heat ejection in orbit and heat protection during the re-entry had to be satisfied. Thermal protection performance of the heat shield system was calculated using an ablation analysis code. Thermal stress analysis of the ablator during the reentry was carried out using a large deformation FEM that considers materials non-linearity. For the development of the heat shield system, various kinds of tests such as the arc-heated tests of the ablator and the strain survey test of the heat shield system were carried out. After the recovery of REV, evaluation of the recovered heat shield system was carried out. The temperature time histories in the ablator obtained during the re-entry and the damage of the heat shield system measured after the recovery were compared with the results of ablation analysis and structural analysis, from which the heating environment during the re-entry and the cause of the damage were estimated. The evaluation of the heat shield system design based on the flight results is also described., JAXA Research and Development Report, 宇宙航空研究開発機構研究開発報告

Thermal protection materials for HOPE (H-2 Orbiting Plane) must be developed by taking account of mechanical and heat resistance properties, light weight, acoustic resistance and chemical resistant properties as well as heat insulating properties against aerodynamic heating during atmospheric reentry. Titanium-base thermal protection materials were to be used in the sections from 300 C to 500 C. For these materials, manufacturing methods such as shaping and binding technologies were studied. By the performance evaluation of multi-wall shaped trial pieces, no problems of heat resistance, heat insulation and mechanical properties were found. Nickel-base thermal protection materials were to be used in the sections from 550 C to 1,100 C. On these materials, manufacturing problems such as bonding technologies and internal adiabatic materials technologies were studied. Trial pieces were evaluated for the density and adiabatic property, and the required performance level was found to be fulfilled. C/C thermal protection materials were to be used in the sections from 550 C to 1,300 C, and were to compose an ACC (Advanced C/C material) Thermal Protection System (TPS). Heat insulating properties of parts bonded by fasteners were studied. Ceramic thermal protection materials were to be used in the sections with temperatures below 1,300 C and in the antenna part. Properties of handling and environmental resistance were discussed., HOPE(H-2ロケット打ち上げ型有翼回収機)の熱防護材は再突入時の空力加熱からの断熱特性の他に、力学特性、耐熱特性、軽量化、耐音響性、および耐化学的特性も考慮に入れて開発されねばならない。チタン合金熱防護材は300度Cから500度Cまでの温度の部位に使用される材料である。この材料について、成型加工技術と接合技術という製造法上の課題を検討した。マルチウォール形状の試作品の性能評価の結果、耐熱性、断熱性および力学特性について問題のないことを確認した。ニッケル合金熱防護材は550度Cから1,100度Cまでの部位に使用される材料である。これについて内部断熱材料の選定、接合技術という製造上の課題を検討した。試作品について密度と断熱性の評価を行い、要求性能を満たすことを確認した。C/C熱防護材は550度Cから1,300度Cまでの部位に使用され、ACC(高性能C/C複合材料)熱防護システム(TPS)を構成するが、そのファスナ結合部の断熱構造を検討した。セラミック熱防護材は1,300度C以下の部位とアンテナ部位に使用される材料であるが、これについて取扱い性と耐環境性を検討した。

In HYFLEX (Hypersonic Flight Experiment) flight, aerodynamic heating to the region of C/C (Carbon/Carbon composite) hot structure and TPS (Thermal Protection System) ceramic tile are measured with some kind of temperature sensors. In this paper, aerodynamic heating measured in windward side will be noticed. In the region near symmetric plane in windward side, abrupt increase of heating rate caused by boundary layer transition are measured, and before the transition, good agreement between measured aerodynamic heating in HYFLEX flight and the value calculated by a simple engineering estimation using results of wind tunnel tests., HYFLEX(極超音速飛行実験)ではC/C(カーボン/カーボン複合材)ノーズキャップ、エレボン部およびタイル部に温度センサを配し、空力加熱計測を行った。ここでは、胴体下面タイル部で計測された加熱率について報告する。胴体下面中心線付近で境界層遷移によると考えられる加熱率の急激な増加が確認された。また境界層遷移以前では、測定された空力加熱は、設計用に取得した風洞試験結果と球淀み点加熱率の推算式とを用いた工学的簡易推算値と比較的良い一致をする事が確認できた。

To predict the characteristics of the ablation thermal protection for the MUSEE-C superorbital re-entry capsule, the trajectory-based analysis of the hypersonic shock layer flow and the ablator interior has been made on the stagnation line in a coupled manner. For the flow analysis, the Viscous Shock-Layer (VSL) equations with the nonequilibrium chemistry of C-O-H-N 26 species are solved. For the ablator, the in-depth thermal response is solved by the Charring Material Ablation (CMA) model. The boundary condition at the surface, which is necessary for both solutions, is determined by the mass and energy balance. Various surface reactions and the surface injection of the pyrolysis gas are considered. The results are compared with the experimental data obtained at the arc-heated wind tunnel. Analysis by using the empirical relations all along the re-entry trajectory is also made and the results are compared with those by the VSL/CMA coupled method. It is shown that such engineering prediction method provides reasonable prediction for the surface temperature and the mass loss of the ablator with sufficient safety margin. The effects of the initial ablator thickness on the ablator temperature are also discussed., The Institute of Space and Astronautical Science Report SP

JAXA Special Publication, 宇宙航空研究開発機構特別資料

The flight range of HOPE (H-2 Orbiting Plane) would be extended to the high altitude of several hundreds km and its speed to the Mach number of 30. In this regard, the aerodynamic characteristics must be estimated for various flight conditions such as low speed during climbs, high speed at low angles of attack, high angles of attack during re-entries, and low speed at landing approach. The aerodynamic characteristics of HOPE were to be estimated in three flight ranges, segmented by different flight Mach number ranges. Items to be estimated were mainly lift to drag ratios, static and dynamic stability in longitudinal and lateral directions, some aerodynamic characteristics particular to the HOPE configuration, jet reaction control systems, boundary layer aerodynamic heating, vortex and flow separation, shock wave interactions, and real gas effects. In respect to numerical simulations, the aerodynamic characteristics became independent of the Mach number in flight as the Mach number increases. A Newtonian flow approximation, however, still held, so that hypersonic aerodynamic characteristics could be dealt with good accuracy. Tools for estimation must be verified by numerical simulations, wind tunnel tests, and flight tests., HOPE(H-2打ち上げ型有翼回収機)の飛行領域は、数百kmの高い高度およびマッハ数30の大きな速度領域に亘っている。このため、上昇時の低速、小迎角の高速、帰還時の高速大迎角、着陸進入時の低速大迎角などの空力特性の推定が必要である。このようなHOPEの空力特性推定は、飛行マッハ数領域を3分割して行われ、また推定項目は、揚抗比、縦横の静安定と動安定、HOPE機体形状による空力特性、ジェット反動コントロール、境界層空力加熱、渦および剥離、衝撃波干渉、および実在気体効果などを含んでいる。数値シミュレーションでは、マッハ数が高くなると空力特性はマッハ数に依存しなくなるが、ニュートン流近似が成立するので、極超音速空力特性を良い精度で取扱うことができた。また、推定ツールは、数値シミュレーション、風洞試験および飛行試験による実証が必要であった。

One of the most important issues in the design of space planes is aerodynamic heating. However, its mechanism is not yet clarified because the pressure and heat flux change in spatially small scales. Measuring techniques of heat fluxes include the following: (1) thin film heat transfer gauge using the non-stationary surface temperature increase of thermal insulators with half-infinite depth, (2) calorimetry using the temperature rise of thermal conductors on a material surface, (3) method using the stationary temperature gradient inside a material, (4) thermal paint method using color changes or phase changes of temperature sensitive paints on a material surface, and (5) measurement of the surface temperature with an infrared thermometer. Methods (2) to (5) have difficulties in spatial resolution and high-speed response, and method (1) is not suitable for measurements with a long flow time. The newly developed measurement technique uses two thin film gauges which measure the temperature gradient between the two films induced by the heat flux entering across a thin heat-resistant film. It has a high spatial resolution and high-speed response. It can be used in a wind tunnel with a long flow time. Among studies concerning aerodynamic heating are studies on the interference between a shock wave and a boundary layer, the interference field induced by sharp fins, turbulent boundary layers, laminar boundary layers, and the interference between two shock waves. There are many forms of aerodynamic heating phenomena. Further improvement in measurement resolution and technique at high temperature and high pressure is to be necessary in the future., 宇宙機の設計において最も重要な問題の1つは、空力加熱であるが、空間的に小さなスケールで圧力や熱流束が変化するため、その詳細なメカニズムは解明されていない。熱流束の計測技術には、(1)半無限深さの熱絶縁性物体の非定常表面温度上昇を利用した薄膜熱伝達ゲージ、(2)物体表面の熱良導体の温度上昇を利用したカロリメータ法、(3)物体内部の定常温度勾配を利用する方法、(4)物体表面に塗布した感温塗料の色彩変化や相変化を利用した熱ペイント法、(5)赤外線温度計で物体の温度上昇を測定して熱流速を求める方法などがある。(2)から(5)の方法は空間分解性と高速応答性に難点があり、(1)の方法は気流持続時間の長い測定には不向きである。新しく開発した熱流速測定法は、薄い熱抵抗膜を横切って流入する熱流速による膜間の温度勾配を2枚の薄膜ゲージで測るものであり、高い空間分解性と高速応答性を備え、気流持続時間の長い風洞にも用いることができる。空力加熱の研究には、衝撃波・境界層干渉、鋭いフィンにより誘起された干渉場、乱流境界層、層流境界層、衝撃波・衝撃波干渉などさまざまなものがある。空力加熱現象には多くの形体が存在するので、精度の改善や、高温高圧下での計測技術が今後必要になる。

Arc-heated wind tunnels are essential ground test facilities for the development and evaluation of Thermal Protection System (TPS) materials which protect the body of space vehicles from severe reentry heating environments. Under an NAL (National Aerospace Laboratory of Japan)/NASDA joint space program, a 450 kW archeated wind tunnel was built and operated to test the TPS materials for a planned Japanese space vehicle, HOPE. In 1993, to improve the operational characteristics of the tunnel, such as aerodynamic flow characteristics and heating rates, a complete reconstruction of the tunnel was done. This included the introduction of a multisegment arc heater, improvements in the evacuating system capacity and so on. As a result, the operating range extended up to 750 kW and a heating rate of 2.2 MW/sq m was obtained for a 50 mm diameter flat faced cylinder model. Due to the inclusion of a channel nozzle, the testing ability of a flat panel model of 100 mm x 100 mm size was added. This paper describes the structure and function of the 750 kW wind tunnel and its components. Operational characteristics of the arc-heater and low density free stream are also delineated., アーク加熱風洞は、宇宙往還機の機体の再突入時の空力加熱から保護する防熱系(Thermal Protection System: TPS)材料の開発、評価に必須の地上試験設備である。日本においても、国産宇宙往還機の開発を目標としたNAL(航空宇宙技術研究所)/NASDAの共同研究のもとで450kWのアーク加熱風洞を1981-1982年に建設、運用して各種の耐熱材料の加熱試験を実施してきた。この風洞の気流特性の改良、高加熱率での試験能力の安定化などの試験能力向上を目的として、1993年にアーク加熱器を大出力化のために交換、ディフューザと排気系の性能向上などを中心とした大改造を実施した。その結果、最大入力750kWでの運用が可能となり、外径50mmの平頭模型で最大加熱率2.2MW/平方メートルで安定した澱点加熱試験が可能となった。また、チャネルノズルの装備により100mm x 100mmの平板供試体の試験も実施できるようになった。本論文では、この750kWアーク加熱風洞の全体構造、各構成要素の性質について詳述する。また、加熱器の放電特性と気流の圧力、加熱率などの運用動作特性を述べる。, Technical Memorandum of National Aerospace Laboratory, 航空宇宙技術研究所資料