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In the late 1950s, a "New Geography" emerged among geography academics that marked not only a change in how geography was studied at universities, but also had a major impact on how geography was taught in schools. In England, geography education was fundamentally changed to adopt the New Geography approach. Its origins are widely recognized as having been the first Madingley Conference, held in 1963. This paper aims to reveal how that conference led to the changes that occurred in the 1960s, with a particular focus on schoolteachers. The conference papers were published as Frontiers in Geographical Teaching: The Madingley Lectures for 1963, in 1965. Some young teachers were keen to build on the innovative ideas in these lectures, and in 1966 they formed a voluntary research group focused on practical applications. This group was remarkably active; some of its results were published as a series of New Geography educational books. These activities led to another type of seminar, the Charney Manor Conference, in 1970, and the outcomes of this conference were published as New Directions in Geography Teaching: Papers from the 1970 Charney Manor Conference, in 1973. This book is quite different from Frontiers in Geographical Teaching. This means that while they were adopting the New Geography paradigm, schoolteachers transfonned its educational content through various activities, and this content ultimately became the New Geography Education.

資料番号: PA1410055000

In a rocket turbopump, an axial thrust balancing system with the balance piston is often applied to reduce the axial thrust. The leakage flow through the balance piston is affected by the configuration of the balance piston and the balancing point in the axial direction. In designing the internal flow system, the balancing point and flow rate have been estimated by the static analysis code of the internal flow. On the other hand, the analysis code for dynamic response of the internal flow through the balance piston was developed. In this study, optimal calculation considering the axial dynamic characteristics of the internal flow system was carried out by using the dynamic analysis code. As the results, the balancing point in the steady state significantly affects the leakage flow rate and the axial vibration., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA0065215000, レポート番号: JAXA-RM-11-017

資料番号: PA1110052000

Research and development to increase reliability of rocket engines, such as LE-7A and LE-5B, were conducted in the present study. Amplitude of turbopump rotor vibration due to rotating cavitation was suppressed by modification of inducer design and optimization of inducer liner geometry. Cryogenic inducer test facility to investigate various aspects of inducer performance in cryogenic fluids was newly constructed. Tests of a cryogenic inducer have been conducted. Excessive lateral thrust of LE-7A engine during start and stop transients was investigated to find a method to suppress it. Low cycle combustion pressure oscillation encountered in the LE-5B engine was also investigated to suppress the oscillation., 資料番号: AA0047944011, レポート番号: JAXA-RM-04-009

資料番号: AA0033392041

航空宇宙技術研究所 16-18 Jun. 1999 東京 日本, National Aerospace Laboratory 16-18 Jun. 1999 Tokyo Japan, この実験的および数値的研究の主な目的は、RJTF(ラムジェット試験設備)を用いて、マッハ数4、6、8の条件下でのエンジン燃焼実験で観測された弱燃焼と強燃焼と呼ばれる2つの代表的な燃焼モード間の遷移過程を明らかにすることである。この目的のために、コールドフロー実験とエンジンノズル内のフロープラグを用いた燃焼圧の数値シミュレーションを行い、燃焼器内の流れの構造と燃焼モード間遷移の際に起こる構造の変化を研究する。これらの結果および燃焼実験の結果から、以下のことが明らかになった。(1)主流中での自発着火は不可能である。すなわち、着火は境界層とストラットの後流の狭い領域内に限られる。(2)弱燃焼は、燃焼器下流部の付着境界層でしかも少し剥離気味の境界層において起こり、燃焼は反応律速下で行われる。(3)強燃焼は、主に、大規模に剥離した境界層および再循環領域で発生する。(4)弱燃焼から強燃焼への遷移は、燃料/空気の混合の促進と境界層内燃焼の相互作用の結果であり、燃焼により境界層剥離域領域は成長する。このような燃焼特性は、現在の燃料/空気混合エンジンでは推力を支える主燃焼が、大規模な境界層剥離と剥離領域内部での亜音速燃焼で見られような乱流混合に強く依存する。混合や燃焼の境界層剥離への依存性は、遷移の初期においてインレット不始動を生ずる。エンジン特性の改善には、燃焼器内の超音速主流内での効率的な超音速混合・着火・燃焼制御がまさに重要である。, Main purpose of this experimental and numerical study is to reveal a transition mechanism between two typical combustion modes called 'weak combustion' and 'intensive combustion' which have been observed in the engine firing tests at Mach four, six, and eight conditions using RJTF (Ramjet Engine Test Facility). For this purpose, flow structures in the combustor and their changes during the transition of combustion mode are investigated by means of cold flow tests as well as numerical simulations in which combustion pressure is simulated using flow plug set in the engine nozzle. Based on these results as well as the firing test results, it is revealed: (1) that self ignition in the core flow is impossible; ignition is possible only within the narrow regions of the boundary layer and the strut wake; (2) that combustion in the weak mode occurs within the attached and slightly separated boundary layer at the downstream part of the combustor and the combustion is kinetically controlled; (3) that combustion in the intensive mode mainly occurs within the largely separated boundary layer and recirculation area; and (4) that transition from 'weak mode' to 'intensive mode' is the consequence of mutual interaction between enhancement of fuel/air mixing and combustion within the boundary layer and the growth of the boundary layer separation area due to the combustion. Such combustion characteristics are due to the fact that in the present engine fuel/air mixing and principal combustion supporting the engine thrust strongly depend on the occurrence of large scale boundary layer separation and on the turbulent mixing as well as subsonic combustion within the separated boundary layer. The dependence of mixing and combustion on the boundary layer separation results in the early transition into inlet unstart. For improvement of engine performance, it is indeed important to carry out efficient supersonic mixing, ignition and combustion control in supersonic core flow in the combustor., 資料番号: AA0001961009, レポート番号: NAL SP-44

20th International Symposium on Space Technology and Science (May 19-25, 1996.), Gifu, Japan, 資料番号: AC1000109000

吐出し圧力が30MPaを越えるLE-7液酸ターボポンプの軸推力調整にセルフバランス方式のバランスピストンを採用した。研究試作機においてバランス室の圧力分布が大幅に予測からずれ、バランス室入口オリフィスの軸方向隙間が過小となった。また、タービンをパーシャルアドミッションからフルアドミッションに変えたことによりタービン軸推力が大きく変った。このため、現象の解明とターボポンプの構造を大きく変更すること無く入口オリフィスの軸方向隙間を十分に確保する方法の確立を試みた。試みた主な方法として2つあげられる。第1の方法は、バランス室から主羽根車入口に直接通じるバランスホールを設けたことである。第2の方法は、主羽根車フロントシュラウドに相対するケーシング壁面に放射状の溝を設けたことである。これまでに十分に存在しない、バランスホール流量係数、ケーシング壁面に設けた溝の圧力分布に及ぼす効果などの基礎的データをまず取得した。このデータに基づいて2つの方法を用いた結果、入口オリフィスの軸方向隙間を十分に確保することができ信頼性の高い原型LE-7液酸ターボポンプの軸推力調整機構を実現した。, A self-balancing type axial-thrust balancing system referred to as 'balance piston' was applied to the LE-7 LOX turbopump, the discharge pressure of which exceeds 30 MPa. A research model of the turbopump showed that pressure distribution in the balancing chamber deviated greatly from the predicted distribution and that the axial clearance of the balancing chamber inlet orifice became too small to maintain safety. Therefore, two methods were employed to ensure sufficient axial clearance of the inlet orifice. In the first method, balancing holes were made directly connecting the balancing chamber and the main impeller inlet. The diameter of the balancing holes was determined by flow coefficient test results of rotating holes. In the second method, grooves were made on the casing wall facing the main impeller front shroud. As a result, proper axial clearance was confirmed, making the development of a reliable axial-thrust balancing system for the LE-7 LOX turbopump prototype possible., 資料番号: AA0004216000, レポート番号: NAL TR-1201

インデューサケーシング入口形状の影響が、3枚翼であるインデューサに発生する非定常なキャビテーションについて調べた。インデューサ入ロ流れ場および生成されたキャビティの構造は、2形状の異なるインデューサケーシングについてPIV(粒子画像流速測定)システムと高速度ビデオを用いて可視化観察を行った。本試験では、非定常キャビテーションの1つである旋回キャビテーションが発生し易い入口形状のケーシングと入口形状径を広げ旋回キャビテーションが抑制されるケーシングを用いて可視化計測した。インデューサ翼端部の漏れ流れ渦とインデューサ上流に発達する逆流の構造は、インデューサケーシング入口形状により旋回キャビテーションの発生に影響していると思われる。, The effects of inlet casing geometry configuration on unsteady cavitation were examined with a 3-bladed inducer. The inlet flow and cavitation pattern in two different casings were visualized using PIV (Particle Image Velocimetry) and high-speed video. It was found that the rotating cavitation was suppressed when the tip clearance at the inducer inlet was increased. The vortex cavitation in tip leakage flow and cavitation in back flow are thought to affect the appearance of the rotating cavitation as a result of the inlet casing geometry configuration., 資料番号: AA0049488000, レポート番号: JAXA-RR-05-046

The Cryogenic Inducer Test Facility (CITF) was constructed in the JAXA Kakuda Space Center and has been in operation since 2003. Liquid nitrogen and liquid hydrogen are used as working fluids for experiments on rocket engine turbopump inducers, enabling the thermodynamic effects on cavitation to be examined. The performance of CITF was verified from the elemental tests and inducer running tests. In this article, the verification test results and future programs are presented., 資料番号: AA0049203000, レポート番号: JAXA-RR-05-042

ラムジェットエンジン試験設備(RJTF)におけるマッハ8条件での始動性および燃焼性能を改善するために、ストラットに加えてランプを取り付けたスクラムジェットエンジン模型の試験を行った。燃料を垂直噴射した時に、エンジン出口での燃焼効率は約90％であった。当量比1.3の条件時に、590Nの推力増加を測定した。しかしながら最適化されていないエンジン形状であったために、十分な推力増加は得られなかった。着火および燃焼に必要な高い温度および圧力が、ランプの装着によって達成された。燃焼器内圧が、エンジンに流入する気流圧力の160倍に達した場合にも、インレットは始動状態を保つことができた。これはランプを用いてインレット内圧力を上昇させた効果による。, To improve combustion efficiency and the inlet start condition, a scramjet model with a ramp combined with a single strut was tested under Mach 8 flight conditions at the Ramjet Engine Test Facility of the Ramjet Propulsion Research Center. The attached ramp shielded some of the fuel injectors. The fuel flow rate from the open injectors to the flow path was designated as the effective fuel flow rate. In the tests, a combustion efficiency of 90 percent was attained with vertical injection of hydrogen fuel. The thrust increase was 590 N at the effective equivalence ratio of 1.3. However, because the engine geometry was not optimized, a sufficient increase in thrust was not attained. The high temperature and high pressure necessary for ignition and combustion were achieved by the ramp. When the pressure in the isolator was 160 times that of the free-stream air due to combustion, the inlet was in the started condition despite the high pressure. This improved started condition was attained by using the ramp to increase the pressure in the inlet., 資料番号: AA0047391000, レポート番号: NAL TR-1471T

航空宇宙技術研究所・角田推進研究所ではラムジェットエンジン試験設備(RJTF)を完成し、水素を燃料とするスクラムジェットエンジンやメタン燃料ラムジェット(HYPR用)エンジンを試験している。このRJTFのエンジン試験を支援するためにRJTFの1/5縮尺の小型風洞を整備し、ラムジェット推進研究部内に横断的な小型風洞実験班を組織した。小型風洞実験班の目的は2つある。第1は、RJTF試験に関わる問題を解決するためのパイロット試験を実施する事である。第2の目的は、RJTFエンジン試験の補完実験を行う事である。本報告書では、この小型風洞を用いた全ての研究とノウハウについてまとめた。最初に小型風洞設備の構造について説明し、次に自由噴流型風洞の始動、風洞およびエンジン不始動時の荷重、風洞ノズル検定について述べる。そしてこれまでのエンジン試験を補完する実験、スクラムジェットエンジンにおける空気捕獲率の測定、オイルフロー実験をまとめた。また抗力測定と壁圧分布測定を実施した。これらのデータからエンジン内部流解析が可能となり、エンジン理論燃焼性能を予測した。また、燃焼によるエンジン不始動移行過程を調べる模擬実験や燃料拡散を調べるためのHe噴射実験、亜音速燃焼ラムジェット用ランプ圧縮型インレット性能も試験した。小型風洞ではエンジン全体の空力測定に加えて超音速燃焼器の要素実験を行い、燃料ジェットと主流との干渉や燃料拡散を調べた。またレーザによるPIVや感圧塗料実験も実施した。計測に関してはRJTFエンジン試験を支援するための6分力や圧力測定における応答性や、ガス採取プローブでの反応凍結を燃焼模擬空気を用いた超音速風洞で検討した。付録にはノズル設計法と形状データ、小型風洞用エジェクタと消音器の設計、第2(K)風洞の整備、蓄熱体加熱器の蓄熱特性、そして最後にRJTF設備とそのエンジン試験の中で蓄積したノウハウについてまとめた。, The National Aerospace Laboratory, Kakuda Propulsion Laboratory completed the RamJet engine Test Facility (RJTF) and has been carrying out tests using H2-fueled scramjet engines and CH4-fueled ramjet engines since 1994. A 1/5-subscale wind tunnel of the RJTF (freestream Mach numbers = 3.4, 5.4 and 6.7 at the nozzle exits, 100 mm x 100 mm) was constructed to supplement the engine tests. A special group was formed in the Ramjet Research Division to conduct these studies and to develop new techniques for evaluation of aerodynamic and combustion performances. In this report, details of the subscale wind tunnel and the studies are summarized. First, the design, calibration of the wind tunnels and transient loads during the wind tunnel start are summarized, then various supplementary studies, i.e., engine alignments, air capture ratios, oil flow on engine internal walls, drag and wall pressure measurements and simulation of engine unstart process are explained. Studies on the mixing process of fuel injected into engines and on the ramp-compression-type inlets for a subsonic combustion ramjet engine are also introduced. A series of combustor experiments using cold flow was conducted to investigate flow structure around the fuel injectors behind backward-facing steps in scramjet combustors. In the component experiments, optical measurements, such as Particle Image Velocimetry (PIV) and Pressure Sensitive Paint (PSP) methods were applied. Measurement techniques for the RJTF engine tests were also developed in these studies. Freezing criteria of chemical reactions in gas sampling probes, six-force balances and response times in low pressure measurements are reviewed. In the appendix, designs of the M3.4-, M5.4- and M6.7-two dimensional, square, and M8-conical nozzles, an air-driven ejector and a noise suppressor are summarized, as are construction of a new wind tunnel and the heat loss in storage air heaters. Finally, the know-how accumulated during RJTF operations is mentioned., 資料番号: AA0046109000, レポート番号: NAL TR-1458

日本機械学会2005年度年次大会(2009年9月19-22日), 調布市, 資料番号: ARDS05066000, レポート番号: 356