Conceptual Design Model of High-Altitude Test Stand for Rocket Engines *
|
|
- Rafe McCarthy
- 5 years ago
- Views:
Transcription
1 Trans. Japan Soc. Aero. Space Sci. Vol. 59, No. 3, pp , 2016 Conceptual Design Model of High-Altitude Test Stand for Rocket Engines * Takeshi KANDA, 1) Yohei OGAWA, 2) Daizo SUGIMORI, 3) and Makoto KOJIMA 1) 1) Research and Development Directorate, JAXA, Kakuda, Miyagi , Japan 2) Research and Development Directorate, JAXA, Tsukuba, Ibaraki , Japan 3) Space Technology Directorate I, JAXA, Tsukuba, Ibaraki , Japan Conceptual design procedures and design models of HATS are revised and renewed. The results calculated using the revised method are compared with the operating conditions of HATS at the JAXA Kakuda Space Center. The previous physical method of test chamber pressure and that of ejector suction are adopted in the present model. The suction model adopts the inviscid momentum exchange mechanism. The deceleration process in the supersonic diffuser is revised using the pseudo-shock model. Physical and thermodynamic models are constructed for condensation and steam saturated flow conditions. The results calculated are in reasonable agreement with the measured values (e.g., pressure of secondary flow at the ejector section and pressure changes during engine shut down). The effect of ejector steam condensation on the operating conditions of HATS is quantitatively presented. Key Words: HATS, Rocket Engine, Ejector, Diffuser, Condensation, Saturation Nomenclature A F: coefficients of vapor pressure equation A: cross-section F: impulse function, force f: streamwise component of force h v : latent heat h: enthalpy L: length M: Mach number _m: mass flow rate p c : critical pressure p v : vapor pressure p: mean pressure s: entropy T: temperature T c : critical temperature T r : normalized temperature by T c u: velocity x: dryness, parameter of temperature Subscripts a: test chamber e: outflow f: friction g: saturated vapor, gas i: entrance of HATS diffuser, inflow l: saturated liquid, liquid p: pseudo-shock r: reaction ref: nominal condition rk: rocket t: total 2016 The Japan Society for Aeronautical and Space Sciences + Received 16 July 2015; final revision received 5 November 2015; accepted for publication 18 January Corresponding author, kanda.takeshi@jaxa.jp x: mixture 1: condition 1, fluid 1, upstream of pseudo-shock 2: condition 2, fluid 2, downstream of pseudo-shock 1. Introduction The High-Altitude Test Stand (HATS) is a facility for testing upper-stage rocket engines. HATS creates a high-altitude environment, and the rocket engines operate under low-pressure conditions created by HATS. JAXA has two HATS facilities at the Kakuda Space Center. 1) One was constructed by National Aerospace Development Agency (NASDA) and is used for development tests and acceptance tests of upper-stage engines. It is called NASDA-HATS in the present study. The other was constructed by the National Aerospace Laboratory (NAL) as the pilot HATS for NAS- DA-HATS and is primarily used for research activities. It is called NAL-HATS. These facilities were designed and constructed about 40 years ago. There are several HATS facilities in the world (e.g., facilities at Stennis Space Center, 2) White Sands Test Facility, 3) Plum Brook, 4) and Lampoldshausen 5) ). In the construction of a new facility, a conceptual study is first conducted to specify requirements. The study progresses by examining exiting physical models, design tools and design data, and shows an outline of the facility. Based on the results of the conceptual study, details of the facility are designed, development tests and simulations are conducted, and the facility is constructed. In the construction of Japan s HATS, the same process was conducted. There are reports of the conceptual study on the JAXA HATSs in the form of a technical report or a technical memorandum of the National Aerospace Laboratory, Japan. The other HATS reports present facility specifications or component study results, not facility conceptual study results. In the conceptual design process of the two JAXA HATSs, several physical models and estimation formulas were used 161
2 Fig. 1. Schematic of a HATS. and the estimated results were verified experimentally. 6,7) In the design process, some parameters were specified experimentally. In the present study, the conceptual design models and design procedures of the HATSs are reviewed and revised. Over the 40 years, several technologies have been constructed, and there are findings on physical mechanisms relating to HATS operation. Some are applied to the present conceptual design model of HATS. The results calculated by the present procedure can be an initial configuration in the HATS design process and in CFD simulation. The present physical models and the procedures will be applied to design other systems (e.g., an exhaust system of wind tunnels, ejector pumps, and saturation of working gases) in which similar or the same physical phenomenon are utilized as in the HATSs. 2. HATS Operation Both the NASDA-HATS and the NAL-HATS are comprised of two steam ejectors and three diffusers. The ejectors are located in the center of the duct. Herein, the operation of a HATS is explained with this configuration. Figure 1 shows a schematic of a HATS. A HATS creates a low-pressure condition in the test chamber through ejector operation. The exhaust-gas pressure is recovered in the diffuser. The combustion gas of a rocket engine expands from the exit of the nozzle. The flow-field of the re-attachment of the combustion gas specifies the pressure in the test chamber. The supersonic rocket engine combustion gas attaching to the HATS wall decelerates in the first diffuser and becomes subsonic, passing through the shock-train, or the pseudoshock. Cooling water is injected to the subsonic combustion gas. The mixture of the combustion gas and the steam of the cooling water is suctioned by the first ejector. Suction is necessary because the pressure of the rocket gas mixture is lower than the ambient pressure. At the same time, the total pressure of the rocket gas mixture is increased by mixing with the high-pressure ejector steam. Pressure is recovered in the second diffuser. The same suction and pressure recovery are conducted in the second ejector and third diffuser. The gas mixture is exhausted to ambient pressure. To match the exit pressure condition, the position of the pseudo-shock changes in the first diffuser. The major phenomena in a HATS are reattachment of the rocket gas to the wall, deceleration of supersonic gas flow, Fig. 2. Schematic of attaching supersonic flow behind the step. suction by the ejector, pressure recovery in the diffuser, and condensation of steam in the ejectors. 3. Physical Models for Calculation 3.1. Test chamber pressure The pressure of the test chamber is the most important specification in the high-altitude tests at a HATS. A two-dimensional, supersonic base flow model 8 10) was applied to attachment of the rocket combustion gas to the HATS duct wall during the design procedure of the two HATSs. Figure 2 shows a schematic of the flow-field. A supersonic fluid flows at the rearward facing step. The fluid expands isentropically at the corner of the step. The accelerated fluid attaches to the wall nonisentropically and creates a shock wave. The primary flow in Fig. 2 corresponds to the rocket combustion gas in the HATS, and the lower wall corresponds to the diffuser duct wall. The vertical wall in Fig. 2 corresponds to the opening between the rocket nozzle and the diffuser duct wall. In designing the HATSs, the pressure of the test chamber was estimated by the pressure in the base region behind the step. The present study also adopts the base flow model to estimate the pressure of the test chamber. There are several flow models and equations for the base pressure. In the present study, the empirical pressure equation is adopted. 9) A one-dimensional flow model can be applied to this process for attaching the rocket combustion gas to the HATS duct wall. Expansion of the rocket exhaust to the HATS duct wall was calculated assuming isentropic change. The rocket combustion gas receives reaction force from the pressure of the test chamber during the expansion process. This pressure is 162
3 Fig. 3. Schematic of pseudo-shock. not equal to the pressure in the isentropic change. In the present study, the one-dimensional flow model is also used for attachment with nonisentropic change. The impulse function of the expanded, attached rocket exhaust is F ¼ F rk þða i A rk Þp a ð1þ 3.2. Pseudo-shock in diffuser The supersonic, rocket engine combustion gas decelerates to subsonic flow through the shock train, or the pseudoshock, in the first diffuser. This phenomenon affects pressure recovery performance in the diffuser. The starting position, or length of the pseudo-shock, affects the required length of the duct. Studies on the pseudo-shock has been conducted. 11) Recently, the starting position of the pseudo-shock was estimated using the momentum balance model. 12) In the model revised for the present study, this estimation method is newly adopted to calculate the position of the pseudo-shock and pressure recovery performance of the first diffuser. This model can be applied to a diffuser of supersonic/hypersonic wind tunnels. The impulse function of the supersonic inflow is not generally equal to that of the subsonic outflow. The boundary condition (e.g., pressure or choking) is applied to the subsonic outflow. On the other hand, there is little friction force in the pseudo-shock region because the boundary layer thickens and separation may appear in the region. Figure 3 shows a schematic of the flow-field with the pseudo-shock. Force balance is written as F e ¼ F i þ f r1 þ f rp þ f r2 f f1 f f2 ð2þ where f r2 is the streamwise component of reaction force, which appears when the end of the pseudo-shock is upstream of the divergent section exit. In the momentum balance model, there is no friction in the pseudo-shock region. By balancing the inflow and outflow impulse functions, the starting position of the pseudo-shock is specified. The length of the pseudo-shock is several times the diameter of the duct. 13) Herein, the length is assumed to be five times larger. Studies on the length of the pseudo-shock are under investigation. When the presumed length of the pseudo-shock is too large, the pseudo-shock starts from the entrance of the first diffuser, especially under the throttled condition of a rocket engine. This pseudo-shock condition interacts with the attachment of rocket engine exhaust to the HATS duct wall, and induces the diffuser break. In the actual operation of a HATS, the diffuser break is measured prior to arrival of the pseudo-shock to the diffuser entrance. As presented later, the results calculated are close to the operating conditions of the HATS. The presumed ratio of the pseudo-shock length would be close to the actual length. When the position of the pseudo-shock is located relatively downstream in the first diffuser, the end position of the shock wave is in the mid of the divergent section in the diffuser. Downstream of the pseudo-shock, pressure is recovered approximately in the isentropic change. As the pseudoshock moves upstream, the Mach number becomes smaller, the shock wave becomes weaker, and total pressure loss becomes smaller. With the momentum balance model, the performance of pressure recovery can be estimated. Since the position of the pseudo-shock changes due to the operating conditions of the rocket engine tested, the required length of the diffuser duct can be confirmed with the model. When the rocket engine decreases its thrust during shutdown, the momentum in the HATS duct decreases and the pseudo-shock is moved toward the entrance of the HATS duct. As the pseudo-shock exists upstream, the momentum and pressure recovery of the HATS are improved. After the shock wave arrives at the entrance, pressure in the test chamber increases, and the rocket combustion gas detaches from the HATS duct wall. The diffuser operates as a subsonic diffuser. This increase in pressure and the change in flow condition is called the diffuser break. The one-dimensional model is used after the diffuser break. In simulating the diffuser break using the one-dimensional model, the rocket gas expands to the pressure of the test chamber, p a, and the gas flow detaches from the HATS duct wall. In the second and third diffusers, the secondary subsonic flow rate is large, and the pseudo-shock model is not applied to the diffusers. The primary, supersonic flow of the ejector steam and the secondary, subsonic flow mix together. The subsonic mixed gas recovers pressure in the divergent section of the diffuser, which is calculated under the assumption of isentropic change. Cooling of the supersonic diffuser has been studied. 14,15) Supersonic combustion gas flow decelerates to subsonic through the pseudo-shock. The general estimation formula for heat transfer is not effective in the pseudo-shock region. Cuffel and Back showed a ratio of heat fluxes was proportional to 0.8 power of a ratio of wall pressures for air in a straight duct. 14) Recently, Kato and Kanda extended the model of Cuffel and Back to a divergent duct and a combustion gas. 16) 3.3. Vaporization of water and condensation of steam Water is sprayed to the combustion gas of the rocket engine for cooling and vaporizes around the exit of the first dif- 163
4 Mach number at exit of nozzle Trans. Japan Soc. Aero. Space Sci., Vol. 59, No. 3, 2016 Table 1. Total pressure, MPa Pressure at the exit of hypersonic wind tunnel. Design pressure (no condensation), kpa Calculated pressure by saturated N2, kpa Measured pressure, kpa fuser. The gas after water-cooling is a mixture of the combustion gas and cooling water. The steam is in a saturated condition and its properties are different from those of a pure gas. The dryness of the saturated steam and flow conditions of the steam are calculated to estimate ejector performance. Mean specific heat is calculated with mean molecular weight. No cooling water momentum, liquid water volume, or effect of liquid water on the sound speed of the gas are presumed. Liquid water has no relation to the ideal gas equation. The velocity of the liquid is the same as that of the gas here. Momentum is conserved prior to and after condensation. Enthalpy is calculated with latent heat. The energy conservation is expressed as Here _m h t ¼ _m g h t;g þ _m l h t;l h t;l ¼ h t;g h v ð4þ Dryness, x, is a ratio of the gaseous mass flow rate to the total flow rate. x ¼ _m g = _m The latent heat of water is calculated using the formula of Watson. 17) Here h v;2 ¼ h v;1 fð1 T r;2 Þ=ð1 T r;1 Þg 0:375 T r ¼ðT=T c Þ The saturated vapor pressure of steam is calculated using the formula of Wagner and Pruss.! 18) ln p v ¼ T c T ðaxþbx1:5 þ C x 3 þ D x 3:5 Here p c þ E x 4 þ F x 7:5 Þ x ¼ 1 ðt=t c Þ A ¼ 7: ; B ¼ 1: ; C ¼ 11: ; D ¼ 22: ; E ¼ 15: ; F ¼ 1: Condensation of steam happens in the ejector nozzles. Its effect was simulated by changing the ratio of specific heats in the design process of the HATSs. 19) In the present study, the dryness and properties in the ejector nozzles are calculated assuming isentropic change. This process is similar to that of the equilibrium flow of combustion gas in rocket nozzles. There are recombination ð3þ ð5þ ð6þ ð7þ ð8þ ð9þ of molecules, change of mean molecular weight and release of heat under the assumption of isentropic change. It is well known that the equilibrium flow calculation of rocket combustion gas reasonably predicts the performance of the rocket nozzles. In the ejector nozzles, the molecular weight of the gas changes due to condensation and latent heat is released. A similar change as in the rocket nozzles progresses in the ejector nozzles. Entropy is calculated using the general thermodynamic relation below. 20) s ¼ s g þfð1 xþh v g=t ð10þ This isentropic flow model with condensation is verified using the results of hypersonic wind tunnel nozzle flows. 21) The Mach numbers of the nozzles are 5.4 and 6.7, respectively. The fluid of the experiments was air at room temperature, whereas the calculation is conducted using the properties of nitrogen. Table 1 lists the test condition and the pressures measured and calculated. The design pressure was calculated using the ideal gas relation with no condensation. The pressure calculated with condensation was higher than the design values and close to the measured values, respectively. One of the reasons for the difference is deviation of the flow conditions in the wind tunnel nozzles from the designed flow conditions due to condensation. According to the present calculation with condensation, the pressure measured corresponds to the Mach 4.7 flow condition, whereas the Mach number calculated is 4.8 for the Mach 6.7 design condition. For the Mach 5.4 condition, both measured and calculated Mach numbers correspond to the Mach 4.5 condition Ejector system The primary, supersonic steam of the ejector interacts with the secondary, subsonic fluid, which is a mixture of the rocket combustion gas and cooling water. In the second ejector system, the secondary mixed flow also contains steam from the first ejector. By the interaction, the secondary fluid is pumped out or suctioned. Figure 4 shows a schematic of the ejector system flow-field. Momentum is exchanged in the interaction of the two flows in an almost inviscid process. This inviscid, momentum exchange mechanism has been presented (e.g., by Fabri and Siestrunck) and verified experimentally. 22) This model was adopted in the HATS design process. The present model also adopts this mechanism. This inviscid model is a general basis for most ejector systems. The conservations of mass, momentum and energy, relation between pressures and relation between cross-sections are _m 0 1 ¼ _m 1 ð11þ _m 0 2 ¼ _m 2 ð12þ 164
5 (a) Fig. 5. Entropy changes due to momentum exchange along the mean pressure, p. _m 0 1 u0 1 þ A0 1 p0 1 ¼ _m 1 u 1 þ A 1 p 1 A p _m 0 2 u0 2 þ A0 2 p0 2 ¼ _m 2 u 2 þ A 2 p 2 þ A ~p p 0 2 ¼ p0 1 A ¼ A 0 1 A 1 ¼ A 2 A 0 2 h 0 t;1 ¼ h t;1 h 0 t;2 ¼ h t;2 (b) Fig. 4. Schematic of the ejector. Secondary flow is subsonic. In (a), pressure of the secondary flow is higher than that of the ejector gas flow. In (b), pressure of the secondary flow is lower than that of the ejector gas flow. ð13þ ð14þ ð15þ ð16þ ð17þ ð18þ The condition after the momentum exchange is expressed with a prime. p 1 is not usually equal to p 2, whereas pressure is equal both in the primary and secondary flows after the momentum exchange. p is the mean pressure during the exchange. A p can be negative. Momentum is exchanged at the dividing streamline. Some of the previous studies calculated changes in flow condition using the method of characteristics. In the present model, the exchange is calculated in a simplified way. 23) Pressure after the interaction, p 0 2 or p0 1, is used for the mean pressure on the dividing streamline, p. Simplification of the pressure was verified by comparison with the experimental results. 24) There are several kinds of inflow boundary conditions for the primary and secondary fluids. For example, the pressure of the supersonic primary flow is presumed to be lower than that of the subsonic secondary flow. The primary flow increases pressure through shock waves and is still supersonic after the interaction. The pressure of the secondary flow decreases even though the flow diverges. It is generally a contradiction to the pressure-cross-section relation. This can happen in the non-isentropic change. As Fabri mentioned, momentum exchange is not an isentropic process. 22) In the present study, this combination of primary and secondary flows is calculated with in a simplified way. In the previous study, this flow condition was not discussed. Figure 5 shows an example calculation result for the relationship between the mean pressure, p, and increases in entropy. Fluid 1 is supersonic steam and fluid 2 is subsonic steam. The total temperature is 400 K and the mass flow rate is 50 kg0s ¹1 for both fluids. The total pressure of fluid 1 is 500 kpa and that of fluid 2 is 40 kpa. The cross-section of fluid 1 is 0.4 m 2 and that of fluid 2 is 2 m 2. In this example calculation, the interacting pressure is equal to the pressure after the interaction, as in the present simplified calculation. Mean entropy after the interaction is calculated with each entropy by weighing each mass flow rate. Mass, momentum and energy are conserved at the mean pressure of 38.4 kpa, indicated as Interacting pressure in the figure. The pressure is slightly lower than the pressure of fluid 2, 38.5 kpa, prior to the interaction. The entropy increases slightly in fluid 2 at the mean pressure, whereas the increase in fluid 1 is extensive. The change of the cross-section is 0.19 m 2 under these conditions. With these conditions, the ejector works to hold a lowpressure environment for the secondary flow and to increase the total pressure of the secondary flow by mixing the primary flow downstream of the momentum exchange. When pressure in the secondary flow is lower than that in the primary flow at the entrance of the ejector section, the primary flow expands and the secondary flow converges. The pressure for both is lower than that of the flows prior to the momentum exchange. Figure 4(b) shows a schematic of these conditions. The secondary flow may choke. This is called Fabri choking and appears when starting and shutting down a rocket engine. It was found that mixing degrades the suction performance of the ejector. 25) Attention is required when designing the ejector Calculation procedure With these physical models and assumptions, flow in HATS is calculated. New physical modes are used for the pseudo-shock in the first diffuser and steam properties in 165
6 the ejector nozzles, and pressure of the test chamber in the 1- D model. The same models as those in the previous study are used for suction performance of the ejectors and pressure of the test chamber with the base-flow model. For the estimation of pressure in the test chamber, a new 1-D model and a conventional model are used. The exit boundary condition is a specified pressure of 1.05 atm, being larger than 1 atm due to the loss of pressure in the silencer downstream of the third diffuser. The upstream boundary condition is the inflow of the rocket combustion gas. The pseudo-shock position is located to balance the inflows and outflows of mass, momentum and energy. The ratio of the specific heat is 1.33 for steam. The properties of steam are calculated under the assumption of invariable specific heat. It is assumed there is no friction in the pseudo-shock region, 12) whereas the friction coefficient is set to be ) Heat transfer is calculated using the Reynolds analogy. Wall temperature is set to 350 K. 4. Results and Discussion Application of the physical models is verified by comparing the calculated and measured operating conditions of the NAL-HATS. 27) The specifications of the NAL-HATS are listed in Table 2. The specifications of the engine tested are listed in Table 3. Though the test results of another engine are also described in the report, most of the HATS operating tests were conducted with the engine listed in Table 3. In the present study, another calculation is conducted under the condition that the outer diameter of the secondary flow duct exit of the second ejector section is reduced to 90%, besides the calculation using the original geometry. Under the original geometry, when the rocket engine thrust is reduced, the pseudo-shock is located at the entrance of the first diffuser and the exit pressure of the HATS is calculated to be lower than the specified exit pressure. By reducing the diameter, the position of the pseudo-shock is still in the mid of the first diffuser even under the reduced-thrust condition. This effect is also attained by reducing the outer diameter of the secondary flow of the first ejector section. Details are described in Section 4.2. This reduction of the cross-section may represent the effect of displacement thickness of the boundary layer. The NAL-HATS was operated by injecting a small amount of air under no rocket exhaust condition for steady operation of the ejector. At the renewal of the NAL-HATS in 2011, the outer diameter of the secondary flow duct of the first ejector section was reduced. The effect of reducing the diameter is shown in Fig. 6. In this example calculation, the mass flow rate and total temperature of the secondary flow are set to be 50 kg0s ¹1 and 500 K, respectively. Those of the primary flow are 122 kg0s ¹1 and 500 K, respectively. The total pressure of the primary flow is 1.30 MPa and the Mach number at the exit of the ejector nozzle is 1.5 or 2.5. The static pressure of the secondary flow is the same as that of the primary flow at the nozzle exit. These are similar to the operating conditions of the NAL- HATS. The static pressure of the secondary flow is equal to that of the primary flow for simplicity. Then, the velocity of the secondary flow is specified by the pressure condition. The cross-section of the secondary flow is a parameter. As the cross-section is decreased, the total pressure of the mixture is increased. The Mach number of the secondary flow becomes large as the cross-section decreases, and the flow chokes at A 2 =A 1 ¼ 1:33 in the case of M 1 ¼ 2:5. As the cross-section of the secondary flow decreases, the velocity of the secondary flow becomes close to that of the primary flow. This reduces the increase in entropy during mixing. In the design process, selection of the secondary flow cross-section of the ejector section is an important parameter. In this section, the results with the 90% outer exit diameter are referred to, as well as those with the original configuration. Table 3. Specifications of the rocket engine tested. Propellants NTO/A-50 Thrust, kn 53 Chamber pressure, MPa 1.2 Area ratio of nozzle 26 Propellant flow rate, kg0s ¹ Diameter at throat, mm 185 Diameter at exit, mm 944 Table 2. Specifications of the NAL-HATS. Diameter, length and exit diameter of 1st 1:1 12 2:0 diffuser, m Sprayed cooling water, kg0s ¹1 35 (50 kn eng.) Mass flow rates of steam ejectors, kg0s ¹1 40 (1st ejector) 120 (2nd ejector) Area ratio of ejector nozzles 11.6 (1st ejector) 5.5 (2nd ejector) Diameter, length and exit diameter of 2nd 1:4 20 2:6 diffuser, m Diameter, length and exit diameter of 3rd 1:4 21 2:4 diffuser, m Fig. 6. Relation between secondary flow cross-section and total pressure of the ejector system mixture. 166
7 Fig. 7. Pressure of secondary flow of the 2nd ejector section. Fig. 8. Changes of pressure and thrust during engine shut-down. 27) In the accumulator, the temperature was higher than the saturated temperature. However, the duct from the accumulator to the ejector has length, and the dryness at the exit of the accumulator was lower than unity. 27) In the present study, the total temperature of steam is the saturated temperature at the stagnation pressure in the ejector manifolds as a reference condition. Calculation with the superheated steam can be conducted using the present model and the results are referred to later Operation of second ejector Figure 7 shows pressure of the secondary flow at the entrance of the second ejector section along the steam stagnation pressure of the second ejector. The calculated pressures are plotted in the figure. In the calculation, the mass flow rate of the cooling water sprayed is 35 kg0s ¹1. The mass flow rate of the secondary flow is 94 kg0s ¹1 and the calculated flow rate is 93 kg0s ¹1. The total temperature of the secondary flow is 445 K. The calculated pressures of the secondary flow reasonably agree with the values measured within the convergence error of the calculation. In the previous design process, the calculated pressure of 60 kpa was attained with the flow rate of 140 kg0s ¹1, which was about 1.5 times larger than the actual flow rate. 27) The pressure of 40 kpa was attained with secondary flow rate of 85 kg0s ¹1, which was 0.9 times the actual flow rate. Qualitative agreement between the calculated and experimental results was addressed. The large difference in the previous calculation would be due to the difference in application of the ejector model, described in Section 3.4. The accuracy is improved in the present HATS design model Test chamber pressure and diffuser break Figure 8 shows changes in the test chamber pressure, wall pressure of the first diffuser and normalized thrust in the engine shut-down process when testing the NAL-HATS. Time measurement started from engine ignition. At 20.81s, the thrust ratio was about 0.6 and the diffuser break happened. Pressure in the test chamber increased after the break. Under the break condition, the rocket engine combustion gas might detach from the HATS duct. The test chamber pressure measured was 1.1 kpa under the design operating conditions. The calculated pressure was 0.4 kpa using the base flow equation and is 1.6 kpa using the one-dimensional, nonisentropic expansion model. There would be measurement and convergence errors. The transient operating condition of the HATS would be different from that in the calculation. This conceptual study calculates steady-state conditions. Under such conditions, the calculations predict approximate pressure. There was no comparison of this pressure history between the measured and the calculated results in the previous study. 27) The diffuser break condition is studied using the one-dimensional model. Until the diffuser break, the rocket exhaust expands nonisentropically to the duct wall and the impulse function of the attached combustion gas flow is calculated using Eq. (1). In the diffuser break condition, the rocket exhaust expands to the test chamber pressure and detaches from the HATS duct wall. The impulse function is the sum of the rocket exhaust momentum and the force of the chamber pressure. The change from attachment to detachment is caused by the momentum balance to attain the HATS exit pressure condition. The diffuser break, which is detachment of the rocket exhaust from the wall, is calculated to happen at the thrust ratio, F=F ref, of 0.55 when the outer diameter is reduced to 85%. In the experiment, the break, which was defined by an abrupt change in pressure, happened around the thrust ratio of 0.6, as shown in Fig. 8. The diffuser break calculated with the present model agreed reasonably well with the experimental break. In the present calculation, the test chamber pressure is 0.9 kpa prior to the diffuser break, whereas it is 4.7 kpa at the thrust ratio of 0.45 after the break. In the experiment, it was 1.1 kpa prior to the break, and 4.7 kpa at the ratio of These pressures agree reasonably with the values measured. There was no comparison on this pressure history between the measured and the calculated results in the previous study. 27) 167
8 4.3. Condensation of steam in ejector nozzle The comparisons in the previous sections show the present model predicts the operating conditions of a HATS reasonably well. In the comparisons, the effect of condensation is already included. In this section, the effect of steam condensation is discussed. In the present study of HATS operation, calculation of steam condensation is conducted in all ducts and sections. As a result, condensation happens only in the ejectors. The discussion is limited to the flow in the ejectors. Condensation in the ejectors was well recognized in previous studies, however, a suitable model for saturated steam flow was not constructed. As described in Section 3.3, the effect of condensation was simulated by changing the ratio of specific heats. Dryness changes due to the local temperature and pressure conditions in the saturated flow, and uncertainty remains in the steam ejector performance. In the present study, the saturated steam flow is calculated using dryness and isentropic flow modeling. Part of the saturated steam condenses during expansion in the ejector nozzle. Steam is assumed to be in a saturated condition at 465 K at the pressure of the ejector manifold. The conditions for ejector steam and operating conditions are listed in Tables 4 and 5. Flow conditions in the superheated condition of 528 K are also listed in the tables. The temperature at the superheated condition is the saturated temperature of steam in the accumulator. For comparison, flow conditions of no gas condensation, in other words, no-liquid condition, are also listed. Total pressures of the mixture downstream of the second ejector (i.e., in the third diffuser) for the no-liquid cases are lower than the exit condition and are shown in parentheses. Results with the outer diameter reduced are listed on the right-hand side in the column. Condensation can even happen at the throat of the ejectors. This makes a difference in the design mass flow rate of the ejectors. At the exit of the ejector nozzles, condensation further progresses and total pressure becomes much lower than the pressure in the no-liquid cases. However, the impulse functions at the exit are approximately the same as those with no condensation in the previous ejectors. A small increase in the impulse function under the saturated condition is similar to increasing the thrust under the equilibrium flow condition. Differences appear in static pressure at the exit of the ejector nozzles, as well as the difference shown in Table 1. In the no-liquid cases, the pressure at the ejector exit is lower than that of the saturated condition. The difference between the pressure of the secondary flow and that at the ejector exit becomes great in no-liquid cases. The pressure in the third diffuser is specified by the pressure exit condition. When the loss of total pressure during mixing is small in the second ejector, high total pressure in the secondary flow is not required. Under the low total pressure condition of the secondary flow in the second diffuser, the pseudo-shock can be positioned in the mid of the first diffuser (i.e., at relatively downstream in the first diffuser). The small total pressure downstream of the first ejector is caused by this mechanism in the saturated steam cases. The effects of condensation in the ejectors should be included in the design procedure for HATS. Table 4. Calculated operating conditions of 1st ejector. 1st ejector Saturated Superheated No liquid (T t ¼ 465 K) No liquid (T t ¼ 528 K) Stagnation pressure, MPa Stagnation temperature, K Mass flow rate, kg0s ¹ Dryness at throat Dryness at exit of nozzle Pressure at nozzle exit, kpa Total pressure ratio at nozzle exit Impulse function, kn Total pressure of secondary flow, kpa 48.7/ / / /44.9 Total pressure of mixed flow, kpa 65.1/ / / /60 Pressure after suction of secondary flow, kpa 47.7/ Table 5. Calculated operating conditions of 2nd ejector. 2nd ejector Saturated Superheated No liquid (T t ¼ 465 K) No liquid (T t ¼ 528 K) Stagnation pressure, MPa Stagnation temperature, K Mass flow rate, kg0s ¹ Dryness at throat Dryness at exit of nozzle Pressure at nozzle exit, kpa Total pressure ratio at nozzle exit Impulse function, kn Total pressure of secondary flow, kpa 65.1/ / / /60 Total pressure of mixed flow, kpa 107/ /109 (105)/109 (105)/109 Pressure after suction of secondary flow, kpa 62.6/
9 5. Conclusion Conceptual design procedures and design models for HATS facilities are reviewed and revised. The results calculated using the present model were compared with the operating conditions of a previous HATS, and the present model was verified. The previous prediction method for the test chamber pressure is also adopted in the present model. The deceleration process was revised using the momentum balance model of pseudo-shock. The inviscid momentum exchange model of the ejector system that calculates operating conditions agreed with the conditions measured. The present model shows that condensation in the ejectors affects the HATS operating conditions. Acknowledgments The authors thank Mr. Kouichiro Tani and Mr. Takeo Tomita of JAXA for useful discussions and advices. References 1) JAXA Kakuda Space Center, index.html (accessed June 16, 2015) 2) Maynard, B. T. and Raines, N.: Altitude Testing of Large Liquid Propellant Engines, AIAA Paper , Seattle, Washington, USA, June ) Harris, D. and Cort, R.: White Sands Test Facility Test Stand 401, AIAA Paper , Cleveland, OH, USA, ) Kenzakowski, D. C. and Brinckman, K. W.: CFD Simulation of NASA B-2 Spray Chamber during Rocket Fire, AIAA Paper , Reno, NV, Jan ) Schäfer, K. and Zimmermann, H.: Simulation of Flight Condition during Lift OFF for Rocket Engine Testing, AIAA Paper , Fort Lauderdale, FL, USA, July ) Kumagai, T., Miyajima, H., Kamata, M., Satoh, M., Abe, N., Sudoh, T., Yamada, A., and Kouchiyama, J.: Simulation Test of Exhaust System of High Altitude Test Stand for LOX/LH 2 Rocket Engine, NAL TM-461, Mar (in Japanese). 7) Miyajima, H., Abe, N., and Kisara, K.: Design Calculation of Diffuser for Rocket Engine High Altitude Test Stand, NAL TM-313, Sep (in Japanese). 8) Korst, H. H.: A Theory for Base Pressure in Transonic and Supersonic Flow, J. Appl. Mech., 23, 4 (1956), pp ) Lamb, J. P. and Oberkampf, W. L.: Review and Development of Base Pressure and Base Heating Correlations in Supersonic Flow, J. Spacecraft Rockets, 32, 1 (1995), pp ) Karashima, K. and Hasegawa, K.: An Approximate Approach to Base Flow behind Two-Dimensional Rearward-Facing Steps Placed in a Uniform Supersonic Stream, ISAS Report, No. 501, 1973, pp ) Matsuo, K., Miyazato, Y., and Kim, H.-D.: Shock Train and Pseudo- Shock Phenomena in Internal Gas Flow, Progress in Aerospace Sciences, 35, 1 (1999), pp ) Kanda, T. and Tani, K.: Momentum Balance Model of Flow Field with Pseudo-Shock, JAXA Report, JAXA-RR E, Mar ) Kanda, T.: Prediction of Pseudo-Shock Position in a Long Duct Using the Momentum Balance Model, Trans. Jpn. Soc. Aeronaut. Space Sci., 55, 3 (2012), pp ) Cuffel, R. F. and Back, L. H.: Flow and Heat Transfer Measurement in a Pseudo-Shock Region with Surface Cooling, AIAA J., 14, 12 (1976), pp ) Zaikovskii, V. N., Trofimov, V. M., and Shtrekalkin, S. I.: Experimental and Computational Investigation of Heat Fluxes in a Supersonic Diffuser, J. Appl. Mechanics Rechnical Physics, 37, 1 (1996), pp ) Kato, K. and Kanda, T.: Calculation of Heat Flux in the Pseudo-Shock Region, 30th ISTS, 2015-a-59, Kobe, Hyogo, Japan, July ) Reid, R. C., Prausnitz, J. M., and Sherwood, T. K.: The Properties of Gases and Liquids, 3rd Ed., McGraw-Hill, New York, 1977, pp ) Wagner, W. and Pruss, A.: International Equations for the Saturation Properties of Ordinary Water Substance. Revised According to the International Temperature Scale of 1990, J. Phys. Chem. Reference Data, 22, 3 (1993), pp ) Miyajima, H., Yamada, A., Kisara, K., Kamata, M., Satoh, M., Ueno, T., Kumagai, T., and Kusaka, K.: Experiments on Steam Ejectors for Rocket Engine Altitude Simulation, NAL TR-566, Apr (in Japanese). 20) Iinuma, K.: Engineering Thermodynamics, 4th Ed., Gakutoh Co. Ltd., Tokyo, 1979, pp (in Japanese). 21) Mitani, T., Hiraiwa, T., Kanda, T., Shimua, T., Tomioka, S., Kobayashi, K., Izumikawa, M., Sakuranaka, N., Watanabe, S., Tarukawa, Y., Kouchi, T., Kitamura, E., and Yatsunami, T.: Subscale Wind Tunnels and Supplemental Studies of Scramjet Engine Tests, NAL TR-1458, Apr (in Japanese). 22) Fabri, J. and Paulon, J.: Theory and Experiments on Supersonic Air-to- Air Ejectors, NACA TM 1410, Jan ) Kanda, T. and Kudo, K.: Conceptual Study of a Combined-Cycle Engine for an Aerospace Plane, J. Propul. Power, 19, 5 (2003), pp ) Aoki, S., Lee, J., Masuya, G., Kanda, T., and Kudo, K.: Aerodynamic Experiments on an Ejector-Jet, J. Propul. Power, 21, 3 (2005), pp ) Tani, K., Hasegawa, S., Ueda, S., Kanda, T., and Nagata, H.: Analytical Method for Prediction of Suction Preformance of Ejector-Jet, Trans. Jpn. Soc. Aeronaut. Space Sci., 58, 4 (2015), pp ) White, F. M.: Viscous Fluid Flow, McGraw-Hill, New York, 1974, pp. 495, ) Ohtsuka, S., et al.: High Altitude Test Facility for Rocket Engine at NAL, National Aerospace Laboratory, NAL TR-454, Apr (in Japanese). T. Shimada Associate Editor 169
SIMULATION OF ENVIRONMENTAL FLIGHT CONDITIONS
SIMULATION OF ENVIRONMENTAL FLIGHT CONDITIONS BY ADVANCED ALTITUDE SIMULATION Klaus SCHÄFER Institute of Space Propulsion, German Aerospace Centre (DLR) Lampoldshausen, 74239 Hardthausen, Germany klaus.schaefer@dlr.de
More informationInvestigation on Divergent Exit Curvature Effect on Nozzle Pressure Ratio of Supersonic Convergent Divergent Nozzle
RESEARCH ARTICLE OPEN ACCESS Investigation on Divergent Exit Curvature Effect on Nozzle Pressure Ratio of Supersonic Convergent Divergent Nozzle Shyamshankar.M.B*, Sankar.V** *(Department of Aeronautical
More informationS.A. Klein and G.F. Nellis Cambridge University Press, 2011
16-1 A flow nozzle is to be used to determine the mass flow rate of air through a 1.5 inch internal diameter pipe. The air in the line upstream of the meters is at 70 F and 95 psig. The barometric pressure
More informationDesign of a small-scaled de Laval nozzle for IGLIS experiment
Design of a small-scaled de Laval nozzle for IGLIS experiment Evgeny Mogilevskiy, R. Ferrer, L. Gaffney, C. Granados, M. Huyse, Yu. Kudryavtsev, S. Raeder, P. Van Duppen Instituut voor Kern- en Stralingsfysika,
More informationTHEORETICAL EVALUATION OF FLOW THROUGH CENTRIFUGAL COMPRESSOR STAGE
THEORETICAL EVALUATION OF FLOW THROUGH CENTRIFUGAL COMPRESSOR STAGE S.Ramamurthy 1, R.Rajendran 1, R. S. Dileep Kumar 2 1 Scientist, Propulsion Division, National Aerospace Laboratories, Bangalore-560017,ramamurthy_srm@yahoo.com
More informationLab # 03: Visualization of Shock Waves by using Schlieren Technique
AerE545 Lab # 03: Visualization of Shock Waves by using Schlieren Technique Objectives: 1. To get hands-on experiences about Schlieren technique for flow visualization. 2. To learn how to do the optics
More informationAIRFLOW GENERATION IN A TUNNEL USING A SACCARDO VENTILATION SYSTEM AGAINST THE BUOYANCY EFFECT PRODUCED BY A FIRE
- 247 - AIRFLOW GENERATION IN A TUNNEL USING A SACCARDO VENTILATION SYSTEM AGAINST THE BUOYANCY EFFECT PRODUCED BY A FIRE J D Castro a, C W Pope a and R D Matthews b a Mott MacDonald Ltd, St Anne House,
More informationThe effect of back spin on a table tennis ball moving in a viscous fluid.
How can planes fly? The phenomenon of lift can be produced in an ideal (non-viscous) fluid by the addition of a free vortex (circulation) around a cylinder in a rectilinear flow stream. This is known as
More informationDEVELOPMENT OF P4.1 ALTITUDE SIMULATION FOR VINCI ENGINE
DEVELOPMENT OF P4.1 ALTITUDE SIMULATION FOR VINCI ENGINE K. Schäfer, C. Böhm, H. Kronmüller, H. Zimmermann, German Aerospace Centre (DLR), Institute of Space Propulsion, Lampoldshausen, 74172 Hardthausen,
More informationPREDICTION OF TOTAL PRESSURE CHARACTERISTICS IN THE SETTLING CHAMBER OF A SUPERSONIC BLOWDOWN WIND TUNNEL
PREDICTION OF TOTAL PRESSURE CHARACTERISTICS IN THE SETTLING CHAMBER OF A SUPERSONIC BLOWDOWN WIND TUNNEL S R Bhoi and G K Suryanarayana National Trisonic Aerodynamic Facilities, National Aerospace Laboratories,
More informationEFFECT OF GEOMETRIC DISTORTION ON THE FLOW FIELD AND PERFORMANCE OF AN AXISYMMETRIC SUPERSONIC INTAKE
Indian J.Sci.Res.() : 909-99, 0 ISSN:50-08(Online) ISSN : 0976-876 (Print) EFFECT OF GEOMETRIC DISTORTION ON THE FLOW FIELD AND PERFORMANCE OF AN AXISYMMETRIC SUPERSONIC INTAKE MOHAMMAD FARAHANI a, ALIREZA
More informationANALYSIS OF AERODYNAMIC CHARACTERISTICS OF A SUPERCRITICAL AIRFOIL FOR LOW SPEED AIRCRAFT
ANALYSIS OF AERODYNAMIC CHARACTERISTICS OF A SUPERCRITICAL AIRFOIL FOR LOW SPEED AIRCRAFT P.Sethunathan 1, M.Niventhran 2, V.Siva 2, R.Sadhan Kumar 2 1 Asst.Professor, Department of Aeronautical Engineering,
More informationTutorial. BOSfluids. Relief valve
Tutorial Relief valve The Relief valve tutorial describes the theory and modeling process of a pressure relief valve or safety valve. It covers the algorithm BOSfluids uses to model the valve and a worked
More informationObserved in Gas Injection
Characteristics into Liquid* of Jetting Observed in Gas Injection By Yasuhisa OZAWA** and Kazum i MOR I * * Synopsis The present study is concerned with jetting behavior of gas jets injected into water
More informationDEVELOPMENT OF HIGH ALTITUDE TEST FACILITY FOR COLD JET SIMULATION
DEVELOPMENT OF HIGH ALTITUDE TEST FACILITY FOR COLD JET SIMULATION Aldin Justin 1, Robin James, Shruti Panicker 2,A.N.Subash,Saravana Kumar 1 Assistant Professor,Karunya University, aldinjustin@karunya.edu
More informationExperimental Verification of Integrated Pressure Suppression Systems in Fusion Reactors at In-Vessel Loss-of -Coolant Events
Experimental Verification of Integrated Pressure Suppression Systems in Fusion Reactors at In-Vessel Loss-of -Coolant Events K. Takase 1), H. Akimoto 1) 1) Japan Atomic Energy Research Institute (JAERI),
More informationInfluence of rounding corners on unsteady flow and heat transfer around a square cylinder
Influence of rounding corners on unsteady flow and heat transfer around a square cylinder S. K. Singh Deptt. of Mech. Engg., M. B. M. Engg. College / J. N. V. University, Jodhpur, Rajasthan, India Abstract
More informationAIR EJECTOR WITH A DIFFUSER THAT INCLUDES BOUNDARY LAYER SUCTION
Engineering MECHANICS, Vol. 20, 2013, No. 3/4, p. 213 220 213 AIR EJECTOR WITH A DIFFUSER THAT INCLUDES BOUNDARY LAYER SUCTION Václav Dvořák* The article deals with axial-symmetric subsonic air-to-air
More informationSteady Analysis of NACA Flush Inlet at High Subsonic and Supersonic Speeds
, July 1-3, 2015, London, U.K. Steady Analysis of NACA Flush Inlet at High Subsonic and Supersonic Speeds Taimur A. Shams 1, Masud J 2 Abstract Essence of this research is to report computational analysis
More informationOptimization of Ejector's Performance With a CFD Analysis. Amanda Mattos Karolline Ropelato Ricardo Medronho
Optimization of Ejector's Performance With a CFD Analysis Amanda Mattos Karolline Ropelato Ricardo Medronho 2 Introduction Ejectors Equipment industrially used and based on the Venturi phenomena; With
More informationEXPERIMENTAL CHARACTERIZATION OF A EJECTOR PUMP
EXPERIMENTAL CHARACTERIZATION OF A EJECTOR PUMP Capela, N. J. S. Department of Mechanical Engineering, Instituto Superior Técnico, Av. Rovisco Pais, 1049-001, Lisbon, Portugal, 2012 ABSTRACT The option
More informationDevelopment of Gasdynamic Probe for Total Temperature Measurement in gases
ASET 20- National Conference on Emerging Trends in Propulsion Technology Development of Gasdynamic Probe for Total Temperature Measurement in gases K Sathiyamoorthy, Souren Misra 2, Srinivas J, Baskaran
More informationPURE SUBSTANCE. Nitrogen and gaseous air are pure substances.
CLASS Third Units PURE SUBSTANCE Pure substance: A substance that has a fixed chemical composition throughout. Air is a mixture of several gases, but it is considered to be a pure substance. Nitrogen and
More informationApplied Fluid Mechanics
Applied Fluid Mechanics 1. The Nature of Fluid and the Study of Fluid Mechanics 2. Viscosity of Fluid 3. Pressure Measurement 4. Forces Due to Static Fluid 5. Buoyancy and Stability 6. Flow of Fluid and
More informationDirect Numerical Simulation on Hydrogen Fuel Jetting from High Pressure Tank
Direct Numerical Simulation on Hydrogen Fuel Jetting from High Pressure Tank Yun-feng Liu 1, Nobuyuki Tsuboi 2, Hiroyuki Sato 1, Fumio Higashino 1 and A. Koichi Hayashi 1 1 Department of Mechanical Engineering,
More informationResearch and optimization of intake restrictor for Formula SAE car engine
International Journal of Scientific and Research Publications, Volume 4, Issue 4, April 2014 1 Research and optimization of intake restrictor for Formula SAE car engine Pranav Anil Shinde Mechanical Engineering,
More informationNUMERICAL INVESTIGATION OF THE FLOW BEHAVIOUR IN A MODERN TRAFFIC TUNNEL IN CASE OF FIRE INCIDENT
- 277 - NUMERICAL INVESTIGATION OF THE FLOW BEHAVIOUR IN A MODERN TRAFFIC TUNNEL IN CASE OF FIRE INCIDENT Iseler J., Heiser W. EAS GmbH, Karlsruhe, Germany ABSTRACT A numerical study of the flow behaviour
More informationDrag Reduction of Hypersonic Vehicles using Aerospike
Pgs: 26-37 Drag Reduction of Hypersonic Vehicles using Aerospike 1 1 B-tech (Aerospace) Student, divyangupta729@gmail.com Department of Aerospace Engineering, University of Petroleum and Energy Studies,
More informationDesign and Numerical Flow Analysis of Expansion Deflection Nozzle
Design and Numerical Flow Analysis of Expansion Deflection Nozzle Shaik Abdul Muwaaz 1, Nazumuddin Shaik 2 Post Graduate Student, Assistant Professor 1 Shaik Abdul Muwaaz, Aerospace Department, Nimra Institute
More informationCover Page for Lab Report Group Portion. Compressible Flow in a Converging-Diverging Nozzle
Cover Page for Lab Report Group Portion Compressible Flow in a Converging-Diverging Nozzle Prepared by Professor J. M. Cimbala, Penn State University Latest revision: 13 January 2012 Name 1: Name 2: Name
More informationMODELLING OF FUME EXTRACTORS C. R.
LD8 19th International Symposium of Ballistics, 7 11 May 21, Interlaken, Switzerland MODELLING OF FUME EXTRACTORS C. R. Woodley WS4 Guns and Warheads Department, Defence Evaluation and Research Agency,
More informationLOW PRESSURE EFFUSION OF GASES revised by Igor Bolotin 03/05/12
LOW PRESSURE EFFUSION OF GASES revised by Igor Bolotin 03/05/ This experiment will introduce you to the kinetic properties of low-pressure gases. You will make observations on the rates with which selected
More informationUNCLASSIFIED AD DEFENSE DOCUMENTATION CENTER FOR TECHNICAL INFORMATION SCIENTIFIC AND VIRGINIA CAMERON STATION, ALEXANDRIA, UNCLASSIFIED
UNCLASSIFIED AD 4 20813 DEFENSE DOCUMENTATION CENTER SCIENTIFIC AND FOR TECHNICAL INFORMATION CAMERON STATION, ALEXANDRIA, VIRGINIA UNCLASSIFIED NOTICE: When government or other drawings, specifications
More informationFLOW CONSIDERATIONS IN INDUSTRIAL SILENCER DESIGN
FLOW CONSIDERATIONS IN INDUSTRIAL SILENCER DESIGN George Feng, Kinetics Noise Control, Inc., 3570 Nashua Drive, Mississauga, Ontario Vadim Akishin, Kinetics Noise Control, Inc., 3570 Nashua Drive, Mississauga,
More informationAnalysis of pressure losses in the diffuser of a control valve
Analysis of pressure losses in the diffuser of a control valve Petr Turecký 1, Lukáš Mrózek 2*, Ladislav Taj 2, and Michal Kolovratník 3 1 ENVIROS, s.r.o., Dykova 53/10, 101 00 Praha 10-Vinohrady, Czech
More informationCh. 11 Mass transfer principles
Transport of chemical species in solid, liquid, or gas mixture Transport driven by composition gradient, similar to temperature gradients driving heat transport We will look at two mass transport mechanisms,
More informationHigh fidelity gust simulations around a transonic airfoil
High fidelity gust simulations around a transonic airfoil AEROGUST Workshop 27 th - 28 th April 2017, University of Liverpool Presented by B. Tartinville (Numeca) Outline of the presentation 1Objectives
More informationMODELING AND SIMULATION OF VALVE COEFFICIENTS AND CAVITATION CHARACTERISTICS IN A BALL VALVE
Proceedings of the 37 th International & 4 th National Conference on Fluid Mechanics and Fluid Power FMFP2010 December 16-18, 2010, IIT Madras, Chennai, India FMFP2010 341 MODELING AND SIMULATION OF VALVE
More informationSTUDIES ON THE OPTIMUM PERFORMANCE OF TAPERED VORTEX FLAPS
ICAS 2000 CONGRESS STUDIES ON THE OPTIMUM PERFORMANCE OF TAPERED VORTEX FLAPS Kenichi RINOIE Department of Aeronautics and Astronautics, University of Tokyo, Tokyo, 113-8656, JAPAN Keywords: vortex flap,
More informationExperimental study of the influence of the model position on the operating conditions of a plasma wind tunnel
Experimental study of the influence of the model position on the operating conditions of a plasma wind tunnel Carlo Purpura, CIRA Capua c.purpura@cira.it, Tel. (+39)0823-623249 - Fax (+39)0823-623947 Federico
More informationCFD Flow Analysis of a Refrigerant inside Adiabatic Capillary Tube
CFD Flow Analysis of a Refrigerant inside Adiabatic Capillary Tube Y Raja Kumar ¹, Dr.P Usha sri ² PG Student¹, Associate professor² Department of Mechanical Engineering, University College of Engineering
More informationExperimental investigation on the supersonic jet impingement
Indian Journal of Engineering & Materials Sciences Vol. 11, April 2004, pp. 100-106 Experimental investigation on the supersonic jet impingement P M Ghanegaonkar a, V Ramanujachari b & S Vijaykant b a
More informationHigh Swept-back Delta Wing Flow
Advanced Materials Research Submitted: 2014-06-25 ISSN: 1662-8985, Vol. 1016, pp 377-382 Accepted: 2014-06-25 doi:10.4028/www.scientific.net/amr.1016.377 Online: 2014-08-28 2014 Trans Tech Publications,
More informationFigure 1 Schematic of opposing air bearing concept
Theoretical Analysis of Opposing Air Bearing Concept This concept utilizes air bearings to constrain five degrees of freedom of the optic as shown in the figure below. Three pairs of inherently compensated
More informationCFD ANALYSIS TO INVESTIGATE THE EFFECT OF AXIAL SPACING IN A SINGLE STAGE TRANSONIC AXIAL FLOW COMPRESSOR
Symposium on Applied Aerodynamics and Design of Aerospace Vehicle (SAROD 2011) November 16-18, 2011, Bangalore, India CFD ANALYSIS TO INVESTIGATE THE EFFECT OF AXIAL SPACING IN A SINGLE STAGE TRANSONIC
More informationAn Impeller Blade Analysis of Centrifugal Gas Compressor Using CFD
An Impeller Blade Analysis of Centrifugal Gas Compressor Using CFD Vivek V. Kulkarni Department of Mechanical Engineering KLS Gogte Institute of Technology, Belagavi, Karnataka Dr. Anil T.R. Department
More informationAERODYNAMIC CHARACTERISTICS OF NACA 0012 AIRFOIL SECTION AT DIFFERENT ANGLES OF ATTACK
AERODYNAMIC CHARACTERISTICS OF NACA 0012 AIRFOIL SECTION AT DIFFERENT ANGLES OF ATTACK SUPREETH NARASIMHAMURTHY GRADUATE STUDENT 1327291 Table of Contents 1) Introduction...1 2) Methodology.3 3) Results...5
More informationCERTIFICATES OF COMPETENCY IN THE MERCHANT NAVY MARINE ENGINEER OFFICER
CERTIFICATES OF COMPETENCY IN THE MERCHANT NAVY MARINE ENGINEER OFFICER EXAMINATIONS ADMINISTERED BY THE SCOTTISH QUALIFICATIONS AUTHORITY ON BEHALF OF THE MARITIME AND COASTGUARD AGENCY STCW 95 CHIEF
More informationNumerical Fluid Analysis of a Variable Geometry Compressor for Use in a Turbocharger
Special Issue Turbocharging Technologies 15 Research Report Numerical Fluid Analysis of a Variable Geometry Compressor for Use in a Turbocharger Yuji Iwakiri, Hiroshi Uchida Abstract A numerical fluid
More informationJ. Szantyr Lecture No. 21 Aerodynamics of the lifting foils Lifting foils are important parts of many products of contemporary technology.
J. Szantyr Lecture No. 21 Aerodynamics of the lifting foils Lifting foils are important parts of many products of contemporary technology. < Helicopters Aircraft Gliders Sails > < Keels and rudders Hydrofoils
More informationNumerical Analysis of the Tip Leakage Flow Field in a Transonic Axial Compressor with Circumferential Casing Treatment
Numerical Analysis of the Tip Leakage Flow Field in a Transonic Axial Compressor with Circumferential Casing Treatment G. Legras 1, N. Gourdain 2 and I. Trebinjac 1 1. LMFA Ecole Centrale de Lyon / Université
More informationEXPERIMENTAL STUDY OF HOT INERT GAS JET IGNITION OF HYDROGEN-OXYGEN MIXTURE
EXPERIMENTAL STUDY OF HOT INERT GAS JET IGNITION OF HYDROGEN-OXYGEN MIXTURE Elhsnawi, M. and Teodorczyk, A. Warsaw University of Technology, ITC, Nowowiejska 21/25, -665 Warszawa, Poland ABSTRACT Experiments
More informationQuantification of the Effects of Turbulence in Wind on the Flutter Stability of Suspension Bridges
Quantification of the Effects of Turbulence in Wind on the Flutter Stability of Suspension Bridges T. Abbas 1 and G. Morgenthal 2 1 PhD candidate, Graduate College 1462, Department of Civil Engineering,
More informationVapor Recovery from Condensate Storage Tanks Using Gas Ejector Technology Palash K. Saha 1 and Mahbubur Rahman 2
37 Journal of Chemical Engineering, IEB Vapor Recovery from Condensate Storage Tanks Using Gas Ejector Technology Palash K. Saha 1 and Mahbubur Rahman 2 Abstract 1 Bibyana Gas Field 2 Department of Petroleum
More informationInfluence of Ambient Temperature on Performance of a Joule-Thomson Refrigerator
Influence of Ambient Temperature on Performance of a Joule-Thomson Refrigerator Y. J. Hong, S.J. Park, J. Ko, H.B. Kim Korea Institute of Machinery & Materials Daejeon, 305-343, Korea ABSTRACT Miniature
More informationPERFORMANCE OF A FLAPPED DUCT EXHAUSTING INTO A COMPRESSIBLE EXTERNAL FLOW
24 TH INTERNATIONAL CONGRESS OF THE AERONAUTICAL SCIENCES PERFORMANCE OF A FLAPPED DUCT EXHAUSTING INTO A COMPRESSIBLE EXTERNAL FLOW P. R. Pratt, J. K. Watterson, E. Benard, S. Hall School of Aeronautical
More informationInvestigation of Suction Process of Scroll Compressors
Purdue University Purdue e-pubs International Compressor Engineering Conference School of Mechanical Engineering 2006 Investigation of Suction Process of Scroll Compressors Michael M. Cui Trane Jack Sauls
More informationThermodynamics ERT 206 Properties of Pure Substance HANNA ILYANI ZULHAIMI
Thermodynamics ERT 206 Properties of Pure Substance HANNA ILYANI ZULHAIMI Outline: Pure Substance Phases of pure substance Phase change process of pure substance Saturation temperature and saturation pressure
More informationCFD Analysis of Supersonic Nozzle with Varying Divergent Profile
CFD Analysis of Supersonic Nozzle with Varying Divergent Profile Kaviya sundar #1, Thanikaivel Murugan. D *2 # UG Degree holder, B.E. Aeronautical Engineering, Jeppiaar Engineering college, Chennai, India,
More informationCitation Journal of Thermal Science, 18(4),
NAOSITE: Nagasaki University's Ac Title Author(s) Noise characteristics of centrifuga diffuser (Noise reduction by means leading tip) Murakami, Tengen; Ishida, Masahiro; Citation Journal of Thermal Science,
More informationAn Investigation of Liquid Injection in Refrigeration Screw Compressors
An Investigation of Liquid Injection in Refrigeration Screw Compressors Nikola Stosic, Ahmed Kovacevic and Ian K. Smith Centre for Positive Displacement Compressor Technology, City University, London EC1V
More informationReduction of Skin Friction Drag in Wings by Employing Riblets
Reduction of Skin Friction Drag in Wings by Employing Riblets Kousik Kumaar. R 1 Assistant Professor Department of Aeronautical Engineering Nehru Institute of Engineering and Technology Coimbatore, India
More informationAIRFOIL PROFILE OPTIMIZATION OF AN AIR SUCTION EQUIPMENT WITH AN AIR DUCT
THERMAL SCIENCE, Year 2015, Vol. 19, No. 4, pp. 1217-1222 1217 AIRFOIL PROFILE OPTIMIZATION OF AN AIR SUCTION EQUIPMENT WITH AN AIR DUCT by Li QIU a,b, Rui WANG a,*, Xiao-Dong CHEN b, and De-Peng WANG
More informationInlet Swirl on Turbocharger Compressor Performance
Inlet Swirl on Turbocharger Compressor Performance Lei Huang, Ying Liu, Hua Chen* National laboratory of Engine Turbocharging Technology, Tianjin, China *corresponding author: Tel.:+86-22-5870-7069; fax:
More informationPerformance Improvements in Boeing/AFOSR Mach 6 Quiet Wind Tunnel Based on CFD Predictions
Performance Improvements in Boeing/AFOSR Mach 6 Quiet Wind Tunnel Based on CFD Predictions Hadassah Naiman, Doyle D. Knight Rutgers University Selin Aradag US Air Force Academy Thomas J. Juliano,, Steven
More informationDesign Review Agenda
Design Review Agenda 1) Introduction, Motivation, and Previous Work a. Previous Work and Accomplishments i. Platform Launches ii. Successful Test Firings 2) More In-Depth Design Overview of the Existing
More informationA & AE 520 Background Information, Adapted from AAE334L, last revised 10-Feb-14 Page 1 1. SUPERSONIC WIND TUNNEL
A & AE 50 Background Information, Adapted from AAE334L, last revised 10-Feb-14 Page 1 1.1 BACKGROUND 1. SUPERSONIC WIND TUNNEL 1.1.1 Objectives: This handout is adapted from the one once used in AAE334L
More informationDesign of a Solid Wall Transonic Wind Tunnel
Design of a Solid Wall Transonic Wind Tunnel David Wall * Auburn University, Auburn, Alabama, 36849 A solid wall transonic wind tunnel was designed with optical access from three sides to allow for flow
More informationDesign and Analysis of a High Pressure Ratio Mixed Flow Compressor Stage
Design and Analysis of a High Pressure Ratio Mixed Flow Compressor Stage Gaurav Giri 1 and Abdul Nassar. 2 SoftInWay Turbomachinery Solutions Pvt Ltd. 70/10, Cunningham Road, Bangalore, KA, 560052, India
More informationEvaluation of Cavitation in a Liquid-Liquid Ejector
ILASS Americas, 24 th Annual Conference on Liquid Atomization and Spray Systems, San Antonio, TX, May 2012 Evaluation of Cavitation in a Liquid-Liquid Ejector Carsten Mehring * Central Engineering, Parker
More informationFlow and Mixing in the Liquid between Bubbles
Excerpt from the Proceedings of the COMSOL Conference 2009 Boston Flow and Mixing in the Liquid between Bubbles Bruce A. Finlayson, Professor Emeritus of Chemical Engineering Department of Chemical Engineering,
More information(Refer Slide Time: 2:16)
Fluid Machines. Professor Sankar Kumar Som. Department Of Mechanical Engineering. Indian Institute Of Technology Kharagpur. Lecture-23. Diffuser and Cavitation. Good morning and welcome you all to this
More informationThe Usage of Propeller Tunnels For Higher Efficiency and Lower Vibration. M. Burak Şamşul
The Usage of Propeller Tunnels For Higher Efficiency and Lower Vibration M. Burak Şamşul ITU AYOC 2014 - Milper Pervane Teknolojileri Company Profile MILPER is established in 2011 as a Research and Development
More informationCFD Simulation of the Flow Through Reciprocating Compressor Self-Acting Valves
Purdue University Purdue e-pubs International Compressor Engineering Conference School of Mechanical Engineering 1994 CFD Simulation of the Flow Through Reciprocating Compressor Self-Acting Valves P. Cyklis
More informationCover Page for Lab Report Group Portion. Compressible Flow in a Converging-Diverging Nozzle
Cover Page for Lab Report Group Portion Compressible Flow in a Converging-Diverging Nozzle Prepared by Professor J. M. Cimbala, Penn State University Latest revision: Prof. Steve Lynch, 14 February 2017
More informationDevelopment of a Shock Loading Simulation Facility
Development of a Shock Loading Simulation Facility K.D. Gardner, A.G. John, and F.K. Lu Aerodynamics Research Center, Mechanical and Aerospace Engineering Department, Box 19018, University of Texas at
More informationResults and Discussion for Steady Measurements
Chapter 5 Results and Discussion for Steady Measurements 5.1 Steady Skin-Friction Measurements 5.1.1 Data Acquisition and Reduction A Labview software program was developed for the acquisition of the steady
More informationCFD Simulation and Experimental Validation of a Diaphragm Pressure Wave Generator
CFD Simulation and Experimental Validation of a Diaphragm Pressure Wave Generator T. Huang 1, A. Caughley 2, R. Young 2 and V. Chamritski 1 1 HTS-110 Ltd Lower Hutt, New Zealand 2 Industrial Research Ltd
More informationFig. 2. M.I. Yaroslavtsev, 2002
SPECIAL FEATURES OF USING N O FOR HEATING OF THE TEST GAS IN A HOT-SHOT WIND TUNNEL M.I. Yaroslavtsev Institute of Theoretical and Applied Mechanics SB RAS, 630090 Novosibirsk Russia 1. The study of heat
More informationMeasurement and simulation of the flow field around a triangular lattice meteorological mast
Measurement and simulation of the flow field around a triangular lattice meteorological mast Matthew Stickland 1, Thomas Scanlon 1, Sylvie Fabre 1, Andrew Oldroyd 2 and Detlef Kindler 3 1. Department of
More informationCFD VALIDATION STUDY OF NEXST-1 NEAR MACH 1
24 TH INTERNATIONAL CONGRESS OF THE AERONAUTICAL SCIENCES CFD VALIDATION STUDY OF NEXST-1 NEAR ACH 1 Keizo Takenaka*, Kazuomi Yamamoto**, Ryoji Takaki** *itsubishi Heavy Industries, Ltd., 10 Oye-cho, inato-ku,
More informationThe Estimation Of Compressor Performance Using A Theoretical Analysis Of The Gas Flow Through the Muffler Combined With Valve Motion
Purdue University Purdue e-pubs International Compressor Engineering Conference School of Mechanical Engineering The Estimation Of Compressor Performance Using A Theoretical Analysis Of The Gas Flow Through
More informationEXPERIMENTAL STUDY OF A STRUT INJECTOR FOR CIRCULAR SCRAMJET COMBUSTORS
EXPERIMENTAL STUDY OF A STRUT INJECTOR FOR CIRCULAR SCRAMJET COMBUSTORS Christopher Rock Graduate Research Assistant and Joseph A. Schetz Advisor, Holder of the Fred D. Durham Chair Department of Aerospace
More informationSingle Phase Pressure Drop and Flow Distribution in Brazed Plate Heat Exchangers
Purdue University Purdue e-pubs International Refrigeration and Air Conditioning Conference School of Mechanical Engineering 2016 Single Phase Pressure Drop and Flow Distribution in Brazed Plate Heat Exchangers
More informationTwo phase discharge flow prediction in safety valves
Dempster W, Elmayyah W/ ICPVT-13 1 Two phase discharge flow prediction in safety valves William Dempster a, Wael Elmayyah b a Department of Mechanical and Aerospace Engineering, University of Strathclyde,
More informationUnsteady airfoil experiments
Unsteady airfoil experiments M.F. Platzer & K.D. Jones AeroHydro Research & Technology Associates, Pebble Beach, CA, USA. Abstract This paper describes experiments that elucidate the dynamic stall phenomenon
More informationDevice Description. Operating Information. CP Q (eq. 1) GT. Technical Bulletin TB-0607-CFP Hawkeye Industries Critical Flow Prover
A compressible fluid traveling at subsonic velocity through a duct of constant cross section will increase velocity when passing through a region of reduced cross-sectional area (in this case, an orifice)
More informationCFD SIMULATIONS OF GAS DISPERSION IN VENTILATED ROOMS
CFD SIMULATIONS OF GAS DISPERSION IN VENTILATED ROOMS T. Gélain, C. Prévost Institut de Radioprotection et de Sûreté Nucléaire (IRSN), Saclay, France Abstract In order to better understand the risks due
More informationSUPERSONIC GAS TECHNOLOGIES
SUPERSONIC GAS TECHNOLOGIES Vladimir Feygin, Salavat Imayev, Vadim Alfyorov, Lev Bagirov, Leonard Dmitriev, John Lacey TransLang Technologies Ltd., Calgary, Canada 1. INTRODUCTION The 3S technology is
More informationAERODYNAMIC CHARACTERISTICS OF SPIN PHENOMENON FOR DELTA WING
ICAS 2002 CONGRESS AERODYNAMIC CHARACTERISTICS OF SPIN PHENOMENON FOR DELTA WING Yoshiaki NAKAMURA (nakamura@nuae.nagoya-u.ac.jp) Takafumi YAMADA (yamada@nuae.nagoya-u.ac.jp) Department of Aerospace Engineering,
More informationLOW PRESSURE EFFUSION OF GASES adapted by Luke Hanley and Mike Trenary
ADH 1/7/014 LOW PRESSURE EFFUSION OF GASES adapted by Luke Hanley and Mike Trenary This experiment will introduce you to the kinetic properties of low-pressure gases. You will make observations on the
More informationChapter 2: Pure Substances a) Phase Change, Property Tables and Diagrams
Chapter 2: Pure Substances a) Phase Change, Property Tables and Diagrams In this chapter we consider the property values and relationships of a pure substance (such as water) which can exist in three phases
More informationOPTIMIZATION OF SINGLE STAGE AXIAL FLOW COMPRESSOR FOR DIFFERENT ROTATIONAL SPEED USING CFD
http:// OPTIMIZATION OF SINGLE STAGE AXIAL FLOW COMPRESSOR FOR DIFFERENT ROTATIONAL SPEED USING CFD Anand Kumar S malipatil 1, Anantharaja M.H 2 1,2 Department of Thermal Power Engineering, VTU-RO Gulbarga,
More informationEffect of Inlet Clearance Gap on the Performance of an Industrial Centrifugal Blower with Parallel Wall Volute
International Journal of Fluid Machinery and Systems DOI: http://dx.doi.org/10.5293/ijfms.2013.6.3.113 Vol. 6, No. 3, July-September 2013 ISSN (Online): 1882-9554 Original Paper (Invited) Effect of Inlet
More informationENGR 292 Fluids and Thermodynamics
ENGR 292 Fluids and Thermodynamics Scott Li, Ph.D., P.Eng. Mechanical Engineering Technology Camosun College Pure Substances Phase-Change Process of Pure Substances Specific Volume Saturation Temperature
More informationMODELLING PIPELINE DECOMPRESSION DURING THE PROPAGATION OF A DUCTILE FRACTURE
MODELLING PIPELINE DECOMPRESSION DURING THE PROPAGATION OF A DUCTILE FRACTURE R.P. Cleaver and P.S. Cumber BG Technology Gas Research and Technology Centre Ashby Road, Loughborough, LE11 3GR There is considerable
More informationNUMERICAL INVESTIGATION OF AERODYNAMIC CHARACTERISTICS OF NACA AIRFOIL WITH A GURNEY FLAP
Int. J. Mech. Eng. & Rob. Res. 2012 MasoudJahanmorad Nouri et al., 2012 Research Paper ISSN 2278 0149 www.ijmerr.com Vol. 1, No. 3, October 2012 2012 IJMERR. All Rights Reserved NUMERICAL INVESTIGATION
More informationPERFORMANCE ANALYSIS OF CYLINDRICAL TYPE AIR EJECTOR
Journal of Engineering Sciences, Assiut University, Vol. 34, No. 3, pp. 733-745, May 2006 PERFORMANCE ANALYSIS OF CYLINDRICAL TYPE AIR EJECTOR Mechanical Power Department, High Institute of Energy, South
More informationAnother convenient term is gauge pressure, which is a pressure measured above barometric pressure.
VACUUM Theory and Applications Vacuum may be defined as the complete emptiness of a given volume. It is impossible to obtain a perfect vacuum, but it is possible to obtain a level of vacuum, defined as
More informationApplication of Simulation Technology to Mitsubishi Air Lubrication System
50 Application of Simulation Technology to Mitsubishi Air Lubrication System CHIHARU KAWAKITA *1 SHINSUKE SATO *2 TAKAHIRO OKIMOTO *2 For the development and design of the Mitsubishi Air Lubrication System
More information