Effects of Disturbances on Quiet Flow in the Mach 4 Ludwieg Tube
|
|
- Beverley Pierce
- 6 years ago
- Views:
Transcription
1 Effects of Disturbances on Quiet Flow in the Mach 4 Ludwieg Tube AAE 50 Experimental Aerodynamics Final Report Matt Borg and Justin Smith May 5, 004
2 Abstract The PQFLT was used to determine the effects of disturbances on flow quietness. Initially, normal runs were performed to ascertain a range of quiet flow Reynold s Numbers. Runs were performed for initial driver tube pressures between 6 psia and 16 psia. While none of the cases tested yielded noise levels low enough to be considered quiet flow by the conventional definition, runs at Unit Reynold s Numbers between 5000 and (driver tube pressures between 8 psia and 1 psia) had noise levels below 0.07% of the mean pressure. Disturbances were then introduced downstream of the test section by bursting the diaphragm asymmetrically. This was done for the same range of driver tube pressures as for the normal bursts. It was determined that asymmetric bursting produces no effect on noise levels. Finally, upstream disturbances were introduced through a jet at the end of the driver tube. Four runs were made at driver tube pressures of 10. psia. With each run, the strength of the jet was adjusted by changing the voltage applied to a pressure regulator controlling the jet. Regulator voltages of 0. V, 0.4 V, 0.6 V, and 0.8 V were used for the four runs and a distinguishable difference in rms noise level between normal and jet bursts was established. Particularly, the increased noise level was found to vary linearly with jet strength (voltage), showing increases from about 0 % at the 0. V level up to about 40% for the 0.8 V setting. Introduction Although supersonic wind tunnels are used in a variety of aerodynamic applications, often the data collected in them cannot be correlated to flight data. Supersonic vehicles are usually exposed to laminar free stream flow. Such flow cannot be properly replicated in typical supersonic wind tunnels. Conventional tunnels have turbulent boundary layers, which radiate noise onto the test object. This noise changes the laminar-turbulent transition Reynold s number on the object, and makes correlation to real data difficult or impossible [1], The noise level in a quiet wind tunnel is on the order of those experienced in flight, ~ 0.05% of mean data [8], typically an order of magnitude less than that in conventional tunnels Quiet supersonic wind tunnels allow proper study of high-speed boundary layers and thus more accurate prediction and control of boundary layer properties such as skin friction and heat transfer. Such flow promotes transition on a model at a Reynold s number comparable to flight []. RMS noise is typically measured as pitot-pressure fluctuations with fast pressure transducers or as massflow fluctuations with hot-wires [1]. Quiet flow is hard to achieve because the mechanisms of transition are many and poorly understood [3]. Below Mach 4, first mode instabilities dominate the flow and initiate transition. The second mode of instability remains stable []. Three modes of disturbance fields in compressible, viscous, and heat - conductive gases have been identified: vorticity mode, entropy mode, and sound-wave mode. The three modes are non-interacting for small fluctuations, but do interact at larger intensities [4].
3 There are several criteria which a quiet facility must meet. There must be a supply of clean, dry air. The tunnel must also minimize disturbances in the settling chamber. Additionally, the nozzle must be smooth, so as to promote the starting of the tunnel and a laminar boundary layer in the nozzle. It is known that the absence of such criteria in conventional supersonic tunnels promotes earlier transition of the boundary layer from laminar to turbulent and thus precludes quiet flow []. The region of quiet flow in a quiet facility is called the test core. This region is defined downstream of the test core by Mach lines traced to the acoustic origin of turbulence in the nozzle-wall boundary layer. The upstream boundary of the test core is defined by Mach lines due to the attainment of the design Mach number for the tunnel. The size of the test core is influenced by free stream Reynold s number. As it increases, the boundary layer transition region moves upstream and the Mach lines from this disturbance move upstream as well, decreasing the test core size []. The PQFLT The Purdue Mach 4 Quiet Flow Ludwieg Tube (PQFLT) was developed between 199 and This facility is a short duration wind tunnel consisting of a driver tube with a nozzle on the end from which flow exits into a test section and a diffuser [5]. The driver tube is 68 feet long with a 1 in diameter. A smooth contraction runs from the driver tube to the first throat (1.368 sq. in with an area ratio of 83 for the driver tube), followed by a 3.8 x 4.3 in Mach 4 rectangular nozzle. The nozzle was fabricated at NASA Langley in the 1970 s, and was donated to Purdue in Although not designed as a quiet nozzle, it was used to achieve quiet flow in the Purdue Ludwieg Tube [6]. Quiet flow has been achieved in the PQFLT for Reynold s numbers up to about based on the length of the quiet-flow test core (approximately 4 in). This corresponds to stagnation pressures up to about 14 psia []. After achieving quiet flow to low Reynold s numbers, the facility was used to develop instrumentation, as well as to study instability waves in quiet flow [7]. A Mylar diaphragm is used to initially isolate pressurized air in the driver tube from vacuum downstream of the test section. Wires are attached to the diaphragm. It is burst by running a large current through these wires to melt the diaphragm at the points of contact. The thickness of Mylar used for the burst diaphragms depends on the initial driver tube pressure. In general, diaphragm thicknesses of.004 in,.007 in, and.010 in are used for driver tube pressures of kpa, kpa, and kpa, respectively [5]. When the diaphragm bursts, a shock wave travels downstream and an expansion wave travels upstream through the test section and into the driver tube. The expansion wave propagates through the driver tube and induces flow through the nozzle. The wave is reflected back through the nozzle from the closed end of the driver tube (with a round trip time of about 10 msec [6]). Additional expansion waves propagate into the driver tube until the stagnation pressure is too low to maintain supersonic flow in the nozzle and test section. A typical run lasts between 3 and 5 seconds [5]. Between expansion wave cycles, the flow is in a quasi-static state. Calculations are thus
4 based on data collected between wave fronts. Motivation It was desired to measure the effects of disturbances on the quiet flow obtained in the Purdue Quiet Flow Ludwieg Tube (PQFLT). An effective range of unit Reynold s numbers providing quiet flow needed to first be established. Additionally, a lower threshold unit Reynold s number still allowing supersonic flow was measured. Disturbances were then introduced into the flow in two forms. It was desired to determine whether downstream disturbances would travel upstream through the subsonic portion of the boundary layer. Thus, the first disturbance investigated was downstream of the quiet test core. This disturbance was formed by bursting a diaphragm asymmetrically. This created a fairly large downstream obstruction to flow, and thus the desired disturbance. Additionally, this study sought to determine whether certain upstream disturbances could preclude quiet flow. An air jet of varying intensity was introduced far upstream of the throat at the back of the driver tube. Since quiet wind tunnel facilities are essential for the further development and testing of hypersonic vehicles, an understanding of quiet facilities is necessary to continue this development. As has been demonstrated, there are a number of factors in any wind tunnel that could preclude quiet conditions. It is desired to ascertain the effects of disturbances in a facility where quiet flow has been consistently demonstrated. The PQFLT facility provided an excellent facility for this study. Procedure Before the tunnel could be run, a cone and sting mount were removed from the test section and a hot wire probe inserted. After cleaning the tunnel and ensuring that no particulate had settled into it, the test section was closed. After several runs had been completed collecting data from the hot wire probe and a Kulite static pressure transducer on the side wall of the tunnel, it was determined that these instruments would not indicate whether the flow was quiet or noisy. The hot wire probe was removed, and a Kulite pitot probe was inserted into the flow. It was placed further forward in the nozzle, near the center of the first window, than the hot wire had been to ensure that it was in the test core of quiet flow. A Mylar diaphragm was then prepared. This involved tracing out a new diaphragm onto a sheet of Mylar, cutting it, and punching holes in it for bolts to go through. The new diaphragm was then inserted into the diaphragm rings. After the bolts were handtightened sufficiently, Nichrome wires were taped across the diaphragm at right angles to each other. At the intersection point, a small loop was made so that when the wires were heated they would not melt through the tape and short each other out. The wire setup can be seen in Figure 1. The diaphragm ring was then inserted into the diaphragm section of the tunnel. The two wire leads from a capacitor bank at the bottom of the diaphragm section were attached to the diaphragm rings. The tunnel was then closed by operating a small hand pump, and sealed by tightening a flange clamp.
5 Figure 1: Diaphragm wire setup Two vacuum pumps were then turned on so as to evacuate the 535 cubic foot vacuum tank on the downstream end of the tunnel. One of two valves were opened which either added filtered, dry air to the driver tube, or applied a vacuum to the driver tube. Using these valves, the driver tube was brought to the desired initial stagnation pressure. After the vacuum tank reached the desired back pressure, -4 torr, the run was initiated. In order to start a run, a bank of capacitors was charged to approximately 40 Volts. A circuit was then closed with an electronic switch, enabling the capacitors to dump their charge into the Nichrome wires in contact with the diaphragm. The wires dissipated approximately 10 Watts. This heated them very quickly and melted the Mylar with which they were in contact. The upstream pressure then burst the diaphragm and the run began. The signal output by the Kulite pitot sensor was run through a circuit that amplified the DC signal by a factor of 100. On a second output, the signal was high-pass filtered the signal to isolate the AC component of the signal and amplified it by a factor of Both of these signals were read by a LeCroy 9104 digital oscilloscope. After the approximately 3.5 second run was finished, a computer was used to save data via a GPIB card for later reduction. The tunnel and vacuum tank were brought up to atmospheric pressure by fully opening the compressed air valve so the tunnel could be opened. After data were saved, the diaphragm section was opened. The used diaphragm rings were removed, and a new diaphragm setup put in its place for the next run. The diaphragms typically petaled into four quadrants. A picture of a typical burst diaphragm can be seen in Figure. Figure : A typical symmetrically burst diaphragm. The pitot probe was then calibrated by pressurizing the driver tube to a known pressure, read from a Paroscientific A pressure gauge, and recording the corresponding output voltage measured by an HP 34401A digital multimeter. The first time the calibration was completed, the lowest driver tube pressure used in the calibration was approximately 1.8 psia. It was then decided that the pitot probe needed to be calibrated to lower pressures because the measured pitot
6 pressures during a run were near this lowest calibrated point. There was also a curious deviation from linearity for the three lowest points on the calibration curve. It was necessary to determine whether the linearity of the voltagepressure relationship broke down at sufficiently low pressures. Thus, a second calibration to lower driver tube pressure, with more data points taken was completed. The deviation from linearity in the case of the second calibration was determined to be too high. Upon attempting a third calibration, it was found that the Paroscientific pressure gauge would not read lower than approximately 1. psia. After connecting the pressure gauge to a different pressure vessel and comparing with a separate pressure gauge, it was determined that something was wrong with the setup of the pressure gauge on the PQFLT facility. The pressure gauge matched almost perfectly in comparison with the other. The Paroscientific was then reconnected to the PQFLT, and the driver tube was pressurized to approximately 1 psia. A bubble leak test confirmed a leak in a connection from the tunnel to the Paroscientific. The connection was tightened, and the leak was successfully stopped. A third pitot calibration was performed, both at high and low pressures. Runs were made with driver tube pressures of approximately 15., 14., 14.1, 13., 1., 11.3, 10.7, 10.1, 9.3, 7.9, 7., and 6.3 psia to determine the maximum driver tube pressure allowing quiet flow in the nozzle, and the minimum pressure allowing supersonic flow. After attempting to run the tunnel to a driver tube pressure of lower than 6.0 psia, it was discovered that there was a significant leak that did not permit driver tube pressures below 5.5 psia. Even with the vacuum line completely open, both vacuum tanks running, and the vacuum tank in the tens of torr range, the tunnel pressure would not drop below approximately 5.5 psia. The tunnel was then pressurized to approximately 3 psia, and a bubble leak test was performed. It was suspected that the leak was from around the windows. The bubble test showed that the windows were air tight. Two leaks were found in other locations, however. One was plugged with a bolt, in place of the RTV that had inadequately sealed before. The second leak was plugged by applying more RTV. A picture of the successful bubble test can be seen in Figure 3. Figure 3: Successful bubble leak test. It was then desired to ascertain whether a disturbance downstream of the test core would propagate upstream through the subsonic region of the boundary layer and cause it to transition from laminar to turbulent, thus upsetting quiet flow. The tunnel was run in the same manner, but the diaphragms were prepared to burst asymmetrically. The diaphragms were burst in such a way that the bottom half
7 remained intact and disturbed the flow. This was accomplished by insulating the bottom half of the vertical Nichrome wire from the Mylar by placing a piece of masking tape between the wire and the Mylar. This method generally worked quite well. A picture of a typical asymmetrically burst diaphragm can been seen in Figure 4. Asymmetric bursts were made at initial driver tube stagnation pressures of 6.8, 10.0, 10.3, 10.5, 10.7, 11., and1. psia. After observing the effects of downstream noise on quiet flow conditions, a disturbance in the form of an air jet was introduced upstream of the test section in the far end of the driver tube. The jet strength was controlled by varying the voltage on the pressure regulator that supplied air to the end of the driver tube. All jet runs were made at an initial driver tube stagnation pressure of approximately 10. psia as that was near the lowest % RMS value achieved during normal tunnel operations. Runs were made with the pressure regulator voltage set to.,.4,.6, and.8 volts. Above that voltage, an overpressure relief valve was tripped, which disallowed higher pressure regulator voltages to be used. Figure 4: Asymmetrically burst diaphragm Theory Pitot Probe Figure 5. Nozzle of the PQFLT (not to scale).
8 A diagram of the nozzle of the PQFLT adapted from [8] is given in Figure 5. Upstream of the test core, the stagnation pressure, p 0,1, is measured by the Paroscientific gauge, and the stagnation temperature, T 0, is measured with a thermocouple at the end of the driver tube. The quiet-flow region is bounded upstream by Mach lines that emanate from the onset of uniform flow, and downstream by Mach lines which radiate from the nozzle walls as the boundary layer transitions from laminar to turbulent []. These boundaries define the quiet-flow test core in the nozzle, prior to the constant-area test section. Stagnation pressure in the quiet flow test core was measured just behind the center of the first window in the tunnel (~1/4" behind center). This location was selected because it is well within the quiet flow core [8]. Since the flow is supersonic in the region where stagnation pressure measurements are made, a bow shock develops in front of the pitot tube, Figure 6. M>1 p 0,1 p 1 Shock p 0,, Pitot Probe Figure 6. Bow shock in front of a pitot tube in supersonic flow. The presence of a shock in front of the Kulite precludes the utility of Bernoulli's equation in determining the velocity, and ultimately the Mach Number of the incoming flow. A normal shock relation, Rayleigh's Pitot Formula [9] must then be used to find an implicit relation for the test-core Mach Number, M, provided that the stagnation pressure behind the shock, p 0, and the driver tube static pressure, p 1, are known, Equation (1). γ p γ 0, M p 1 = 4γ 1 ( γ + 1) M 1 γ + γ M ( γ 1) γ + 1 (1) The driver tube static pressure is obtained from the isentropic relation [10] of Equation () and must be solved simultaneously with Equation (1). This is done iteratively, i.e. a guess is made for M, and plugged into both Equations (1) and (). Dividing the two to get p 0, / p 0,1, this answer is then checked against the measured ratio. A new guess is then made on M and the process is repeated until the desired accuracy is reached. For this analysis, M was solved to three significant figures. p 1 p0,1 1 M γ γ 1 γ 1 = + () Unit Reynold s Numbers are calculated from Equation (3). ρ Re = U 1 (3) µ For this analysis, density is given by the isentropic relation [3], 1 γ 1 γ 1 ρ = ρ0 1+ M (4) where the driver tube density, ρ 0, is obtained from p 0,1, T 0, and the ideal gas law (with R=87 J/Kg/K for air at STP).
9 Dynamic viscosity, µ, was taken from Sutherland's Law [], run, the prerun fluctuations are smaller than those during the run. T + S T 3 / ref µ = µ ref T + S T (5) ref A For air, µ ref = x10-5 kg/(m.s) for T ref = K, and S = 110 K. Finally, the flow speed, U, is calculated from the speed of sound, c, in the following manner: U = cm = γrtm (6) with temperature computed from an additional isentropic relation [3]: Pressure [psia] B Time [s] Figure 7. Stagnation pressure in the test core for an initial driver tube pressure of 8.6 psia. C D Pressure [kpa] 1 γ 1 T = T0 1+ M (7) Quiet Flow Measurements A typical DC pitot trace for an initial driver tube pressure of 8.6 psia is given in Figure 3. The accompanying AC trace follows in Figure 8. Data were collected for 5 seconds at 50 khz. The oscilloscope record of the data consists of four main sections. These sections, labeled in Figure 7, are A Prerun B Startup C Expansion Wave Cycles D Subsonic Flow The prerun portion of the DC trace gives the driver tube pressure before the run has started. The AC data during the prerun is used to determine the electronic noise level when the tunnel is not running, the prerun noise. Notice from Figure 8 that for a typical Figure 8. AC pitot data for an initial driver tube pressure of 8.6 psia. During the startup, pressure fluctuates greatly as uniform supersonic flow begins to establish itself in the test core. No useful data can be taken during this time. Following the startup, after quiet flow has been established, expansion waves emanate through the driver tube and reflect back towards the test core. During each wave cycle, quasistatic conditions of uniform supersonic flow exist within the test core. It is during the first expansion wave cycle that
10 measurements of "flow quietness" are obtained. RMS pressure fluctuations are obtained from Equation (8): p rms ( p p ) = (8) ac ac In this equation, symbols with bars denote average quantities. The condition for quiet flow is [11] p rms p p dc rms,prerun (9) In this calculation, the prerun noise is subtracted so that only fluctuations inherent to the tunnel operation are considered. After approximately 30 expansion wave cycles have passed, the driver tube pressure is no longer great enough to sustain supersonic flow. This event is marked by a dramatic increase in rms pressure fluctuation because the turbulent subsonic boundary layer radiates noise into the test core. the leak pressure was not negligible and skewed the results. Since for higher pressures the calibration remained the same, all initial stagnation pressure data recorded in previous runs was still accurate and usable. A linear fit of the third data set was used to determine the following formula for conversion from Kulite voltage to pressure: p 0,1 = V (10) A sample calculation for a run with an initial driver tube pressure of 11.3 psia (77910 kpa) and temperature of 74.5 o F (96.8 K) is now given. For this run, the diaphragm was burst normally and data were sampled at 500 khz for 0.5 seconds. The collected data include 0.1 seconds of prerun and about 3 1/ expansion cycles of same duration. Figure 10 shows both the AC and DC pressure traces. Results Several calibrations of the pitot probe were performed and are shown in Figure 9. As can be seen, all three calibrations matched quite well for driver tube pressures above 4 psia. Below 4 psia, the third calibration deviated from the other two, but remained linear with the higher pressure data. This sort of behavior is to be expected. Since the leak was so small, at higher pressures, the leak pressure difference was very small compared to the mean pressure. It thus had a negligible effect. When the mean pressure dropped sufficiently, however, Figure 9. Kulite pressure transducer calibration curves in the PQFLT.
11 Plugging this and the Mach Number into Equation (6), U = 1.4 = 669 m/s ( 87 J / Kg K)( 74 K)( 3.88) Finally, viscosity is obtained from Equation (5) with variables for air: Figure 10. Sample pressure data for a driver tube pressure of 11.3 psia. For the first pressure step after startup, the mean pressure was computed to be p 0,1 = 1.73 psia. Using Equations (1) and (), this corresponds to a Mach Number of From the ideal gas law, the stagnation density is computed in the following manner: ρ 0 p = RT 0,1 0 = Pa ( 87J/Kg K)( 96.8K ) 3 = 0.915kg / m Then, from Equation (4), ρ = 0.915kg / m 1+ = kg/m 3 ( 3.88) From Equation (7), with T 0 = 96.8 K, T = 96.8 K 1+ = 74 K ( 3.88) 1 µ = 3/ kg/m s = 5.039x10-6 kg/m.s Plugging all of these values into Equation (3), the unit Reynold's Number is obtained for this run: 3 ( 0.084kg / m )( 669m / s) Re1 = kg / m s = /cm From Equation (8), the rms pressure fluctuation was computed to be , while the prerun rms was Using these values and the mean stagnation pressure on the first step, Equation (9) yields a fractional rms pressure fluctuation of: = This is a noise level of % of the mean signal. While this level is typically an order of magnitude below that of conventional tunnels, it still does not meet the requirements of Equation (9) for a quiet run. Figures 11 and 1 summarize Unit Reynold s Number and Mach Number calculations for all tunnel runs, respectively. Reynold s Number is shown to vary linearly with initial driver
12 tube pressure. A range of to was found for operation with pressures from 6 psia to 16 psia. Figure 1. Mach Number vs. driver tube pressure. Figure 11. Unit Reynold s Number vs. driver tube pressure. Mach Numbers around 3.9 were obtained for the entire range of pressures examined. Higher Mach Numbers were obtained at higher driver tube pressures. This is expected because the boundary layer is thinner at the point of pressure measurement. Thus the expansion ratio in the nozzle is effectively larger and the Mach Number increases as well. Near driver tube pressures of 6-7 psia, three seemingly anomalous Mach Numbers were found. These points were also found to have the highest noise levels computed. For these runs, the first pressure steps were around 0.7 psia. It is unknown why these points differ from expected values. It is doubtful that the flow dropped to subsonic because the pressure drop was larger than would be expected in subsonic flow, and the oscilloscope trace showed clearly defined expansion wave steps. Runs at these conditions should be repeated to ensure that these results are anomalous. Flow noise levels for all of the runs performed in the PQFLT are summarized in Figure 13. These runs include normal bursts, asymmetric burst, jet bursts, and combinations of asymmetric and jet bursts. From the figure, it can be seen that none of the normal burst runs have rms noise levels below 0.05% of the mean. However, a relatively wide range of driver tube pressures (~8-1 psia) has noise levels below 0.07%. This result is consistent with [11]. This pressure range corresponds to Reynold s Numbers between 5000 and Figure 13. Flow noise levels in the PQFLT. The horizontal scale is Reynold s Number per cm.
13 It can be seen that asymmetric diaphragm bursting has no effect on noise levels. If anything, asymmetric bursting decreases the noise level near the optimum Reynold s Number (i.e. near the Reynold s Numbers with lowest noise levels for normal bursting). The application of jets increases noise levels by approximately 30-40% on average. This is evident from Figure 14, as jet runs were performed at Reynold s Numbers around 35000, p 0,1 ~ 10. psia. For these runs, the jet pressure was varied with a regulator. Four settings were used from 0. V to 0.8 V. For regulator voltages higher than this, the relief valve would no longer let air into the tunnel. A plot of the trend in rms noise vs. regulator voltage is shown in Figure 14. This indicates an approximate linear relationship between voltage and noise level, but more settings would need to be used to verify this. Conclusions 1) Quiet flow, according to convention (<.05%), was not achieved. This is possibly because the nozzle was slightly dirty and accumulated particulate increased noise levels. However, noise levels of <.07% were found for a significant range of Unit Reynold s Numbers, keeping with previous results. ) Downstream disturbances in the form of asymmetrically burst diaphragms did not propagate upstream with sufficient strength to trip the boundary layer to early turbulent transition. The disturbance caused by the asymmetric burst was damped out by the time it reached the test core. 3) Upstream disturbances caused by jets introduced into the plenum (driver tube) increased the RMS percent noise linearly with voltage applied to the pressure regulator. Thus, upstream jet disturbances do effectively cause the boundary layer to transition earlier and to radiate noise into the test core. 4) Quiet flow is not influenced as heavily by downstream disturbances as upstream. Upstream jets have a strong effect on tunnel noise level. Nozzle dirtiness (accumulate particulate) also influences boundary layer transition. Figure 14. RMS Noise vs. Regulator Voltage for jet runs in the PQFLT. 5) It is suspected that the lowest unit Reynold s Number providing supersonic flow is approximately 15000/cm as RMS noise increased markedly near this value. Schlieren imaging would help to ascertain whether this is the case.
14 References [1] Schneider, S. P. and Skoch, C. 001: Mean Flow and Noise Measurements in the Purdue Mach-6 Quiet-Flow Ludwieg Tube, AIAA [] Haven, C. E., 1995: Measurements of Laminar-Turbulent Transition on Supersonic Wind-Tunnel Walls, MS Thesis, Purdue University, West Lafayette, IN. [3] Blaisdell, G. 003: Class Notes: Introduction to Aerodynamics, Purdue University, West Lafayette, IN. [4] Kovasznay, L. 1953: Turbulence in Supersonic Flow, Journal of the Aeronautical Sciences, 0(10), [5] Ladoon, D. W., 1998: Wave Packets Generated by a Surface Glow Discharge on a Cone at Mach 4, Ph D Dissertation, Purdue University, West Lafayette, IN. [6] Schneider, S. P., Haven, E. 1994: Mean Flow and Noise Measurements in the Purdue Quiet-Flow Ludwieg Tube, AIAA [7] Schneider, S. P., Skoch, C., Rufer, S., Swanson, E., Borg, M. 004: Bypass Transition on the Nozzle Wall of the Boeing/AFOSR Mach-6 Quiet Tunnel, AIAA [8] Salyer, T. R., Laser Differential Interferometry for Supersonic Blunt Body Receptivity Experiments, Dissertation, Purdue University, West Lafayette, 00, pp. 11. [9] Anderson, J. D., Fundamentals of Aerodynamics, 3 rd Edition, McGraw- Hill, New York, 001, pps. 448, [10] Currie, I. G., Fundamental Mechanics of Fluids, 3 rd Edition, Marcel Dekker, New York, 003, pp [11] Schneider, S. P., and Haven, C. E., Quiet-Flow Ludwieg Tube for High- Speed Transition Research, AIAA Journal, Vol. 33, No. 4, 1995, pp
Refining the Transonic Capability of Purdue s 12-inch Ludwieg Tube. AAE520 Project Report Draft April 16, Lynn Hendricks Matt Schmitt
Refining the Transonic Capability of Purdue s 12-inch Ludwieg Tube AAE520 Project Report Draft April 16, 2007 Lynn Hendricks Matt Schmitt School of Aeronautics and Astronautics Purdue University Abstract
More informationLaminar-Turbulent Transition Measurements in the Boeing/AFOSR Mach-6 Quiet Tunnel
Laminar-Turbulent Transition Measurements in the Boeing/AFOSR Mach-6 Quiet Tunnel Steven P. Schneider and Thomas J. Juliano School of Aeronautics and Astronautics Purdue University West Lafayette, IN 47907-18
More informationFundamentals of Gas Dynamics (NOC16 - ME05) Assignment - 8 : Solutions
Fundamentals of Gas Dynamics (NOC16 - ME05) Assignment - 8 : Solutions Manjul Sharma & Aswathy Nair K. Department of Aerospace Engineering IIT Madras April 5, 016 (Note : The solutions discussed below
More informationAEROSPACE ENGINEERING DEPARTMENT. Second Year - Second Term ( ) Fluid Mechanics & Gas Dynamics
AEROSPACE ENGINEERING DEPARTMENT Second Year - Second Term (2008-2009) Fluid Mechanics & Gas Dynamics Similitude,Dimensional Analysis &Modeling (1) [7.2R*] Some common variables in fluid mechanics include:
More informationTransition Research with Temperature-Sensitive Paints in the Boeing/AFOSR Mach-6 Quiet Tunnel
1st AIAA Fluid Dynamics Conference and Exhibit 7-3 June 11, Honolulu, Hawaii AIAA 11-387 Transition Research with Temperature-Sensitive Paints in the Boeing/AFOSR Mach-6 Quiet Tunnel Amanda Chou, Christopher
More informationIn which of the following scenarios is applying the following form of Bernoulli s equation: steady, inviscid, uniform stream of water. Ma = 0.
bernoulli_11 In which of the following scenarios is applying the following form of Bernoulli s equation: p V z constant! g + g + = from point 1 to point valid? a. 1 stagnant column of water steady, inviscid,
More informationOptical and Pressure Measurements of Tunnel Noise in the Ludwieg Tube
California Institute of Technology Optical and Pressure Measurements of Tunnel Noise in the Ludwieg Tube Ae104c Project Thomas Vezin Stacy Levine Neal Bitter June 3, 2011 Abstract Caltech s Ludwieg tube,
More information6.1 According to Handbook of Chemistry and Physics the composition of air is
6. Compressible flow 6.1 According to Handbook of Chemistry and Physics the composition of air is From this, compute the gas constant R for air. 6. The figure shows a, Pitot-static tube used for velocity
More informationPlease welcome for any correction or misprint in the entire manuscript and your valuable suggestions kindly mail us
Problems of Practices Of Fluid Mechanics Compressible Fluid Flow Prepared By Brij Bhooshan Asst. Professor B. S. A. College of Engg. And Technology Mathura, Uttar Pradesh, (India) Supported By: Purvi Bhooshan
More informationBoundary-Layer Instability Measurements in a Mach-6 Quiet Tunnel
Boundary-Layer Instability Measurements in a Mach- Quiet Tunnel Dennis C. Berridge, Christopher A.C. Ward, Ryan P.K. Luersen, Amanda Chou, Andrew D. Abney, and Steven P. Schneider School of Aeronautics
More informationOblique Shock Visualization and Analysis using a Supersonic Wind Tunnel
Oblique Shock Visualization and Analysis using a Supersonic Wind Tunnel Benjamin M. Sandoval 1 Arizona State University - Ira A. Fulton School of Engineering, Tempe, AZ, 85281 I. Abstract In this experiment,
More informationTHEORETICAL AND EXPERIMENTAL INVESTIGATIONS ON CHOKING PHENOMENA OF AXISYMMETRIC CONVERGENT NOZZLE FLOW
8 TH INTERNATIONAL CONGRESS OF THE AERONAUTICAL SCIENCES THEORETICAL AND EXPERIMENTAL INVESTIGATIONS ON CHOKING PHENOMENA OF AXISYMMETRIC CONVERGENT NOZZLE FLOW Ryuta ISOZUMI*, Kazunori KUBO*, Daisuke
More informationShock tunnel operation and correlation of boundary layer transition on a cone in hypervelocity flow
Shock tunnel operation and correlation of boundary layer transition on a cone in hypervelocity flow J.S. Jewell 1, J.E. Shepherd 1, and I.A. Leyva 2 1 Introduction The Caltech T reflected shock is used
More informationBoundary-Layer Transition Measurements in a Mach-6 Quiet Tunnel
th Fluid Dynamics Conference and Exhibit June - 1 July 1, Chicago, Illinois AIAA 1-71 Boundary-Layer Transition Measurements in a Mach-6 Quiet Tunnel Christopher A.C. Ward, Bradley M. Wheaton, Amanda Chou,
More informationEffect of Freestream Noise on Roughness-Induced Transition for the X-51A Forebody
46th Annual Fluid Dynamics Conference, January, Reno, NV 2008-0592 Effect of Freestream Noise on Roughness-Induced Transition for the X-51A Forebody Matthew P. Borg, Steven P. Schneider, and Thomas J.
More informationHypersonic Boundary-Layer Transition Experiments in a Mach-6 Quiet Tunnel
th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Eposition - 7 January 1, Orlando, Florida AIAA 1-11 Hypersonic Boundary-Layer Transition Eperiments in a Mach- Quiet Tunnel
More informationSupersonic air and wet steam jet using simplified de Laval nozzle
Proceedings of the International Conference on Power Engineering-15 (ICOPE-15) November 30- December 4, 2015, Yokohama, Japan Paper ID: ICOPE-15-1158 Supersonic air and wet steam jet using simplified de
More informationExperimental Investigation of a Long Rise- Time Pressure Signature through Turbulent Medium
Experimental Investigation of a Long Rise- Time Pressure Signature through Turbulent Medium Boeing Executive Seminar November 20 th 2014 Tohoku University Takahiro Ukai Institute of Fluid Science, Tohoku
More informationthe pitot static measurement equal to a constant C which is to take into account the effect of viscosity and so on.
Mechanical Measurements and Metrology Prof. S. P. Venkateshan Department of Mechanical Engineering Indian Institute of Technology, Madras Module -2 Lecture - 27 Measurement of Fluid Velocity We have been
More informationFluctuating Pressure Inside/Outside the Flow Separation Region in High Speed Flowfield
Journal of Aerospace Science and Technology 1 (2015) 18-26 doi: 10.17265/2332-8258/2015.01.003 D DAVID PUBLISHING Fluctuating Pressure Inside/Outside the Flow Separation Region in High Speed Flowfield
More informationHypersonic Flight Effects on Optical Sensors
A Tutorial Of: Hypersonic Flight Effects on Optical Sensors Matt Salem The University of Arizona: OPTI 521 12/4/2016 Background: In recent years hypersonic vehicles have received a lot of attention from
More informationTransient Temperature Probe Measurements. in a Mach 4 Nitrogen Jet
Transient Temperature Probe Measurements in a Mach 4 Nitrogen Jet D R BUTTSWORTH Faculty of Engineering and Surveying University of Southern Queensland Toowoomba, Qld, 435 Australia Email: buttswod@usq.edu.au
More informationCompressible Gas Flow
Compressible Gas Flow by Elizabeth Adolph Submitted to Dr. C. Grant Willson CHE53M Department of Chemical Engineering The University of Texas at Austin Fall 008 Compressible Gas Flow Abstract In this lab,
More informationIntroduction to Aerodynamics. Dr. Guven Aerospace Engineer (P.hD)
Introduction to Aerodynamics Dr. Guven Aerospace Engineer (P.hD) Aerodynamic Forces All aerodynamic forces are generated wither through pressure distribution or a shear stress distribution on a body. The
More informationComparison of drag measurements of two axisymmetric scramjet models at Mach 6
16th Australasian Fluid Mechanics Conference Crown Plaza, Gold Coast, Australia 2-7 December 27 Comparison of drag measurements of two axisymmetric scramjet models at Mach 6 Katsuyoshi Tanimizu, D. J.
More informationIntroduction to Fluid Mechanics - Su First experiment: Flow through a Venturi
530.327 - Introduction to Fluid Mechanics - Su First experiment: Flow through a Venturi 1 Background and objectives. In this experiment, we will study the flow through a Venturi section using both flow
More informationTransition and Instability Measurements in a Mach 6 Hypersonic Quiet Wind Tunnel
AIAA SciTech 13-17 January 2014, National Harbor, Maryland 52nd Aerospace Sciences Meeting AIAA 2014-0074 Transition and Instability Measurements in a Mach 6 Hypersonic Quiet Wind Tunnel Brandon C. Chynoweth,
More informationSignature: (Note that unsigned exams will be given a score of zero.)
Neatly print your name: Signature: (Note that unsigned exams will be given a score of zero.) Circle your lecture section (-1 point if not circled, or circled incorrectly): Prof. Dabiri Prof. Wassgren Prof.
More informationfor what specific application did Henri Pitot develop the Pitot tube? what was the name of NACA s (now NASA) first research laboratory?
1. 5% short answers for what specific application did Henri Pitot develop the Pitot tube? what was the name of NACA s (now NASA) first research laboratory? in what country (per Anderson) was the first
More informationIX. COMPRESSIBLE FLOW. ρ = P
IX. COMPRESSIBLE FLOW Compressible flow is the study of fluids flowing at speeds comparable to the local speed of sound. This occurs when fluid speeds are about 30% or more of the local acoustic velocity.
More informationSPC 407 Sheet 2 - Solution Compressible Flow - Governing Equations
SPC 407 Sheet 2 - Solution Compressible Flow - Governing Equations 1. Is it possible to accelerate a gas to a supersonic velocity in a converging nozzle? Explain. No, it is not possible. The only way to
More informationA STUDY ON THE BEHAVIOR OF SHOCK WAVE AND VORTEX RING DISCHARGED FROM A PIPE
A STUDY ON THE BEHAVIOR OF SHOCK WAVE AND VORTEX RING DISCHARGED FROM A PIPE S. KITAJIMA 1, J. IWAMOTO 2 and E. TAMURA 3 Corresponding author S. KITAJIMA ABSTRACT In this paper, the behavior of shock wave
More informationOne-Dimensional Isentropic Flow
Cairo University Second Year Faculty of Engineering Gas Dynamics AER 201B Aerospace Department Sheet (1) 2011-2012 One-Dimensional Isentropic Flow 1. Assuming the flow of a perfect gas in an adiabatic,
More informationIntroduction to Aerospace Engineering
Introduction to Aerospace Engineering Lecture slides Challenge the future 3-0-0 Introduction to Aerospace Engineering Aerodynamics 5 & 6 Prof. H. Bijl ir. N. Timmer Delft University of Technology 5. Compressibility
More informationReview of Fundamentals - Fluid Mechanics
Review of Fundamentals - Fluid Mechanics Introduction Properties of Compressible Fluid Flow Basics of One-Dimensional Gas Dynamics Nozzle Operating Characteristics Characteristics of Shock Wave A gas turbine
More informationP 1 P * 1 T P * 1 T 1 T * 1. s 1 P 1
ME 131B Fluid Mechanics Solutions to Week Three Problem Session: Isentropic Flow II (1/26/98) 1. From an energy view point, (a) a nozzle is a device that converts static enthalpy into kinetic energy. (b)
More information1. (20 pts total 2pts each) - Circle the most correct answer for the following questions.
ME 50 Gas Dynamics Spring 009 Final Exam NME:. (0 pts total pts each) - Circle the most correct answer for the following questions. i. normal shock propagated into still air travels with a speed (a) equal
More informationAerodynamic Optimization of the Expansion Section in a Hypersonic Quiet Nozzle Based on Favorable Pressure Effect
Journal of Applied Mathematics and Physics, 2014, 2, 443-448 Published Online May 2014 in SciRes. http://www.scirp.org/journal/jamp http://dx.doi.org/10.4236/jamp.2014.26054 Aerodynamic Optimization of
More informationEXPERIMENTAL STUDY OF JET FLOW FIELD BY DUAL HOLOGRAM INTERFEROMETRY
7 TH INTERNATIONAL CONGRESS OF THE AERONAUTICAL SCIENCES EXPERIMENTAL STUDY OF JET FLOW FIELD BY DUAL HOLOGRAM INTERFEROMETRY Peng Lv*, Zhimin Chen, Xing Wang *Northwestern Polytechnical University, Xian,
More informationWALL ROUGHNESS EFFECTS ON SHOCK BOUNDARY LAYER INTERACTION FLOWS
ISSN (Online) : 2319-8753 ISSN (Print) : 2347-6710 International Journal of Innovative Research in Science, Engineering and Technology An ISO 3297: 2007 Certified Organization, Volume 2, Special Issue
More informationAerodynamics. Basic Aerodynamics. Continuity equation (mass conserved) Some thermodynamics. Energy equation (energy conserved)
Flow with no friction (inviscid) Aerodynamics Basic Aerodynamics Continuity equation (mass conserved) Flow with friction (viscous) Momentum equation (F = ma) 1. Euler s equation 2. Bernoulli s equation
More informationConcentration Probe Measurements in a Mach 4 Nonreacting Hydrogen Jet
Concentration Probe Measurements in a Mach 4 Nonreacting Hydrogen Jet D R Buttsworth Faculty of Engineering and Surveying University of Southern Queensland Toowoomba, Qld, 435 Australia Email: buttswod@usq.edu.au
More informationMCE380: Measurements and Instrumentation Lab
MCE380: Measurements and Instrumentation Lab Chapter 8: Flow Measurements Topics: Basic Flow Equations Flow Obstruction Meters Positive Displacement Flowmeters Other Methods Holman, Ch. 7 Cleveland State
More informationSPC 407 Sheet 5 - Solution Compressible Flow Rayleigh Flow
SPC 407 Sheet 5 - Solution Compressible Flow Rayleigh Flow 1. Consider subsonic Rayleigh flow of air with a Mach number of 0.92. Heat is now transferred to the fluid and the Mach number increases to 0.95.
More informationOverview of Boundary-Layer Transition Research at AEDC Tunnel 9
Arnold Engineering Development Complex Overview of Boundary-Layer Transition Research at AEDC Tunnel 9 Eric Marineau AEDC White Oak June 29 2016 AEDC Hypervelocity Wind Tunnel No. 9 Unique hypersonic blowdown
More informationDESIGN AND CONSTRUCTION OF A WATER TUNNEL. Stephen C. Ko
i DESGN AND CONSTRUCTON OF A WATER TUNNEL By Stephen C. Ko This work has been carried out as a part of a grant from the National Science Foundation for the development of fluid mechanics laboratory equipments
More informationPART 1B EXPERIMENTAL ENGINEERING. SUBJECT: FLUID MECHANICS & HEAT TRANSFER LOCATION: HYDRAULICS LAB (Gnd Floor Inglis Bldg) BOUNDARY LAYERS AND DRAG
1 PART 1B EXPERIMENTAL ENGINEERING SUBJECT: FLUID MECHANICS & HEAT TRANSFER LOCATION: HYDRAULICS LAB (Gnd Floor Inglis Bldg) EXPERIMENT T3 (LONG) BOUNDARY LAYERS AND DRAG OBJECTIVES a) To measure the velocity
More informationTECHNICAL NOTE D ABLATING UNDER CONSTANT AERODYNAMIC CONDITIONS. By Robert R. Howell. Langley Research Center Langley Station, Hampton, Va.
-. NASA TN D-1635 TECHNICAL NOTE 1 D- 1635 AN EXPERIMENTAL STUDY OF THE BEHAVIOR OF SPHERES ABLATING UNDER CONSTANT AERODYNAMIC CONDITIONS By Robert R. Howell Langley Research Center Langley Station, Hampton,
More informationInternational Conference on Methods of Aerophysical Research, ICMAR 2008
International Conference on Methods of Aerophysical Research, ICMAR 8 EXPERIMENTAL STUDY OF UNSTEADY EFFECTS IN SHOCK WAVE / TURBULENT BOUNDARY LAYER INTERACTION P.A. Polivanov, А.А. Sidorenko, A.A. Maslov
More informationChapter 5 Phenomena of laminar-turbulent boundary layer transition (including free shear layers)
Chapter 5 Phenomena of laminar-turbulent boundary layer transition (including free shear layers) T-S Leu May. 3, 2018 Chapter 5: Phenomena of laminar-turbulent boundary layer transition (including free
More informationarxiv: v1 [physics.flu-dyn] 31 May 2017
arxiv:1705.11011v1 [physics.flu-dyn] 31 May 2017 Combined free-stream disturbance measurements and receptivity studies in hypersonic wind tunnels by means of a slender wedge probe and DNS Alexander Wagner,
More informationSUPERSONIC JET CONTROL WITH INTERNAL GROOVES
Proceedings of the International Conference on Mechanical Engineering 2005 (ICME2005) 28-30 December 2005, Dhaka, Bangladesh ICME05- SUPERSONIC JET CONTROL WITH INTERNAL GROOVES Shafiqur Rehman 1, M. Jamil
More informationCHAPTER 5 CONVECTIVE HEAT TRANSFER COEFFICIENT
62 CHAPTER 5 CONVECTIVE HEAT TRANSFER COEFFICIENT 5.1 INTRODUCTION The primary objective of this work is to investigate the convective heat transfer characteristics of silver/water nanofluid. In order
More informationCFD Simulation of Internal Flowfield of Dual-mode Scramjet
CFD Simulation of Internal Flowfield of Dual-mode Scramjet C. Butcher, K. Yu Department of Aerospace Engineering, University of Maryland, College Park, MD, USA Abstract: The internal flowfield of a hypersonic
More informationThe E80 Wind Tunnel Experiment the experience will blow you away. by Professor Duron Spring 2012
The E80 Wind Tunnel Experiment the experience will blow you away by Professor Duron Spring 2012 Objectives To familiarize the student with the basic operation and instrumentation of the HMC wind tunnel
More informationMechanisms of Hypersonic Boundary Layer Transition on Two Generic Vehicle Geometries
Mechanisms of Hypersonic Boundary Layer Transition on Two Generic Vehicle Geometries Steven P. Schneider Purdue University, School of AAE AFOSR Contractors Meeting 12 September 2003 San Destin, Florida
More informationSUPERSONIC WIND TUNNEL Project One. Charles R. O Neill School of Mechanical and Aerospace Engineering Oklahoma State University Stillwater, OK 74078
41 SUPERSONIC WIND UNNEL Project One Charles R. O Neill School of Mechanical and Aerospace Engineering Oklahoma State University Stillwater, OK 74078 Project One in MAE 3293 Compressible Flow September
More informationEXPERIMENTS OF CLOSED-LOOP FLOW CONTROL FOR LAMINAR BOUNDARY LAYERS
Fourth International Symposium on Physics of Fluids (ISPF4) International Journal of Modern Physics: Conference Series Vol. 19 (212) 242 249 World Scientific Publishing Company DOI: 1.1142/S211945128811
More informationTutorial Materials for ME 131B Fluid Mechanics (Compressible Flow & Turbomachinery) Calvin Lui Department of Mechanical Engineering Stanford University Stanford, CA 94305 March 1998 Acknowledgments This
More informationFUNDAMENTAL STUDIES OF THE CATALYTIC IGNITION PROCESS
FUNDAMENTAL STUDIES OF THE CATALYTIC IGNITION PROCESS Final Report KLK411 N08-03 National Institute for Advanced Transportation Technology University of Idaho Robert Lounsbury; Katrina Leichliter; and
More informationRapid Prototyping for Aerospace Launch Vehicles
Rapid Prototyping for Aerospace Launch Vehicles K. Siva Prasad *, E.Rathakrishnan +, Sanjay.G.Dhande * * Department of Mechanical engineering, IIT-Kanpur, India + Department of Aerospace Engineering, IIT-Kanpur,
More informationLecture 22. Mechanical Energy Balance
Lecture 22 Mechanical Energy Balance Contents Exercise 1 Exercise 2 Exercise 3 Key Words: Fluid flow, Macroscopic Balance, Frictional Losses, Turbulent Flow Exercise 1 It is proposed to install a fan to
More informationChapter 17. For the most part, we have limited our consideration so COMPRESSIBLE FLOW. Objectives
Chapter 17 COMPRESSIBLE FLOW For the most part, we have limited our consideration so far to flows for which density variations and thus compressibility effects are negligible. In this chapter we lift this
More informationEXTERNAL-JET (FLUID) PROPULSION ANALOGY FOR PHOTONIC (LASER) PROPULSION By John R. Cipolla, Copyright February 21, 2017
EXTERNAL-JET (FLUID) PROPULSION ANALOGY FOR PHOTONIC (LASER) PROPULSION By John R. Cipolla, Copyright February 21, 2017 ABSTRACT External-jet propulsion uses a narrow jet of high velocity water or conceptually
More informationEFFECT OF WALL JET ON OSCILLATION MODE OF IMPINGING JET
EFFECT OF WALL JET ON OSCILLATION MODE OF IMPINGING JET Y. Sakakibara 1, M. Endo 2, and J. Iwamoto 3 ABSTRACT When an axisymmetric underexpanded jet impinges on a flat plate perpendicularly, the feedback
More information2 Internal Fluid Flow
Internal Fluid Flow.1 Definitions Fluid Dynamics The study of fluids in motion. Static Pressure The pressure at a given point exerted by the static head of the fluid present directly above that point.
More informationIntroduction and Basic Concepts
Topic 1 Introduction and Basic Concepts 1 Flow Past a Circular Cylinder Re = 10,000 and Mach approximately zero Mach = 0.45 Mach = 0.64 Pictures are from An Album of Fluid Motion by Van Dyke Flow Past
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 informationCHAPTER (13) FLOW MEASUREMENTS
CHAPTER (13) FLOW MEASUREMENTS 09/12/2010 Dr. Munzer Ebaid 1 Instruments for the Measurements of Flow Rate 1. Direct Methods: Volume or weight measurements. 2. Indirect Methods: Venturi meters, Orifices
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 informationWilliam В. Brower, Jr. A PRIMER IN FLUID MECHANICS. Dynamics of Flows in One Space Dimension. CRC Press Boca Raton London New York Washington, D.C.
William В. Brower, Jr. A PRIMER IN FLUID MECHANICS Dynamics of Flows in One Space Dimension CRC Press Boca Raton London New York Washington, D.C. Table of Contents Chapter 1 Fluid Properties Kinetic Theory
More informationHOMEWORK ASSIGNMENT ON BERNOULLI S EQUATION
AMEE 0 Introduction to Fluid Mechanics Instructor: Marios M. Fyrillas Email: m.fyrillas@frederick.ac.cy HOMEWORK ASSIGNMENT ON BERNOULLI S EQUATION. Conventional spray-guns operate by achieving a low pressure
More information1. Mark the correct statement(s)
1. Mark the correct statement(s) Figure to the right shows a mass measurement scale using a spring. 1.1 The span of the scale is a) 16 kg b) 21 kg c) 11 kg d) 5-16 kg 1.2 The range of the scale is a) 16
More information2 Navier-Stokes Equations
1 Integral analysis 1. Water enters a pipe bend horizontally with a uniform velocity, u 1 = 5 m/s. The pipe is bended at 90 so that the water leaves it vertically downwards. The input diameter d 1 = 0.1
More information1. For an ideal gas, internal energy is considered to be a function of only. YOUR ANSWER: Temperature
CHAPTER 11 1. For an ideal gas, internal energy is considered to be a function of only. YOUR ANSWER: Temperature 2.In Equation 11.7 the subscript p on the partial derivative refers to differentiation at
More informationPart A: 1 pts each, 10 pts total, no partial credit.
Part A: 1 pts each, 10 pts total, no partial credit. 1) (Correct: 1 pt/ Wrong: -3 pts). The sum of static, dynamic, and hydrostatic pressures is constant when flow is steady, irrotational, incompressible,
More informationApplied Gas Dynamics Flow With Friction and Heat Transfer
Applied Gas Dynamics Flow With Friction and Heat Transfer Ethirajan Rathakrishnan Applied Gas Dynamics, John Wiley & Sons (Asia) Pte Ltd c 2010 Ethirajan Rathakrishnan 1 / 121 Introduction So far, we have
More informationBackground Information for Use of Pitot Tube, Manometer, Hot Wires, and Hot Films
AAE 50 Notes, 9-Jan-04 Page 1 Background Information for Use of Pitot Tube, Manometer, Hot Wires, and Hot Films 1 Background The following is adapted from the handout in AAE333L. 1.1.1 Specific Applications:
More informationHypersonic Boundary-Layer Transition Experiments in the Boeing/AFOSR Mach-6 Quiet Tunnel
5th AAA Aerospace Sciences Meeting including the New orizons orum and Aerospace xposition 9-1 January 1, Nashville, Tennessee AAA 1-8 ypersonic Boundary-Layer Transition xperiments in the Boeing/AOSR Mach-6
More informationSyllabus for AE3610, Aerodynamics I
Syllabus for AE3610, Aerodynamics I Current Catalog Data: AE 3610 Aerodynamics I Credit: 4 hours A study of incompressible aerodynamics of flight vehicles with emphasis on combined application of theory
More informationCHAPTER 7 SEVERAL FORMS OF THE EQUATIONS OF MOTION
CHAPTER 7 SEVERAL FORMS OF THE EQUATIONS OF MOTION 7.1 THE NAVIER-STOKES EQUATIONS Under the assumption of a Newtonian stress-rate-of-strain constitutive equation and a linear, thermally conductive medium,
More informationA Shock Tube Driven Transonic Wind Tunnel
J. Nash a, b and B.W. Skews b Received 28 November 2012, in revised form and accepted 20 March 2013 A small transonic wind tunnel was designed, constructed and calibrated in order to provide a tool for
More informationACCOUNTING FOR FRICTION IN THE BERNOULLI EQUATION FOR FLOW THROUGH PIPES
ACCOUNTING FOR FRICTION IN THE BERNOULLI EQUATION FOR FLOW THROUGH PIPES Some background information first: We have seen that a major limitation of the Bernoulli equation is that it does not account for
More informationNumerical Study of Boundary-Layer Receptivity on Blunt Compression-Cones in Mach-6 Flow with Localized Freestream Hot-Spot Perturbations
Blunt Compression-Cones in Mach-6 Flow with Localized Freestream Hot-Spot Perturbations Yuet Huang 1 and Xiaolin Zhong 2 Mechanical and Aerospace Engineering Department, University of California, Los Angeles,
More informationME332 FLUID MECHANICS LABORATORY (PART I)
ME332 FLUID MECHANICS LABORATORY (PART I) Mihir Sen Department of Aerospace and Mechanical Engineering University of Notre Dame Notre Dame, IN 46556 Version: January 14, 2002 Contents Unit 1: Hydrostatics
More informationLAMINAR FLOW CONTROL OF A HIGH-SPEED BOUNDARY LAYER BY LOCALIZED WALL HEATING OR COOLING
LAMINAR FLOW CONTROL OF A HIGH-SPEED BOUNDARY LAYER BY LOCALIZED WALL HEATING OR COOLING Fedorov A.V.*, Soudakov V.G.*, Egorov I.V.*, Sidorenko A.A.**, Gromyko Y.*, Bountin D.** *TsAGI, Russia, **ITAM
More informationMeasurements using Bernoulli s equation
An Internet Book on Fluid Dynamics Measurements using Bernoulli s equation Many fluid measurement devices and techniques are based on Bernoulli s equation and we list them here with analysis and discussion.
More informationHypersonic flow and flight
University of Stuttgart, Aerospace Engineering and Geodesy Dept. - Lecture - Hypersonic flow and flight Master Level, Specialization 4 lecture hours per week in WS, 3-6 LPs/ECTS Lecturer: Dr. Markus J.
More informationAE 3051, Lab #16. Investigation of the Ideal Gas State Equation. By: George P. Burdell. Group E3
AE 3051, Lab #16 Investigation of the Ideal Gas State Equation By: George P. Burdell Group E3 Summer Semester 000 Abstract The validity of the ideal gas equation of state was experimentally tested for
More informationIntroduction to Aerospace Engineering
4. Basic Fluid (Aero) Dynamics Introduction to Aerospace Engineering Here, we will try and look at a few basic ideas from the complicated field of fluid dynamics. The general area includes studies of incompressible,
More informationGALCIT Ludwieg Tube Ae 104b
Ae 104b TA: Greg Smetana Lab: Guggenheim 026 Phone: gsmetana@caltech.edu Winter 2014 1 Principles of supersonic wind tunnels Supersonic wind tunnels, like their subsonic counterparts, operate by accelerating
More informationAnemometry Anemometer Calibration Exercise
Atmospheric Measurements and Observations II EAS 535 Anemometry Anemometer Calibration Exercise Prof. J. Haase http://web.ics.purdue.edu/~jhaase/teaching/eas535/index.html Class Objectives How is wind
More informationAerodynamics of the reentry capsule EXPERT at full modeling viscous effect conditions
ISTC-STCU WORKSHOP FOR AEROSPACE TECHNOLGIES Aerodynamics of the reentry capsule EXPERT at full modeling viscous effect conditions A.M. Kharitonov ITAM SB RAS Ljubljana, Slovenia 10-12 March 2008 CONTENTS
More informationPREDICTION OF SOUND PRESSURE LEVELS ON ROCKET VEHICLES DURING ASCENT Revision E
PREDICTION OF SOUND PRESSURE LEVELS ON ROCKET VEHICLES DURING ASCENT Revision E By Tom Irvine Email: tomirvine@aol.com July 0, 011 Figure 0. Schlieren Photo, Wind Tunnel Test Engineers conducted wind tunnel
More informationDepartment of Mechanical Engineering ME 96. Free and Forced Convection Experiment. Revised: 25 April Introduction
CALIFORNIA INSTITUTE OF TECHNOLOGY Department of Mechanical Engineering ME 96 Free and Forced Convection Experiment Revised: 25 April 1994 1. Introduction The term forced convection refers to heat transport
More informationUnified Quiz: Thermodynamics
Unified Quiz: Thermodynamics October 14, 2005 Calculators allowed. No books or notes allowed. A list of equations is provided. Put your ID number on each page of the exam. Read all questions carefully.
More informationSPC Aerodynamics Course Assignment Due Date Monday 28 May 2018 at 11:30
SPC 307 - Aerodynamics Course Assignment Due Date Monday 28 May 2018 at 11:30 1. The maximum velocity at which an aircraft can cruise occurs when the thrust available with the engines operating with the
More informationInstitute of Aerodynamics and Gas Dynamics, University of Stuttgart, 70569, Stuttgart, Germany
ICMAR 01 STATIC CALIBRATION OF WEDGE HOT-FILM PROBE AND DETERMINATION OF SENSITIVITIES FOR MODAL ANALYSIS IN T-35 SUPERSONIC WIND TUNNEL M. Krause 1, U. Gaisbauer 1, E. Krämer 1, Y.G. Yermolaev, A.D. Kosinov
More information1. Introduction Some Basic Concepts
1. Introduction Some Basic Concepts 1.What is a fluid? A substance that will go on deforming in the presence of a deforming force, however small 2. What Properties Do Fluids Have? Density ( ) Pressure
More informationDevelopment of Flow over Blunt-Nosed Slender Bodies at Transonic Mach Numbers
Development of Flow over Blunt-Nosed Slender Bodies at Transonic Mach Numbers Gireesh Yanamashetti 1, G. K. Suryanarayana 2 and Rinku Mukherjee 3 1 PhD Scholar & Senior Scientist, 2 Chief Scientist, 3
More information