M.S.

Two steel pipe piles in place in abutments for two different bridge constructions sites were instrumented with strain gauges to measure the magnitude of negative skin friction. The piles were monitored before, during and up to 19 months after construction was completed. The load versus depth and time in each pile is discussed. Maximum observed dragloads ranged from 98 to 127 kips. A comparison with two methods for calculating dragloads is presented. Both comparison methods were found to be conservative, with the Briaud and Tucker (1997) approach more closely estimating the observed load versus depth behavior.

A direct, simple, and accurate way to measure skin friction by oil-film laser interferometry has been developed by various researchers. Equations and methods were developed to correct measurement errors arising from three-dimensional effects and pressure gradients. The oil-film, dual-beam laser interferometer was constructed to measure the skin friction around a wing-body junction in a three-dimensional, turbulent boundary layer with pressure gradients. The flow was dominated by the formation of a junction vortex generated at the nose of the wing-body. The oil-film skin friction results were compared with previous skin-friction measurements for the flow obtained by hot-wire measurements. The skin friction values agreed within approximately 8% between the two methods. The effects and benefits of scanning laser interferometry and alternative beam directions were investigated and discussed. The effect of dirt contamination on the data is also discussed. Methods to improve the data quality are presented. / Master of Science

This study explores the potential of the naphthalene sublimation technique to be used to analyze the manner in which surfaces regress, specifically focusing on the effect that skin friction has on regression and vice versa. For this experimentation, a flat steel plate installed with three skin friction sensors was coated with naphthalene via a mechanized sprayer and was installed in the wall of the Ahmic Aerospace Turbulent Boundary Layer Research supersonic wind tunnel. The plate and sensor configuration was subjected to three subsequent tunnel runs at Mach 2.31. This process was repeated at plenum pressures of 0.35 MPa and 0.69 MPa, which correspond to Reynolds Numbers of 1.4x10^7/m and 2.8x10^7/m. Between the first and final run, a -4.7% and -3.7% percent change in the coefficient of friction was seen at the 0.35 MPa and 0.69 MPa plenum conditions, respectively. Images of the plate taken before and after each run qualitatively indicate continual naphthalene regression with each subsequent tunnel run. This decrease in the coefficient of friction was attributed in part to the regression of the naphthalene coating, indicating that this method has the potential to be used to study the interaction between skin friction and regressing surfaces. Additionally, this study showed that it is certainly possible to measure skin friction with sensors where both the head of the movable sensor element and the surrounding wall is coated with sprayed naphthalene. / Master of Science / This study explores the potential of using an experimental method to better understand how surfaces recede when subjected to moving air. A chemical called naphthalene was used because naphthalene naturally recedes very easily, unlike many other materials which a researcher would want to study the receding behavior of, making it an ideal choice for this study. Sensors made to detect the amount of force generated by flowing air over a surface were installed into a flat metal plate. The plate was coated with naphthalene and installed into the wall of a wind tunnel, which generates airflow over the plate to simulate flight conditions. Three separate and sequential airflow runs over the plate were conducted. This process was repeated under two different air conditions. In both conditions, a decrease in the amount of force due to the air moving over the plate was detected as subsequent runs were conducted and the naphthalene receded. This indicates the method has the potential to be used to study how other surfaces recede when subjected to moving air.

The purpose of the project was to examine fiber optic sensors for the measurement of pressure, skin friction, temperature, and heat flux in high Reynolds number, supersonic flow. Using a standard fiber optic signal conditioning unit (specifically a broadband interferometric system using spectra), the work centered around determining under what conditions these sensors will work effectively and quantifying the total system limitations. An interferometric-based, fiber optic skin friction sensor was developed for the measurement of wall shear stress in complex, supersonic flows. This sensor type was tested successfully in laminar, incompressible flow, and supersonic flow up to Mach 1.92, Mach 2.4 and 3.0 flow, in which the sensor operated with varying success. A micromachined, fiber optic pressure sensor was also tested in these supersonic conditions, also with varying success. The accurate operation of these sensors was found to be tied to the flow conditions and the fiber optic, signal processing system. A correlation was found to exist between the energy of the flow, either through its dynamic pressure or through external disturbances such as shocks or separation, and the noise in the signals, expressed by the variance of the gap estimate, for the pressure and skin friction sensors in these flows. The energy of the flow couples with the mechanical properties of the sensor reducing the fringe contrast of the signal used by the optical signal processing system to determine a gap estimate. As the energy of the flow is increased and the sensor is excited, the fringe contrast is reduced. A practical limit of a normalized fringe contrast of 0.10 was found for producing accurate gap estimates in real flows. A consequence is that there is a limit to the dynamic pressure of the flow for the sensors to operate accurately, which is demonstrated by the varying success of the supersonic wind tunnel tests. This correlation is sensor specific, meaning that sensors can be designed to operate successfully in any flow. Also, the signal processing system, which forms the other end of the total system, could be improved to allow accurate measurements with the current sensors. / Ph. D.

This research produced a new class of skin friction gages that measures wall shear even in shock environments. One test specimen separately measured wall shear and variable-pressure induced moment. Through the investigation of available computational modeling methods, techniques for accurately predicting gage physical responses were developed. The culmination of these model combinations was a design optimization procedure. This procedure was applied to three disparate test conditions: 1) short-duration, high-enthalpy testing, 2) blow-down testing, and 3) flight testing. The resulting optimized gage designs were virtually tested against each set of nominal load conditions. The finalized designs each successfully met their respective test condition constraints while maximizing strain output due to wall shear. These gages limit sources of apparent strain: inertia, temperature gradient, and uniform pressure. A unique use of bellows provided a protective shroud for surface strain gages. Oil fill provided thermal and dynamic damping while eliminating uniform pressure as a source of output voltage. Two Wheatstone bridge configurations were developed to minimize temperature effects first from temperature gradient and then from spatially varying heat flux induced gradient. An inertia limiting technique was developed that parametrically investigated mass and center of gravity impact on strain output. Multiple disciplinary computational simulations of thermal, dynamic, shear, moment, inertia, and instrumentation interaction were developed. Examinations of instrumentation error, settling time, filtering, multiple input dynamic response, and strain gage placement to avoid thermal gradient were conducted. Detailed mechanical drawings for several gages were produced for fabrication and future testing. / Ph. D.

A series of direct-measuring skin friction gages were developed for a high-speed, high-temperature environment of the turbulent boundary layer in flows such as that in supersonic combustion ramjet (scramjet) engines, with a progression from free-jet ground tests to a design for an actual hypersonic scramjet-integrated flight vehicle. The designs were non-nulling, with a sensing head that was flush with the model wall and surrounded by a small gap. Thus, the shear force due to the flow along the wall deflects the head, inducing a measurable strain. Strain gages were used to detect the strain. The gages were statically calibrated using a direct force method. The designs were verified by testing in a well-documented Mach 2.4 cold flow. Results of the cold-flow tests were repeatable and within 15% of the value of Cf estimated from simple theory. The first gage design incorporated a cantilever beam with semiconductor strain gages to sense the shear on the floating head. Cooling water was routed both internally and around the external housing in order to control the temperature of the strain gages. This first gage was installed and tested in a rocket-based-combined-cycle (RBCC) engine model operating in the scramjet mode. The free-jet facility provided a Mach 6.4 flow with P0 = 1350 psia (9310 kPa) and T0 = 2800 °R (1555 °K). Local wall temperatures were measured between 850 and 900 °R (472-500 °K). Output from the RBCC scramjet tests was reasonable and repeatable. A second skin friction gage was designed for and tested in a wind tunnel model of the Hyper-X flight vehicle scramjet engine. These unsuccessful tests revealed the need for a radically different skin friction gage design. The third and final skin friction gage was specifically developed to be installed on the Hyper-X flight vehicle. Rather than the cantilever beam and semiconductor strain gages, the third skin friction gage made use of a flexure ring and metal foil strain gages to sense the shear. The water-cooling and oil-fill used on the previous skin friction sensors were eliminated. It was qualified for flight through a rigorous series of environmental tests, including pressure, temperature, vibration, and heat flux tests. Finally, the third skin friction gage was tested in the Hyper-X Engine Model (HXEM), a full-scale-partial-width wind tunnel model of the flight vehicle engine. These tests were conducted at Mach 6.5 enthalpy with P0 = 555 psia (3827 kPa) and h0 = 900 Btu/lbm in a freejet facility. The successful testing in the wind tunnel scramjet model provided the final verification of the gage before installation in the flight vehicle engine. The development, testing, and results of all three skin friction gages are discussed. / Ph. D.

This investigation concerns the design, analysis, and initial testing of a new, two-component wall shear gage for 3D, high-temperature flows. This gage is a direct-measuring, non-nulling design with a round head surrounded by a small gap. Two flexure wheels are used to allow small motions of the floating head. Fiber-optic displacement sensors measure how far the polished faces of counterweights on the wheels move in relation to a fixed housing as the primary measurement system. No viscous damping was required. The gage has both fiber-optic instrumentation and strain gages mounted on the flexures for validation of the newer fiber optics. The sensor is constructed of Haynes 230, a high-temperature nickel alloy. The gage housing is made of 316 stainless steel. All components of the gage in pure fiber-optic form can survive to a temperature of 1073 K. The bonding methods of the backup strain gages limit their maximum temperature to 473 K. The dynamic range of the gage is from 0-500 Pa (0-10g) and higher shears can be measured by changing the floating head size. Extensive use of finite element modeling was critical to the design and analysis of the gage. Static structural, modal, and thermal analyses were performed on the flexures using the ANSYS finite element package. Static finite element analysis predicted the response of the flexures to a given load, and static calibrations using a direct force method confirmed these results. Finite element modal analysis results were within 16.4% for the first mode and within 30% for the second mode when compared with the experimentally determined modes. Vibration characteristics of the gage were determined from experimental free vibration data after the gage was subjected to an impulse. Uncertainties in the finished geometry make this level of error acceptable. A transient thermal analysis examined the effects of a very high heat flux on the exposed head of the gage. The 100,000 W/m2 heat flux used in this analysis is typical of a value in a scramjet engine. The gage can survive for 10 minutes and operate for 3 minutes before a 10% loss in flexure stiffness occurs under these conditions. Repeated cold-flow wind tunnel tests at Mach 2.4 with a stagnation pressure from 3.7-8.2 atm (55-120 psia) and ambient stagnation temperature (Re=6.6x107/m) and Mach 4.0 with a stagnation pressure from 10.2-12.2 atm (150-180 psia) and ambient stagnation temperature (Re=7.4x107/m) were performed in the Virginia Tech Supersonic Wind Tunnel. Some of these tests had the gage intentionally misaligned by 25o to create a virtual 3D flow in this nominally 2D facility. Experimental results gave excellent agreement with semi-empirical prediction methods for both the aligned and 25o experiments. This fiber-optic skin friction gage operated successfully without viscous damping. These tests in the supersonic wind tunnel validated this wall shear gage design concept. / Ph. D.

An instrument for the direct measurement of skin friction in complex supersonic and hypersonic flows was developed. The flows were complex because they were of very short duration, with high temperature and shocks and often injection, mixing, and combustion. A wall-mounted, miniature cantilevered beam device measured the small tangential shear force on the non-intrusive floating element. Semiconductor strain gages mounted at the beam’s base measure the small strains that are generated. By modifying the geometry of the sensing unit, this design can be adapted for a variety of test flows. The use of engineering plastics and short beam length provide high frequency response and make the beam stiff so that the floating head’s deflection due to the shear is negligible, allowing for a non-nulling design. Measurements were made in scramjet models at the NASA Ames 16-inch Shock Tunnel and the General Applied Science Laboratory HYPULSE facility. Test flow conditions were harsh with the facilities simulating Mach 14 enthalpy conditions (320 atm and 10000 R total temperatures) for 0.3-2 milliseconds. The use of engineering plastics reduces heat transfer, so that measurements can be made in these very hot impulsive flows without thermal contamination of the data. Skin friction data in agreement with other correlations and measurements were obtained at both facilities. Mach 2.4 cold flow tests were also performed in the Virginia Tech Supersonic Tunnel. These helped verify the concept and to establish pressure gradient sensitivity in the case of a shock wave impacting directly on the sensing head. Analysis of the measurement uncertainty in the cold supersonic flow tests showed that an uncertainty of approximately 10 percent is achievable. An uncertainty of 15-20% is estimated for the most severe hot cases. An assortment of variations were applied to the gage to extend gage life. The most significant was the replacement of the oil in the sensing gap with a silicon rubber, eliminating service requirements. Tests at all of the facilities confirmed that the rubber-filled gages provided approximately the same level of accuracy as was achieved with the original oil-filled gage design, except when shocks impacted the gage head. / Ph. D.

An investigation was conducted to design instruments to directly measure skin friction along the chamber walls of supersonic combustor models. Measurements were made in a combustor at the General Applied Science Laboratory (GASL) and in the Direct Connect Arcjet Facility (DCAF) supersonic combustor at the NASA AMES Research Center. Flow conditions in the high enthalpy combustor models ranged from total pressures of 275-800 psia (1900-5550 kPa) and total temperatures from 5800-8400 R (3222-4667 K). This gives enthalpies in the range of 1700-3300 BTU/Ib_m (3950-7660 KJ/kg) and simulated flight Mach number from 9 to 13. A direct force measurement device was used to measure the small tangential shear force resulting from the flow passing over a non-intrusive floating element. The floating head is mounted to a stiff cantilever beam arrangement with deflection due to the shear force on the order of 0.0005 in (0.0125 mm). This small deflection allows the balance to be a non-nulling type. Several measurements were conducted in cold supersonic flows to verify the concept and establish accuracy and repeatability. This balance design includes actively controlled cooling of the floating sensor head temperature through an internal cooling system to eliminate nonuniform temperature effects between the head and the surrounding chamber wall. This enabled the device to be suitable for shear force measurement in very hot flows. The key to this device is the use of a quartz tube cantilever with strain gages bonded at orthogonal positions directly on the surface at the base. A symmetric fluid flow was developed inside the quartz tube to provide cooling to the backside of the floating head. Bench tests showed that this did not influence the force measurement. Numerical heat transfer calculations were conducted for design feasibility and analysis, and to determine the effectiveness of the active cooling of the floating head. Analysis of the measurement uncertainty in cold supersonic flow tests show that uncertainty under 8% is achievable, but variations in the balance cooling during a particular test raised uncertainty up to 20% in these very hot flows during the early tests. Improvements to the strain gages and balance cooling reduced uncertainty for the later tests to under 15%. / Ph. D.