Results for “shock loading” — Page 29

NASA NTRS1994

Microelectronics Instrument Products Shock and Vibration Electro-optics: C-Qualification Test Report

In this test report all measurements made during testing are recorded in ATP 20049 DS data sheets and are included in the log. The motor/encoder (henceforth referred to as the UUT) test sequence began with a baseline functional evaluation, which demonstrated that the motor satisfied the operating torque, cogging torque, winding resistance, and mechanical requirements of SOW. In addition, the encoder electrical requirements were verified, as well as the alignment of the encoder outputs relative tc, the motor shaft position. There were no discrepancies observed in this portion of the test. The UUT was then exposed to a number of environments, including thermal vacuum, thermal cycling, random and sine vibration, and mechanical shock. During the thermal environments, the performance of the UUT under load was verified at specified points in the cycles, as described in ATP 20049. In addition, the UUT was bench tested between the two thermal environments. No anomalies were observed during the thermal tests. The load attachment method was subsequently corrected, and vibration of S/N 0002 began while 0003 was being repaired.

Source record↗

NASA NTRSDate not supplied

Cruise Slotted Wing Design with Natural Laminar Flow for Transonic Commercial Transport Aircraft

The present computational study investigates the aerodynamic design and analysis of cruise slotted wings with natural laminar flow for transonic transport aircraft. The cruise slotted wing is a multielement wing concept that features an intermediate slot to achieve greater aft loading relative to supercritical wings for the potential benefit of reduced shock strength and pressure drag. Transonic near-cruise, off-design assessments have also shown improved drag rise characteristics due to the ability of the slot to mitigate boundary layer separation on the aft flap component. However, due to the decreased Reynolds number of the flap, the cruise slotted wing has historically incurred a skin-friction drag penalty relative to a conventional supercritical wing. To offset this penalty, a cruise slotted wing with the forward main element and aft flap element shaped to achieve natural laminar flow is desired. Toward this effort, a knowledge-based aerodynamic design method, CDISC, has been leveraged to design a partial-span cruise slotted wing with natural laminar flow for a Mach-0.8 variant of the Common Research Model. Drag comparisons will be provided at cruise and near-cruise, off-design conditions relative to both fully turbulent and natural laminar flow conventional wing designs. It is anticipated that pairing natural laminar flow technology with the cruise slotted wing architecture will allow for cruise drag performance similar to conventional laminar flow wings with improved drag rise characteristics and more limited laminar-flow sensitivity at off-design conditions. Preliminary results for a cruise slotted wing with laminar flow on the outboard wing section only have shown a 10-ct cruise drag reduction relative to a conventional supercritical wing, but a 8-ct penalty relative to the conventional NLF wing. The final paper will include results for a cruise slotted wing design with laminar flow on both the inboard and outboard wing sections.

CFD↗

NASA NTRSDate not supplied

Cruise Slotted Wing Design with Natural Laminar Flow for Transonic Commercial Transport Aircraft

The present computational study investigates the aerodynamic design and analysis of cruise slotted wings with natural laminar flow for transonic transport aircraft. The cruise slotted wing is a multielement wing concept that features an intermediate slot to achieve greater aft loading relative to supercritical wings for the potential benefit of reduced shock strength and pressure drag. Transonic near-cruise, off-design assessments have also shown improved drag rise characteristics due to the ability of the slot to mitigate boundary layer separation on the aft flap component. However, due to the decreased Reynolds number of the flap, the cruise slotted wing has historically incurred a skin-friction drag penalty relative to a conventional supercritical wing. To offset this penalty, a cruise slotted wing with the forward main element and aft flap element shaped to achieve natural laminar flow is desired. Toward this effort, a knowledge-based aerodynamic design method, CDISC, has been leveraged to design a partial-span cruise slotted wing with natural laminar flow for a Mach-0.8 variant of the Common Research Model. Drag comparisons will be provided at cruise and near-cruise, off-design conditions relative to both fully turbulent and natural laminar flow conventional wing designs. It is anticipated that pairing natural laminar flow technology with the cruise slotted wing architecture will allow for cruise drag performance similar to conventional laminar flow wings with improved drag rise characteristics and more limited laminar-flow sensitivity at off-design conditions. Preliminary results for a cruise slotted wing with laminar flow on the outboard wing section only have shown a 10-ct cruise drag reduction relative to a conventional supercritical wing, but a 8-ct penalty relative to the conventional NLF wing. The final paper will include results for a cruise slotted wing design with laminar flow on both the inboard and outboard wing sections.

Natural Laminar Flow↗

DOE OSTI2023

In situ measurement of damage evolution in shocked magnesium as a function of microstructure

Accurate modeling and prediction of damage induced by dynamic loading in materials have long proved to be a difficult task. Examination of postmortem recovered samples cannot capture the time-dependent evolution of void nucleation and growth, and attempts at analytical models are hindered by the necessity to make simplifying assumptions, because of the lack of high-resolution, in situ, time-resolved experimental data. We use absorption contrast imaging to directly image the time evolution of spall damage in metals at ~1.6-μm spatial resolution. We observe a dependence of void distribution and size on time and microstructure. The insights gained from these data can be used to validate and improve dynamic damage prediction models, which have the potential to lead to the design of superior damage-resistant materials.

36 MATERIALS SCIENCE↗

NASA NTRSDate not supplied

OVERFLOW Analysis of Supersonic Retropropulsion Testing on a Blunt Mars Entry Vehicle Concept

Supersonic retropropulsion (SRP) flow over a Hypersonic Inflatable Aerodynamic Decelerator (HIAD) blunt-body vehicle was simulated using the Overflow Computational Fluid Dynamics (CFD) solver. Two nozzle configurations were tested (1E and 1F) as a subset of the seven total configurations in the Descent System Study (DSS) testing campaign. A generalized, Adaptive Mesh Refinement (AMR) shock capturing and plume refinement technique was developed and calibrated for producing automatic and unique grid systems optimized for any given testing condition. Solution independence from grid resolution was determined by successively increasing grid refinement until asymptotic convergence of the mean aerodynamic loads was observed. Overflow Reynolds-averaged Navier-Stokes (RANS) solutions produced realistic SRP flow phenomena, including Mach disk normal shocks contained in high-thrust plumes and bow shock triple-points. Solutions for the 1E nozzle configuration were steady. Conversely, a subset of the 1F test conditions were unsteady and exhibited periodic, non-sinusoidal expansions and contractions of the streamwise-oriented plume, resulting in unsteady aerodynamic loads on the vehicle. Initial investigations into the effect of turbulence modeling fidelity demonstrated significant differences between the aerodynamic loads simulated with RANS and Detached Eddy Simulation (DES). Multiple flow mechanisms were identified as root causes of these differences, including bow shock shape augmentation and reduced entrainment due to reduced turbulence in the DES simulations. Initial comparisons of pitching moment control authority and aerodynamic drag performance between the tested nozzle configurations demonstrated that the 1E configuration may be more advantageous for the Mars Entry, Descent, and Landing (EDL) task.

Supersonic Retropropulsion↗

NASA NTRSDate not supplied

OVERFLOW Analysis of Supersonic Retropropulsion Testing on a Blunt Mars Entry Vehicle Concept

Supersonic retropropulsion (SRP) flow over a Hypersonic Inflatable Aerodynamic Decelerator (HIAD) blunt-body vehicle was simulated using the Overflow Computational Fluid Dynamics (CFD) solver. Two nozzle configurations were tested (1E and 1F) as a subset of the seven total configurations in the Descent System Study (DSS) testing campaign. A generalized, Adaptive Mesh Refinement (AMR) shock capturing and plume refinement technique was developed and calibrated for producing automatic and unique grid systems optimized for any given testing condition. Solution independence from grid resolution was determined by successively increasing grid refinement until asymptotic convergence of the mean aerodynamic loads was observed. Overflow Reynolds-averaged Navier-Stokes (RANS) solutions produced realistic SRP flow phenomena, including Mach disk normal shocks contained in high-thrust plumes and bow shock triple-points. Solutions for the 1E nozzle configuration were steady. Conversely, a subset of the 1F test conditions were unsteady and exhibited periodic, non-sinusoidal expansions and contractions of the streamwise-oriented plume, resulting in unsteady aerodynamic loads on the vehicle. Initial investigations into the effect of turbulence modeling fidelity demonstrated significant differences between the aerodynamic loads simulated with RANS and Detached Eddy Simulation (DES). Multiple flow mechanisms were identified as root causes of these differences, including bow shock shape augmentation and reduced entrainment due to reduced turbulence in the DES simulations. Initial comparisons of pitching moment control authority and aerodynamic drag performance between the tested nozzle configurations demonstrated that the 1E configuration may be more advantageous for the Mars Entry, Descent, and Landing (EDL) task.

Supersonic Retropropulsion↗

NASA NTRS1994

Experimental studies of transpiration cooling with shock interaction in hypersonic flow, part B

This report describes the result of experimental studies conducted to examine the effects of the impingement of an oblique shock on the flowfield and surface characteristics of a transpiration-cooled wall in turbulent hypersonic flow. The principal objective of this work was to determine whether the interaction between the oblique shock and the low-momentum region of the transpiration-cooled boundary layer created a highly distorted flowfield and resulted in a significant reduction in the cooling effectiveness of the transpiration-cooled surface. As a part of this program, we also sought to determine the effectiveness of transpiration cooling with nitrogen and helium injectants for a wide range of blowing rates under constant-pressure conditions in the absence of shock interaction. This experimental program was conducted in the Calspan 48-Inch Shock Tunnel at nominal Mach numbers of 6 and 8, for a Reynolds number of 7.5 x 10(exp 6). For these test conditions, we obtained fully turbulent boundary layers upstream of the interaction regions over the transpiration-cooled segment of the flat plate. The experimental program was conducted in two phases. In the first phase, we examined the effects of mass-addition level and coolant properties on the cooling effectiveness of transpiration-cooled surfaces in the absence of shock interaction. In the second phase of the program, we examined the effects of oblique shock impingement on the flowfield and surface characteristics of a transpiration-cooled surface. The studies were conducted for a range of shock strengths with nitrogen and helium coolants to examine how the distribution of heat transfer and pressure and the characteristics of the flowfield in the interaction region varied with shock strength and the level of mass addition from the transpiration-cooled section of the model. The effects of the distribution of the blowing rate along the interaction regions were also examined for a range of blowing rates through the transpiration-cooled panels. The regions of shockwave/boundary layer interaction examined in these studies were induced by oblique shocks generated with a sharp, flat plate, inclined to the freestream at angles of 5 degrees, 7.5 degrees, and 10 degrees. It was found that, in the absence of an incident shock, transpiration cooling was a very effective method for reducing both the heat transfer and the skin friction loads on the surface. The helium coolant was found to be significantly more effective than nitrogen, because of its low molecular weight and high specific heat. The studies of shock-wave/transpiration-cooled surface interaction demonstrated that the interaction region between the incident shock and the low-momentum transpiration-cooled boundary layer did not result in a significant increase in the size of attached or separated interaction regions, and did not result in significant flowfield distortions above the interaction region. The increase in heating downstream of the shock-impingement point could easily be reduced to the values without shock impingement by a relatively small increase in the transpiration cooling in this region. Surprisingly, this increase in cooling rate did not result in a significant increase in size of the region ahead of the incident shock or create a significantly enlarged interaction region with a resultant increase in the distortion level in the inviscid flow. Thus, transpiration cooling appears to be a very effective technique to cool the internal surfaces of scramjet engines, where shocks in the engine would induce large local increases in wall heating and create viscous/inviscid interactions that could significantly disturb the smooth flow through the combustor. However, if hydrogen is used as the coolant, burning upstream of shock impingement might result in localized hot spots. Clearly, further research is needed in this area.

Holden, Michael S.↗

NASA NTRS2006

Observations of Shock Diffusion and Interactions in Supersonic Freestreams with Counterflowing Jets

One of the technical challenges in long-duration space exploration and interplanetary missions is controlled entry and re-entry into planetary and Earth atmospheres, which requires the dissipation of considerable kinetic energy as the spacecraft decelerates and penetrates the atmosphere. Efficient heat load management of stagnation points and acreage heating remains a technological challenge and poses significant risk, particularly for human missions. An innovative approach using active flow control concept is proposed to significantly modify the external flow field about the spacecraft in planetary atmospheric entry and re-entry in order to mitigate the harsh aerothermal environments, and significantly weaken and disperse the shock-wave system to reduce aerothermal loads and wave drag, as well as improving aerodynamic performance. To explore the potential benefits of this approach, we conducted fundamental experiments in a trisonic blow down wind tunnel to investigate the effects of counterflowing sonic and supersonic jets against supersonic freestreams to gain a better understanding of the flow physics of the interactions of the opposing flows and the resulting shock structure.

Daso, Endwell O.↗

DOE OSTI2022

Time-resolved velocity and ion sound speed measurements from simultaneous bow shock imaging and inductive probe measurements

We present a technique to measure the time-resolved velocity and ion sound speed in magnetized, supersonic high-energy-density plasmas. We place an inductive (“b-dot”) probe in a supersonic pulsed-power-driven plasma flow and measure the magnetic field advected by the plasma. As the magnetic Reynolds number is large (R M > 10), the plasma flow advects a magnetic field proportional to the current at the load. This enables us to estimate the flow velocity as a function of time from the delay between the current at the load and the signal at the probe. The supersonic flow also generates a hydrodynamic bow shock around the probe, the structure of which depends on the upstream sonic Mach number. By imaging the shock around the probe with a Mach–Zehnder interferometer, we determine the upstream Mach number from the shock Mach angle, which we then use to determine the ion sound speed from the known upstream velocity. We use the sound speed to infer the value of Z̄T e , where Z̄ is the average ionization and T e is the electron temperature. We use this diagnostic to measure the time-resolved velocity and sound speed of a supersonic (M S ~ 8), super-Alfvénic (M A ~ 2) aluminum plasma generated during the ablation stage of an exploding wire array on the Magpie generator (1.4 MA, 250 ns). The velocity and Z̄T e measurements agree well with the optical Thompson scattering measurements reported in the literature and with 3D resistive magnetohydrodynamic simulations in GORGON.

47 OTHER INSTRUMENTATION↗

NASA NTRS1966

Torus elements used in effective shock absorber

Energy absorbing device forces torus elements to revolve annularly between two concentric tubes when a load is applied to one tube. Interference forces can be varied by using torus elements of different thicknesses. The device operates repeatedly in compression or tension, and under problems of large onset rate tolerance or structural overload.

Cunningham, P.↗

NASA NTRS1979

High speed interference heating loads and pressure distributions resulting from elevon deflections

Effects of elevon-induced three-dimensional shock-wave turbulent boundary-layer interactions on hypersonic aircraft surfaces are analyzed. Detailed surface pressure and heating rate distributions obtained on wing-elevon-fuselage models representative of aft portions of hypersonic aircraft are compared with analytical and experimental results from other sources. Examples are presented that may be used to evaluate the adequacy of current theoretical methods for estimating the effects of three-dimensional shock-wave turbulent boundary-layer interactions on hypersonic aircraft surfaces.

Johnson, C. B.↗

NASA NTRS2001

Fluctuating Pressure Data from 2-D Nozzle Cold Flow Tests (Dual Bell)

Rocket engines nozzle performance changes as a vehicle climbs through the atmosphere. An altitude compensating nozzle, ACN, is intended to improve on a fixed geometry bell nozzle that performs at optimum at only one trajectory point. In addition to nozzle performance, nozzle transient loads are an important consideration. Any nozzle experiences large transient toads when shocks pass through the nozzle at start and shutdown. Additional transient toads will occur at transitional flow conditions. The objectives of cold flow nozzle testing at MSFC are CFD benchmark / calibration and Unsteady flow / sideloads. Initial testing performed with 2-D inserts to 14" transonic wind tunnel. Recent review of 2-D data in preparation for nozzle test facility 3-D testing. This presentation shows fluctuating pressure data and some observations from 2-D dual-bell nozzle cold flow tests.

Nesman, Tomas E.↗

NASA NTRS2016

Failure Analysis and Recovery of a 50-mm Highly Elastic Intermetallic NiTi Ball Bearing for an ISS Application

Ball bearings used inside the ISS Distillation Assembly centrifuge require superior corrosion and shock resistance to withstand acidic wastewater exposure and heavy spacecraft launch related loads. These requirements challenge conventional steel bearings and provide an ideal pathfinder application for 50-mm bore, deep-groove ball bearings made from the corrosion immune and highly elastic intermetallic material 60NiTi. During early ground testing in 2014 one 60NiTi bearing unexpectedly and catastrophically failed after operating for only 200 hr. A second bearing running on the same shaft was completely unaffected. An investigation into the root cause of the failure determined that an excessively tight press fit of the bearing outer race coupled with NiTi's relatively low elastic modulus were key contributing factors. The proposed failure mode was successfully replicated by experiment. To further corroborate the root cause theory, a successful bearing life test using improved installation practices (selective fitting) was conducted. The results show that NiTi bearings are suitable for space applications provided that care is taken to accommodate their unique material characteristics.

DellaCorte, Christopher↗

NASA NTRS1996

Force limited vibration testing of Cassini spacecraft cosmic dust analyser

The testing of the cosmic dust analyzer for the Cassini mission using the force limited method in order to avoid overtesting and to verify the ability of the specimen design to withstand the loads during launch and cruise, is reported on. In order to implement the method, force gages, fixtures and a test controller are required and the test specimen is subjected to sine vibration, random vibration and half sine shock. The practical aspects of the use of the force limited method are described. Due to the high loads and the weak design of the structural element, a notching method is used which provides the possibility of limiting the excitation to flight expected levels.

Jahn, Heiko↗

NASA NTRS1990

Spring-Loaded-Bolt Locking Device

Locking device designed to clamp small objects temporarily onto object or vehicle that accelerates. Intended to be used in place of toggle clamp, which can snap out of lock during excessive shock or vibration or because of accidental contact of person or object with toggle locking handle. Device looks and operates somewhat like spring-loaded door bolt. Moderate vibrations do not cause accidental unlocking.

Calco, Frank S.↗

NASA NTRS1987

High enthalpy detailed aerothermal loads

The topics are presented in viewgraph form and include the following: motivation, approach, planned experimental activities, shock-on-lip studies, and mass addition cooling studies.

Wieting, Allan R.↗

DOE OSTI2020

Hands-on Introduction to Shock Physics (2020 Report, RR2020-IPD-1)

Los Alamos National Laboratory (LANL) researchers have played key roles in the development of high-pressure science and shock physics since the 1950s. Central to the Laboratory’s experimental capability are multiple impact systems and explosive loading capabilities that span impact velocities in excess of 8 km/s. More recently, our efforts have taken us to 3rd and 4th generation light sources where these loading platforms are coupled to X-ray diagnostics that are now providing new insights into matter-at-extremes. As the availability of these new diagnostics and loading techniques increases, there is an increased need within the high pressure community for researchers with the knowledge base and expertise to use them. Few institutions offer any level of theoretical or experimental coursework designed to train students for work in the field of shock physics or dynamic compression science, and those that do cannot meet the current demand of the national laboratories.

71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗

NASA NTRS2022

Design of Hypervelocity Flow Generator (HFG) and Its Diagnostics

Ground facilities for hypersonic research are a key element for successful development of aerodynamically proven hypervelocity vehicles. Design concepts and diagnostics of a hypervelocity flow generator (HFG) were made as a test platform for hypersonic/hypervelocity spacecraft models at the NASA Langley Research Center. The HFG is a hypersonic flow field generator using optically heated gas which is blown into an 80 cu m vacuum chamber. The vacuum chamber is kept at a stable vacuum pressure with a combination of three large vacuum pumps, while the HFG is in the test mode. The HFG provides a relatively small test section with approximately a 20 cm window. This facility was designed to generate 2.45 km/sec of flow speed, and potentially generate a continuous flow with the nozzle and vacuum system. The window in test section provides a direct view of the shock wave around a model in order to measure temperature, pressure, and density profile within the shock layer. One of the key test goals under this project is to understand why the emission spectra from the standing shockwave plasma predicted by Lora-Loran codes are significantly different from the measured emission spectra from the Flight Investigation Reentry Environment (FIRE II) Flight. The correct estimate of the thermal loading on the leading edge of hypersonic vehicles greatly affects the aerodynamic design, the material selection for the vehicle, and the cooling requirement and can be obtained by the precise modeling of emission spectra from shockwave plasma. However, the estimation of thermal loading is not an easy task due to complex non-equilibrium radiative process within high temperature shock layers that still falls into a category of cold plasma. Direct flight experiments are the most desirable, but not a cost-effective approach. Analysis by computational fluid dynamics (CFD) offers many test flexibilities. However, the CFD codes must be fully tested and validated with experimental data before the codes are effectively used for practical design. Large discrepancies between experiments and codes appear in hypersonic/hypervelocity flow regimes at high altitudes of 60 km ~ 90 km. This HFG facility offers some important parameters for CFD code validation, such as collision cross-sections, relaxation times, reaction rate coefficients and transportation coefficients. The HFG test facility is based on the ejection flow of high temperature gas heated over to 3500 K through a nozzle. The tungsten gas chamber of the HFG is heated up to a desired temperature by a 60 kW optical power beam source. This system consists of an optical power source, a thermal chamber, an expansion nozzle, a test section, and an 80 cu m vacuum tank. 60 kW optical input power is obtained from the 150 kW Vortek arc lamp system (by Vortek Industries, Vancouver, Canada). This optical beam is focused to heat the gas chamber within which a flow media is heated. The maximum achievable temperature of the flow medium reached approximately 3500 K or even higher but is limited by the melting point of the chamber material used. The exhaust velocity through the nozzle was determined by the stagnation temperature and the molecular weight of the working medium at the test section. To provide design parameters, a NASA Chemical Equilibrium with Application (CEA) computer program is used for the simulation of aerothermal data. This CEA program can calculate chemical equilibrium and properties of complex mixtures using shock tube parameters. For nitrogen gas at 2666 K stagnation temperature, the maximum achievable velocity at the test section is approximately 2.45 km/sec which is within the range of the thermal velocity of 8000 K shock layer. Based on the calculation through the CEA program, the design parameters of a HFG were determined and implemented for the test section that includes an expansion nozzle. The installed test section of the experimental facility can sustain a condition of a re-entry vehicle from the Space at an altitude of 60-90 km. After installation of the HFG, the system was fully tested and its operational parameters were measured. An 80 cu m vacuum chamber of HFG was set at 1 torr level to keep a stable downstream condition. The pumping time to reach the minimum vacuum pressure (~ 1 torr) at the test chamber from atmospheric pressure was approximately an hour using Kinney (MBV-14000/MB -1600/KT-300) and Stokes (Model 1772 and 412) pumps. Such a setup condition allowed a continued stable operation of the HFG experiment with flowrates through 1-, 3-, or 5-mm diameters of nozzle throat. The flow characteristics of the HFG for various operating conditions were performed using a focused Schlieren method. At 0.8 torr chamber pressure, a barrel shock was observed at the test section. With a sphere obstacle of 12 mm diameter in the flow, a bow shock (~ 2-mm thickness) was observed by a focused Schlieren visualization method.

Hypervelocity flow↗