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NASA NTRS2001

Characterization of Pressure Sensitive Paint Intrusiveness Effects on Aerodynamic Data

One effect of using pressure sensitive paint (PSP) is the potential intrusiveness to the aerodynamic characteristics of the model. The paint thickness and roughness may affect the pressure distribution. and therefore, the forces and moments on the wind tunnel model. A study of these potential intrusive effects was carried out at NASA Langley Research Center where a series of wind tunnel tests were conducted using the Modem Design of Experiments (MDOE) test approach. The PSP effects on the integrated forces were measured on two different models at different test conditions in both the Low Turbulence Pressure Tunnel (LTPT) and the Unitary Plan Wind Tunnel (UPWT) at Langley. The paint effect was found to be very small over a range of Reynolds numbers, Mach numbers and angles of attack. This is due to the very low surface roughness of the painted surface. The surface roughness, after applying the NASA Langley developed PSP, was lower than that of the clean wing. However, the PSP coating had a localized effects on the pressure taps, which leads to an appreciable decrease in the pressure tap reading.

Amer, Tahani R.↗

NASA NTRS1946

Plane and Three-Dimensional Flow at High Subsonic Speeds

For two- and three-dimensional flow in a compressible medium, a simple relation is given by which, to a first approximation, the quantitative influence of compressibility upon the velocities and pressures can be understood in a clear manner. In the application of this relation the distinct behaviors of two-dimensional and axially symmetric three-dimensional flow with increasing Mach number are brought out. For slender elliptic cylinders and ellipsoids of revolution, calculations are made of the critical Mach number; that is, the Mach number at which local sonic velocity is achieved on the body. As a further example, the lifting wing of finite span is considered, and it is shown that the increase of wing lift with Mach number at a given angle of attack is greatly dependent upon the aspect ratio b(exp 2)/F.

Gothert, B.↗

NASA NTRS2015

Description, Usage, and Validation of the MVL-15 Modified Vortex Lattice Analysis Capability

MVL-15 is the most recent version of the Modified Vortex-Lattice (MVL) code developed within the Aerodynamics Systems Analysis Branch (ASAB) at NASA LaRC. The term "modified" refers to the primary modification of the core vortex-lattice methodology: inclusion of viscous aerodynamics tables that are linked to the linear solution via iterative processes. The inclusion of the viscous aerodynamics inherently converts the MVL-15 from a purely analytic linearized method to a semi-empirical blend which retains the rapid execution speed of the linearized method while empirically characterizing the section aerodynamics at all spanwise lattice points. The modification provides a means to assess non-linear effects on lift that occur at angles of attack near stall, and provides a means to determine the drag associated with the application of design strategies for lift augmentation such as the use of flaps or blowing. The MVL-15 code is applicable to the analyses of aircraft aerodynamics during cruise, but it is most advantageously applied to the analysis of aircraft operating in various high-lift configurations. The MVL methodology has been previously conceived and implemented; the initial concept version was delivered to the ASAB in 2001 (van Dam, C.), subsequently revised (Gelhausen, P. and Ozoroski, T. 2002 / AVID Inc., Gelhausen, P., and Roberts, M. 2004), and then overhauled (Ozoroski, T., Hahn, A. 2008). The latest version, MVL-15 has been refined to provide analysis transparency and enhanced to meet the analysis requirements of the Environmentally Responsible Aviation (ERA) Project. Each revision has been implemented with reasonable success. Separate applications of the methodology are in use, including a similar in-house capability, developed by Olson, E. that is tailored for structural and acoustics analyses. A central premise of the methodology is that viscous aerodynamic data can be associated with analytic inviscid aerodynamic results at each spanwise wing section, thereby providing a pathway to map viscous data to the inviscid results. However, a number of factors can sidetrack the analysis consistency during various stages of this process. For example, it should be expected that the final airplane lift curve and drag polar results depend strongly on the geometry and aerodynamics of the airfoil section; however, flap deflections and flap chord extensions change the local reference geometry of the input airfoil, the airplane wing, the tabulated non-dimensional viscous aerodynamics, and the spanwise links between the linear and the viscous aerodynamics. These changes also affect the bound circulation and therefore, calculation and integration of the induced angle of attack and induced drag. MVL-15 is configured to ensure these types of challenges are properly addressed. This report is a comprehensive manual describing the theory, use, and validation of the MVL-15 analysis tool. Section 3 summarizes theoretical, procedural, and characteristic features of MVL-15, and includes a list of the files required to setup, execute, and summarize an analysis. Section 4, Section 5, Section 6, and Section 7 combine to comprise the User's Guide portions of this report. The MVL-15 input and output files are described in Section 4 and Section 5, respectively; the descriptions are supplemented with example files and information about the file formats, parameter definitions, and typical parameter values. Section 6 describes the Wing Geometry Setup Utility and the 2d-Variants Utility files that simplify and assist setting up a consistent set of MVL-15 geometry and aerodynamics input parameters and input files. Section 7 describes the use of the 3d-Results Presentation Utility file that can be used to automatically create summary tables and charts from the MVL-15 output files. Section 8 documents the Validation Results of an extensive and varied validation test matrix, including results of an airplane analysis representative of the ERA Program. A start-to-finish example of the airplane analysis procedure is described in Section 7.

Ozoroski, Thomas A.↗

NASA NTRS1982

Experimental investigation of gamma effects on heat transfer to a 0.006 scale Shuttle Orbiter at Mach 6

Real-gas effects on heating to the windward centerline of the Shuttle Orbiter and shock shape were simulated by testing a 0.006-scale Orbiter model in Mach 6 air and CF4 flows. The range of free-stream Reynolds number, based on model length, was 0.4 x 10 to the 6th to 5.3 x 10 to the 6th for air, resulting in laminar and turbulent heating. The Reynolds number for CF4 was 0.3 x 10 to the 6th and gamma was equal to 1.13, corresponding to a normal shock density ratio of 12 as compared to 5.25 in air. Angle of attack was varied from 15 deg to 45 deg in 5 deg increments. Centerline heating distributions were compared for the two gases, and to Mach 8 wind-tunnel data, predictions, and flight (STS-1 and STS-2) data. The present results revealed a significant increase in heating with increasing density ratio and decreasing Reynolds number at the lower angles of attack; the effects of density ratio and Reynolds number diminished with increasing angle of attack and were relatively small at the angles of attack for the hypersonic portion of Space Shuttle Orbiter flights STS-1 and STS-2. Wind-tunnel and flight heating distributions were in relatively good agreement, and the present heating rates in air were accurately predicted by a local infinite swept cylinder analysis.

Miller, C. G.↗

NASA NTRS1984

Locally linearized longitudinal and lateral-directional aerodynamic stability and control derivaties for the X-29A aircraft

The locally linearized longitudinal and lateral-directional aerodynamic stability and control derivatives for the X-29A aircraft were calculated for altitudes ranging from sea level to 50,000 ft, Mach numbers from 0.2 to 1.5, and angles of attack from -5 deg to 25 deg. Several other parameters were also calculated, including aerodynamic force and moment coefficients, control face position, normal acceleration, static margin, and reference angle of attack.

Budd, G. D.↗

NASA NTRS1987

Experimental study of three-dimensional separated flow surrounding a hemisphere-cylinder at incidence

Detailed experimental surface pressure coefficient measurements, obtained for a hemisphere-cylinder-flare model in a low supersonic flow (freestream M = 1.2) at various angles of attack (0 deg to 27.5 deg), have been analyzed. The pressure values for each angle of attack were smoothed and checked against their respective oil-flow photographs. The smoothed results were then used to validate a theory which relates the number and type of singular points observed in the oil-flow patterns with local surface-pressure extrema.

Meade, Andrew J., Jr.↗

NASA NTRS1992

Flow visualizations of perpendicular blade vortex interactions

Helium bubble flow visualizations have been performed to study perpendicular interaction of a turbulent trailing vortex and a rectangular wing in the Virginia Tech Stability Tunnel. Many combinations of vortex strength, vortex-blade separation (Z(sub s)) and blade angle of attack were studied. Photographs of representative cases are presented. A range of phenomena were observed. For Z(sub s) greater than a few percent chord the vortex is deflected as it passes the blade under the influence of the local streamline curvature and its image in the blade. Initially the interaction appears to have no influence on the core. Downstream, however, the vortex core begins to diffuse and grow, presumably as a consequence of its interaction with the blade wake. The magnitude of these effects increases with reduction in Z(sub s). For Z(sub s) near zero the form of the interaction changes and becomes dependent on the vortex strength. For lower strengths the vortex appears to split into two filaments on the leading edge of the blade, one passing on the pressure and one passing on the suction side. At higher strengths the vortex bursts in the vicinity of the leading edge. In either case the core of its remnants then rapidly diffuse with distance downstream. Increase in Reynolds number did not qualitatively affect the flow apart from decreasing the amplitude of the small low-frequency wandering motions of the vortex. Changes in wing tip geometry and boundary layer trip had very little effect.

Rife, Michael C.↗

NASA NTRS1993

Flight and wind-tunnel calibrations of a flush airdata sensor at high angles of attack and sideslip and at supersonic Mach numbers

A nonintrusive airdata-sensing system was calibrated in flight and wind-tunnel experiments to an angle of attack of 70 deg and to angles of sideslip of +/- 15 deg. Flight-calibration data have also been obtained to Mach 1.2. The sensor, known as the flush airdata sensor, was installed on the nosecap of an F-18 aircraft for flight tests and on a full-scale F-18 forebody for wind-tunnel tests. Flight tests occurred at the NASA Dryden Flight Research Facility, Edwards, California, using the F-18 High Alpha Research Vehicle. Wind-tunnel tests were conducted in the 30- by 60-ft wind tunnel at the NASA LaRC, Hampton, Virginia. The sensor consisted of 23 flush-mounted pressure ports arranged in concentric circles and located within 1.75 in. of the tip of the nosecap. An overdetermined mathematical model was used to relate the pressure measurements to the local airdata quantities. The mathematical model was based on potential flow over a sphere and was empirically adjusted based on flight and wind-tunnel data. For quasi-steady maneuvering, the mathematical model worked well throughout the subsonic, transonic, and low supersonic flight regimes. The model also worked well throughout the angles-of-attack and -sideslip regions studied.

Moes, Timothy R.↗

NASA NTRS1993

Flight and wind-tunnel calibrations of a flush airdata sensor at high angles of attack and sideslip and at supersonic Mach numbers

Moes, Timothy R.↗

NASA NTRS1996

Subsonic Analysis of 0.04-Scale F-16XL Models Using an Unstructured Euler Code

The subsonic flow field about an F-16XL airplane model configuration was investigated with an inviscid unstructured grid technique. The computed surface pressures were compared to wind-tunnel test results at Mach 0.148 for a range of angles of attack from 0 deg to 20 deg. To evaluate the effect of grid dependency on the solution, a grid study was performed in which fine, medium, and coarse grid meshes were generated. The off-surface vortical flow field was locally adapted and showed improved correlation to the wind-tunnel data when compared to the nonadapted flow field. Computational results are also compared to experimental five-hole pressure probe data. A detailed analysis of the off-body computed pressure contours, velocity vectors, and particle traces are presented and discussed.

Lessard, Wendy B.↗

NASA NTRS1995

Acoustic flight tests of rotorcraft noise-abatement approaches using local differential GPS guidance

This paper presents the test design, instrumentation set-up, data acquisition, and the results of an acoustic flight experiment to study how noise due to blade-vortex interaction (BVI) may be alleviated. The flight experiment was conducted using the NASA/Army Rotorcraft Aircrew Systems Concepts Airborne Laboratory (RASCAL) research helicopter. A Local Differential Global Positioning System (LDGPS) was used for precision navigation and cockpit display guidance. A laser-based rotor state measurement system on board the aircraft was used to measure the main rotor tip-path-plane angle-of-attack. Tests were performed at Crows Landing Airfield in northern California with an array of microphones similar to that used in the standard ICAO/FAA noise certification test. The methodology used in the design of a RASCAL-specific, multi-segment, decelerating approach profile for BVI noise abatement is described, and the flight data pertaining to the flight technical errors and the acoustic data for assessing the noise reduction effectiveness are reported.

Chen, Robert T. N.↗

NASA NTRS1981

Aerodynamic heating on the corrugated surface of a 10.2 deg half-angle blunted cone at Mach 6.7

A 10.2 deg half-angle blunted cone with corrugated surfaces was tested in the Langley 8-foot high-temperature structures tunnel to measure the aerodynamic heating of its surfaces. The tests were made in a turbulent boundary layer at angles of attack of 0 deg, 5 deg, and 10 deg. Heating of the windward side was in reasonable agreement with theoretical turbulent predictions for a smooth cone, while heating on the leeward side was between laminar and turbulent predictions as a result of local transitional flow or flow separation produced by high lee-side pressures. Localized heating measurements indicated a significant increase in heating at large cross-flow angles, with the maximum heating rates occurring where the flow reattaches on the upstream side of the corrugation crest and the minimum occurring on the downstream side where the flow is separated.

Weinstein, I.↗

NASA NTRS1990

Discrete-vortex model for the symmetric-vortex flow on cones

A relatively simple but accurate potential flow model was developed for studying the symmetric vortex flow on cones. The model is a modified version of the model first developed by Bryson, in which discrete vortices and straight-line feeding sheets were used to represent the flow field. It differs, however, in the zero-force condition used to position the vortices and determine their circulation strengths. The Bryson model imposed the condition that the net force on the feeding sheets and discrete vortices must be zero. The proposed model satisfies this zero-force condition by having the vortices move as free vortices, at a velocity equal to at the local crossflow velocity at their centers. When the free-vortex assumption is made, a solution is obtained in the form of two nonlinear algebraic equations that relate the vortex center coordinates and vortex strengths to the cone angle and angle of attack. The vortex center locations calculated using the model are in good agreement with experimental values. The cone normal forces as well as center locations are in good agreement with the vortex cloud method of calculating symmetric flow fields.

Gainer, Thomas G.↗

NASA NTRS1979

Aircraft wake flow effect and horizontal tail buffet

As part of a program to investigate the fluctuating pressure distribution and response behavior of a fighter aircraft in transonic maneuver, an F-5A scale model has previously been tested in an 11-ft transonic wind tunnel. The model, with a number of static and dynamic pressure transducers imbedded in the lifting surfaces was tested at various angles of attack up to 16 deg. In this paper, test results of particular interest to wake flow and horizontal tail buffet are described. It is shown that the dynamic pressure data on the tail surface at the specified flight conditions serve to determine the local dynamic loads. They also influence the control performance of the aircraft under maneuver conditions where buffet is encountered. The data presented demonstrate a number of contributing factors that affect the tail dynamic pressures in the transonic regime.

Hwang, C.↗

DOE OSTI2023

Precursor Analysis Report: SQL Slammer Worm Infection of Davis-Besse Nuclear Power Plant 2003

The SQL Slammer Worm Infection of Davis-Besse Nuclear Power Plant 2003 Precursor Analysis Report leverages publicly available information about Davis-Besse’s 2003 cyber attack and catalogs anomalous observables for each technique employed in the attack. This analysis is based upon the methodology of the Cybersecurity for the Operational Technology Environment (CyOTE) program. On 25 January 2003, the SQL Slammer worm infected more than 90% of vulnerable hosts and crashed the internet in 10 to 15 minutes, making it one of the fastest spreading worms in history. SQL Slammer is a fileless, memory-resident worm that remotely exploits a stack-based buffer overflow vulnerability on local hosts to intensively scan and rapidly self-propagate across the internet. The worm infected approximately 300,000 unpatched hosts running Microsoft Structured Query Language (SQL) Server 2000 or Microsoft Desktop Engine (MSDE) 2000 with SQL Server Resolution Service. The SQL Slammer worm indirectly infected FirstEnergy’s Davis-Besse nuclear power plant by first infecting a consultant’s company network server and then propagating through an external misconfigured connection into Davis-Besse’s site network. The infection caused major network congestion, slow performance, data overloads, and the inability of local hosts to communicate with each other, which eventually caused a loss of availability and a loss of view when the Safety Parameter Display System (SPDS) and Plant Process Computer (PPC) crashed. At the time of the infection, the plant was already offline, the digital monitoring systems had redundant analog backups, and the plant control and safety functions were not affected, so there were no concerns of a safety breach. However, this incident resulted in many lessons learned and spawned important discussions about cybersecurity’s role in nuclear safety and electric power reliability regulation, policy, and guidance. Researchers and analysts identified 10 unique techniques utilized during the attack with a total of 640 observables using MITRE ATT&CK® for Industrial Control Systems. The CyOTE program assesses observables accompanying techniques used prior to the triggering event to identify opportunities to detect malicious activity. If observables accompanying the attack techniques are perceived and investigated prior to the triggering event, earlier comprehension of malicious activity can take place. Eight of the identified techniques used during Davis-Besse cyber attack were precursors to the triggering event. Analysis identified 596 observables associated with these precursor techniques, 428 of which were assessed to have an increased likelihood of being perceived in the 331 days preceding the triggering event. The response and comprehension time could have been reduced if the observables had been identified earlier. The information gathered in this report contributes to a library of observables tied to a repository of artifacts, data sources, and technique detection references for practitioners and developers to support the comprehension of indicators of attack. Asset owners and operators can use these products if they experience similar observables or to prepare for comparable scenarios.

45 MILITARY TECHNOLOGY, WEAPONRY, AND NATIONAL DEF↗

NASA NTRS2001

Afterbody Heating Characteristics of a Proposed Mars Sample Return Orbiter

Aeroheating wind-tunnel tests were conducted on a 0.028 scale model of an orbiter concept considered for a possible Mars sample return mission. The primary experimental objectives were to characterize hypersonic near wake closure and determine if shear layer impingement would occur on the proposed orbiter afterbody at incidence angles necessary for a Martian aerocapture maneuver. Global heat transfer mappings, surface streamline patterns, and shock shapes were obtained in the NASA Langley 20-inch Mach 6 Air and CF4 Tunnels for post-normal shock Reynolds numbers (based on forebody diameter) ranging from 1,400 to 415,000, angles of attack ranging from -5 to 10 degrees at 0, 3, and 6 deg sideslip, and normal-shock density, ratios of 5 and 12. Laminar, transitional, and turbulent shear layer impingement on the cylindrical afterbody was inferred from the measurements and resulted in a localized heating maximum that ranged from 40 to 75% of the reference forebody stagnation point heating. Comparison of laminar heating prediction to experimental measurement along the orbiter afterbody highlight grid alignment challenges associated with numerical simulation of three-dimensional separated wake flows.

Horvath, Thomas J.↗

NASA NTRS2002

Wake Closure Characteristics and Afterbody Heating on a Mars Sample Return Orbiter

Aeroheating wind-tunnel tests were conducted on a 0.028 scale model of an orbiter concept considered for a possible Mars sample return mission. The primary experimental objectives were to characterize hypersonic near wake closure and determine if shear layer impingement would occur on the proposed orbiter afterbody at incidence angles necessary for a Martian aerocapture maneuver. Global heat transfer mappings, surface streamline patterns, and shock shapes were obtained in the NASA Langley 20-Inch Mach 6 Air and CF4 Tunnels for post-normal shock Reynolds numbers (based on forebody diameter) ranging from 1,400 to 415,000, angles of attack ranging from -5 to 10 degrees at 0, 3, and 6 degree sideslip, and normal-shock density ratios of 5 and 12. Laminar, transitional, and turbulent shear layer impingement on the cylindrical afterbody was inferred from the measurements and resulted in a localized heating maximum that ranged from 40 to 75 percent of the reference forebody stagnation point heating. Comparison of laminar heating prediction to experimental measurement along the orbiter afterbody highlight grid alignment challenges associated with numerical simulation of three- dimensional separated wake flows. Predicted values of a continuum breakdown parameter revealed significant regions of non-continuum flow downstream of the flow separation at the MSRO shoulder and in the region of the reattachment shock on the afterbody. The presence of these regions suggest that the Navier-Stokes predictions at the laminar wind-tunnel condition may encounter errors in the numerical calculation of the wake shear layer development and impingement due to non-continuum effects.

Horvath, Thomas J.↗

DOE OSTI2023

Reinforcement Learning Approach to Cybersecurity in Space (RELACSS)

Securing satellite groundstations against cyber-attacks is vital to national security missions. However, these cyber threats are constantly evolving. As vulnerabilities are discovered and patched, new vulnerabilities are discovered and exploited. In order to automate the process of discovering existing vulnerabilities and the means to exploit them, a reinforcement learning framework is presented in this report. We demonstrate that this framework can learn to successfully navigate an unknown network and detect nodes of interest despite the presence of a moving target defense. The agent then exfiltrates a file of interest from the node as quickly as possible. This framework also incorporates a defensive software agent that learns to impede the attacking agents progress. This setup allows for the agents to work against each other and improve their abilities. We anticipate that this capability will help uncover unforeseen vulnerabilities and the means to mitigate them. The modular nature of the framework enables users to swap out learning algorithms and modify the reward functions in order to adapt the learning tasks to various use cases and environments. Several algorithms, viz., tabular Q learning, deep Q networks, proximal policy optimization, advantage actor-critic, generative adversarial imitation learning, are explored for the agents and the results highlighted. The agent learns to solve the tasks in a light-weight abstract environment. Once the agent learns to perform sufficiently well, it can be deployed in a minimega virtual machine environment (or a real network) with wrappers that map abstract actions to software commands. The agent also uses a local representation of the actions called a ‘slot-mechanism’. This allows the agent to learn in a certain network and generalize it to different networks. The defensive agent learns to predict the actions taken by an offensive agent and uses that information to anticipate the threat. This information can then either be used to raise an alarm or to take actions to thwart the attack. We believe that with the appropriate reward design, a representative environment, and action set, this framework can be generalized to tackle other cybersecurity tasks. By sufficiently training these agents, we can anticipate vulnerabilities leading to robust future designs. We can also deploy automated defensive agents that can help secure satellite groundstation and their vital national security missions.

45 MILITARY TECHNOLOGY, WEAPONRY, AND NATIONAL DEF↗