Unsteady Surface Pressures on the SLS Block 1B Cargo Liftoff Configuration
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L. J. Mears↗
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A Flexible Multibody Framework for Space Launch System Liftoff with Umbilical Separations
ICPSU clearances to the lifting SLS vehicle were computed utilizing a coupled system nonlinear simulation approach in order to accurately capture the impact of system flexibilities, SLS forcing functions, pad separation (twang), ML dynamics and timings • i.e., integrate ICPSU model into a SLS nonlinear Coupled Loads Analysis (CLA) This required utilizing a flexible multibody framework for the CLA and the integration of large displacements/rotations capabilities into that CLA • Gantry rotation • Disconnect and retraction of flexible hoses
Arya Majed↗
Experimental and Computational Examination of the Coandă Effect on the Space Launch System at Liftoff Conditions
During development of an aerodynamic database to cover ground wind loads, uncertaintyquantification to account for the Coandă effect forced a closer look into how this phenomenonmanifests on the Space Launch System. Aerodynamic data collected across the life of the pro-gram is explored to look for trends and the ability to characterize not just the bounds of forcesand moments but also better understand their distributions. Experimental data collected inthe NASA Langley 14- by 22-Foot Subsonic Tunnel is used to explore integrated forces andmoments. This is followed up by a similar exploration using computational data generatedusing the Kestrel flow solver. After a survey of the data at large which confirms the existenceof the Coandă states throughout the history of the program, a few highlighted cases are usedto characterize the flow physics. This characterization is used to summarize how each Coandăstate is predicted to load the vehicle
Coandă Effect↗
Experimental and Computational Examination of the Coandă Effect on the Space Launch System at Liftoff Conditions
During development of an aerodynamic database to cover ground wind loads, uncertainty quantification to account for the Coandă effect forced a closer look into how this phenomenon manifests on the Space Launch System. Aerodynamic data collected across the life of the pro-gram is explored to look for trends and the ability to characterize not just the bounds of forces and moments but also better understand their distributions. Experimental data collected in the NASA Langley 14- by 22-Foot Subsonic Tunnel is used to explore integrated forces and moments. This is followed up by a similar exploration using computational data generated using the Kestrel flow solver. After a survey of the data at large which confirms the existence of the Coandă states throughout the history of the program, a few highlighted cases are used to characterize the flow physics. This characterization is used to summarize how each Coandă state is predicted to load the vehicle.
Coandă Effect↗
Development of a Data Fusion Methodology for Lineload Aerodynamic Databases for a Launch Vehicle during Liftoff and Transition
The need for databases for the distributed loading on launch vehicles during the early portion of flight necessitates the use of expensive computational flows in regimes where wake effects dominate. While also being expensive, this is a regime that computational tools tend to historically have problems simulating accurately. To help tackle this problem, a method of data fusion to combine computational results to wind tunnel derived force and moment data is developed. Using this method, significant reduction in computational costs and increases in confidence of the final product is possible and has been used to generate several databases for the Space Launch System (SLS) at NASA. While the full details of database generation are not part of this work, the crucial method at its core is developed here. Two SLS geometries are used throughout the work to demonstrate the techniques. These are two of the larger geometries and represent both planned crewed missions to the Moon as well as potential cargo missions to deep space. The method uses principal component analysis (PCA) to generate a reduced ordered model (ROM) to help fill in the full parameter space. Other similar techniques are explored, but were not found to have a significant result on the predictions of the ROM. Because the full number of components are kept to generate the model, this lack of difference is expected. This method is then extended to ensure that predicted surfaces match trusted force and moment data derived from wind tunnel testing. This extension is done by setting up a constrained optimization problem in order to minimize the deviation from the surface resolved computational data while still integrating to the desired values. When generating the constrained optimization problem, a weighting factor to balance these competing needs is introduced. The work compares previously introduced weighting terms from similar work to the proposed terms and shows that the previously used terms do not have as desirable behavior in this flow regime. This method is then expanded by developing a technique to incorporate uncertainty quantification into the developed data fusion methodology. This expansion takes a two pronged approach. One examines transferring the uncertainties in the force and moment database and characterizes how those adjustments change the predicted lineloads. The second looks at model form error and looks how rebuilding the model using slightly different data changes the predictions. These two terms are then combined in order to create an uncertainty model that takes both effects into account. The limitations of the proposed methods is then discussed as well as possible techniques to address these shortcomings.
Launch Vehicles↗
Liftoff Environments Overview
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Matt Casiano↗
Unsteady Surface Pressure Characterization on Launch Vehicle in Liftoff Configuration
Acquisition of unsteady pressure using remote scanners can provide practical advantages over making the same measurements directly using flush-mounted pressure transducers. However, pneumatic distortion of the signal is introduced by the tubing between the sensing module and surface port. This typically limits the usefulness of such measurements to averaged, steady data only, restricting the application of pressure scanners for unsteady pressure measurements in the wind tunnel and flight-testing applications. In this work, unsteady pressure on the surface of a 1.75%-scale model of the NASA Space Launch System Block 1B Cargo vehicle was remotely measured at the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel using an ESP pressure scanner. Subsequent reconstruction using an empirically-tuned, Wiener-filtered inverse transfer function produced excellent agreement compared to nearby surface-mounted transducers within a usable bandwidth of approximately 500 Hz. Interactions between the launch vehicle and the support tower produced strong oscillations that were absent in the vehicle-alone configuration, and it is likely that the wake interaction leads to a resonant flow response when the launch tower is immediately downstream of the vehicle.
Unsteady pressure↗
Liftoff Environments
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Brandon Riley Williams↗
Dickinsonia liftoff: Evidence of current derived morphologies
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Scott D. Evans↗
River Plume Liftoff Dynamics and Surface Expressions
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R. A. Branch↗
Forward Skirt Structural Testing on the Space Launch System (SLS) Program
Structural testing was performed to evaluate heritage forward skirts from the Space Shuttle program for use on the Space Launch System (SLS) program. One forward skirt is located in each solid rocket booster. Heritage forward skirts are aluminum 2219 welded structures. Loads are applied at the forward skirt thrust post and ball assembly. Testing was needed because SLS ascent loads are roughly 40% higher than Space Shuttle loads. Testing objectives were to determine margins of safety, demonstrate reliability, and validate analytical models. Two forward skirts were structurally tested using the test configuration. The test stand applied loads to the thrust post. Four hydraulic actuators were used to apply axial load and two hydraulic actuators were used to apply radial and tangential loads. The first test was referred to as FSTA-1 (Forward Skirt Structural Test Article) and was performed in April/May 2014. The purpose of FSTA-1 was to verify the ultimate capability of the forward skirt subjected to ascent ultimate loads. Testing consisted of two liftoff load cases taken to 100% limit load followed by an ascent load case taken to 110% limit load. The forward skirt was unloaded to no load after each test case. Lastly, the forward skirt was tested to 140% limit and then to failure using the ascent loads. The second test was referred to as FSTA-2 and performed in July/August of 2014. The purpose of FSTA-2 was to verify the ultimate capability of the forward skirt subjected to liftoff ultimate loads. Testing consisted of six liftoff load cases taken to 100% limit load followed by the six liftoff cases taken to 140% limit load. Two ascent load cases were then tested to 100% limit load. The forward skirt was unloaded to no load after each test case. Lastly, the forward skirt was tested to 140% limit and then to failure using the ascent loads. The forward skirts on FSTA-1 and FSTA-2 successfully carried all applied liftoff and ascent load cases. Both FSTA-1 and FSTA-2 were tested to failure by increasing the ascent loads. Failure occurred in the forward skirt thrust post radius. The forward skirts on FSTA-1 and FSTA-2 had nearly identical failure modes. FSTA-1 failed at 1.72 times limit load and FSTA-2 failed at 1.62 times limit load. This difference is primarily attributed to variation in material properties in the thrust post region. Test data were obtained from strain gages, deflection gages, ARAMIS digital strain measurement, acoustic emissions, and high-speed video. Strain gage data and ARAMIS strain were compared to finite element (FE) analysis predictions. Both the forward skirt and tooling were modeled. This allows the analysis to simulate the loading as close as possible to actual test configuration. FSTA-1 and FSTA-2 were instrumented with over 200 strain gages to ensure all possible failure modes could be captured. However, it turned out that three gages provided critical strain data. One was located in the post bore and two on the post radius. More gages were not specified due to space limitations and the desire to not interfere with the use of the ARAMIS system on the post radius. Measured strains were compared to analysis results for the load cycle to failure. Note that FSTA-1 gages were lost before failure was reached. FSTA-2 gages made it to the failure load but one of the radius gages was lost before testing began. This gage was not replaced because of the time and cost associated with disassembly of the test structure. Correlation to analysis was excellent for FSTA-1. FSTA-2 was not quite as good because there was more residual strain from previous load cycles. FSTA-2 was loaded and unloaded with 12 liftoff cases and two ascent cases before taking the skirt to failure. FSTA-1 only had two liftoff cases and one ascent case before taking the skirt to failure. The ARAMIS system was used to determine strain at the post radius by processing digital images of a speckled paint pattern. Digital cameras recorded images of the speckled paint pattern. ARAMIS strain results for FSTA-2 just prior to failure. Note a high strain location develops near the left side. This high strain compares well to analysis prediction for both FSTA-1 and FSTA-2. The strain at this location was also plotted versus limit load. Both FSTA-1 and FSTA-2 had excellent correlation between ARAMIS and analysis strains. Acoustic emission (AE) sensors were used to monitor for damage formation that may occur during testing (e.g., crack formation and growth or propagation). AE was very important because after disassembly of FSTA-1, a crack was observed in the ball fitting radius. The ball fitting did not crack on FSTA-2. AE data was used to reconstruct when the crack occurred. The AE energy versus time plot for FSTA. The energy increased considerably at 850 seconds (152% limit load), indicating a crack could have formed at this point. The only visual evidence found that could have corresponded to this was the crack that initiated in the ball fitting. The cracks in the forward skirt aluminum structures would likely have been lower energy due to a lower modulus and all that were found after failure correlated to occurring after the initial crack in the post radius. This was verified by high-speed cameras used to record the failure.
Lohrer, J. D.↗
Flight motor set 360L001 (STS-26R). (Reconstructed dynamic loads analysis)
A transient analysis was performed to correlate the predicted versus measured behavior of the Redesigned Solid Rocket Booster (RSRB) during Flight 360L001 (STS-26R) liftoff. Approximately 9 accelerometers, 152 strain gages, and 104 girth gages were bonded to the motors during this event. Prior to Flight 360L001, a finite element model of the RSRB was analyzed to predict the accelerations, strains, and displacements measured by this developmental flight instrumentation (DFI) within an order of magnitude. Subsequently, an analysis has been performed which uses actual Flight 360L001 liftoff loading conditions, and makes more precise predictions for the RSRB structural behavior. Essential information describing the analytical model, analytical techniques used, correlation of the predicted versus measured RSRB behavior, and conclusions, are presented. A detailed model of the RSRB was developed and correlated for use in analyzing the motor behavior during liftoff loading conditions. This finite element model, referred to as the RSRB global model, uses super-element techniques to model all components of the RSRB. The objective of the RSRB global model is to accurately predict deflections and gap openings in the field joints to an accuracy of approximately 0.001 inch. The model of the field joint component was correlated to Referee and Joint Environment Simulation (JES) tests. The accuracy of the assembled RSRB global model was validated by correlation to static-fire tests such DM-8, DM-9, QM-7, and QM-8. This validated RSRB global model was used to predict RSRB structural behavior and joint gap opening during Flight 360L001 liftoff. The results of a transient analysis of the RSRB global model with imposed liftoff loading conditions are presented. Rockwell used many gage measurements to reconstruct the load parameters which were imposed on the RSRB during the Flight 360L001 liftoff. Each load parameter, and its application, is described. Also presented are conclusions and recommendations based on the analysis of this load case and the resulting correlation between predicted and measured RSRB structural behavior.
Call, V. B.↗
Launch Vehicle Propulsion Parameter Design Multiple Selection Criteria
The optimization tool described herein addresses and emphasizes the use of computer tools to model a system and focuses on a concept development approach for a liquid hydrogen/liquid oxygen single-stage-to-orbit system, but more particularly the development of the optimized system using new techniques. This methodology uses new and innovative tools to run Monte Carlo simulations, genetic algorithm solvers, and statistical models in order to optimize a design concept. The concept launch vehicle and propulsion system were modeled and optimized to determine the best design for weight and cost by varying design and technology parameters. Uncertainty levels were applied using Monte Carlo Simulations and the model output was compared to the National Aeronautics and Space Administration Space Shuttle Main Engine. Several key conclusions are summarized here for the model results. First, the Gross Liftoff Weight and Dry Weight were 67% higher for the design case for minimization of Design, Development, Test and Evaluation cost when compared to the weights determined by the minimization of Gross Liftoff Weight case. In turn, the Design, Development, Test and Evaluation cost was 53% higher for optimized Gross Liftoff Weight case when compared to the cost determined by case for minimization of Design, Development, Test and Evaluation cost. Therefore, a 53% increase in Design, Development, Test and Evaluation cost results in a 67% reduction in Gross Liftoff Weight. Secondly, the tool outputs define the sensitivity of propulsion parameters, technology and cost factors and how these parameters differ when cost and weight are optimized separately. A key finding was that for a Space Shuttle Main Engine thrust level the oxidizer/fuel ratio of 6.6 resulted in the lowest Gross Liftoff Weight rather than at 5.2 for the maximum specific impulse, demonstrating the relationships between specific impulse, engine weight, tank volume and tank weight. Lastly, the optimum chamber pressure for Gross Liftoff Weight minimization was 2713 pounds per square inch as compared to 3162 for the Design, Development, Test and Evaluation cost optimization case. This chamber pressure range is close to 3000 pounds per square inch for the Space Shuttle Main Engine.
Shelton, Joey Dewayne↗
Time-Accurate Computational Fluid Dynamics Simulation of a Pair of Moving Solid Rocket Boosters
Since the Columbia accident, the threat to the Shuttle launch vehicle from debris during the liftoff timeframe has been assessed by the Liftoff Debris Team at NASA/MSFC. In addition to engineering methods of analysis, CFD-generated flow fields during the liftoff timeframe have been used in conjunction with 3-DOF debris transport methods to predict the motion of liftoff debris. Early models made use of a quasi-steady flow field approximation with the vehicle positioned at a fixed location relative to the ground; however, a moving overset mesh capability has recently been developed for the Loci/CHEM CFD software which enables higher-fidelity simulation of the Shuttle transient plume startup and liftoff environment. The present work details the simulation of the launch pad and mobile launch platform (MLP) with truncated solid rocket boosters (SRBs) moving in a prescribed liftoff trajectory derived from Shuttle flight measurements. Using Loci/CHEM, time-accurate RANS and hybrid RANS/LES simulations were performed for the timeframe T0+0 to T0+3.5 seconds, which consists of SRB startup to a vehicle altitude of approximately 90 feet above the MLP. Analysis of the transient flowfield focuses on the evolution of the SRB plumes in the MLP plume holes and the flame trench, impingement on the flame deflector, and especially impingment on the MLP deck resulting in upward flow which is a transport mechanism for debris. The results show excellent qualitative agreement with the visual record from past Shuttle flights, and comparisons to pressure measurements in the flame trench and on the MLP provide confidence in these simulation capabilities.
Strutzenberg, Louise L.↗
Acoustics Research of Propulsion Systems
The liftoff phase induces some of the highest acoustic loading over a broad frequency for a launch vehicle. These external acoustic environments are used in the prediction of the internal vibration responses of the vehicle and components. Thus, predicting these liftoff acoustic environments is critical to the design requirements of any launch vehicle but there are challenges. Present liftoff vehicle acoustic environment prediction methods utilize stationary data from previously conducted hold-down tests; i.e. static firings conducted in the 1960's, to generate 1/3 octave band Sound Pressure Level (SPL) spectra. These data sets are used to predict the liftoff acoustic environments for launch vehicles. To facilitate the accuracy and quality of acoustic loading, predictions at liftoff for future launch vehicles such as the Space Launch System (SLS), non-stationary flight data from the Ares I-X were processed in PC-Signal in two forms which included a simulated hold-down phase and the entire launch phase. In conjunction, the Prediction of Acoustic Vehicle Environments (PAVE) program was developed in MATLAB to allow for efficient predictions of sound pressure levels (SPLs) as a function of station number along the vehicle using semiempirical methods. This consisted, initially, of generating the Dimensionless Spectrum Function (DSF) and Dimensionless Source Location (DSL) curves from the Ares I-X flight data. These are then used in the MATLAB program to generate the 1/3 octave band SPL spectra. Concluding results show major differences in SPLs between the hold-down test data and the processed Ares IX flight data making the Ares I-X flight data more practical for future vehicle acoustic environment predictions.
Gao, Ximing↗
Liquid oxygen sloshing in Space Shuttle External Tank
This paper describes a numerical simulation of the hydrodynamics within the liquid oxygen tank of the Space Shuttle External Tank during liftoff. Before liftoff, the tank is filled with liquid oxygen (LOX) to approximately 97 percent with the other 3 percent containing gaseous oxygen (GOX) and helium. During liftoff, LOX is drained from the bottom of the tank, and GOX is pumped into the tank's ullage volume. There is a delay of several seconds before the GOX reaches the tank which causes the ullage pressure to decrease for several seconds after liftoff; this pressure 'slump' is a common phenomenon in rocket propulsion. When four slosh baffles were removed from the tank, the ullage gas pressure dropped more rapidly than in all previous flights. The purpose of this analysis was to determine whether the removal of the baffles could have caused the increased pressure 'slump' by changing the LOX surface dynamics. The results show that the LOX surface undergoes very high vertical accelerations (up to 5 g) and, therefore, splashing almost certainly occurs. The number of baffles does not affect the surface if the structural motion is assumed; but, the number of baffles may affect the structural motion of the tank.
Kannapel, M. D.↗
Apollo 11 Launch
This NASA Kennedy Space Center video release presents the countdown and liftoff of Apollo 11, the first manned journey to the Moon which began at Pad A, Launch Complex 39, Kennedy Space Center, Florida at 9:32 a.m. EDT on July 16, 1969. The crew of Apollo 11 included Commander Neil A. Armstrong, Command Module pilot Michael Collins, and Lunar Module pilot Edwin E. Aldrin, Jr. Several different camera viewpoints of the spacecraft as well as over-head shots of the Kennedy launch control center are presented prior to liftoff. Other footage includes shots of President Lyndon B. Johnson and his wife among the Florida audience viewing liftoff. During the countdown several audio updates from Kennedy launch control are presented as to the status of pre-launch testing and system readiness. Captivating footage from liftoff to the spacecraft nearing the outer Earth atmosphere is shown as the video ends with Neil Armstrong's confirmation of engine skirt separation and launch escape tower separation from the spacecraft.
Source record↗
Analysis of Parallel Burn Without Crossfeed TSTO RLV Architectures and Comparison to Parallel Burn With Crossfeed and Series Burn Architectures
There are currently three dominant TSTO class architectures. These are Series Burn (SB), Parallel Burn with crossfeed (PBw/cf), and Parallel Burn without crossfeed (PBncf). The goal of this study was to determine what factors uniquely affect PBncf architectures, how each of these factors interact, and to determine from a performance perspective whether a PBncf vehicle could be competitive with a PBw/cf or SB vehicle using equivalent technology and assumptions. In all cases, performance was evaluated on a relative basis for a fixed payload and mission by comparing gross and dry vehicle masses of a closed vehicle. Propellant combinations studied were LOX: LH2 propelled orbiter and booster (HH) and LOX: Kerosene booster with LOX: LH2 orbiter (KH). The study conclusions were: 1) a PBncf orbiter should be throttled as deeply as possible after launch until the staging point. 2) a detailed structural model is essential to accurate architecture analysis and evaluation. 3) a PBncf TSTO architecture is feasible for systems that stage at mach 7. 3a) HH architectures can achieve a mass growth relative to PBw/cf of < 20%. 3b) KH architectures can achieve a mass growth relative to Series Burn of < 20%. 4) center of gravity (CG) control will be a major issue for a PBncf vehicle, due to the low orbiter specific thrust to weight ratio and to the position of the orbiter required to align the nozzle heights at liftoff. 5 ) thrust to weight ratios of 1.3 at liftoff and between 1.0 and 0.9 when staging at mach 7 appear to be close to ideal for PBncf vehicles. 6) performance for all vehicles studied is better when staged at mach 7 instead of mach 5. The study showed that a Series Burn architecture has the lowest gross mass for HH cases, and has the lowest dry mass for KH cases. The potential disadvantages of SB are the required use of an air-start for the orbiter engines and potential CG control issues. A Parallel Burn with crossfeed architecture solves both these problems, but the mechanics of a large bipropellant crossfeed system pose significant technical difficulties. Parallel Burn without crossfeed vehicles start both booster and orbiter engines on the ground and thus avoid both the risk of orbiter air-start and the complexity of a crossfeed system. The drawback is that the orbiter must use 20% to 35% of its propellant before reaching the staging point. This induces a weight penalty in the orbiter in order to carry additional propellant, which causes a further weight penalty in the booster to achieve the same staging point. One way to reduce the orbiter propellant consumption during the first stage is to throttle down the orbiter engines as much as possible. Another possibility is to use smaller or fewer engines. Throttling the orbiter engines soon after liftoff minimizes CG control problems due to a low orbiter liftoff thrust, but may result in an unnecessarily high orbiter thrust after staging. Reducing the number or size of engines size may cause CG control problems and drift at launch. The study suggested possible methods to maximize performance of PBncf vehicle architectures in order to meet mission design requirements.
Smith, Garrett↗