Future Launch Vehicles Research Articles

Towards robotic on-orbit assembly of large space telescopes: Mission architectures, concepts, and analyses

Imaging systems larger than the James Webb Space Telescope will enable unprecedented astronomy in the future. Larger aperture space-based imagers will also support next-generation Earth Observation missions for national security and disaster monitoring. However, launching and operating such large telescopes in the poses several practical challenges. A key challenge is that mirrors with apertures larger than 3 m cannot be monolithically manufactured so a segmented design is utilized to achieve primary mirror apertures. Even if larger mirrors could be made, it would be impossible to stow them in fairings of most existing and future launch vehicles. Folded-wing designs to deploy segmented primary mirrors, as done with the James Webb Space Telescope, are one approach to overcoming this volumetric challenge but inappropriate for 25 m apertures considered in this study. This paper presents the concept of operations and mission architectures/analysis for on-orbit assembly of a 25 m aperture telescope operating in the visible waveband of the electromagnetic spectrum capable of 1 m spatial resolution from geostationary orbit. Further, a technology demonstration roadmap towards maturing the robotic assembly technology stack is then presented as precursors to the 25 m imaging system.

Experimental Investigation on Transmission of Sound in To Metallic Enclosures During Static Firing Tests Of Solid Rocket Motor

Satellite launch vehicle during lift-off generates intense dynamic environment mainly the acoustic noise, which in turn acts on the nearby launch pad systems. Reliable operation of these launch pad ground systems needs to be ensured by safeguarding against these dynamic loads with adequate protection measures. For future launch vehicle programs where the launch turnaround time is a crucial aspect, sensitive systems like ground check-out electronic packages are planned to be placed adjacent to the launch vehicle mounting pedestal on the launch pad to ensure minimal time for preparation. The enclosures housing these electronic packages needs to be configured to minimize the transmission of acoustic levels being imposed from the lift-off induced noise. In the present work, an experimental investigation is carried out on acoustic transmission characteristics of metallic enclosures of different material and thickness during the static firing tests of solid rocket motors. Acoustics generated from the supersonic jet noise has a wide frequency band from 100 to 10,000 Hz with sound power peaking near 1,000 Hz. The electronic components and packages, which are generally vulnerable to the high frequency noise, need to be protected from this lift-off induced acoustics. The enclosures are placed in the far field of this jet noise and the transmission loss measurements are obtained. Variation of transmission loss with material and thickness of the enclosures are studied with specific attention to the internal acoustic modes and the coincidence phenomena. Subsequently, the enclosures treated with acoustic blankets of different thickness are investigated to verify its influence on internal acoustic modes and also the additional reduction in transmission of sound in the dominant frequency band of supersonic jet noise.

ARMA time series prediction model for fault detection of launch vehicle

Spaceflight is a high-risk activity, especially for launch vehicles, and the losses caused by launch failures are considerable. Real-time fault detection is a prerequisite to ensure that future launch vehicles can detect and isolate faults in a timely manner, and ultimately reduce the hazard of fault. In this article, a time series prediction model of axial overload and three-axis angular rate is established using ARMA (Auto Regressive Moving Average) time series prediction model, and the residuals of the prediction model are analyzed and combined with the fault determination decision strategy to determine whether the system has a fault, and a fault tree can be built to further locate the fault location.

Cost estimation for launch vehicle families considering uncertain market scenarios

When optimizing the staging of a launch vehicle, usually a reference mission with a fixed payload mass is used to determine the specific launch cost (or a surrogate thereof, e.g. total mass or dry mass). While this approach neglects suboptimal loading, it delivers sufficiently good results for the evaluation of an individual launcher if the design mission is well known. However, this approach does not capture the full range of payloads that a launch vehicle is expected to launch during its operational lifetime. In order to estimate the cost and subsequently optimize the staging of such launch vehicle families, a full launch scenario has to be considered. The cost of an individual launch vehicle loses importance and the deciding factor is the total cost associated with launching the specified scenario. However, estimates of the future market inevitably include large uncertainties. These uncertainties and their impact are ideally considered and quantified when assessing the specific recurring cost of a single launch vehicle. The consideration of a whole launch market scenario is especially relevant when evaluating a family of launch vehicles, an option currently being discussed for future European launchers. In this case, the distribution of the payloads onto the different members of the launch vehicle family has to be optimized in order to reach the lowest possible cost. Within this manuscript, a method that allows the determination of the recurring cost of such a launcher family based on a launch market scenario including uncertainties is proposed. Based on the launch numbers derived from the combinatorial optimization, parametric cost models for recurring and non-recurring cost are adapted and applied to three possible future launch vehicle families.

The roar of the rocket: A hot topics discussion

An increasing number of companies and countries are building and launching space vehicles of different sizes and configurations for diverse purposes. The goal of this Hot Topics presentation is two-fold: first, to introduce a wide-ranging audience of Acoustical Society of America members to principles of rocket noise radiation, propagation, and reception; second, to discuss needs and challenges related to future launch vehicle acoustics research. If attendees leave both mildly entertained and better informed, the presentation will have been a success.

Energy efficient large-scale storage of liquid hydrogen

The world’s largest liquid hydrogen storage tanks were constructed in the mid-1960s at the NASA Kennedy Space Center. These two vacuum-jacketed, perlite powder insulated tanks, still in service today, have 3,200 m3 of useable capacity. In 2018, construction began on an additional storage tank at Launch Complex 39B. This new tank will give an additional storage capacity of 4,700 m3 for a total on-site storage capacity of roughly 8,000 m3. NASA’s new Space Launch System (SLS) heavy lift rocket for the Artemis program includes an LH2 tank that makes up the bulk of the vehicle, holding 2,033 m3 of LH2 in its 8.4-m diameter by 40-m height. The new storage tank includes two new energy-efficient technologies: a glass bubbles insulation system in lieu of perlite, and an Integrated Refrigeration and Storage (IRAS) heat exchanger for controlled storage capability. The evacuated glass bubbles insulation system is based on the prior two decades of research to prove the thermal performance benefits as well as the mechanical and vacuum integrity; and has been shown to reduce LH2 boiloff by 46% versus perlite in field demonstrations. The IRAS capability is centered on a heat exchanger system that is built within the inner vessel to reject heat from the bulk liquid through the future implementation of an external helium refrigerator. Controlled storage via IRAS, when fully implemented, will provide full control of the ullage pressure, zero boiloff, and even production of densified LH2 pending its adoption on future launch vehicles. The design basics are described along with main construction and testing processes involved. The key features of the new technology items and implications on simplified operations and long-term energy savings are addressed.

Analysis for design optimization of high thrust liquid engine hot test facility

A high thrust liquid oxygen-kerosene fuelled engine is being developed for future launch vehicle applications. Comprehensive hot testing of engine at sea level conditions is required for qualification and acceptance of flight engine and launch vehicle stage. The engine and stage will be tested in a specialized rocket engine and stage hot test facility which deflects the flame safely to avoid any hazard to engine and adjoining mechanical structures. The facility is equipped with water jets to sufficiently cool the rocket plume before impinging on deflector pit walls. The high temperature and high-pressure rocket exhaust jet is cooled with sufficient quantity of water to maintain the impingement wall temperature within safe working limits.In this study, a multi-phase computational fluid dynamics (CFD) methodology is developed for design of water-cooling jet configuration of deflector facility. Multi-phase model validation study using full phase treatment and droplet assumption for water injection is performed prior to water jet optimization analysis of high thrust LOX-kerosene test facility. Validation studies based on simulation and analysis of existing test stand is developed. A discrete particle (DPM) based two phase methodology is chosen to model water injection in RANS framework. The optimum cooling configuration is determined iteratively through numerous multi-phase simulations. The multi-phase analysis showed ineffectiveness of initial water injection scheme which results in facility wall temperature above safe limit. A multi-plane water injection scheme is devised there on with provision of water injection at two axial locations from the nozzle exit. A detailed parametric multi-phase analysis is carried out for design optimization of multi-plane water injection scheme and to determine thermal condition of the test facility. Analysis showed effective rocket plume cooling and lower facility wall temperature with two plane water injection schemes in comparison to initial single plane design configuration. The study highlights the role of CFD based methodology to formulate an engineering solution for design optimization of LOX-kerosene engine hot test facility.

Validation of Puck\u2019s failure criterion for CFRP composites in the cryogenic regime

For future launch vehicles, lightweight cryogenic pressure vessels are required for storage of the liquid hydrogen fuel. For their structural assessment, reliable and validated failure criteria are required. The present contribution provides an overview over the results of an ongoing research activity concerned with the validation of Puck’s composite failure criterion in the cryogenic regime. In a first step, an experimental investigation on unidirectionally fiber reinforced materials on coupon level has been performed. This test campaign has been complemented by tests on small breadboard-type specimens with an angle-ply stacking sequence. The specimens were featuring holes and tapered sections to provide stress gradients and concentrations. Test were performed at ambient temperature and in a liquid Helium environment. Puck’s failure criterion has been applied and found to provide a good prediction of first ply failure in both environments.

Guidance command generation and nonlinear dynamic inversion control for reusable launch vehicles

Future launch vehicle concepts and technologies for expendable and reusable launch vehicles are currently investigated by the DLR research projects Akira and X-tras. In particular, the winged Liquid Fly-back Booster concept Lfbb based on an LOX/LH2 propellant combination for vertical takeoff and vertical landing (VTVL), as well as the delta-winged horizontal takeoff and horizontal landing (HTHL) concept Aurora based on an LOX/Kerosene propellant combination are considered in these projects. Because of the complexity and risks involved in on-line trajectory optimization, off-line reference trajectories are still considered important for tracking purposes. In that sense, the goal of this paper is to investigate an off-line and general-purpose guidance and control (G&C) architecture for preliminary studies of reusable launch vehicles. This is done by using trajectory optimization combined with Modelica models for the generation of optimal guidance commands, and then trajectory tracking is performed by means of inner-loop feedback controls in terms of nonlinear dynamic inversion with prescribed desired dynamics. We showcase the advantages of this baseline G&C architecture in terms of early stability and controllability aspects during the preliminary design studies of an example configuration of a reusable launch vehicle investigated in the context of the research projects above mentioned.

A novel aspect of composite sandwich fairing structure optimization of a two-stage launch vehicle (Safir) using multidisciplinary design optimization independent subspace approach

In this article, a novel composite sandwich structure analysis of launch vehicle fairing is considered by a new multidisciplinary design optimization methodology. Among the most important roles of this method, in addition to the convergence of optimization process, is tackling with complicated composite structure discipline. The bidirectional coupling existing between composite structure and trajectory disciplines is one of the complex problems of this multidisciplinary design optimization method. Accordingly, multidisciplinary design optimization based on independent subspaces is employed using the fixed point iteration method to achieve the best convergence at system level and segregate the disciplines. Therefore, the two proposed subspaces overcome the difficulties of common mentioned multidisciplinary design optimization of launch vehicles as the main novelty of this study. The first subspace is a multidisciplinary design optimization which includes propulsion, aerodynamics, weight and trajectory disciplines. The second one includes the novel composite fairing structure optimization as the other single discipline optimization which is analytically and numerically considered as a compact problem and is regarded as the other novelty of this work. In a case study, by applying the proposed architecture on Safir launch vehicle and considering propulsion, trajectory and also composite sandwich fairing structure design as the variables and then performing an optimization process, the Safir fairing mass is reduced from 100 kg to 57.8 kg. This causes the launch vehicle gross mass to decrease from 26 tons to 25.2 tons due to the payload nature of fairing. This 3% mass decrease of an operational launch vehicle, despite the preservation of mission performance, can be called an industrial novelty which leads to the cost reduction of space transportation. The proposed system engineering demonstrates the great importance of using multidisciplinary design optimization in complicated designs using independent subspaces to be employed for the design of future launch vehicles. It can also be a road map for future designers of space vehicles, especially those who want to consider structure optimization in the design loop of launch vehicles.

Status and Trends of Smallsats and their Launch Vehicles — An Up-to-date Review

This paper presents an analysis of the scenario of small satellite and its correspondent launch vehicles. The miniaturization of electronics, together with reliability and performance increase as well as reduction of cost, have allowed the use of commercials-off-the-shelf in the space industry, fostering the Smallsat use. An analysis of the launched Smallsats during the last 20 years is accomplished and the main factors for the Smallsat (r)evolution, outlined. Based on historic data, future scenarios for different mass categories of Smallsats are presented. An analysis of current and future launch vehicles reveals that we are currently in a phase of transition, where old launch vehicles get retired and new ones enter the market. However, the satellite launch vehicle business has been established to carry payloads of thousands of kilos into low Earth orbit and has not adjusted itself to the market of Smallsats. As a result, there is only 1 launch vehicle for dedicated Smallsat launches commercially available, but it carries a high price tag. Several small lowcost launch vehicles under development are identified and the challenges to overcome, discussed. Since these small launch vehicles have similar complexity as huge launch vehicles, high development costs are intrinsic, leading to a high specific price (USD/kg payload).

A novel non-probabilistic reliability-based design optimization algorithm using enhanced chaos control method

In this study, an efficient and robust algorithm of non-probabilistic reliability-based design optimization (NRBDO) is proposed based on convex model. In this double-nested optimization model, the inner loop concerns a Min–max problem for the evaluation of reliability index, where the target performance approach is applied to substitute the Min–max problem. To improve the convergence rate, an enhanced chaos control (ECC) method is developed on the basis of chaotic dynamics theory, which can check and re-update the control factor by the Wolfe–Powell criterion. To further enhance the optimization efficiency, a novel NRBDO algorithm is developed based on the proposed ECC, where HL–RF algorithm is applied at the initial stage of this algorithm, while ECC is used to improve the robustness once the oscillation or chaotic behavior is identified. Three mathematical examples, one numerical example and one complex engineering problem, i.e. axially compressed stiffened shells in launch vehicles, are utilized to demonstrate the robustness and efficiency of the proposed method by comparison with other existing methods. Results indicate that the proposed method is particularly suitable for complicated engineering problems without prior knowledge of uncertainty distributions, which is expected to be utilized in the structural design of future launch vehicles.

Numerically and experimentally predicted knockdown factors for stiffened shells under axial compression

Stiffened shells in launch vehicles are very sensitive to various forms of imperfections. In this study, the imperfection sensitivity of a 4.5mdiam isogrid stiffened shell under axial compression is investigated. The measured imperfection, NASA SP-8007 and several types of assumed imperfections, including eigenmode-shape imperfection and dimple-shape imperfections (produced by the single perturbation load approach (SPLA) and worst multiple perturbation load approach (WMPLA)), are introduced into FE model to predict the knockdown factors (KDFs), respectively. Then, the buckling test of this full-scale stiffened shell under axial compression is carried out to validate the above numerical approaches. It can be found that the KDF predicted by the WMPLA is very close to the test results, while the ones predicted by eigenmode-shape imperfection and NASA SP-8007 are extremely conservative. Besides, the measured imperfection and other assumed imperfections are proven to be risky, because these methods overestimate the actual load-carrying capacity. Finally, it can be concluded that the WMPLA is a potential and efficient approach to predict KDFs in the design stages for future launch vehicles.

Ecological aspect of launch vehicles development by criterion of minimal cost

One of the topical problems in modern aerospace engineering is accordance between ecological requirements and performance of the vehicle. On the other hand, problem of economical efficiency leads to change of the main criterion of designing to the minimization of costs (instead of maximal performance). According to modern trends of “low-cost” vehicles, different concepts of the future cost-effective launch vehicles are considered. It is necessary to validate these concepts according to requirements of ecological safety for the purpose of detection of the dominant launch vehicle configuration. Typical configurations of the future 'low-cost' launch vehicle are presented by 6 conceptual groups (Koelle, 2001). Conceptual group 1 (CG1) is presented by the Ballistic “Single stage to orbit” (SSTO) reusable vehicle. All vehicles which use classical rocketry scheme of the propulsion trajectory are called “Ballistic” i.e. the ballistic vehicle is lifted to orbit under the impact of rocket engines thrust. CG1-vehicle is able to reach the low earth orbit (LEO) without stage separation reducing the number of required rocket engines. Technological feasibility of SSTO concepts is proven by numerous studies (Koelle, 2001). CG2 representatives are ballistic “Two stages to orbit” (TSTO) reusable vehicles. The difference between CG1 and CG2 consists in application of vacuum rocket engines in the second stage and, consequently, stage separation. CG2 are the most mass-effective vehicles. CG3 is presented by the winged SSTO vehicles with rocket propulsion by “Lifting body” aerodynamic scheme. Ascensional force is provided by the aerodynamic shape of the vehicle’s structure at high speeds. Winged TSTO vehicles with rocket propulsion and parallel or tandem staging form the CG4. The winged configuration provides wide landing capability for both stages. CG5 is presented by winged TSTO vehicles with airbreathing propulsion in the first stage and rocket-propelled second stage. Airbreathing jet engines provide high reusability ratio comparing with other concepts as well as the widest landing capability. Aerospace Plane with scramjet-rocket propulsion forms CG6. The vehicle is able to reach near-cosmic speed in rarefied layers of the atmosphere and then accelerate with rocket engines. The most ecologically important resemblance of represented concepts is reusability. This reduces space debris formation (due to lack of waste hardware). Reusable launch vehicles can also be used to return the spent satellites. Structural differences between the concepts form 3 criterions of comparison by ecological impact: 1) propellant toxicity; 2) safety of surface facilities (vehicle damage inside the atmosphere); 3) probability of space debris formation (vehicle damage outside the atmosphere). Comparison of the concepts by these criterions allows substantiating the most ecologically acceptable direction of research. Results of the comparison demonstrate that the most ecologically acceptable low-cost launch vehicle configuration is: Ballistic SSTO or TSTO reusable launch vehicle with “LOX+LH2” propellant. The results can be explained by following way: combustion products of the propellant “liquid oxygen + liquid hydrogen” are absolutely safe for environment. It also provides maximal performance of rocket engine (due to the highest specific impulse). Ballistic ascent scheme allows using relatively simple technologies and provides high reliability level. In combination with minimal time of atmospheric flight this provides high level of safety for surface facilities. These results may be used for substantiation of dominant research direction.

Experimental studies for semi-cryo engine development

Development of Cryogenic engine is essential requirement in order to improve the pay load capability of rocket launching systems. Cryogenic stage is more suitable for the upper stage of the launch vehicle mainly due to the high specific impulse and relatively low structural weight. As the pay load increases a high thrust booster engine is also required for the launch vehicle. Semi cryogenic propulsion is ideally suitable for booster engine, in terms of higher density impulse and low cost and easy handling. LPSC has undertaken the development of Semi Cryo Engine to power the booster stages of India's future launch vehicles. The Engine works on Oxidizer rich staged combustion cycle, using LOX and Isrosene propellant combination. Preburner (PB) and Thrust Chamber (TC) are the major subsystems of the engine. PB produces the hot gases at lower temperature for driving the turbo pump which raises the propellant pressure and TC produces hot gas in the chamber at high temperature which expands through the nozzle for generating thrust. Design of injector and arrangement of injector are critical in the design of pre burner and Thrust chamber so as to admit and meter the propellants into the chamber with proper mixture to ensure complete and stable combustion. The propellants require ignition source for initiating ignition. A mixture of Tri-ethyl Aluminium and Tri-ethyl Boron (TEA/TEB) is used as a source of ignition which is hypergolic with oxygen. In order to understand the behavior of the different systems and technology involved, it is planned to design and realize injector and evaluate the characteristics of its performance during cold flow as well as hot test. The development programme is planned in two phases. Initially Single element pre burner and thrust chamber hardware are realized, cold flow test are carried out and hot test are being done. It is also planned to realize subscale thrust chamber and pre burner as a phase-2 activity to study the behavior of multi element interaction and its performance. A lot of critical technologies are to be developed and demonstrated prior to the development of full scale engine. The characterization of injector elements, performance validation at near operating conditions, Ignition characteristics of hypergolic liquid, validation of material, evaluation of different coating are set for the objectives of the initial phase of tests. This paper discuss the different experimental studies carried out, experimental results and features of hardware, design aspects, cold flow characteristics, hot test result for the initial phase of development and also future plan of action.

Postures and Motions Library Development for Verification of Ground Crew Human Factors Requirements

Spacecraft and launch vehicle ground processing activities require a variety of unique human activities. These activities are being documented in a primitive motion capture library. The library will be used by human factors engineering analysts to infuse real to life human activities into the CAD models to verify ground systems human factors requirements. As the primitive models are being developed for the library, the project has selected several current human factors issues to be addressed for the Space Launch System (SLS) and Orion launch systems. This paper explains how the motion capture of unique ground systems activities is being used to verify the human factors engineering requirements for ground systems used to process the SLS and Orion vehicles, and how the primitive models will be applied to future spacecraft and launch vehicle processing.

Rocket Noise Sources Localization Through a Tailored Beam-Forming Technique

A source localization technique is applied to analyze wall pressure measurements undertaken on the surface of a space launcher mock-up. The technique is based on the combined use of a standard beam-forming technique and elementary acoustic solutions tailored to the space launcher geometry. Two classes of acoustic elementary solutions are considered: plane waves impinging on the surface of an in nite cylinder computed analytically trough a literature formula, and plane waves impinging on the real space launcher computed numerically through a computational acoustic technique. In both cases, the 3D acoustic eld resulting from the impingement of an arbitrarily oriented plane wave is obtained by a truncated series summation of elementary axial-symmetric solutions. These two classes of basis functions are proven to provide consistent source localization results in the addressed frequency regime. However, although the analytical functional basis enables a faster localization of the noise sources that take place during the rocket ring, the numerical functional basis is expected to enable, at an acceptable computational cost, a more reliable reconstruction of the acoustic loads on the space launcher surface in a higher frequency regime. The proposed CAA-based beam-forming and source reconstruction technique is therefore a useful tool for the vibro-acoustic design of future launch vehicles.

Cryogenic Moisture Uptake in Foam Insulation for Space Launch Vehicles

Rigid polyurethane foams and rigid polyisocyanurate foams (spray-on foam insulation), like those flown on Shuttle, Delta IV, and will be flown on Ares-I and Ares-V, can gain an extraordinary amount of water when under cryogenic conditions for several hours. These foams, when exposed for eight hours to launch pad environments on one side and cryogenic temperature on the other, increase their weight from 35 to 80 percent depending on the duration of weathering or aging. This effect translates into several thousand pounds of additional weight for space vehicles at lift-off. A new cryogenic moisture uptake apparatus was designed to determine the amount of water/ice taken into the specimen under actual-use propellant loading conditions. This experimental study included the measurement of the amount of moisture uptake within different foam materials. Results of testing using both aged specimens and weathered specimens are presented. To better understand cryogenic foam insulation performance, cryogenic moisture testing is shown to be essential. The implications for future launch vehicle thermal protection system design and flight performance are discussed.

Spray-on foam insulations for launch vehicle cryogenic tanks

Spray-on foam insulation (SOFI) has been developed for use on the cryogenic tanks of space launch vehicles beginning in the 1960s with the Apollo program. The use of SOFI was further developed for the Space Shuttle program. The External Tank (ET) of the Space Shuttle, consisting of a forward liquid oxygen tank in line with an aft liquid hydrogen tank, requires thermal insulation over its outer surface to prevent ice formation and avoid in-flight damage to the ceramic tile thermal protection system on the adjacent Orbiter. The insulation also provides system control and stability throughout the lengthy process of cooldown, loading, and replenishing the tank. There are two main types of SOFI used on the ET: acreage (with the rind) and closeout (machined surface). The thermal performance of the seemingly simple SOFI system is a complex array of many variables starting with the large temperature difference of 200–260K through the typical 25-mm thickness. Environmental factors include air temperature and humidity, wind speed, solar exposure, and aging or weathering history. Additional factors include manufacturing details, launch processing operations, and number of cryogenic thermal cycles. The study of the cryogenic thermal performance of SOFI under large temperature differentials is the subject of this article. The amount of moisture taken into the foam during the cold soak phase, termed Cryogenic Moisture Uptake, must also be considered. The heat leakage rates through these foams were measured under representative conditions using laboratory standard liquid nitrogen boiloff apparatus. Test articles included baseline, aged, and weathered specimens. Testing was performed over the entire pressure range from high vacuum to ambient pressure. Values for apparent thermal conductivity and heat flux were calculated and compared with prior data. As the prior data of record was obtained for small temperature differentials on non-weathered foams, analysis of the different methods is provided. Recent advancements and applications of SOFI systems on future launch vehicles and spacecraft are also addressed.

Potential Composite Structures for NASA Future Launch Vehicles and Crew Spacecraft

Because of their well known specific strength, stiffness and excellent durability properties, advanced composites are being considered for primary structures in launch vehicles, crew modules and various components for increased performance and cost reduction. Therefore, it is prudent and beneficial to review the engineering practices and lessons learned in connection with their use in related aeronautic and energy applications, where they are already replacing formerly used metallic materials. Examples of composite components will be shown for the reusable space shuttle orbiter where a number of different composite systems performed very satisfactorily. In addition, very large potential polymer composite designs for future launch vehicles will be discussed. Among them are payload shrouds, interstage structures and the typical intertank shell, wherein thrust booster rockets are often attached between the core stage propellant and oxidizer tanks. In addition, cryogenic propellant composite vessels of different sizes and shapes were fabricated with mostly excellent results, although some spectacular failures were also observed. High pressure composite overwrapped vessels, with and without metallic liners, will be described. Compared to widely used metallic materials, some special features of composites are listed. Relevant design allowables, depending on mission requirements, will be summarized and currently used design practice for aircraft and spacecraft in the US will be reviewed. The well-known “building - block” approach, which is often used to design military as well as civilian aircraft, will be summarized. Some of the most popular micromechanics and macromechanics computer programs used to analyze composite structures, especially with finite elements, will be listed. Although very high temperature composites like carbon/carbon, carbon/ceramic and ceramic/ceramic fiber/matrix systems are also selectively used, the focus in our discussion will be on advanced polymer matrix, carbon/glass fiber systems.